What is fuel in a rocket? Solid rocket fuels. Blended rocket fuels

One of the youngest, fastest growing and most powerful components of condensed energy systems (ECS) is mixed solid rocket fuel (MSRP).

SRTT is a multicomponent heterogeneous coarse highly filled explosive system consisting of an oxidizer, a fuel binder and special additives (energy, technological and operational) and obtained by mechanical mixing of components with subsequent transformation into a monoblock capable of natural combustion.

Table 3 - Recipes and properties of ballistic-based colored lights compositions

The ancestor of SRTT was black powder (DP). The Chinese were the first to use it as solid fuel for rockets. The rocket had a 2.5 m long pole as a stabilizer. Bamboo tubes were used as the shell-hull. The Indians already used an iron body as a shell body. In 1799, the Indians used missiles in combat against the British during the defense of the city of Seringapatam. For the massive use of missiles, a corps of missile shooters numbering up to 5,000 people was created there. The mass of the missiles ranged from 3 to 6 kt.

In Europe, the first rockets also appeared with the invention of gunpowder. The British mastered the technology of making black powder rockets in 1804. The missiles' flight range was 2.5 km. The rockets had an iron body, and in order to increase the combustion area, the charge had a channel. They were adopted for service in 1806 (they were used during the siege of Boulogne and in 1807 during the shelling of Copenhagen). The mass of the rocket ranged from 3 to 17 kg. Following England, missiles are being adopted into service in Austria, France, and Prussia.

Russian rocket technology followed its own independent path, and there is evidence that Russia was far ahead Western Europe. Already at the beginning of the 17th century. methods for making combat missiles were well known. In 1680, the first “rocket establishment” was founded in Moscow, consisting of several laboratories involved in the preparation of special rocket powders and individual rocket parts.

In 1807, a forty-four-millimeter DP signal flare was developed, which was in service for more than 100 years. Widely used powder rockets developed by Russian scientists A.D. Zasyadko and K.I. Konstantinov, were found during the Russian-Turkish War in 1828-1829, in military operations in the Caucasus in 1850 and during the defense of Sevastopol from foreign invaders in 1854-1855. .

DP rockets lost their importance for two reasons:

Due to the unsatisfactory energy characteristics of gunpowder;

Due to the low accuracy of the missiles.

The advent of rifled artillery, which made it possible to significantly increase the accuracy of hits, finally negated the interest in DP.

During the Second World War, due to the fact that ballistic gunpowders were in short supply, and some of their properties did not allow the use of these gunpowders as a source of missile energy, efforts scientific workers Many countries have been aimed at developing mechanically strong CPTTs.

In 1942 at the Artillery Academy named after. F.E. Dzerzhinsky developed casting compositions SRTT based on ammonium nitrate and organic combustible binders such as polyvinyl acetate, and in 1946 A.A. Schmidt was the first to substantiate the possibility of producing solid fuels based on polymerizing substances. He predicted the real paths of this direction and its prospects. The earliest works in this direction include studies
G.V. Kalabukhova. In 1948, he proposed CPTT based on ammonium and potassium perchlorates and a flammable high-polymer binder consisting of colloxylin, polystyrene and rubber. However, in terms of energy characteristics and strength, the developed compositions were inferior to ballistic powders. The charges were made by blind and through pressing.

The first American CPTTs were obtained in the laboratory of the California Institute of Technology.

They included:

potassium perchlorate or ammonium nitrate – 75%;

bitumen - 18%;

petroleum oil - 7%.

Subsequently, in order to increase energy, ammonium perchlorate (AP) and metallic aluminum began to be used as an oxidizer, and rubber-like combustible binders were used to improve the physical and mechanical characteristics of the fuel. Thus, based on thiokol (polysulfide rubber) and PCA, SRTTs were developed for the Sergent operational-tactical missile weighing about 4 tons and with a flight range of up to 150 km. Then, based on polyurethane and PCA, fuel was created for the Pershing operational missile with a flight range of up to 700 km, as well as the Polaris strategic missile weighing about 13 tons and a flight range of up to 4000 km. Subsequently, based on PCA and a copolymer of polybutadiene with acrylic acid, fuel was developed that was used to manufacture charges for the Minuteman intercontinental missile with a flight range of up to 10,000 km.

All these missiles were developed and put into service in the period from 1953-1963. At the end of 1970 the army, Navy and US aviation had 600 Polaris missiles at submarines and 1,000 Minuteman missiles installed in silos at combat positions.

In the USSR, the development and use of SRTT in a broad sense began in 1958. In 1959, at the Artillery Academy
them. F.E. Dzerzhinsky produced and studied polyurethane fuel on a laboratory scale. In the same year it was developed in industrial scale CPTT based on thiokol and PCA. Somewhat later, CPTTs were created based on polyesters and polyesters, acrylonitrile rubbers, butyl rubber and carboxyl rubbers.

Since 1961, the efforts of researchers have been aimed at increasing the specific impulse of the CPTT, increasing the level of physical and mechanical characteristics and stabilizing the combustion process.

