How do rocket engines work? Rocket fuel saga
Federal State Budgetary Educational Institution of Higher Professional Education
(FGBOU VPO)
Astrakhan State Technical University (ASTU)
"Institute of Marine Technology, Energy and Transport" (IMTEiT)
Department "Heat power engineering" (TEN)
Course work
in the discipline "Fuel"
on the topic " Rocket propellants»
Fulfilled
student of the TET-21 group
Prikazchikov A.A.
Reviewers:
students of the TET-21 group
Putyatin S.S., Zhidkov S.M.
Teacher:
Doctor of Chemistry, Professor Ryabukhin Yu.I.
Astrakhan - 2012
1. Historical background
Main types of rocket fuel
1 Liquid propellants
1.1 Oxidizers
1.2 Fuel
1.3 Comparison of the most common liquid propellants
2 Solid propellants
2.1 Rocket propellants
2.2 Mixed propellants
Bibliography
. History reference
Solid fuel rockets appeared much earlier than rockets with liquid rocket engines (LRE). The latter have become so familiar to us that we forget about when they began to be used to conquer space and in the combat operations of the belligerents. And this happened just some 50 years ago. Prior to this, solid-propellant rockets, or rockets with powder engines, had been successfully operated and used by the troops for several centuries. The possibility of using liquids, including liquid hydrogen H2 and oxygen O2, as fuel for rockets was pointed out by K. E. Tsiolkovsky<#"justify">2. MAIN TYPES OF ROCKET FUEL
The choice of propellant depends on many factors. There is no ideal fuel, each has its pros and cons. Factors such as price, specific impulse, burn rate, burn rate versus pressure function, safety and manufacturability, and others can influence the choice of fuel.
2.1 LIQUID PROPELLANTS
Oxidizerand fuelbicomponent fuels are contained in separate containers - tanks and, using various devices, are separately fed into the engine chamber for combustion. Two-propellant liquid fuels are currently the most widely used, as they provide the highest specific engine thrust, easily allow you to adjust the amount and direction of thrust in flight, as well as turn off the engine and start it again. The disadvantage of these fuels is a complex engine device with a large number of parts and assemblies with a complex control and regulation system.
To self-ignitinginclude such two-component fuels, the combustion of which begins by itself when the oxidizer and fuel are mixed in the engine chamber.
Non self-ignitingfuel to start burning when starting engines require the use of additional means of ignition. Self-igniting fuels provide more reliable engine start and stable operation.
Liquid one-componentfuels are pre-prepared non-self-igniting mixture of oxidizer and fuelin the ratio necessary for combustion, or such a liquid substance that, under certain conditions, decomposes with the release of heat and the formation of gases. Single-component propellants are placed on the rocket in one tank and are fed into the combustion chamber through nozzles through one line.
advantageof such fuels before two-component ones is simplification of the engine designas only one supply line is needed. But these fuels have not been widely used in liquid-propellant rocket engines, since they cannot provide the necessary specific thrust. Those single-component propellants that allow obtaining sufficient specific thrust are unsuitable for use due to their great tendency to spontaneous explosion. Single-component fuels are also dangerous for their use in order to cool the combustion chamber. These fuels are used for the most part only for auxiliary purposes: for low-thrust engines, which are used to control and stabilize aircraft, as well as for the rotation of the turbines of the LRE turbopump units.
Table 1. The main characteristics of two-component liquid fuels at the optimal ratio of components (pressure in the combustion chamber 100 kgf/cm 2, at the nozzle exit 1 kgf / cm2 ).
OxidizerFuelThermal value of the fuel*, kcal/kgDensity*, g/cm2Temperature in the combustion chamber, KV specific impulse in vacuum, sNitric acid (98%) 20 %)14201,393050313Жидкий кислородСпирт (94 %)20200,393300255Водород20200,323250391Керосин22001,043755335НДМГ 22001,023670344Гидразин22301,073446346Аммиак22000,843070323АТКеросин15501,273516309НДМГ22001,203469318Гидразин22301,233287322Жидкий фторВодород23000,624707412Гидразин22301,314775370
In two-component fuels, for the complete combustion of both components, for each unit mass of one of them, a strictly defined amount of the other is required. So, to burn 1 kg of kerosene, 15 kg of air, or 5.5 kg of nitric acid, or 3.4 kg of liquid oxygen are needed. AT practically completed LRE the oxidizer is fed into the chamber in a slightly smaller amountthan required for complete combustion.
It turns out that in this case the highest value of specific thrust is obtained. The reason is that with a decrease in the consumption of the oxidizer, the composition of the combustion products changes somewhat. As a result, the process of thermal decomposition of gas molecules - combustion products - into atoms and ions, which occurs with a large absorption of heat and its useless entrainment outside the nozzle, is reduced, and the conditions for energy conversion in the nozzle are also improved.
For the operation of liquid rockets, the boiling point of the fuel is of great importance. All fuel components are divided into high-boilingand low-boiling.
To high-boilinginclude oxidizers and combustibles that can be contained in a liquid state at normal missile operating temperatures (up to +150 0C) under atmospheric or elevated pressure, the rest refer to low-boiling.
2.1.1 Oxidizers
in liquid rockets the amount of oxidizing agent by mass exceeds the amount of fuelon average 3-6 times, and the mass of fuel is 9 times greater than the mass of the engine structure.
Fuel properties largely depend on the nature of the oxidizer. For example, according to the most important characteristic - specific thrust - the fuel "liquid oxygen and kerosene" differs from the fuel "nitric acid and kerosene" by about 15%.
Of the low-boiling oxidizers, the most widely used in common engines is liquid oxygen. The possibility of using liquid fluorine, its connections with oxygen and ozone.
Of the high-boiling ones, they are widely used Nitric acidand its mixtures with nitrogen tetroxide. Can be applied nitrogen tetroxide, hydrogen peroxide. Compounds under investigation fluorineWith chlorineand tetranitromethane.
Consider some types of oxidizers.
1. LIQUID OXYGEN (O 2 ). It is a mobile liquid of a bluish color, slightly heavier than water.
Peculiarities : oxygen is one of the most powerful oxidizers, since its molecule does not contain atoms that are not involved in the oxidation process, as is the case, for example, in nitric acid. Fuels are more efficient than with oxygencan only be obtained from ozone, fluorineor fluoride oxygen.
Main property, which determines the features of working with liquid oxygen, lies in its low boiling point. Because of this, it evaporates very quickly, which causes its large losses during storage and refueling of the rocket. The rocket tank is filled with liquid oxygenjust before the rocket launch. Evaporation losses during refueling are up to 50%, and when contained in a rocket up to 3% per hour. Liquid oxygenstored and transported in special containers - metal tanks with good thermal insulation.
Liquid oxygen not poisonous. Briefly contact him in not large quantities with open areas of the human body is not dangerous: the resulting gaseous layer does not allow freezing of the skin.
Liquid oxygen- one of the most cheap oxidizers, which is explained by the ease of production and the abundance of raw materials. It is 89% by mass in water and 23% in air. Usually receive oxygenfrom the air, by liquefaction and separation in liquid form from nitrogenand other gases earth's atmosphere.
2. NITRIC ACID (HNO 3 ) . Chemically pure 100% nitric acid is a colorless, easily mobile, heavy liquid that strongly smokes in the air.
Peculiarities : 100% nitric acid unstable and easily decomposedon the water oxygenand nitrogen oxides.
HNO 3 - Powerful oxidizerbecause its molecule contains
% oxygen. During the oxidation of various combustibles, it decomposes into water, oxygenand nitrogen. It compares favorably with all commonly used oxidizing agents large specific gravity. Due to high heat capacityit can be used as a cooling component of the LRE chamber.
Under normal operating conditions Nitric acid- liquid, which is one of its advantages. rockets,in which it is used as an oxidizing agent, can be stored refilled for a long time, in constant readiness for launch. Operational disadvantages include a significant increase in pressurein hermetically sealed containers nitric acid,due to the process of its decomposition. Main disadvantage nitric acid - high corrosivityfor most materials. Aggressiveness nitric acidmakes it much more difficult to handle. Its storage and transportation is carried out using special containers.
Flaws : Nitric acidhas poisonousproperties. Contact with it on human skin causes the appearance of painful, long-term non-healing ulcers. Vapors are also harmful to health nitric acid. They are more venomous carbon monoxide 10 times.
Price nitric acidsmall. Main receiving method nitric acidinvolved in the oxidation ammonia oxygenair in the presence platinumand dissolving the resulting nitrogen oxides in water.
N 2+ 2O2 => 2 NO 2
. DINITROGEN TETRAOXIDE (N 2 O 4 ) . It is a yellow liquid at normal temperature.
Peculiarities : with increasing temperature, it decomposes into nitrogen dioxide, painted in red-brown color, the so-called "brown gas".
Is somewhat more efficient oxidant, how Nitric acid. Fuels based on it have a specific thrust of about 5% more than nitric acid.
Flaws : relative to materials dinitrogen tetroxideh much less aggressive, how Nitric acid, but not less poisonous.
The main disadvantage is low boiling pointand heat hardening, which sharply reduces the possibility of its use in rocket fuels in its pure form. The conditions for its use are improved in mixtures with other nitrogen oxides.
4. HYDROGEN PEROXIDE (H 2 O 2 ). Colorless transparent heavy liquid.
Peculiarities: hydrogen peroxide is an unstable chemical compound that easily decomposes into water and oxygen. The tendency to decompose increases with increasing concentration. During decomposition, a significant amount of heat is released.
The most widespread are aqueous solutions of 80% and 90% concentration of hydrogen peroxide. The chemical stability of solutions and the safety of working with them can be achieved by introducing stabilizing substances. These include phosphoric, aceticand oxalic acid. Mandatory stabilization conditionhydrogen peroxide - purity. Minor impuritiesand pollution sharply accelerate its decompositionand may even lead to an explosion.
Compared with nitric acid hydrogen peroxidehas low corrosivity, but it oxidizes some metals.
Flaws : Hydrogen peroxide is flammable and explosive. Organic substances in contact with it easily ignite. At a temperature of +175 0C it explodes. Contact with the skin causes severe burns.
At present, hydrogen peroxide is little used, since fuels based on it give a relatively low thrust.
5. LIQUID FLUORINE (F 2 ). It is a heavy liquid of bright yellow color.
Peculiarities: fluorine has best oxidizing properties, how oxygen. Of all chemical elements he is the most active, entering into compounds with almost all oxidizing substances at ordinary room temperature. In this case, ignition often occurs. Even oxygenoxidized fluorineburning up in its atmosphere.
Due to its exceptionally high chemical activity fluorinewith all combustible forms self-igniting fuels. However, fluorine fuels give higher specific thrust than oxygen, only if the fuel is rich hydrogen. Combustible containing many carbon, form with fluorinemuch less efficient fuels.
Flaws : fluorinevery poisonous. It is highly corrosive to the skin, eyes, and respiratory tract. In rocket technology, it is still used only in experimental engines.
2.1.2 Fuel
As a fuel in liquid fuels, mainly substances are used in which the oxidized atoms of chemical elements are atoms carbonand hydrogen. In nature, there is an extremely large number of chemical compounds of these elements. Most of them are organic.
Currently, rocket technology uses a wide variety of fuels. Despite the fact that fuel makes up only 15-25% of the mass of fuel, it right choice is of great importance. Only with a successful combination of oxidizer and fuel can be satisfied, if not all, then at least the most important requirements for fuel. Most types of rocket fuel are high-boiling. Their common flaw - low specific gravity , one and a half to two times less than that of oxidizing agents.
In practice as rocket fuel most commonly used hydrocarbon, which is a product of oil refining (kerosene), amines, ammonia, hydrazineand its derivatives.
Consider some types of fuel.