S.P. Korolev created the first solid-fuel rocket RT-1 using ballistic powder with a flight range of 2500 km with a launch weight of 34 tons, using insert charges with a diameter of 800 mm. Only by switching to SRTT, he was able to create a second solid-fuel rocket
RT-2 (8K-98), which has a flight range of already 9500 km with a launch weight of 51 tons. Its first launch took place on November 4, 1966, and it was put into service in 1968.

The charge, firmly attached to the body, is made by pouring the fuel mass directly into the combustion chamber. Bonding of the charge to the body is carried out using special protective and fastening (adhesive) layers (Figure 44).

TRT – solid rocket fuel; TZP – heat-protective coating;

ZKS – protective and fastening layer; SOK - nozzle block

Figure 44 - Fastening scheme using protective and fastening layers

The dimensions and design form of the charge are selected from the condition of ensuring the required value of second fuel consumption, time and traction characteristics, loads, temperature conditions operation and application. The required dependence of the current value of the combustion surface on the size of the burnt arch is ensured by the shape of the channel (cylindrical, star-shaped, slotted, cylindrical-conical, etc.), as well as by the introduction of special combustion compensators in the form of grooves for partial or complete opening of the ends, etc.

The perfection of the charge is largely determined by the coefficient of volumetric filling of the combustion chamber, minimizing the ratio of the current value of the combustion surface to the average integral value, manufacturability, resistance to external factors. The mass parameters of charges vary widely: from fractions of a gram to several hundred tons.

The use of SRTT is not limited to the armed forces. At the same time, they began to be widely used for space exploration and in national economy.

Use of SRTT for peaceful purposes. Solid propellant rocket engines (solid propellant rocket motors) are found wide application for peaceful purposes in the national economy as auxiliary engines for solving a wide variety of problems in rocket and space technology.

Solid propellant rocket motors are most often used in the emergency rescue system for cosmonauts and pilots, for braking and accelerating a spacecraft, separating launch vehicle stages, dropping payloads, stabilizing and correcting the trajectory of a spacecraft, correcting the orbit of a spacecraft, landing spacecraft on planets, and launching rockets. -carriers and return spacecraft in the Shuttle systems, as meteorological rocket engines used to lift equipment into the upper layers of the atmosphere, anti-hail and anti-avalanche.

The advantages of solid propellant rocket motors, which ensure their widespread use in rocket and spacecraft, are high reproducibility of parameters, including accuracy of meeting the requirements for the full thrust impulse, high coefficient of mass perfection, long warranty periods application and relative safety during storage and operation.

To separate rocket stages Small-sized solid propellant rocket motors of a wide variety of designs are used, the type of which is determined by the task being performed. A charge from the SRTT, the inserted or rigidly attached version of the equipment, is presented in Figure 45.

1 – igniter; 2 – chamber shell; 3 – SRTT charge;

4 - nozzle block

Figure 45 - Small-sized solid propellant rocket engine

Brake motors are used for braking during the descent of a wide variety of spacecraft. For these purposes, spherical solid propellant rocket engines are mainly used, for example, spherical solid propellant rocket engines TE-M series(USA) from Thiokol Powder were used for braking during the descent of the Gemeni spacecraft, during the landing of the Surveyor spacecraft on the moon, etc. The design of the TK-M-385 type braking engine is shown in Figure 46.

1 – protective casing; 2 – centering mirror block; 3− charge
solid fuel; 4 – thermal insulation coating; 5 – body;
6 – liner; 7 – expanding part of the nozzle; 8 − rubber plug;

9 − ignition device

Figure 46 – Braking solid propellant rocket motor type TK-M-385

The solid fuel charge is made in the form of an eight-rayed star made of polysulfide fuel, consisting of PCA and a binder with the addition of 2% aluminum.

Correction of the spacecraft's orbit is necessary to ensure its inter-orbital transitions and perform various maneuvers in orbit. Engines of this type include apogee and pyrogenous solid propellant rocket engines, which ensure the transition of a spacecraft from one orbit to another.

The apogee solid rocket engine of the FW-5 type used in the United States is shown in Figure 47.

Figure 47 - Solid propellant rocket motor type FW-5

The case is made of titanium alloy. The engine uses a CPTT based on a polyurethane binder, PCA and aluminum.
A composition based on phenolic resin is used as a heat-shielding material in the housing.

Figure 48 shows a similar-purpose solid propellant rocket motor, MAGE-1. Its body is made of the Kevlar-49 composite material, the charge is made of aluminized fuel.

1 – thermal insulation; 2 – charge of solid fuel; 3 – nozzle block;
4 – body; 5 - ignition device

Figure 48 - MAGE-1 type solid propellant rocket motor

Figure 49 shows a promising apogee solid propellant rocket engine (USA) of the STAR-48 series from Thiokol Chemical, which allows increasing the specific thrust impulse in vacuum by 59.0–88.5 kN · s/kg with a volumetric fill factor of up to 0.935.

1 – body; 2 – heat-protective coating; 3 − toroidal
igniter; 4 – nozzle block; 5 − graphite liner

Figure 49 - STAR-48 series apogee solid propellant rocket engine

These engines have the following advantages:

A charge made of CPTT based on polybutadiene rubber has a cylindrical shape with radial slot cuts and fills the entire front part of the body;

The body is made of titanium alloy with thermal protection made of carbon-carbon composite material.

Particular importance in the design of small solid propellant rocket engines is given to the choice of fuel. The most complete requirements are met by SRTTs, in which polyurethanes or hydrocarbon rubbers are used as a combustible binder, and aluminum as highly thermally conductive additives. The thermodynamic characteristics of CPTT can be increased by using both more powerful oxidizing agents and metal hydrides, for example, aluminum.

Some characteristics of the SRTT used in small-sized solid propellant rocket engines in the USA are given in Table 4.

A serious disadvantage of PCA-based CPTTs is their toxicity, because When it burns, a large amount of toxic chlorine and hydrogen chloride is released. For example, during the launch of the Space Shuttle, about 2 tons of chlorine and 210 tons of hydrogen chloride are released into the atmosphere during the operation of solid fuel accelerators, which have a harmful effect on environment. Therefore, in order to facilitate the use of CPTT for peaceful purposes, a lot of work is being done both here and abroad to replace PCA with environmentally friendly oxidizing agents: ammonium salt of dinitroazic acid (ADNA), ammonium nitrate.

Table 4 - Main characteristics of fuels for solid propellant rocket engines

In the USA, a cheap and environmentally friendly CPTT has been developed for the engines of large space accelerators, in which ammonium nitrate, hexogen, octogen and a binder based on polyglycedyl azide plasticized with nitroesters are used as the main components.

FSUE "Soyuz" has created an environmentally friendly SRTT "Center", the unfavorable properties of which, in particular, the phase instability of ammonium nitrate, are eliminated by introducing a modifying additive into the crystals. It uses an active binder with a crystallization temperature of minus 50 °C based on a eutectic mixture with nitroesters. The use of ammonium nitrate and nitrile butadiene rubber reduces the cost of fuel.

However, the use of ammonium nitrate instead of PCA significantly reduces the energy of the CPTT and limits its use in products where the value of a single impulse plays a decisive role. In addition, the use of ammonium nitrate is limited by its increased hygroscopicity.

The developed environmentally friendly fuels are used as charges for meteorological rockets, in gas-dynamic drilling devices, and powder pressure accumulators.

Currently, an increasing number of launch vehicles are used to launch various types satellites are used as solid propellant rocket engines as accelerators. For example, in the Titan-3C rocket (USA), for launch, in addition to the main liquid rocket engines (LPRE), two powerful solid propellant rocket engines with a diameter of 3 m and a length of 25.8 m are used as accelerators, developing thrust in a vacuum of up to 540 10 4 n with an operating time of 110 s. Their use made it possible to increase the mass of the load put into orbit to 11.4 tons. The launch weight of the rocket is 700 tons.

Powerful accelerators operating on SRTT with a charge mass of 100 to 200 tons began to be used to launch the French Ariane-2 and Ariane-4 rockets, which are used to launch artificial Earth satellites into orbit.

The companies IIS (USA) and SEP (France) have developed an improved version of this type of engine, providing a specific thrust impulse of 2970 kN at an average pressure in the chamber of 33.9 10 5 Pa · s/kg.

The charge is firmly attached to the engine housing and has a channel that does not reach the front bottom of the housing. This design makes it possible to increase the volumetric fill factor to 0.92 and provide a fairly small combustion surface area. The charge is made from high-pulse solid fuel based on PCA and polybutadiene rubber (10%), aluminum (20%) and HMX (12%).

Effective direction The conversion of many SRTT production plants is the manufacture on their basis of launch accelerators for powerful launch vehicles and spacecraft that launch various spacecraft into orbit. Accelerators have a very large mass (from 150 to 400 tons), their production ensures the loading of SRTT production plants in Peaceful time. Typically, two such boosters are attached to the sides of the main body of the rocket and provide its lift, and after the fuel is exhausted, they are separated from the rocket using special ejection solid propellant rocket engines and fall to the ground.

A typical starting solid propellant rocket engine has five to six interchangeable sections, assembled one above the other and forming common building engine

A diagram of the launch engine for the Titan-3C launch vehicle, with the help of which various US satellites are launched into orbit, is shown in Figure 50.

It consists of five sections with a diameter of 3.0 m and a length of 3.0 m. The mass of each section is 33.0 tons. The charge is firmly attached to the body of each section and is made of CPTT containing PCA, aluminum and a binder based on polybutadiene, methacrylic acid and acrylonitrile. This SRTT at a chamber pressure of 6.0-6.2 MPa provides a specific thrust impulse of 2480. The engine housing is welded (made of bridge steel). A heat-protective coating made of synthetic rubber with silicon filler is applied to the inner surface.