1. HYDROCARBONS (petroleum products) are mixtures of chemical compounds carbonWith hydrogen. Their energy performance is lower than that of hydrogen, but higher than carbon. Kerosene is the most widely used.
Kerosene features: it is a light liquid with a high boiling point, which is highly resistant to decomposition when heated. Kerosene is not a substance of a strictly defined compositionwith an unambiguous chemical formula, which makes it impossible to accurately determine its properties. Depending on the oil field, the composition and properties of kerosene may vary. Rocket kerosene contains increased contentsuch hydrocarbons, which give less depositsduring engine cooling.
Disadvantages of kerosene: it does not ignite when in contact with common oxidizing agents, therefore requires a special ignition source.
Kerosene is widely used in rocket propellants with liquid oxygen, nitric acidoxidizers and hydrogen peroxide.
2. AMINES - compounds that are obtained if in the molecule ammoniaone, two or three atoms hydrogenreplace hydrocarbon groups. In rocket technology, triethylamine, aniline, xylidine, etc. have found application.
Peculiarity : amines actively interact withnitric acidand dinitrogen tetroxideleading to self-ignition. In terms of efficiency, fuel based aminesclose to kerosene. Ability amines cause corrosion of metals is small. They are stored and transported in containers made of ordinary ferrous metals.
Flaws: amines significantly higher costcompared to kerosene , as well as toxicity, which manifests itself both by inhalation of vapors and by contact with the skin.
To improve the physico-chemical properties, aminesused as a fuel in mixture with other substances, including other amines.
fuel based aminesfound application in self-igniting fuels with nitric acid, nitrogen tetroxide and their mixtures.
3. HYDRAZINE . During the combustion of hydrazine, only atoms participate in the oxidation reaction hydrogen, a nitrogenreleased in free form, increasing the amount of gas.
Hydrazine is a colorless, transparent liquid (in about the same temperature range as water) and has an ammoniacal smell. Usually used in mixtures with other substances.
Peculiarities: hydrazine is an efficient fuel. This is facilitated by the fact that its molecule is formed with the absorption of heat, which is released during combustion in addition to the heat of oxidation. Another positive feature is large specific gravity.
Flaws: hydrazine has high solidification temperaturewhich is very inconvenient to use. Its vapors explode when heated and struck. When exposed oxygenair, it oxidizes. Hydrazine corrosive. Resistant to it are aluminumand its alloys, stainless steels, polyethylene, polyfluoroethylene, fluoroplast. Hydrazine poisonous, irritating to the mucous membrane of the eyes and can cause temporary blindness.
4. ASYMMETRIC DIMETHYLHYDRAZINE It is a colorless transparent liquid with a pungent odor.
Peculiarities : compared to hydrazine, it is much more convenient to use, as it remains a liquid in a larger temperature range. It has good heat resistance. Unlike hydrazine, its vapors do not explode from external influences. main feature- high chemical activity. It is easily oxidized by atmospheric oxygen, and with carbonic acid forms salts that precipitate.
Flaws : dimethylhydrazine (compared to hydrazine) has a worse efficiency as a fuel, since its molecule contains less effective carbon atoms in addition to hydrogen atoms. Self-ignites in air at 250 0C, mixtures of dimethylhydrazine vapor with air explode easily, and it poisonous.
2.1.3 Comparison of the most common liquid propellants
. Fuels based on liquid oxygen provide the highest specific thrustof all currently used rocket fuels. Their main disadvantage is low boiling pointoxidizer. This makes it difficult to use them in combat missiles, which must be ready for launch for a long time.
With liquid oxygen, combustibles such as kerosene, asymmetric dimethylhydrazine, ammonia. Special placetakes fuel oxygen+ hydrogen, which provides a specific thrust 30-40% greater than other common fuels. This fuel is most suitable for use in large rockets.
2. Fuels based on nitric acid in a mixture of 20-30% nitrogen oxidesmuch inferior oxygenfuels by specific thrust, but have weight advantage. In addition, these fuels are high-boiling long-termsubstances, which allows you to keep combat missiles fully equipped and fueled for a long time.
Nitric acid oxidants have good cooling properties. But due to the relatively low temperatures in the combustion chamber, the cooling of engines of medium and large thrusts can be provided with fuel, although the composition of the fuel contains less fuel than the oxidizer.
Combustible as a mixture amines, unsymmetrical dimethylhydrazineand some other substances formwith nitric acid oxidants self-igniting fuels. Kerosene and others hydrocarbons require forced ignition.
3. Fuels based on nitrogen tetroxide give slightly higher specific thrustthan nitric acid, but have reduced specific gravity. Despite such operational disadvantages as high oxidizer solidification temperature, they find use in long-range missiles. These fuels have been replaced oxygenfuel, because they make it possible to store the rocket in a refueled state, ready to launch.
The advantage of fuel based on nitrogen tetroxide is also autoignition.
2.2 Solid propellants
By appearance all charges of solid fuel are dense solidsmostly dark colors. Rocket powders are usually dark brown in color and look like a horn-like substance. If they contain additives (in the form of soot, for example), then their color is black. Blended fuels come in black and black-gray colors depending on the color of the fuel and additives, and are usually similar to heavily vulcanized rubber, but less elastic and more brittle.
Solid fuels are practically safeboth in terms of the impact on the human body and in relation to various structural materials. When stored under normal conditions, they do not emit aggressive substances. Rocket gunpowder due to the volatile properties of the solvent - nitroglycerin (Fig. 1) - can cause short-term, not very severe headaches.
Fig.1. Structural formula of nitroglycerin
2.2.1 Rocket propellants
Rocket gunpowder is a complex multicomponent system in which each substance has its own role in order to obtain the desired properties of a particular type of gunpowder. The main components of gunpowder are cellulose nitrates,which, when burned, release the greatest amount of heat energy. They also determine the physicochemical properties of gunpowder. Consider some of the components of gunpowder.
1. CELLULOSE NITRATE , or nitrocellulose, are obtained by treating cellulose with a mixture of nitric and sulfuric acids. This processing is called nitration. Raw material - cellulose(fiber) - a substance widespread in nature, of which flax, hemp, cotton, etc. are almost entirely composed.
Cellulose nitrates are loose mass. They are flammableeven from a weak spark. Combustion occurs due to the oxygen contained in the nitro groups, and no external oxygen supply required. However, direct use nitrocelluloseas rocket fuel is excluded, since it is impossible to make a charge from it that burns according to a strictly defined law. Even after strong pressing, it has many pores. Its combustion occurs not only outside but also inside, because the combustible gas penetrates through the pores inside. Thereby an explosion may occurcapable of destroying the engine. To prevent this, they produce plasticization nitrocellulose, i.e., a solid solution of a homogeneous composition is prepared from it, without pores.
2. SOLVENTS-PLASTICIZERS nitrocellulose - nitroglycerine, nitroglycoland some other substances. They are the second main component of gunpowder both in terms of mass and energy reserve. They are often called non-volatile solvents, since they are not removed from the solution during the production process, but remain completely in the composition of the gunpowder.
NITROGLYCERINE - a substance formed during nitration trihydric alcohol glycerine- mixture nitricand sulfuric acid. It is a colorless oily liquid.
Nitroglycerine - powerful explosive. It explodes easily on impact or friction. Its combustion occurs due to the oxygen contained in the nitro groups. Since there is an excess of oxygen in its molecule, part of the oxygen goes to the additional oxidation of nitrocellulose, which leads to an overall increase in the energy reserve of solid fuel. With an increase in the content of nitroglycerin in gunpowder are growingnot only them energy indicators, but also explosivenessand shock sensitivity. Rocket powders with a high content of nitroglycerin provide high specific thrust.
For plasticizing nitrocellulosein order to facilitate production technology, increase the time and allowable temperature Other solvents are also used for storage of charges.
NITROGLYCOL like an explosive less sensitive to mechanical stress. It is obtained by nitration ethylene glycol. stock oxygenless in its molecule than in the molecule nitroglycerin, so use as a solvent worsens energy performance gunpowder.
Except nitroglycerinand nitroglycolsometimes a solvent is used nitrocellulose, how nitroguanidine.
3. ADDITIONAL PLASTICIZERS and substances that regulate the energy properties of the fuel are well combined with basic solvents. They contain no or very little active oxygenand therefore are introduced into the composition of gunpowder in small quantities, so as not to reduce their energy characteristics. These include substances such as dinitroluene,dibutyl phthalate, diethyl phthalate.
4. STABILIZERS are introduced into the composition of gunpowders to increase their chemical resistance. Decomposition occurs during storage of gunpowder. nitrocellulosewith education nitrogen oxides, which accelerate its further decomposition, making it explosive. Stabilizers slow down decomposition nitrocellulose, connecting with the prominent nitrogen oxides, they bind them, turning them into chemically inactive substances.
5. COMBUSTION IMPROVING SUBSTANCES GUNPOWDER , provide acceleration, slowdownor stabilizationcombustion process in the chamber of solid rocket engines. These include a large number of salts or oxides of various metals ( tinsn , manganeseMn , zincZn , chromeCr , leadPb , titaniumTi , potassiumK , bariumBa etc.).
6. TECHNOLOGICAL ADDITIVES ? substances that facilitate the process of making gunpowder are introduced in the most critical operations for reduce friction and stress on machines. They play the role of lubricants both inside the fuel mass and between the mass and the tool. For this, chalk is used, which reduces internal friction, petroleum jelly and transformer oil, graphite, stearate leadand other substances reducing pressing pressure. They are introduced in small quantities.
The production of rocket powders is carried out according to a complex technological scheme using high temperatures and pressures. The task of production includes the manufacture of solid homogeneous powder charges that meet a number of stringent requirements, from a large number of substances that are heterogeneous in chemical and physical properties, as well as the state of aggregation.
2.2.2 Mixed propellants
Mixed fuels are much simpler in composition compared to gunpowder. They include two or three, rarely four components. Let's consider some of them.
1. AS OXIDIZERS MIXED FUELS usually salts of inorganic acids are used - nitricand chloride. Their feature is a large percentage of oxygen in the molecule. All of them, by weight, are about half composed of oxygen. Under normal conditions, they have chemical resistance, but with strong heating with able to decompose with the release of free oxygen.All solid oxidizing agents contain, in addition to oxygen, atoms of chemical elements capable of oxidation. Therefore, during the decomposition of these oxidizing agents, a part oxygenturns out to be associated with these elements and free oxygenmuch less is released than is available in the molecule.
The most common oxidizing agent for solid fuels is PERCHLORATE AMMONIA . This salt is a white (colorless) crystalline powder and decomposes when heated above 150 0C. In the air slightly moistened. Sensitive to impact and friction, especially in the presence of organic impurities. Can burn without fuel and explode. When burning, it does not emit solid substances, but its combustion products contain an aggressive and rather toxic gas - hydrogen chloride (HCl), which, in the presence of moisture, forms with it hydrochloric acid. The advantages of ammonium perchlorate are that it has a low decomposition temperature and decomposes only into gaseous products with a small molecular weight, has low hygroscopicity, is available, and cheap.
Another oxidizing agent is POTASSIUM PERCHLORATE . This salt decomposes at temperatures above 440 0C, does not moisten in air (non-hygroscopic), does not burn and does not explode. All the oxygen contained in its composition is active. When burned, it releases a solid substance - potassium chloride, which creates a dense smoke cloud. The presence of potassium chloride in combustion products sharply worsens the properties of rocket fuels, i.e., the conditions for the transition of thermal energy into kinetic energy in the nozzle of a rocket engine.
Another widely used oxidizing agent is AMMONIUM NITRATE (ammonium nitrate), also used as a nitrogen fertilizer. It is a colorless (white) crystalline powder. Decomposes at a temperature of 243 0C. Capable of burning and exploding. During combustion, a large amount of only gaseous products is released. Mixtures with organic substances are capable of spontaneous combustion, so the storage of rocket fuels based on it is a serious problem. Has poisonous properties.