Solid propellant rocket motor for the Shuttle system. The Shuttle space system (weighing more than 2000 tons), designed to launch into orbit the manned and reentry spacecraft Challenger, Atlantic, Discovery, Columbia and others, is a bundle that includes an orbital vehicle with crew, two solid fuel boosters to accelerate the ship in the initial part of the trajectory and a disposable fuel tank. The latter is the main element of the system, to which the descent vehicle and boosters are attached, returned to Earth and reused.

The main feature of this system, in contrast to the domestic Energia-Buran system, is that it has two accelerators running on solid fuel. In the Energia-Buran system, the starting boosters run on liquid fuel.

The solid fuel accelerator is a sectional solid propellant rocket motor, has a diameter of 3.7 m, a length of 45.5 m and a curb weight of about 590 tons, and after fuel burnout - 78 tons. The mass of solid fuel is approximately 500 tons. Its operating time is more than 2 minutes, the total thrust is over 26 Mn.

CPTT, which contains PCA, powdered aluminum, polybutadiene binder, iron oxide and other additives, is used as a solid fuel. The shape of the charge, rigidly attached to the body, is cylindrical, with an internal channel, ensuring compliance with the necessary law of thrust growth, which creates the most favorable overload regime (no more than 3) for astronauts. After the fuel burns out, the accelerator bodies are separated from the ship at an altitude of 70-90 km, and then, upon reaching the dense layers of the atmosphere, the parachute system is activated, ensuring their splashdown. The booster bodies raised from the water are restored and refilled with fuel.

Solid propellant boosters are also used in France on a powerful launch vehicle to launch space objects, including manned reusable spacecraft Hermes and Ariane 5.

The use of solid propellant boosters, which have a fairly significant mass of fuel, both in the Shuttle system and in rocket launchers of the Titan-3C type, has created good prerequisites for the conversion of plants producing solid rocket fuel in the United States, ensuring their loading in peacetime without changing the technological process and equipment.

Engines of the cosmonaut emergency rescue system. All launch vehicles used to launch space objects into orbit with astronauts on board are equipped with emergency rescue systems for people at the time of launch and active flight.

The basis of this system is a specially designed solid propellant rocket engine that uses ballistic and mixed type solid rocket propellants. For example, in a three-stage Soyuz launch vehicle, the third stage is a block 8 m long and 2.6 m in diameter, to which a spacecraft is docked through an adapter, closed on top by a fairing with a diameter of 3.0 m. At the top of the fairing there is an emergency rescue propulsion system ship, shaped like a large mushroom (Figure 51).

1 – propulsion system; 2 − Soyuz rocket

Figure 51 - Emergency rescue propulsion system
on the Soyuz ship

The purpose of the installation in the event of a failure of a rocket that has not yet consumed a huge mass of fuel is to instantly take the cosmonauts away from the source of the inevitable fire and explosion to a distance from which a parachute descent to a safe place is possible.

The emergency rescue system (ESS) of the Soyuz spacecraft is equipped as follows: an emergency propulsion system consisting of three types of solid propellant rocket engines is mounted in the bow of the rocket.

The main engine is installed directly on the head fairing, which turns on in the event of an accident and quickly removes the upper part of the head fairing with the compartment and the salvage apparatus of the ship from the rocket.

The twelve nozzles of this powerful engine are arranged in a circle at its top and oriented at an angle of 30 degrees from the longitudinal axis. Above them there is a small fairing in the form of a hemisphere, under which four control engines are hidden. They turn on after the main one, ensuring a turn and removal of the rescued part away from the danger zone. Even higher is the separation engine, which, when turned on last, ensures the separation of the head fairing and its removal from the descent vehicle. After this, the main parachute is inserted, and the descent vehicle descends and soft lands in the same way as when returning from a normal flight. Braking during landing is carried out by brake solid propellant rocket engines running on solid fuel.

Heat-resistant fuels for SRTT gas generators. To intensify oil production, the method of torpedoing wells with special charges has become widely used. Powder gases create channels and cracks in the rock, facilitating the flow of oil. But ballistic gunpowders used for these purposes have certain limitations: for example, they can only be used in those wells where the temperature does not exceed 110 °C (i.e. to the depth
3 km). Developed compositions based on PCA and inactive hydrocarbon binders eliminate this drawback. They remain functional after being held at a temperature of 150 °C for 6 hours and can be stored for 10 years at a temperature of 50 °C. The critical temperature for a block diameter of 150–200 mm is 170–200 °C. The hydrochloric acid released during the combustion of this fuel, entering the formation and reacting with the rock, can contribute to the intensive development of cracks. The production of charges from these fuels can be done on existing equipment using the technology of SRTT production plants.

CPTT is a source of aerosols. One of the promising methods for extinguishing fires in premises for storing alcohol, kerosene, acetone, food in stores, wine cellars, and in ship compartments is aerosol, i.e. instantaneous filling of the room with an aerosol medium containing almost no oxygen, as a result of which combustion stops.

This method, patented by Kühn at the end of the 19th century, was subsequently significantly improved and became widespread. “Kühn cans” were filled with a pyrotechnic composition, which had a number of significant disadvantages: for example, caking, insufficient level of physical and mechanical characteristics, etc. To replace it, new types of gunpowders were developed - sources of aerosols, specifically designed for fire extinguishing systems and preventing the explosion of gas-air mixtures. This new class gunpowder was called PAS (powder, aerosol, mixed). A feature of these compositions is their high economic efficiency; consumption of fire extinguishing agent is 20-90 g/m 3 instead of 200-700 g/cm 3 used previously, environmental friendliness, high reliability and constant readiness for use, the presence of advanced technology using the free casting method (the viscosity of the mass is within
(2-8) 10 4, survivability more than 24 hours).

Several compositions have been developed (for example, PAS-8, PAS-11), which include as the main component nitrates K, Na and carbon dioxide K and Na, NaCl, KCl, K 2 Cr 2 O 7, perchlorates K, Na, NH 4 , and as a binder - nitrocellulose, rubbers, polyester, epoxy or resol resins. Their combustion temperature ranges from 910–1495 K, the mass fraction of the solid phase is 13–39%.

In addition to solid propellant rocket engines, solid fuel as a source of gas can be used in other areas of technology: to rotate a turbine, drive pneumatic systems, fill elastic shells, etc. But their widespread use is hampered by the high combustion temperature. The lowest calorie solid fuels produce gas with a temperature of 1400–1500 K, while traditional materials for technology (metal, plastic, rubber) can withstand temperatures of 300–400 K. Therefore, it is necessary to reduce the temperature of the fuel combustion products. According to V.A. Shandakov and V.F. Komarov, the temperature of gases can be reduced if a charge is created in the form of a material with through porosity. The combustion zone is located at the blind end of the combustion chamber (Figure 52).

1 – blind end of the combustion chamber; 2 – CT charge; 3 – filter; 4 - nozzle

Figure 52 - Scheme of combustion of a porous charge of a CT in a combustion chamber

The pressure developing in it pushes out gas through the pores in the charge and propels liquid combustion products through the body of the porous fuel block, heating it to the gasification temperature, i.e. The heated body is the combustion products of the TT. With complete heat exchange, the gas in front of the thermal wave front will have a temperature equal to the initial temperature of the charge. In practice, it is 300–330 K.

Another advantage of such solid fuels is that individual gases, for example, N2, O2, H2 with a purity of 98.0–99.0% can be obtained as combustion gases. The scope of application of such devices is very wide: means of human rescue on land and water, emergency pneumatic systems, flame suppression and fire extinguishing means, lifting devices and displacement devices, and then medical assistance.

High temperatures can also be used in technology, for example, in the oil and gas industry.

An oil well fades over time due to clogging of the pores of the oil reservoir with solid particles, paraffin carbons and resinous substances carried out by the oil. There was a method of influencing the oil-bearing formation with water pressure, but it was expensive. If, in a liquid-filled well in the zone of an oil reservoir, a pressure higher than the pressure of the rocks is created for a short time during the liquefaction of the TT, then it is possible not only to clear the clogged pores by the pressure and temperature of the gases pressed into the reservoir, but also to create new pores. You just need to burn the CT very quickly, taking advantage of the inertia of the liquid column above it.

To increase the well's flow rate, hydro-reactive compounds are used during thermochemical treatment.

Solid fuels can be used as a chemical reactor for the synthesis of various substances. For example, if we take a mixture of aluminum nitrate Al(NO 3) 3 with cobalt, chromium, and iron nitrates as an oxidizing agent, we obtain a mixed oxide Al x O y of blue, green and red color - a light-resistant pigment for paints.

If we take mixed zirconium and yttrium nitrates, we get the basis of heat-resistant ceramics - stabilized zirconium dioxide. Using mixed barium, copper and yttrium nitrates, superconducting ceramics are obtained.

Hydroreacting compounds are used to pressurize pontoons when lifting sunken objects. The main characteristics of hydro-reacting compositions are the amount of heat released during the combustion of charges when interacting with water, the amount of water required for the combustion of one composition and gas productivity.

Powder pressure accumulators. Powder pressure accumulators (PAD) − solid fuel energy devices used to convert the chemical energy of solid fuel into the energy of compressed gas.

A typical PAD design includes a housing consisting of a high-strength shell, a bottom, a nozzle outlet device and support elements for the charge, the solid propellant charge itself, an igniter and launch initiation means.

PAD has a number of significant advantages compared to cold gas compression systems:

Compactness;

Performance;

Smaller mass-dimensional characteristics;

Good performance properties under various atmospheric influences;

High operational reliability.

They have found wide application in various pneumatic displacement systems for civil and special purposes. For example, the release of missile signals from launch silos, pressurization of various containers, rapid opening and closing of covers, hatches, shutters, pressurization of oil wells, emergency braking.

annotation

The educational manual is intended to help specialists of Baiterek JSC in consolidating knowledge in mastering their functional responsibilities.

The work examines rocket fuels, compressed gases,

on missile systems, their properties, a choice of rocket fuel is proposed.