The examples given do not exhaust the list of possible oxidizers for solid rocket engines, which can be used, for example, lithium perchlorates, nitrosyland nitronium, dinitrate hydrazine and etc.
2. FUEL-BINDING SUBSTANCES of mixed fuels - this is high molecular weight organic compounds, or polymers. Polymerssuch compounds are called, the molecules of which consist of a very large number of elementary units of the same structure. Elementary links are interconnected into long chains of a linear or branched structure. The properties of the polymer depend on the chemical structure of elementary units, their number and mutual arrangement.
Many solid polymers are obtained from liquid substances - monomers, whose molecules consist of a relatively small number of atoms. Monomers are able to spontaneously combine into long chains - polymers? this process is called polymerization.
To speed up polymerization, or curing, some special substances are used, called initiators, or hardeners.
Many high-molecular compounds are able to mix well and stick together with powders (with a crystalline oxidizer and metal powder), and then turn into a solid monolithic mass after polymerization. When heated, some polymers soften, become viscous, and in this form can mix with fillers, holding them firmly. At the same time, they can be poured into molds and receive fuel charges. given sizes and shapes.
Synthetic compounds of the type rubbers, resins and plastics, as well as heavy oil products - asphalt and bitumen. The composition and properties of petroleum products vary over a very wide range, and the desired mechanical properties are retained only in a small temperature range. That's why synthetic substances are used more oftenhaving a more constant composition and better mechanical properties. In practice, rubbers are used - POLYURETHANE , BUTADIENE andPOLYSULFIDE , resin - POLYESTER , EPOXY AndUREA , as well as some plastics, which include atoms nitrogen, oxygen, sulfuror chlorine.
Main limitationspolymer resins and plastics as fuel-binding substances - low elasticityand increased brittleness at low temperatures. Synthetic rubbers are largely free from these shortcomings.
3. POWDER METALS can be introduced into the composition of mixed fuels as an additional combustible component. Suitable for this are metal beryllium, lithium, aluminum, magnesium, as well as some of their compounds. As a result of the introduction of these metals, energy boostfuel, i.e. increased specific thrustengines. In addition, metal additives increase the specific gravity of the fuel, which improves the performance of the engine and the rocket as a whole. It should be taken into account that the higher the content of metal-containing fuel, the higher the temperature of their combustion products. Almost all modern composite fuels contain metals as components.
The most efficient metallic fuel is BERYLLIUM , however, the prospects for the use of beryllium are very limited, because it reserves insignificant, and the products of combustion are very poisonous. The next most efficient metal is LITHIUM . Its use is hampered very low melting point (+186 0C) and self-ignition in airin a molten state. The most common and cheapest metallic fuel is ALUMINUM . The use of finely ground aluminum powder in mixed fuels is not only increases specific thrustengines, but improves reliabilitythem launchand increases the stability of fuel combustion. MAGNESIUM It is rarely used, since it gives low specific thrust in fuels.
In addition to pure metals, the use of their compounds with hydrogen (hydrides) as additional combustible substances is being studied.
4. CATALYSTS AND OTHER ADDITIVES are introduced into mixed fuels in small quantitiesfor improving the combustion process(soot, salts of some metals), givingfuel plastic properties(vegetable, mineral and synthetic oils), improved storage stability and formulation stability ( diethyl phthalate, ethyl centralite), facilitating production technology.
The technology for manufacturing charges from mixed propellants includes mixing propellant components, casting, and curing. In general, the process of manufacturing mixed propellants is simpler than that of gunpowder, however, in the manufacture of large-sized charges, great technological difficulties must be overcome.
Bibliography
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The powerful space rocket is propelled by the same force as the festive fireworks in the park of culture and recreation - the force of the reaction of gases flowing from the nozzle. Breaking out like a column of fire from the rocket engine, they push the engine itself and everything that is structurally connected with it in the opposite direction.
The main fundamental difference of any jet engine (rocket engines are a powerful branch of an extensive family of jet engines, direct reaction engines) is that it directly generates motion, sets in motion the vehicle associated with it without the participation of intermediate units called propulsors. In an aircraft powered by piston or turboprop engines, the motor drives the propeller, which, crashing into the air, throws a mass of air back and makes the aircraft fly forward. In this case, the propeller is the propeller. The ship's propeller works in a similar way: it throws off a lot of water. A car or train is driven by a wheel. And only a jet engine does not need support in the environment, in the mass from which the apparatus would be repelled. The mass that the jet engine throws back and thereby receives forward movement is located in itself. It is called the working fluid, or the working substance of the engine.
Usually, hot gases operating in an engine are formed during the combustion of fuel, i.e., during a chemical reaction of the rapid oxidation of a combustible substance. The chemical energy of the burning substances is converted into the thermal energy of the 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 forward 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 released during the reaction of nuclear fission or fusion. In some types of electric rocket engines, the working substance is accelerated even without the participation of heat 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 spaceflight. A large class of these machines, the so-called jet engines, use ambient air to oxidize fuel. Naturally, they can work only within the limits of the earth's atmosphere.
For work in space, two types of rocket thermochemical engines are used: solid propellant rocket engines (SRM) and liquid propellant rocket engines (LRE). 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. The solid propellant is a solid mixture of essential substances. In an LRE, fuel and oxidizer are stored in liquid form, usually in separate tanks, and ignition takes place in a combustion chamber where the fuel mixes with the oxidizer.
Rocket motion occurs when the working substance is discarded. It is far from indifferent to the speed with which the working fluid flows out of the nozzle of a jet engine. The physical law of conservation of momentum says that the momentum of the rocket (the product of its mass by the speed with which it flies) will be equal to the momentum of the working body. This means that the greater the mass of gases ejected from the nozzle and the speed of their outflow, the greater the thrust of the engine, the greater the speed can be given to the rocket, the greater its mass and payload can be.
In a large rocket engine, in 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, in addition to dividing it into stages, there is only one way - to increase the thrust of the engines. And to increase thrust without increasing fuel consumption, it is possible only by increasing the rate of outflow of gases from the nozzle.
There is a concept in rocket technology of the specific thrust of a rocket engine. Specific thrust is the thrust obtained in the engine at the expense of one kilogram of fuel in one second.
Specific thrust is identical to the specific impulse - the impulse developed by a rocket engine for every kilogram of fuel (working fluid) consumed. The specific impulse is determined by the ratio of engine thrust to the mass of fuel consumed in one second. Specific impulse is the most important characteristic of a rocket engine.
The specific impulse of the engine is proportional to the speed of the outflow of gases from the nozzle. Increasing the exhaust rate allows you to reduce fuel consumption per kilogram of thrust developed by the engine. The greater the specific thrust, the greater the speed of the expiration of the working fluid, the more economical the engine, the less fuel the rocket needs to complete the same flight.
And the velocity of the outflow 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 caloric content, energy efficiency 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 a decrease in the 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 the combustion of fuel, the greater the rate of their expiration. Therefore, hydrogen as a propellant component is doubly beneficial due to its high calorific value and low molecular weight.
A very important characteristic of a rocket engine is its specific gravity, that is, the mass of the engine per unit of its thrust. A rocket engine must develop a lot of thrust and at the same time be very light. After all, lifting each kilogram of load into space is given at a high price, and if the engine is heavy, then it will lift mainly only itself. Most jet engines generally have a relatively small specific gravity, but this indicator is especially good for LRE and solid propellant rocket engines. This is due to the simplicity of their device.
solid propellant rocket engine and rocket engine
Solid propellant 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. The engine is difficult to turn off until all the fuel has burned out. Its work is extremely difficult to regulate. The fuel must burn slowly, at a more or less constant rate, regardless of changes in pressure and temperature. It is possible to regulate the value of solid propellant thrust 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 supply of fuel. Solid propellant rockets with rotary nozzles have appeared, they are structurally quite complex, but this allows us to solve the problem of controlling the direction of thrust.
However, solid propellant 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 solid propellant rocket engines is the charge of solid fuel. The characteristics of the engine depend on the elements of the fuel, and on the structure and device of the charge. There are two main types of solid rocket propellants: dibasic, or colloidal, and mixed. Colloidal fuels are a solid homogeneous solution organic matter, whose molecules contain oxidizing and combustible elements. The most widely used solid solution of nitrocellulose and nitroglycerin.
Mixed fuels are mechanical mixtures of fuel and oxidizer. As an oxidizer in these fuels, inorganic crystalline substances - ammonium perchlorate, potassium perchlorate, etc. are usually used. Typically, such a fuel consists of three components: in addition to the oxidizer, it includes a polymer fuel that serves as a binder, and a second fuel in the form of powdered metal additives, which significantly improve the energy characteristics of the fuel. The binder fuel can be polyester and epoxy resins, polyurethane and polybutadiene rubber, etc. The second fuel is most often powdered aluminum, sometimes beryllium or magnesium. Blended fuels usually have a higher specific impulse than colloidal ones, greater density, greater stability, better stored, more manufacturable.
Charges of solid fuel are fastened to the body of the engine chamber (they are made by pouring fuel directly into the body) and loose, which are made separately and inserted into the body in the form of one or more checkers.
The geometric shape of the charge is very important. By changing it and using armor coatings on charge surfaces that should not burn, they 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 burning area remains unchanged. This happens if, for example, a block of solid fuel burns from the end or simultaneously from the outer and inner surfaces (for this, a cavity is made inside the charge). In regressive combustion, the combustion surface decreases. The flow is obtained if the cylindrical checker burns from the outer surface. And, finally, for progressive combustion, which provides an increase in pressure in the combustion chamber, an increase in the burning area is necessary. Most simple example 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 dispense with special external cooling. In astronautics, solid propellant rocket engines are currently used to a limited extent. Powerful solid propellant rocket engines are used on some American launch vehicles, such as the Titan rocket.
Large modern solid propellant rocket engines develop hundreds of tons of thrust, even more powerful engines with thousands of tons of thrust are being developed, solid fuels are being improved, and thrust control systems are being designed. And yet, in astronautics, rocket engines certainly dominate. The main reason for this is the lower efficiency of solid propellants. The best solid propellant rocket engines have a speed of outflow of gases from a nozzle of 2500 meters per second. LREs have a higher specific thrust and an exhaust velocity (for the best modern engines) of 3500 meters per second, and using fuel with a very high calorific value (for example, liquid hydrogen as fuel and liquid oxygen as an oxidizer), one can obtain an exhaust velocity of four s half a kilometer per second.
For the device and operation of the rocket engine great value has the fuel that the engine runs on.
Known fuels that release energy during the decomposition reaction, for example, hydrogen peroxide, hydrazine. They naturally consist of one component, one liquid. However, the most widely used in rocket technology are chemical propellants that release energy during the combustion reaction. They consist of an oxidizer and a fuel. Such fuels can also be one-component, that is, they can be one liquid. This may be a substance, the molecule of which includes both oxidizing and combustible elements, for example, nitromethane, or a mixture of an oxidizing agent and a fuel, or a solution of a fuel in an oxidizing agent. However, such fuels are usually prone to explosion and are of little use. The vast majority of liquid propellant rocket engines run on bipropellant. The oxidizer and fuel are stored in separate tanks and mixed in the engine chamber. The oxidizer usually makes up a large part of the mass of the fuel - it is consumed two to four times more than the fuel. The most commonly used oxidants are liquid oxygen, nitrogen tetroxide, nitric acid, and hydrogen peroxide. Kerosene, alcohol, hydrazine, ammonia, liquid hydrogen, etc. are used as fuel.