The educational and methodological manual allows you to consolidate knowledge on the components of rocket fuel and compressed gases, which largely determine the technical appearance of the rocket launcher.

Abstract 2

Accepted abbreviations 4

1 Rocket propellants 5

2 Compressed gases and their properties 13

3 Selecting propellant 16

4 Practical part 17

Test questions 21

Literature 22

Accepted abbreviations

G – fuel

Remote control - propulsion system

ITO – testing technological equipment

KA – spacecraft

KGCH - space warhead

KRT – rocket fuel component

O – oxidizing agent

RAS – Russian Academy of Sciences

RB – accelerating block

RD - rocket engine

Solid propellant rocket engine - solid fuel rocket engine

RKK - rocket and space complex

RN – launch vehicle

RT - rocket fuel

SG – compressed gases

TNT – trinitrotoluene

TTZ - tactical and technical task

Rocket fuels

Rocket fuel largely determines the technical appearance, tactical, technical and operational characteristics of the entire RSC, and also forms the operating system and the system for ensuring the safety of personnel.

In the propulsion systems of modern launch vehicles, spacecraft and boosters, the energy of chemical reactions of components 1 of chemical rocket fuel 2 is used as energy sources. Chemical rocket fuel is not only currently, but in the near future it will be the main type of RT.

Rocket fuels consist of two fundamentally different components: an oxidizer (O) and a fuel (G).

Oxidizer - RT component, consisting mainly of oxidizing elements and serving for the oxidation of fuel in the rocket engine.

Fuel - RT component, consisting mainly of flammable elements and entering into a chemical oxidation (combustion) reaction when interacting with the oxidizing agent in the RD.

Chemical rocket fuels are classified according to the following criteria:

a) by state of aggregation: liquid and solid;

b) by the number of components: one-component (unitary); two-room and multi-component;

c) according to ignition ability: non-self-igniting and self-igniting;

d) by boiling point: low-boiling (cryogenic) and high-boiling.

______________________

1 Component[lat. sotropepz - component] rocket fuel (KRT) -separately stored and brought to the taxiway, differing in composition, part of the missiles legs about fuel.

2 Chemical rocket fuel- rocket fuel, which, as a result of thermal reactions of oxidation, decomposition or recombination, forms high-temperature products that create jet thrust when expelled from the taxiway

Liquid RT make it possible to obtain the highest specific impulse, perform thrust control and multiple launches of the rocket engine. Liquid RT can be unitary (one-component), but, most often, two-component.

Solid RT (SRT) according to their physical nature, they are divided into two classes. oallistite (gunpowder) and mixed RT.

Ballistic TRT They are solid solutions of homogeneous substances, the molecules of which contain flammable and oxidizing elements.

They are used for auxiliary rocket engines (stage separation systems, brake propulsion systems descent vehicles, etc.).

Ballistic TRTs are ignited from a low-power energy source - a spark from a squib, igniter, etc. is enough.

Mixed TRTs are mechanical heterogeneous mixtures of oxidizer and fuel. Inorganic substances are used as an oxidizing agent.

compounds, for example, ammonium perchlorate NH 4 CIO 4, as fuel - synthetic polymeric organic compounds, for example, polyurethane -

new rubber. To improve energy characteristics, powdered metal, for example, aluminum, magnesium, etc., is added as fuel.

Mixed TRTs ignite only from a powerful energy source (igniter) and burn stably only in the presence of pressure in the combustion chamber (at least 2-3 MPa).

The presence of solid rocket fuels on board the launch vehicle and spacecraft imposes

increased requirements for the protection of launch vehicles and spacecraft from static electricity and from mechanical shocks of the solid propellant rocket motors themselves.

Unitary RT - single-component rocket fuel or a homogeneous mixture (solution) of several chemically non-interacting components.

Unitary RTs include hydrogen peroxide, hydrazine, etc. The decomposition reaction of unitary RTs occurs in reactors in the presence of a catalyst. Unitary RTs are used only in auxiliary devices, for example, in gas generators for driving TPU turbines and in the control systems of spacecraft orientation and stabilization systems.

Self-igniting fuel - two-component liquid RT, flammable at normal temperatures in the event of contact of the oxidizer and fuel. The ignition delay period is no more than 3 - 8 ms.

Cryogenic RT[Greek krios-cold; genes- giving birth] - liquid RT, at least one of the components of which is cryogenic.

Cryogenic component RT- low-boiling CRT in the form of a liquefied gas with a boiling point at normal pressure in the cryogenic temperature range (below 120 K or -153 ° C). Liquid oxygen and liquid hydrogen are currently used as cryogenic CRTs.

The main physicochemical properties of liquid MCTs are given in Table 2. RT components have a number of properties that require compliance not only with specific measures and safety rules when working with them, but also the creation special conditions operation. These properties include:

toxicity;

fire hazard (fire hazard) and explosion safety;

aggressiveness;

boiling and freezing temperatures.