The Soviet carrier rocket Vostok operated on a fuel consisting of liquid oxygen and kerosene, which ensured the launch of many of our spacecraft with cosmonauts on board. The engines of the American Atlas and Titan rockets, the first stage of the Saturn-5 rocket, with the help of which the Apollo spacecraft were launched to the Moon, ran on the same fuel. Fuel consisting of liquid oxygen and kerosene is well mastered in production and operation, reliable and cheap. It is widely used in LRE.
Unsymmetrical dimethylhydrazine has been used as a fuel. This fuel paired with an oxidizing agent - liquid oxygen - is used in the RD-119 engine, which is widely used in launching Kosmos satellites. This engine achieved the highest specific impulse for liquid-propellant rocket engines operating on oxygen and high-boiling fuels.
The most effective of the currently widely used rocket fuels 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, ever more efficient rocket fuels is ongoing. Scientists and designers are working hard to use fluorine in LRE, which has a stronger oxidizing effect than oxygen. The fuels formed with the use of fluorine make it possible to obtain the highest specific impulse for a liquid-propellant rocket engine and have a high density. However, its use in LRE is hampered by the high chemical aggressiveness and toxicity of liquid fluorine, high combustion temperature (more than 4500 ° C) and high cost.
Nevertheless, a number of countries are developing and bench testing LRE on fluorine. For the first time, F. A. Tsander proposed the use of liquid fluorine for LRE in 1932, and in 1933, V. P. Glushzho proposed a mixture of liquid fluorine and liquid oxygen as an oxidizer.
Many fluorine-based fuels spontaneously ignite when an 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 allows to simplify the design of the LRE and increase its reliability. Some fuels become self-igniting when a catalyst is added. Thus, 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 fuel (if it is not self-igniting, then pyrotechnic or electric ignition is used, or injection of a portion of self-igniting starting fuel) occurs in the engine chamber. The chamber is the main unit of the rocket engine. It is in the chamber that the fuel components are mixed, it is burned, and as a result, gas is formed with a very high temperature (2000-4500 ° C) and under high pressure (tens and hundreds of atmospheres). Flowing out of the chamber, this gas creates a reactive force, the thrust of the engine. LRE chamber consists of a combustion chamber with a mixing head and a nozzle. Mixing of fuel components occurs in the mixing head, combustion takes place in the combustion chamber, and gases flow out through the nozzle. Usually, all chamber units are made as a single unit. 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 rocket engine. It is the so-called mixture formation-injection, spraying and mixing of fuel components. The fuel components - oxidizer and fuel - enter the mixing head of the chamber separately. Through the nozzles of the head, they are introduced into the chamber 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 to 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 fastest and most complete education of the fuel mixture burning in the chamber, to ensure that each oxidizer particle meets with a fuel particle.
The formation of a fuel mixture prepared for combustion consists of three processes that pass one into another - atomization of liquid components, their evaporation and mixing. When spraying - crushing the liquid into drops - its surface increases significantly and the evaporation process accelerates. Very important is the fineness and uniformity of spraying. The fineness of this process is characterized by the diameter of the resulting droplets: the smaller each droplet, the better. The next step in preparing the fuel for combustion after spraying 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 LRE chamber takes only two to eight thousandths of a second.
As a result of atomization and evaporation of the fuel components, oxidizer and fuel vapors are formed, from which the mixture burning in the engine chamber is obtained. The mixing of the 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 even in the liquid phase, during fuel atomization. With non-self-igniting fuels, combustion begins in the gas phase when heat is supplied from external source.
Liquid fuel components are fed into the chamber through nozzles located in the head. The most commonly used nozzles are of two types: jet or centrifugal. But now the fuel is sprayed, mixed, ignited. When it burns, a large amount of heat energy is released in the combustion chamber. Further energy conversion takes place in the nozzle. The successful design of the mixing head primarily determines the perfection of the engine - it ensures the completeness of fuel combustion, combustion stability, etc.
Nozzle - part of the combustion chamber in which the thermal energy of the compressed working fluid (mixture of gases) is converted into the kinetic energy of the gas flow, i.e., it accelerates to the speed of outflow from the engine. The nozzle usually consists of tapering and expanding parts, which are connected in the critical (minimum) section.
A very difficult task is to ensure the cooling of the LRE chamber. Typically, the chamber consists of two shells - an inner fire wall and an outer jacket. A liquid flows through the space between the shells, cooling the inner wall of the LRE chamber. Usually 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 returned to the chamber. Such cooling (regenerative) was first proposed by K. E. Tsiolkovsky and is widely used in rocket technology.
In most modern LREs, special turbopump units are used 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 formed when the combustion chamber of the engine is cooled. There are other pump drive systems.
The creation of modern liquid-propellant rocket engines requires a high level of development of science and technology, the perfection of design ideas, and advanced technology. The fact is that very high temperatures are reached in a liquid-propellant rocket engine, 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 rocket engine must have, especially during launches spacecraft with astronauts on board, a very high degree of reliability. It is high reliability and many other advantages that distinguish the liquid-propellant 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 guidance of the chief rocket engine designer V P. Glushko. These are the world's first mass-produced engines running on high-calorie fuel; liquid oxygen and kerosene. They have a high specific thrust, which made it possible to obtain huge power with relatively moderate fuel consumption. In the void, 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 the fuel, the other the oxidizer. Both fuel and oxidizer are fed 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, that is, it is used for cooling. Reliable cooling keeps the temperature inside the combustion chambers high. Oscillating steering combustion chambers, similar in design to the main ones, were first used in this engine to control the direction of thrust.
The engine of the second stage of the rocket "Vostok" RD-108 has a similar design. True, it has four steering cameras and some other differences. Its thrust in the void is 96 tons. Interestingly, it is launched on Earth at the same time as the first stage engines. The RD-107 and RD-108 engines of various modifications have been used for many years to launch spacecraft, artificial earth satellites, spacecraft to the Moon, Venus and Mars.
The second stage of the two-stage launch vehicle "Cosmos" is equipped with the RD-119 liquid-propellant rocket engine developed in 1958-1962 (also in the GDL-OKB), which has a thrust of 11 tons; The fuel of this engine is asymmetric dimethylhydrazine, the oxidizer is liquid oxygen. Titanium and other modern construction materials. Along with high reliability, a distinctive feature of this engine is 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 useful power of the Proton rocket propulsion systems is three times the power of the Vostok rocket engines and amounts to 60 million horsepower. These engines provide high combustion efficiency, significant pressure in the system, uniform and balanced outflow of combustion products from the nozzles.
At present, rocket engines have reached high degree perfection and their development continues. LREs of various classes have been created - from micro-rocket engines for attitude control and stabilization systems of aircraft with very little thrust (several kilograms or less) to huge powerful rocket engines with hundreds of tons of thrust (for example, the American G- 1 for the first stage of the Saturn-5 launch vehicle has a thrust of 690 tons (five such engines are installed on the rocket).
Liquid-propellant rocket engines are being developed on highly efficient fuels - mixtures of liquid hydrogen (fuel) and liquid oxygen or liquid fluorine as oxidizers. Long-storable propellant engines have been created that can operate during long-term space flights.
There are projects of combined rocket engines - turbojet and rocket-ramjet engines, which should be an organic combination of liquid-propellant rocket engines with air-jet ones. The creation of such engines makes it possible to use atmospheric oxygen as an oxidizing agent at the initial and final stages of a space flight and thereby reduce the fuel supply on board the rocket. Work is also underway to create the first stages of reuse. Such stages, equipped with air-jet 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 advanced models will be created. However, the possibilities of thermochemical rockets are limited by the very nature of the fuel, oxidizer, and reaction products. With the limited energy efficiency of rocket fuels, which does not allow to obtain a very high speed of the expiration 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 not only a matter of saving, but sometimes the most possible! and space flight.
Even to solve a relatively simpler task in the field of space flights - launching artificial Earth satellites, the starting 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 put into orbit. To achieve the second cosmic 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 flying 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 correct the flight orbit, slow down the spacecraft before landing on the target planet, take off to return to Earth, etc. The starting mass of thermochemical rockets for such flights becomes incredibly large - several million tons!
Scientists and engineers have long been thinking about what should be the rocket engines of the future? The eyes of scholars naturally turned 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. So, for example, 1 kilogram of uranium in a fission reaction can release as much energy as 1,700 tons of gasoline when burned. Reaction nuclear fusion gives more energy.
The use of nuclear energy makes it possible to drastically reduce the stock of fuel on board the 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 nuclear missile carries 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 separation of fuel and working substance comes into play. It becomes possible to use the working substance with the lowest molecular weight - hydrogen.
The chemical rocket also uses a combination of the relatively high energy efficiency of hydrogen with a low molecular weight. But there, the working substance is the product of hydrogen combustion with a molecular weight of 18. And the molecular weight of pure hydrogen, which can serve as the working body of a nuclear rocket engine, is 2. Reducing the molecular weight of the working substance by 9 times at a constant temperature allows you to increase the outflow rate by 3 times . Here it is, a tangible advantage of an atomic rocket engine!
We are talking about atomic rocket engines that use the energy of nuclear fission of heavy elements. The nuclear fusion reaction has so far been artificially carried out only in a hydrogen bomb, and a controlled thermonuclear fusion reaction is still a dream, despite the intensive work of many scientists in the world.
So, in an atomic rocket engine, it is possible to obtain a significant increase in the rate of outflow of gases 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 of the proposed schemes of atomic rocket engines, the working fluid is heated, washing the fuel elements of the reactor, then it expands in the nozzle and is ejected from the engine. The temperature is about the same as in chemical rocket engines. True, the engine itself is much more complex and heavy. Especially when you consider the need for a screen to protect astronauts from radiation on manned spacecraft. And yet, a nuclear rocket promises a considerable gain.
In the United States, under the so-called Rover program, intensive work is underway to create an atomic rocket engine. Projects of nuclear rocket engines have also arisen, in which the active zone 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 greatly increase the speed of the expiration of the working fluid. But the creation of such reactors is an extremely complicated matter: the nuclear fuel here is mixed with the working substance, and it is necessary to somehow separate it before ejecting the working substance from the engine nozzle. Otherwise there will be continuous losses. nuclear fuel, a deadly plume of high radiation will stretch behind the rocket. Yes, and the critical mass of nuclear fuel necessary to maintain reactions, in the 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 counterpart - the Shuttle reusable rocket system, leaves much to be desired in terms of its characteristics.
They turned out to be not so reusable. Launch boosters withstand the entire 3-4 flight, and the winged vehicle itself burns 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 the Earth, going into outer space and returning back. True, the main problem remains unresolved - the engine. Air-jet engines (WJ) of known types are capable of operating only up to a speed of 4-5 M (M is the speed of sound), and the first space speed, as you know, is 24 M. But even here, it seems, the first steps to success have already been outlined.
At the Aviadvigatele-Build-92 exhibition, held in Moscow, among all sorts of exhibits - from ancient steam engines for airships to giant turbines of ultra-modern transport aircraft - a small barrel modestly stood on the stand - the world's first and only hypersonic model (Hypersonic - from 6M and above) air-jet engine (scramjet). It was created at the Central Institute of Aviation Motors (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 ... Indeed, 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, a few decades ago, proposed the basic principle underlying all modern scramjet engines. The engine he developed was already capable of operating at hypersonic (above 5-6 Mach) speeds at that time. These people have created a miracle of technology, which, perhaps, will revolutionize space propulsion in the near future.
But let's not rush to "fit" a new engine to a space plane, whether it's Buran or Spiral, let's turn to theory. The fact is that each engine can only operate in a certain range, which is too narrow for space tasks, and it is far from easy to make it master hypersound. Let's see why.
In any WFD for successful work three must be met essential conditions. First of all, you need to compress the air as much as possible. Then burn the fuel without loss in the combustion chamber. And finally, with the help of a nozzle, the combustion products must expand to atmospheric pressure. Only then the efficiency will be high enough.
Look at the picture. Here is a diagram of the world's first hypersonic ramjet engine (scramjet). His first task - air compression - he solves in a very original way - on the principle of ... a cleaver. Imagine: a cleaver crashes into a soft dense log, the layers of wood in front of it remain unchanged, and compacted on the sides. The boundary between normal and denser layers is what scientists call a "compression shock." This is what happens in the engine. A pointed central body is located along its axis. Crashing into the air, it creates such a "jump" - a zone of high pressure. There is a "reflection" of air from the central body to the walls of the body. At the same time, it is repeatedly compressed additionally. The air speed decreases, and the temperature rises, the kinetic energy is converted into internal, thermal.
Now, in order for the fuel injected into the stream to completely burn out, it is desirable to get the speed as low as possible. But then the air temperature can reach 3-5 thousand degrees. It would seem good - the fuel will flare up like gunpowder. But even if there is real gunpowder here, the flash will not work. The thing is that at such high temperatures, along with the oxidation process, molecules also break down into individual atoms. If in the first energy is released, then in the second it is absorbed. And the paradox is that as the temperature rises, there may come a moment when more will be absorbed than released. In other words, the furnace will turn into ... a refrigerator.
The original way out of the situation back in 1956 was suggested by Professor Shchetinkov. He suggested compressing the air only until its supersonic speed is about the same as that of ... a bullet. As it is now recognized all over the world, only under these conditions is the operation of a scramjet possible.
But even here there are difficulties: even a mixture of hydrogen and air, known to us in the course of chemistry under the name "explosive gas", in such conditions will hardly have time to catch fire. And although liquid hydrogen was chosen as fuel for the engine, we had to resort to tricks. Hydrogen first cools the walls. By heating itself from -256 ° C to + 700 ° C, it saves the metal from melting. Part 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 are burning here, capable of instantly burning through a sheet of steel. They ignite the hydrogen-air mixture. The one that under normal conditions explodes from a spark dropped from a nylon shirt.
And here, perhaps the main task, which we and the Americans have spent about 30 years on. How to get complete combustion, having a chamber of acceptable length - 3-5 m? It is known that a theory without a test 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 planes, however, there are wind tunnels, but they are very, very expensive. For the final test of the scramjet, the designers installed their device in the nose of the rocket and accelerated it to the desired speed.
Let us clarify that this was not about creating a new type of rocket, but only about checking the quality of hydrogen combustion in the engine. She 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 happens if we want to increase this small exhibition model, making it suitable for lifting an aircraft into the air. Apparently, it will acquire the features of a heavy thirty-meter pipe with a huge diffuser and nozzle and a very modest combustion chamber. And who needs such an engine? Dead end? No, there is a way out and 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 into the air, it creates a series of shock waves, and all of them directly fall on the inlet of the combustion chamber. Hot gases emerging from it, expanding to atmospheric pressure, slide over the surface of the aft part of the aircraft, creating thrust, as in a good nozzle. At hypersonic speeds, this is possible! Surprisingly, theoretically, you can even do without a camera, and “simply” inject fuel near the protrusion on the belly of the VKS! You get an engine that doesn't seem to exist. It is called "external combustion" scramjet. True, its “simplicity” in research work is so expensive that so far no one has taken it seriously.
Therefore, let us return to the aerospace aircraft with a classic-type scramjet. Its start and acceleration to b M should take place using conventional turbojet engines. In the figure you see a unit consisting of a traditional turbojet engine and a scramjet located nearby. At "small" speeds, the scramjet is separated by a streamlined bulkhead and does not interfere with flight.
And on large ones, the partition blocks the air flow going into the turbojet engine, and the scramjet engine is turned on.
At first, everything will go well, but then, as speed increases, engine thrust will begin to fall, and appetites - fuel consumption - will increase. At this moment, his insatiable womb must be fed with liquid oxygen. Like it or not, you still have to take it with you. True, in quantities much smaller than on a conventional rocket. Somewhere about 60 kilometers from the Earth, the scramjet will stall from lack of air. This is where a small liquid-propellant rocket engine comes into play. The speed is already high, and the fuel with the oxidizer will “eat” quite a bit before entering orbit. With the launch weight equal to that of the rocket, the aerospace plane was launched into orbit with a payload 5-10 times greater. And the cost of launching each kilogram will be ten times lower than missiles. This is exactly what scientists and designers are striving for today.
How a liquid-propellant engine works and works
Liquid-propellant engines are currently used as engines for heavy rocket projectiles. air defense, long-range and stratospheric missiles, rocket planes, rocket air bombs, aerial torpedoes, etc. Sometimes rocket engines are also used as starting engines to facilitate the take-off of aircraft.
Keeping in mind the main purpose of LRE, we will get acquainted with their design and operation using two engines as examples: one for a long-range or stratospheric rocket, the other for a rocket aircraft. These particular engines are by no means typical and, of course, inferior in their data to the latest engines of this type, but they are still characteristic in many ways and give a fairly clear idea of \u200b\u200bthe modern liquid-propellant engine.
LRE for long-range or stratospheric rocket
Rockets of this type were used either as a long-range super-heavy projectile or for exploring the stratosphere. For military purposes, they were used by the Germans to bomb London in 1944. These missiles had about a ton of explosive and a flight range of about 300 km. When exploring the stratosphere, the rocket head carries various research equipment instead of explosives and usually has a device for separation from the rocket and parachute descent. Rocket lift height 150–180 km.
The appearance of such a rocket is shown in Fig. 26, and its section in Fig. 27. The figures of people standing next to the rocket give an idea of the impressive size of the rocket: its total length is 14 m, diameter about 1.7 m, and plumage about 3.6 m, the weight of an equipped rocket with explosives is 12.5 tons.
Fig. 26. Preparing to launch a stratospheric rocket.
The rocket is propelled by a liquid-propellant engine located at its rear. General form engine is shown in Fig. 28. The engine runs on two-component fuel - ordinary wine (ethyl) alcohol 75% strength and liquid oxygen, which are stored in two separate large tanks, as shown in Fig. 27. The stock of fuel on the rocket is about 9 tons, which is almost 3/4 of the total weight of the rocket, and in terms of volume, the fuel tanks make up most of the entire volume of the rocket. Despite such a huge amount of fuel, it is only enough for 1 minute of engine operation, since the engine consumes more than 125 kg fuel per second.
Fig. 27. A section of a long-range missile.
The amount of both fuel components, alcohol and oxygen, is calculated so that they burn out simultaneously. Since for combustion 1 kg alcohol in this case consumes about 1.3 kg oxygen, the fuel tank holds approximately 3.8 tons of alcohol, and the oxidizer tank holds about 5 tons of liquid oxygen. Thus, even in the case of using alcohol, which requires much less oxygen for combustion than gasoline or kerosene, filling both tanks with fuel alone (alcohol) using atmospheric oxygen would increase the duration of the engine by two to three times. This is where the need to have an oxidizer on board a rocket comes in.
Fig. 28. Rocket engine.
The question involuntarily arises: how does a rocket cover a distance of 300 km if the engine runs for only 1 minute? This is explained in Fig. 33, which shows the trajectory of the rocket, as well as the change in speed along the trajectory.
The launch of the rocket is carried out after placing it in a vertical position using a light launcher, as can be seen in Fig. 26. After launch, the rocket initially rises almost vertically, and after 10–12 seconds of flight, it begins to deviate from the vertical and, under the action of rudders controlled by gyroscopes, moves along a trajectory close to an arc of a circle. Such a flight lasts all the time while the engine is running, that is, for about 60 seconds.
When the speed reaches calculated value, control devices turn off the engine; by this time, there is almost no fuel left in the rocket tanks. The height of the rocket at the end of the engine is 35–37 km, and the axis of the rocket makes an angle of 45° with the horizon (point A in Fig. 29 corresponds to this position of the rocket).
Fig. 29. The flight path of a long-range missile.
This elevation provides maximum range in the subsequent flight when the missile is coasting, similar to artillery shell, which would fly out of the gun, the sawn-off barrel of which is at a height of 35–37 km. The trajectory of further flight is close to a parabola, and total time flight is approximately 5 minutes. The maximum height that the rocket reaches in this case is 95-100 km, stratospheric rockets reach much higher altitudes, more than 150 km. In photographs taken from this height by a device mounted on a rocket, the sphericity of the earth is already clearly visible.
It is interesting to see how the flight speed along the trajectory changes. By the time the engine is turned off, i.e. after 60 seconds of flight, the flight speed reaches its highest value and is approximately 5500 km/h, i.e. 1525 m/s. It is at this moment that the power of the engine also becomes the greatest, reaching for some rockets almost 600,000 l. With.! Further, under the influence of gravity, the speed of the rocket decreases, and after reaching the highest point of the trajectory, for the same reason, it begins to grow again until the rocket enters the dense layers of the atmosphere. During the entire flight, except for the very initial phase - acceleration - the speed of the rocket significantly exceeds the speed of sound, average speed along the entire trajectory is approximately 3500 km/h and even on the ground, the rocket falls at a speed two and a half times the speed of sound and equal to 3000 km/h. This means that the powerful sound from the flight of the rocket is heard only after it has fallen. Here it will no longer be possible to catch the approach of a rocket with the help of sound pickups, usually used in aviation or navy, this will require completely different methods. Such methods are based on the use of radio waves instead of sound. After all, a radio wave propagates at the speed of light - the highest speed possible on earth. This speed of 300,000 km/sec is, of course, more than sufficient to mark the approach of the fastest rocket.
Another problem is related to the high speed of rocket flight. The fact is that at high flight speeds in the atmosphere, due to braking and compression of the air running on the rocket, the temperature of its body rises greatly. The calculation shows that the temperature of the walls of the rocket described above should reach 1000–1100 °C. Tests showed, however, that in reality this temperature is much lower due to the cooling of the walls by thermal conduction and radiation, but nevertheless it reaches 600–700 ° C, i.e., the rocket heats up to a red heat. As the rocket's flight speed increases, the temperature of its walls will rise rapidly and may become a serious obstacle to a further increase in flight speed. Recall that meteorites (heavenly stones) bursting at a tremendous speed, up to 100 km/s, into the earth's atmosphere, as a rule, "burn out", and what we take for a falling meteorite ("shooting star") is in reality only a clot of hot gases and air, formed as a result of the movement of a meteorite at high speed in the atmosphere. Therefore, flights with very high speeds are possible only in the upper layers of the atmosphere, where the air is rarefied, or outside it. The closer to the ground, the lower the permissible flight speeds.
Fig. 30. Scheme of the rocket engine.
The rocket engine diagram is shown in Fig. 30. Noteworthy is the relative simplicity of this scheme compared to conventional piston aircraft engines; especially characteristic of LRE almost complete absence in the power circuit of the engine moving parts. The main elements of the engine are a combustion chamber, a jet nozzle, a steam generator and a turbopump unit for fuel supply and a control system.
Fuel combustion occurs in the combustion chamber, i.e., the conversion of the chemical energy of the fuel into thermal energy, and in the nozzle, the thermal energy of the combustion products is converted into the high-speed energy of the gas jet flowing from the engine into the atmosphere. How the state of gases changes during their flow in the engine is shown in Fig. 31.