Toxicity of CRT - the ability of CRT to have harmful effects on humans, animals and plants. An indicator of toxicity can be

maximum permissible concentration (MPC) 1 MPC in the air of the working area

According to the degree of toxicity, substances, including CRT, are divided into four classes

1st class - extremely dangerous maximum permissible concentrations< 0,0001 мг/л (г/м 3);

2nd class - highly hazardous MPC = (0.0001-0.001) mg/l (g/m3);

3rd class - moderately hazardous MPC = (0.0011-0.01) mg/l (g/m3)

4th class - low-hazard MPC > 0.01 mg/l (g/m).

Table 2.

Physicochemical properties of MCT

LIQUID ROCKET FUEL- chemical rocket fuel, all components of which are in a liquid state under operating conditions. Modern liquid-propellant rocket engines are based on the use of two-component rocket fuel, which releases energy as a result of the interaction of the oxidizer and the fuel.

Depending on the type of components involved in the oxidation reaction, such fuel can be self-igniting rocket fuel and non-self-igniting rocket fuel. In the latter case, for chemical ignition of the main fuel, starting fuel. Found application and unitary rocket fuels.

Part liquid rocket fuel to improve performance and rocket fuel efficiency various additives are introduced, fine powders of some metals are added (see. metal-containing fuel). For this purpose, we are also studying multicomponent rocket fuels(incl. ternary rocket fuel), capable of providing a higher specific impulse value.

Basic requirements for liquid rocket fuels: ensuring a given specific impulse; good chemical stability; explosion safety under operating conditions; suitability and sufficiency of one of the components for cooling the rocket engine (see. fuel cooling capacity); preservation of the liquid state under operating conditions without unjustified costs; compatibility with construction materials; possibly high density; minimal viscosity and toxicity; provision of raw materials.

The most widely used in rocket technology are: from oxidizers - liquid oxygen, nitrogen tetroxide, nitric acid rocket oxidizers, hydrogen peroxide; from flammable - kerosene, monomethylhydrazine, unsymmetrical dimethylhydrazine, hydrazine, liquid hydrogen, amines etc. (specific impulse 2500-4500 m/s). How promising fuel components are studied fluorine oxidizers, borohydrides, as well as combinations of liquid fuel components with light metals (lithium, beryllium, aluminum) etc. (specific impulse 3500-5000 m/s).

Solid rocket fuels are used in rocket engines, propulsion engines, ramjet and ramjet engines, and hydropropellants. They can be divided into two groups: ballistic (homogeneous), for example, N and NM-2 (Table 1.8) and mixed (heterogeneous).

Mixed solid fuels contain 20...30% of a binder rubbery or resinous substance, 60...80% oxidizer and up to 20% aluminum; There are also compositions containing components of ballistic and mixed fuels. It is also possible to use hydrides of light and heavy metals as fuel. Ammonium perchlorate is usually used as an oxidizing agent; it is possible to use other solid salts of perchloric and nitric acids rich in oxygen (Table 1.9).

Rubbers (polysulfide, polyurethane, etc.), polymers (polyester, phenolic and epoxy resins, polyisobutylene, etc.), and heavy petroleum products (asphalt, bitumen, etc., Table 1.10) are used as fuel binders. HMX and RDX are sometimes also added to mixed solid fuels. Some compositions (with a certain degree of convention) of mixed solid fuels in the USA and their characteristics are given in Table. 1.11.

Conventional ballistic and mixed fuels do not meet the requirements for gas generator fuels. Therefore, special gas-generating fuel compositions are being developed with a low combustion temperature (see the last column of Table 1.11), limited from above (by the heat resistance of the materials of valves, turbine blades and other elements of the flow part) and from below (by the stability of fuel combustion). In addition, GGs must sometimes operate for a long time, and the fuel must have a low burning rate. For controlled gas generators, a fuel composition has been proposed in which the combustion rate decreases with increasing pressure (<0). Дополнительные требования могут предъявляться и к составу продуктов сгорания топлив для ГГ: отсутствие конденсированной фазы, коэффициент избытка окислителя должен быть не более единицы (обычно). Смесевые топлива применяют и в воспламенительных ГГ (двигателях запуска).

Mixed solid fuels include pyrotechnic compositions. Pyrotechnic compositions are used as fillers for ignition devices and pyroenergy sensors; it is possible to use them in GG.

The main components included in pyrotechnic compositions can be divided into the following groups (Table 1.12):

1.Oxidizing agents – potassium perchlorate KCIO, sodium nitrates NaNO, potassium KNO, barium Ba ( NO, peroxide and barium chromate BaO, etc.

2.Combustible – metals (aluminum, magnesium, zirconium, boron, titanium) and alloys (aluminum-magnesium, zirconium-nickel), non-metals (phosphorus, carbon and sulfur), inorganic compounds (sulfides, phosphides, silicides, etc.), organic compounds .

Table 1.9

Characteristics of solid oxidizing agents

Table 1.10

3.Cementants (binders) are organic polymers that provide mechanical strength to pyrotechnic compositions (iditol, rosin, epoxy resins, rubbers, ethylcellulose).