The pressure in the combustion chamber is 20–21 ata, and the temperature reaches 2,700 °C. Characteristic of the combustion chamber is a huge amount of heat that is released in it during combustion per unit time or, as they say, the heat density of the chamber. In this regard, the LRE combustion chamber is significantly superior to all other combustion devices known in the art (boiler furnaces, engine cylinders). internal combustion and others). In this case, the amount of heat released per second in the combustion chamber of the engine is enough to boil more than 1.5 tons of ice water! So that the combustion chamber with such huge number the heat released in it has not failed, it is necessary to intensively cool its walls, as well as the walls of the nozzle. For this purpose, as seen in FIG. 30, the combustion chamber and nozzle are cooled by fuel - alcohol, which first washes their walls, and only then, heated, enters the combustion chamber. This cooling system, proposed by Tsiolkovsky, is also beneficial because the heat removed from the walls is not lost and returns to the chamber again (this is why such a cooling system is sometimes called regenerative). However, only external cooling of the engine walls is not enough, and cooling of their inner surface is simultaneously applied to lower the temperature of the walls. For this purpose, the walls in a number of places have small holes located in several annular belts, so that through these holes alcohol enters the chamber and nozzle (about 1/10 of its total consumption). The cold film of this alcohol, flowing and evaporating on the walls, protects them from direct contact with the flame of the torch and thereby reduces the temperature of the walls. Despite the fact that the temperature of the gases washing from the inside of the walls exceeds 2500 °C, the temperature of the inner surface of the walls, as tests have shown, does not exceed 1000 °C.
Fig. 31. Change in the state of gases in the engine.
Fuel is supplied to the combustion chamber through 18 prechamber burners located on its end wall. Oxygen enters the prechambers through the central nozzles, and alcohol leaving the cooling jacket through a ring of small nozzles around each prechamber. In this way, a sufficiently good mixing of the fuel is ensured, which is necessary for the implementation of complete combustion in the very short time while the fuel is in the combustion chamber (hundredths of a second).
The jet nozzle of the engine is made of steel. Its shape, as can be clearly seen in Fig. 30 and 31, is first a narrowing and then expanding pipe (the so-called Laval nozzle). As mentioned earlier, nozzles and powder rocket engines have the same shape. What explains this shape of the nozzle? As you know, the task of the nozzle is to ensure the complete expansion of the gas in order to obtain the highest exhaust velocity. To increase the speed of gas flow through a pipe, its cross section must first gradually decrease, which also occurs with the flow of liquids (for example, water). The velocity of the gas will increase, however, only until it becomes equal to the velocity of sound in the gas. A further increase in velocity, in contrast to a liquid, will only be possible with the expansion of the pipe; this difference between gas flow and liquid flow is due to the fact that the liquid is incompressible, and the volume of the gas increases greatly during expansion. In the throat of the nozzle, i.e., in its narrowest part, the gas flow velocity is always equal to the speed of sound in the gas, in our case, about 1000 m/s. The outflow velocity, i.e., the velocity in the outlet section of the nozzle, is 2100–2200 m/s(thus the specific thrust is approximately 220 kg sec/kg).
The supply of fuel from the tanks to the combustion chamber of the engine is carried out under pressure by means of pumps driven by a turbine and arranged together with it into a single turbopump unit, as can be seen in Fig. 30. In some engines, the fuel supply is carried out under pressure, which is created in sealed fuel tanks using some kind of inert gas - for example, nitrogen, stored under high pressure in special cylinders. Such a supply system is simpler than a pumping one, but, with a sufficiently large engine power, it turns out to be heavier. However, even when pumping fuel in the engine we are describing, the tanks, both oxygen and alcohol, are under some excess pressure from the inside to facilitate the operation of the pumps and protect the tanks from collapse. This pressure (1.2–1.5 ata) is created in the alcohol tank with air or nitrogen, in the oxygen tank - with vapors of evaporating oxygen.
Both pumps are centrifugal type. The turbine that drives the pumps runs on a steam-gas mixture resulting from the decomposition of hydrogen peroxide in a special steam-gas generator. Sodium permanganate, which is a catalyst that accelerates the decomposition of hydrogen peroxide, is fed into this steam and gas generator from a special tank. When a rocket is launched, hydrogen peroxide under nitrogen pressure enters the steam-gas generator, in which a violent reaction of peroxide decomposition begins with the release of water vapor and gaseous oxygen (this is the so-called "cold reaction", sometimes used to create thrust, in particular, in launch rocket engines). Vapor-gas mixture having a temperature of about 400 °C and pressure over 20 ata, enters the turbine wheel and then is released into the atmosphere. The power of the turbine is spent entirely on the drive of both fuel pumps. This power is not so small already - at 4000 rpm of the turbine wheel, it reaches almost 500 l. With.
Since a mixture of oxygen and alcohol is not a self-reactive fuel, some kind of ignition system must be provided to start combustion. In the engine, ignition is carried out using a special fuse, which forms a flame torch. For this purpose, a pyrotechnic fuse (a solid igniter such as gunpowder) was usually used, and a liquid igniter was less commonly used.
Rocket launch is carried out as follows. When the ignition torch is ignited, the main valves are opened, through which alcohol and oxygen enter the combustion chamber by gravity from the tanks. All valves in the engine are controlled by compressed nitrogen stored on the rocket in a battery of high-pressure cylinders. When the combustion of the fuel begins, an observer located at a distance, using an electrical contact, turns on the supply of hydrogen peroxide to the steam and gas generator. The turbine begins to work, which drives the pumps that supply alcohol and oxygen to the combustion chamber. The thrust grows and when it becomes more than the weight of the rocket (12-13 tons), the rocket takes off. From the moment the ignition flame is ignited to the moment the engine develops full thrust, only 7-10 seconds pass.
When starting, it is very important to ensure a strict order of entry into the combustion chamber of both fuel components. This is one of the important tasks of the engine control and regulation system. If one of the components accumulates in the combustion chamber (because the intake of the other is delayed), then an explosion usually follows this, in which the engine often fails. This, along with random interruptions in combustion, is one of the most common causes of accidents during LRE testing.
Noteworthy is the negligible weight of the engine compared to the thrust it develops. When the engine weight is less than 1000 kg thrust is 25 tons, so that the specific gravity of the engine, i.e., the weight per unit of thrust, is only
For comparison, we indicate that a conventional piston aircraft engine running on a propeller has a specific gravity of 1–2 kg/kg, i.e., several tens of times more. It is also important that the specific gravity of a rocket engine does not change with a change in flight speed, while the specific gravity of a piston engine increases rapidly with increasing speed.
LRE for rocket aircraft
Fig. 32. Project LRE with adjustable thrust.
1 - mobile needle; 2 - mechanism for moving the needle; 3 - fuel supply; 4 - oxidant supply.
The main requirement for an aircraft liquid-propellant engine is the ability to change the thrust it develops in accordance with the flight modes of the aircraft, up to stopping and restarting the engine in flight. The simplest and most common way to change the thrust of an engine is to regulate the supply of fuel to the combustion chamber, as a result of which the pressure in the chamber and thrust change. However, this method is unfavorable, since with a decrease in pressure in the combustion chamber, which is lowered in order to reduce thrust, the proportion of thermal energy of the fuel that passes into the high-speed energy of the jet decreases. This results in an increase in fuel consumption by 1 kg thrust, and consequently, by 1 l. With. power, i.e., the engine starts to work less economically. To reduce this shortcoming, aircraft rocket engines often have two to four combustion chambers instead of one, which makes it possible to turn off one or more chambers when operating at reduced power. Thrust control by changing the pressure in the chamber, i.e., by supplying fuel, is retained in this case as well, but is used only in a small range up to half the thrust of the chamber being switched off. The most advantageous way to regulate the thrust of a liquid-propellant rocket engine would be to change the flow section of its nozzle while reducing the fuel supply, since in this case a decrease in the per second amount of escaping gases would be achieved while maintaining the same pressure in the combustion chamber, and, hence, the exhaust velocity. Such regulation of the nozzle flow area could be carried out, for example, using a movable needle of a special profile, as shown in Fig. 32, depicting the design of a liquid-propellant rocket engine with thrust regulated in this way.
In FIG. 33 shows a single-chamber aircraft rocket engine, and Fig. 34 - the same rocket engine, but with an additional small chamber, which is used in cruise flight when little thrust is required; the main camera is turned off completely. Both chambers work at maximum mode, and the large one develops a thrust of 1700 kg, and small - 300 kg, so the total thrust is 2000 kg. The rest of the engines are similar in design.
The engines shown in Fig. 33 and 34 operate on self-igniting fuel. This fuel consists of hydrogen peroxide as the oxidizer and hydrazine hydrate as the fuel, in a weight ratio of 3:1. More precisely, the fuel is a complex composition consisting of hydrazine hydrate, methyl alcohol and copper salts as a catalyst that ensures a fast reaction (other catalysts are also used). The disadvantage of this fuel is that it causes corrosion of engine parts.
The weight of a single chamber engine is 160 kg, the specific gravity is
per kilogram of thrust. Engine length - 2.2 m. The pressure in the combustion chamber is about 20 ata. When operating at the minimum fuel supply to obtain the least thrust, which is 100 kg, the pressure in the combustion chamber decreases to 3 ata. The temperature in the combustion chamber reaches 2500 °C, the gas flow rate is about 2100 m/s. Fuel consumption is 8 kg/s, and the specific fuel consumption is 15.3 kg fuel per 1 kg thrust per hour.
Fig. 33. Single-chamber rocket engine for rocket aircraft
Fig. 34. Two-chamber aircraft rocket engine.
Fig. 35. Scheme of fuel supply in an aviation LRE.
The scheme of fuel supply to the engine is shown in Fig. 35. As in a rocket engine, the supply of fuel and oxidizer stored in separate tanks is carried out at a pressure of about 40 ata impeller driven pumps. A general view of the turbopump unit is shown in Fig. 36. The turbine runs on a steam-gas mixture, which, as before, is obtained as a result of the decomposition of hydrogen peroxide in a steam-gas generator, which in this case is filled with a solid catalyst. Before entering the combustion chamber, the fuel cools the walls of the nozzle and the combustion chamber, circulating in a special cooling jacket. The change in the fuel supply necessary to control the engine thrust during the flight is achieved by changing the supply of hydrogen peroxide to the steam-gas generator, which causes a change in the speed of the turbine. The maximum speed of the impeller is 17,200 rpm. The engine is started using an electric motor that drives the turbopump unit.
Fig. 36. Turbopump unit of an aviation rocket engine.
1 - gear drive from the starting electric motor; 2 - pump for the oxidizer; 3 - turbine; 4 - fuel pump; 5 - turbine exhaust pipe.
In FIG. 37 shows a diagram of the installation of a single-chamber rocket engine in the rear fuselage of one of the experimental rocket aircraft.
The purpose of aircraft with liquid-propellant engines is determined by the properties of liquid-propellant rocket engines - high thrust and, accordingly, high power at high flight speeds and high altitudes and low efficiency, i.e., high fuel consumption. Therefore, rocket engines are usually installed on military aircraft - interceptor fighters. The task of such an aircraft is, upon receiving a signal about the approach of enemy aircraft, to quickly take off and gain a high altitude at which these aircraft usually fly, and then, using their advantage in flight speed, impose on the enemy air battle. The total duration of the flight of an aircraft with a liquid-propellant engine is determined by the fuel supply on the aircraft and is 10-15 minutes, so these aircraft can usually perform combat operations only in the area of their airfield.
Fig. 37. Scheme of the installation of rocket engines on the plane.
Fig. 38. Rocket fighter (view in three projections)
In FIG. 38 shows an interceptor fighter with the LRE described above. The dimensions of this aircraft, like other aircraft of this type, are usually small. The total weight of the aircraft with fuel is 5100 kg; the fuel reserve (over 2.5 tons) is only enough for 4.5 minutes of engine operation at full power. Maximum flight speed - over 950 km/h; the ceiling of the aircraft, i.e. the maximum height that it can reach, is 16,000 m. The rate of climb of an aircraft is characterized by the fact that in 1 minute it can rise from 6 to 12 km.
Fig. 39. The device of a rocket aircraft.
In FIG. 39 shows the device of another aircraft with a rocket engine; this is an experimental aircraft built to achieve flight speeds in excess of the speed of sound (i.e. 1200 km/h at the ground). On the plane, in the rear of the fuselage, an LRE is installed, which has four identical chambers with a total thrust of 2720 kg. Engine length 1400 mm, maximum diameter 480 mm, weight 100 kg. The stock of fuel on the plane, which is used as alcohol and liquid oxygen, is 2360 l.
Fig. 40. Four-chamber aircraft rocket engine.
The external view of this engine is shown in Fig. 40.
Other applications of LRE
Along with the main use of liquid-propellant rocket engines as engines for long-range missiles and rocket aircraft, they are currently used in a number of other cases.
LREs have been widely used as engines for heavy rocket projectiles, similar to the one shown in Fig. 41. The engine of this projectile can serve as an example of the simplest rocket engine. Fuel (gasoline and liquid oxygen) is supplied to the combustion chamber of this engine under the pressure of neutral gas (nitrogen). In FIG. 42 shows a diagram of a heavy rocket used as a powerful anti-aircraft projectile; the diagram shows the overall dimensions of the rocket.
Liquid-propellant rocket engines are also used as starting aircraft engines. In this case, a low-temperature hydrogen peroxide decomposition reaction is sometimes used, which is why such engines are called "cold".
There are cases of using LRE as boosters for aircraft, in particular, aircraft with turbojet engines. In this case, fuel supply pumps are sometimes driven from the turbojet engine shaft.
Liquid-propellant rocket engines are also used, along with powder engines, for launching and accelerating aircraft (or their models) with ramjet engines. As you know, these engines develop very high thrust at high flight speeds, high speeds of sound, but do not develop thrust at all during takeoff.
Finally, one more application of LRE, which has recently taken place, should be mentioned. To study the behavior of an aircraft at high flight speeds approaching and exceeding the speed of sound requires a serious and costly research work. In particular, it is required to determine the resistance of aircraft wings (profiles), which is usually carried out in special wind tunnels. In order to create in such pipes the conditions corresponding to the flight of an aircraft at high speed, it is necessary to have power plants of very high power to drive the fans that create a flow in the pipe. As a result, the construction and operation of tubes for testing at supersonic speeds require huge costs.
Recently, along with the construction of supersonic tubes, the task of studying various wing profiles of high-speed aircraft, as well as testing ramjet engines, by the way, is also being solved with the help of liquid-propellant
Fig. 41. Rocket projectile with rocket engine.
engines. According to one of these methods, the investigated profile is installed on a long-range rocket with a liquid-propellant rocket engine, similar to the one described above, and all readings of instruments that measure the resistance of the profile in flight are transmitted to the ground using radio telemetry devices.
Fig. 42. Scheme of the device of a powerful anti-aircraft projectile with a rocket engine.
7 - combat head; 2 - cylinder with compressed nitrogen; 3 - tank with oxidizer; 4 - fuel tank; 5 - liquid-propellant engine.
According to another method, a special rocket trolley is being built, moving along rails with the help of a liquid-propellant rocket engine. The results of testing a profile installed on such a trolley in a special weight mechanism are recorded by special automatic devices also located on the trolley. Such a rocket cart is shown in Fig. 43. The length of the rail track can reach 2–3 km.
Fig. 43. Rocket trolley for testing aircraft wing profiles.
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Rocket fuelA LITTLE THEORY From the school physics course (the law of conservation of momentum) it is known that if a mass m separates from a body at rest with a mass M with a speed V, then the remaining part of the body with a mass M-m will move with a speed m / (M-m) x V in the opposite direction. This means that the greater the discarded mass and its speed, the greater the speed acquired by the rest of the mass i.e. the greater will be the force that sets it in motion. For the operation of a rocket engine (RD), as well as any jet engine, an energy source (fuel), a working fluid (RT) is needed, which ensures the accumulation of the energy of the source, its transfer and transformation), a device in which the energy is transferred to the RT and a device in which the internal energy RT is converted into the kinetic energy of the gas jet and transferred to the rocket in the form of thrust. Chemical and non-chemical fuels are known: in the first (liquid-propellant rocket engines - LRE and solid propellant rocket engines - solid propellant rocket engines) the energy necessary for engine operation is released as a result of chemical reactions, and the resulting gaseous products serve as a working fluid, in the second for heating the working fluid. bodies use other sources of energy (for example, nuclear energy). The efficiency of the RD, as well as the efficiency of the fuel, is measured by its specific impulse. Specific thrust impulse (specific thrust), defined as the ratio of thrust force to the second mass flow rate of the working fluid. For LRE and solid propellant rocket engines, the consumption of the working fluid coincides with the fuel consumption, and the specific impulse is the reciprocal of the specific fuel consumption. The specific impulse characterizes the efficiency of the RD - the larger it is, the less fuel (in the general case, the working fluid) is spent to create a unit of thrust. In the SI system, the specific impulse is measured in m/sec and practically coincides in magnitude with the velocity of the jet. In the technical system of units (its other name is MKGSS, which means: Meter - KiloGram Force - Second), widely used in the USSR, the kilogram of mass was a derived unit and was defined as the mass of which a force of 1 kgf imparts an acceleration of 1 m / s per sec. It was called the "technical unit of mass" and amounted to 9.81 kg. Such a unit was inconvenient, so weight was used instead of mass, specific gravity instead of density, and so on. In rocket technology, when calculating the specific impulse, not mass but weight fuel consumption was also used. As a result, the specific impulse (in the MKGSS system) was measured in seconds (it is 9.81 times less in magnitude than the specific “mass” impulse). The value of the specific impulse of the RD is inversely proportional to the square root of the molecular weight of the working fluid and directly proportional to the square root of the working fluid temperature in front of the nozzle. The temperature of the working fluid is determined by the calorific value of the fuel. Its maximum value for the beryllium+oxygen pair is 7200 kcap/kg. which limits the value of the maximum specific impulse of the LRE to no more than 500 sec. The value of the specific impulse depends on the thermal efficiency of the RD - the ratio of the kinetic energy imparted in the engine to the working fluid to the entire calorific value of the fuel. The conversion of the calorific value of the fuel into the kinetic energy of the outgoing jet in the engine occurs with losses, since part of the heat is carried away with the outgoing working fluid, part is not released at all due to incomplete combustion of the fuel. Electrojet engines have the highest specific impulse. For a plasma electric propulsion engine, it reaches 29000 sec. The maximum impulse of serial Russian RD-107 engines is 314 seconds. The characteristics of the RD are 90% determined by the fuel used. Rocket fuel - a substance (one or more), which is a source of energy and RT for RD. It must meet the following basic requirements: it must have a high specific impulse, high density, the required aggregate state of the components under operating conditions, it must be stable, safe to handle, non-toxic, compatible with structural materials, have raw materials, etc. The thruster engine runs on chemical fuel. The main energy characteristic (sp. impulse) is determined by the amount of heat released (calorific value of the fuel) and the chemical composition of the reaction products, which determines the completeness of the conversion of thermal energy into kinetic energy of the flow (the lower the molecular weight, the higher the sp. pulse). According to the number of separately stored components, chemical rocket propellants are divided into one- (unitary), two-, three- and multi-component ones, according to the aggregate state of the components - into liquid, solid, hybrid, pseudo-liquid, jelly-like ones. Single-component fuels - compounds such as hydrazine N 2 H 4, hydrogen peroxide H 2 O 2 in the RD chamber decompose with release a large number heat and gaseous products, have low energy properties. For example, 100% hydrogen peroxide has a beat pulse of 145 s. and is used as auxiliary fuel for control and orientation systems, turbopump drives RD. Gel fuels are fuels usually thickened with salts of macromolecular organic acids or special additives (rarely an oxidizing agent). An increase in the specific impulse of rocket fuels is achieved by adding metal powders (Al, etc.). For example, "Saturn-5" burns 36 tons during the flight. aluminum powder. Two-component liquid and solid fuels have received the greatest application. LIQUID FUEL A two-component liquid fuel consists of an oxidizer and a fuel. The following specific requirements are imposed on liquid fuels: the widest possible temperature range of the liquid state, the suitability of at least one of the components for cooling liquid RD (thermal stability, high boiling point and heat capacity), the possibility of obtaining high efficiency, minimum viscosity of the components and its low dependence on temperature. To improve the characteristics, various additives are introduced into the composition of the fuel (metals, for example, Be and Al to increase the specific impulse, corrosion inhibitors, stabilizers, ignition activators, substances that lower the freezing point). Kerosene (naphtha-kerosene and kerosene-gas oil fractions with a boiling range of 150-315°C), liquid hydrogen, liquid methane (CH 4), alcohols (ethyl, furfuryl) are used as fuel; hydrazine (N 2 H 4), and its derivatives (dimethylhydrazine), liquid ammonia (NH 3), aniline, methyl-, dimethyl- and trimethylamines, etc. The following are used as an oxidizing agent: liquid oxygen, concentrated nitric acid (HNO 3), nitrogen tetroxide (N 2 O 4), tetranitromethane; liquid fluorine, chlorine and their compounds with oxygen, etc. When fed into the combustion chamber, the fuel components may spontaneously ignite (concentrated nitric acid with aniline, nitrogen tetroxide with hydrazine, etc.) or not. The use of self-igniting propellants simplifies the design of the RD and makes it possible to carry out reusable launches in the simplest way. Hydrogen-fluorine pairs (412s), hydrogen-oxygen (391s) have the maximum specific impulse. From the point of view of chemistry, the ideal oxidizing agent is liquid oxygen. It was used in the first ballistic missiles of the FAA, its American and Soviet copies. But its boiling point (-183 0 C) did not suit the military. The required operating temperature range is from -55 0 C to +55 0 C. Nitric acid, another obvious oxidizing agent for rocket engines, suited the military more. It has a high density, low cost, is produced in large quantities, is quite stable, including at high temperatures, and is fire and explosion safe. Its main advantage over liquid oxygen is its high boiling point and, consequently, its ability to be stored indefinitely without any thermal insulation. But nitric acid is such an aggressive substance that it continuously reacts with itself - hydrogen atoms are split off from one acid molecule and attached to 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, formed at the bottom of the tank). To reduce corrosivity, various substances began to be added to nitric acid; only 0.5% hydrofluoric (hydrofluoric) acid reduces the corrosion rate of stainless steel tenfold. Nitrogen dioxide (NO 2) is added to the acid to increase the impulse. It is a brown gas with a pungent odor. When cooled below 21 0 C, it liquefies, and nitrogen tetroxide (N 2 O 4), or nitrogen tetroxide (AT), is formed. At atmospheric pressure AT boils at +21 0 С, and freezes at –11 0 С. The gas consists mainly of NO 2 molecules, the liquid consists of a mixture of NO 2 and N 2 O 4, and only tetroxide molecules remain in the solid. Among other things, the addition of AT to the acid binds the water that enters the oxidizer, which reduces the corrosive activity of the acid, increases the density of the solution, reaching a maximum at 14% of the dissolved AT. This concentration was used by the Americans for their combat missiles. Ours to get the maximum beat. pulse used 27% AT solution. Such an oxidizer received the designation AK-27. In parallel with the search for the best oxidizer, there was a search for the optimal fuel. The first widely used fuel was alcohol (ethyl), which was used on the first Soviet rockets R-1, R-2, R-5 ("legacy" of FAU-2). In addition to low energy indicators, the military was obviously not satisfied with the low resistance of personnel to “poisoning” by such fuel. The military was most satisfied with the product of oil distillation, but the problem was that such fuel does not spontaneously ignite when in contact with nitric acid. This disadvantage was bypassed by the use of starting fuel. Its composition was found by German rocket scientists during the Second World War, and it was called "Tonka-250" (in the USSR it was called TG-02). Substances that contain nitrogen in addition to carbon and hydrogen are best ignited with nitric acid. Such a substance with high energy characteristics was hydrazine (N 2 H 4). In terms of physical properties, it is very similar to water (the density is several percent higher, the freezing point is +1.5 0 C, the boiling point is +113 0 C, the viscosity and everything else is like that of water). But the military was not satisfied with the high freezing temperature (higher than that of water). The USSR developed a method for producing unsymmetrical dimethylhydrazine (UDMH), while the Americans used a simpler process for producing monomethylhydrazine. Both of these liquids were extremely poisonous, but less explosive, absorbed less water vapor, and were thermally more stable than hydrazine. But the boiling point and density are lower compared to hydrazine. Despite some shortcomings, the new fuel suited both the designers and the military quite well. UDMH also has another, "unclassified" name - "heptyl". "Aerozine-50" used by the Americans on their liquid rockets is a mixture of hydrazine and UDMH, which was the result of the invention technological process, in which they were received at the same time. After ballistic missiles began to be placed in mines, in a sealed container with a temperature control system, the requirements for the operating temperature range of rocket fuel were reduced. As a result, nitric acid was abandoned, switching to pure AT, which also received an unclassified name - "amyl". The boost pressure in the tanks raised the boiling point to an acceptable value. Corrosion of tanks and pipelines with the use of AT decreased so much that it became possible to keep the rocket refueled throughout the entire period of combat duty. The first missiles to use AT as an oxidizer were the UR-100 and the heavy R-36. They could stay refueled for up to 10 years in a row. The main characteristics of two-component liquid fuels with an optimal ratio of components (pressure in the combustion chamber, 100 kgf/cm2, at the nozzle exit 1 kgf/cm2) Oxidizer Fuel , kcal/kg of combustion, K s Nitrogen Kerosene 1460 1.36 2980 313 k-ta (98%) TG-02 1490 1.32 3000 310 Aniline (80%) + furfuryl 1420 1.39 3050 313 alcohol (20%) Oxygen Alcohol (94%) 2020 0.39 3300 255 (Liquid) Hydrogen l. 0.32 3250 391 Kerosene 2200 1.04 3755 335 UDMH 2200 1.02 3670 344 Hydrazine 1.07 3446 346 Ammonia l. 0.84 3070 323 AT Kerosene 1550 1.27 3516 309 UDMH 1.195 3469 318 Hydrazine 1.23 3287 322 Fluorine Hydrogen l. 0.62 4707 412 (liquid) Hydrazine 2230 1.31 4775 370 * the ratio of the total mass of the oxidizer and fuel to their volume. SOLID FUEL Solid propellant is subdivided into pressed ballistic propellant - nitroglycerin powders), which is a homogeneous mixture of components (not used in modern powerful rocket engines) and mixed propellant, which is a heterogeneous mixture of oxidizer, fuel-binder (facilitating the formation of a monolithic fuel block) and various additives (plasticizer , powders of metals and their hydrides, hardener, etc.). Solid propellant charges are made in the form of channel blocks burning on the outer or inner surface. The main specific requirements for solid fuels are: the uniformity of the distribution of components and, consequently, the constancy of the physicochemical and energy properties in the block, the stability and regularity of combustion in the RD chamber, as well as a set of physical and mechanical properties that ensure the engine's performance in conditions of overloads, variable temperature, vibrations. According to the specific impulse (about 200 s.), solid fuel is inferior to liquid fuel, because due to chemical incompatibility, it is not always possible to use energy-efficient components in solid fuels. The disadvantage of solid fuels is their susceptibility to "aging" (an irreversible change in properties due to chemical and physical processes occurring in polymers). American rocket scientists quickly abandoned liquid fuel and preferred solid mixed fuel for combat missiles, work on the creation of which in the United States had been carried out since the mid-40s, which made it possible already in 1962. to adopt the first solid-propellant ICBM "Minuteman-1". In our country, large-scale research began with a significant delay. Decree of November 20, 1959. It was envisaged to create a three-stage rocket RT-1 with solid rocket motors (RDTT) and a range of 2500 km. Since by that time there were practically no scientific, technological and production bases for mixed charges, there was no alternative to the use of solid ballistic propellants. The maximum allowable diameter of the powder cartridges produced by the method of continuous pressing did not exceed 800 mm. Therefore, the engines of each stage had a package layout of 4 and 2 blocks at the first and second stages, respectively. The loose powder charge burned along the inner cylindrical channel, the ends and the surface of 4 longitudinal slots located in the front part of the charge. Such a shape of the combustion surface provided the required pressure diagram in the engine. The rocket had unsatisfactory characteristics, for example, with a launch weight of 29.5 tons. The Minuteman-1 had a maximum range of 9300 km, while for the RT-1 these characteristics were, respectively, 34 tons. and 2400 km. The main reason for the lag of the RT-1 rocket was the use of ballistic gunpowder. To create a solid-fuel ICBM with characteristics approaching the Minuteman-1, it was necessary to use mixed fuels that provide higher energy and better mass characteristics engines and rockets in general. In April 1961 a Government Decree was issued on the development of ICBMs on solid fuel - RT-2, a kick-off meeting was held and the Nylon-S program was prepared for the development of mixed fuels with an impulse pulse of 235 s. These propellants were supposed to make it possible to manufacture charges weighing up to 40 tons. casting method into the engine housing. At the end of 1968 the rocket was put into service, but required further improvement. Thus, mixed fuel was molded in separate molds, then the charge was put into the body, and the gap between the charge and the body was filled with a binder. This created certain difficulties in the manufacture of the engine. The RT-2P rocket had a PAL-17/7 solid propellant based on butyl rubber, which has high ductility, does not have noticeable aging and cracking during storage, while the fuel was poured directly into the engine case, then it was polymerized and molded required charge burning surfaces. By their own flight performance RT-2P approached the Minuteman-3 missile. Mixed fuels based on potassium perchlorate and polysulfide were the first to be widely used in solid propellant rocket engines. A significant increase in beats. The impulse of the solid propellant rocket engine occurred after ammonium perchlorate was used instead of potassium perchlorate, and instead of polysulfide - polyurstane, and then polybutadiene and other rubbers, and additional fuel was introduced into the fuel - powdered aluminum. Almost all modern solid propellant rocket engines contain charges made from ammonium perchlorate, aluminum and butadiene polymers (CH 2 =CH-CH=CH 2). The finished charge looks like hard rubber or plastic. It is subjected to careful control for the continuity and uniformity of the mass, strong adhesion of the fuel to the hull, etc. Cracks and pores in the charge, as well as delaminations from the body, are unacceptable, as they can lead to an undesigned increase in solid propellant thrust (due to an increase in the burning surface), burnouts of the body and even explosions. The characteristic composition of the mixed fuel used in modern powerful solid propellant rocket engines: oxidizer (usually ammonium perchlorate NH 4 C1O 4) 60-70%, fuel-binder (butyl rubber, nitrile rubbers, polybutadienes) 10-15%, plasticizer 5-10%, metal (powders of Al, Be, Mg and their hydrides) 10-20%, hardener 0.5-2.0% and combustion catalyst 0.1-1.0%. and modified dibasic or blended dibasic fuel. In composition, it is intermediate between the usual ballistic dibasic (dual-base powders - smokeless powders in which two main components: nitrocellulose - most often in the form of pyroxylin, and a non-volatile solvent - most often nitroglycerin) fuel and mixed. The dual-base mixed fuel usually contains crystalline ammonium perchlorate (oxidant) and powdered aluminum (fuel) bound by a nitrocellulose-nitroglycerium mixture. Here is a typical composition of a modified dual-base fuel: ammonium perchlorate - 20.4%, aluminum - 21.1%, nitrocellulose - 21.9%, nitroglycerin - 29.0%, triacetin (solvent) - 5.1%, stabilizers - 2.5%. At the same density as the mixed polybutadiene fuel, the modified two-base fuel is characterized by a slightly higher specific impulse. Its disadvantages are a higher combustion temperature, high cost, increased explosiveness (tendency to detonation). In order to increase the specific impulse, highly explosive crystalline oxidizers, such as hexogen, can be introduced into both mixed and modified dual-base fuels. HYBRID FUEL In a hybrid fuel, the components are in different states of aggregation. Fuels can be: solidified petroleum products, N 2 H 4, polymers and their mixtures with powders - Al, Be, BeH 2, LiH 2, oxidizing agents - HNO 3, N 2 O 4, H 2 O 2, FC1O 3, C1F 3, O 2 , F 2 , OF 2 . In terms of specific impulse, these fuels occupy an intermediate position between liquid and solid ones. Fuels have the maximum specific impulse: BeH 2 -F 2 (395s), VeH 2 -H 2 O 2 (375s), VeH 2 -O 2 (371s). The hybrid fuel developed by Stanford University and NASA is based on paraffin. It is non-toxic and environmentally friendly (during combustion, it forms only carbon dioxide and water), its thrust is regulated over a wide range, and a restart is also possible. The engine has a fairly simple device, an oxidizer (gaseous oxygen) is pumped through a paraffin tube located in the combustion chamber, during ignition and further heating, the surface layer of the fuel evaporates, supporting combustion. The developers have achieved high speed combustion and thus solve the main problem that previously hampered the use of such engines in space rockets. Good prospects may have the use of metallic fuel. One of the most suitable metals for this purpose is lithium. When burning 1 kg. This metal releases 4.5 times more energy than when kerosene is oxidized with liquid oxygen. Only beryllium can boast of greater calorific value. US patents have been published for solid rocket fuel containing 51-68% lithium metal.
- traction is uncontrollable
- after ignition, the engine cannot be turned off or restarted
Drawbacks mean solid rockets are useful for short duration missions (missiles) or boost systems. If you need to control the engine, you will have to turn to the liquid fuel system.
Liquid rockets
In 1926, Robert Goddard tested the first liquid fuel engine. Its engine used gasoline and liquid oxygen. He also tried and solved a number of fundamental problems in rocket engine design, including pumping mechanisms, cooling strategies, and steering gears. It is these problems that make liquid propellant rockets so difficult.
The main idea is simple. In most liquid propellant rocket engines, fuel and an oxidizer (such as gasoline and liquid oxygen) are pumped into the combustion chamber. There they burn to create a stream of hot gases at high speed and pressure. These gases pass through a nozzle that accelerates them even more (from 8,000 to 16,000 km / h, as a rule), and then exits. Below you will find a simple circuit.
This diagram does not show the actual complexities of a conventional engine. For example, normal fuel is a cold liquid gas like liquid hydrogen or liquid oxygen. One of the big problems with such an engine is the cooling of the combustion chamber and nozzle, so the cold liquid first circulates around the overheated parts to cool them. The pumps must generate extremely high pressure to overcome the pressure that the burning fuel creates in the combustion chamber. All this pumping and cooling makes a rocket engine look more like failed attempt plumbing self-realization. Let's look at all kinds of fuel combinations used in liquid rocket motors:
- Liquid hydrogen and liquid oxygen (primary space shuttle engines).
- Gasoline and liquid oxygen (Goddard's first rockets).
- Kerosene and liquid oxygen (used in the first stage of the Saturn V in the Apollo program).
- Alcohol and liquid oxygen (used in German V2 rockets).
- Nitrogen tetroxide/monomethylhydrazine (used in Cassini engines).
The future of rocket engines
We are used to seeing chemical rocket engines that burn propellant to produce thrust. But there are many other ways to get traction. Any system that is capable of pushing mass. If you want to accelerate a baseball to incredible speed, you need a viable rocket engine. The only problem with this approach is the exhaust, which will be dragged through space. It is this small problem that causes rocket engineers to prefer gases over burning products.
Many rocket engines are extremely small. For example, attitude thrusters on satellites don't generate much thrust at all. Sometimes satellites use almost no fuel - pressurized nitrogen gas is ejected from the tank through a nozzle.
New designs must find a way to accelerate ions or atomic particles to high speeds to make thrust more efficient. In the meantime, we will try to do and wait for what else Elon Musk will throw out with his SpaceX.
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