4.Other additives that act as combustion accelerators or retardants or reduce the sensitivity of compositions to friction (phlegmatizers).

To ignite mixed solid fuels with a high NH content, pyrotechnic mixtures are used: KCIO - 26...50%, Ba ( NO - 15...17%, zirconium-nickel alloy (50/50) - 32...54%, ethylcellulose - 3% (US patent).

Pyrotechnic compositions in the form of pressed tablets are used in ignition devices. Density is largely determined by pressing pressure and fluctuations within 1.3...2.8 g/cm. specific heat capacity – 0.8…1.25 J/(kg*K), thermal conductivity – 62.8…104.7 W/(m*K).

Table 1.12

Calorific value of pyrotechnic compositions

at a stoichiometric ratio of components

The burning rate of pyrotechnic compositions under the conditions of their operation in an ignition device when blowing tablets with high-temperature combustion products is presented in the form u=map , where m ,a ,v– empirical coefficients.

Pyrotechnic solid fuels are also called compositions with a large amount of metal fuel (more than 50%) and salts of inorganic acids as an oxidizer; they are intended for ramjet engines (RPD).

The charge of mixed TRT can be made in the form of a block (blocks), tablets or powders.

Aluminum, double aluminum decaborane, boron and zirconium diboride, polyethylene, etc. were used as experimental powdered fuels, and ammonium perchlorate, ammonium nitrate, etc. were used as an oxidizing agent. The particles had a size from 2 to 2000 microns. Inert (nitrogen), oxidizing (air, oxygen) and flammable (hydrogen, methane) gases were used as fluidizing gases.

The following methods of supplying pseudo-liquid from the tank to the combustion chamber are possible: using compressed gas, a piston, a screw pump and a jet pump. Powdered fuels are used in combined bench GGs, which allow the pressure, temperature and composition of combustion products to be varied within wide limits in order to study the effect of multiphase flows on materials.

Powdered fuel is black gunpowder (DRP) with a grain diameter of 0.15...1.25 mm and coarse-grained black powder (KZDP) with a grain diameter of 5.1...10.2 mm; composition in%: potassium nitrate – 74; charcoal – 15.6; sulfur – 10.4; combustion temperature 2600K; flow complex 1200 m/s.

DRP grain density 1.75 g/cm, bulk density of DRP 0.9...1.15 g/cm, minimum stable combustion pressure 0.1 MPa, temperature sensitivity = 0.005 1/ C.

The dependence of the burning rate on pressure has the form

u =1.37*(p /98100) .

Ignition of solid rocket fuel occurs when exposed to:

1.thermal energy flow (radiation, contact and convective heating);

2.flow of chemically highly active gases or liquids causing a heterogeneous exothermic reaction upon contact with the surface of solid fuel;

3.mechanical shock and friction.

The actual ignition process in a real solid propellant rocket engine is complex. The main difficulties in its study include the problems of determining the control mechanism, choosing an ignition criterion, determining the chemical kinetics of reactions preceding combustion, as well as the heterogeneous nature of mixed solid fuels. When conducting experiments, the beginning of ignition is taken to be:

1.the first appearance of a flame, recorded photographically or by a photocell;

2.sudden change in thermocouple readings;

3.the onset of loss of fuel mass.

Table 1.13

Mechanical characteristics of TRT

The performance properties of solid fuels are determined by their physical, mechanical (Table 1.13), thermophysical (Table 1.14), chemical characteristics, as well as the physical and chemical characteristics of combustion products. Along with energy, strength, and thermophysical indicators, solid rocket fuel is characterized by explosion safety, sensitivity to impact and friction, degree of toxicity and smokiness of combustion products, manufacturability and equipment, stability of physical and chemical characteristics throughout the entire volume of the charge (especially at the boundaries) under all operating conditions .

Table 1.14

Thermophysical characteristics of TRT

1.3 MAIN CONSTRUCTION ELEMENTS

Rocket launch mass m having n stages, is related to the maximum flight range L approximate ratio m,

Where m- payload mass; / m;I- average value of void specific impulse; A And A– coefficients, the values ​​of which in a first approximation are A=407,A=1/3 at 300 km 6000 km; A=825, a=1/4 at 6000 km 12000 km.

Moreover, in the range L 500 usuallyn =1, in the range of 500 km 5000 kmn =2, in the range of 5000 km 12000 kmn =3.

Optimal relative fuel reserve)).

Taking into account the speed losses due to overcoming gravitational forces and passing through dense layers of the atmosphere in a first approximation leads to the relations ( n=2; 3):

; =(1,08…1,12) ;

The operating time of the stage is related to the initial, specified thrust-to-weight ratio n t= (provided m const).

For each stage according to known and m the main design parameters are found, which for multi-stage rockets are considered to be the stage diameter, fuel mass, engine pressure, degree of nozzle expansion, length of the supersonic part, length of the recessed part, operating time (Table 1.15).

Table 1.15

Parameters of multistage rocket stages

*D– engine diameter.

The engines account for 80...90% of the mass of the entire solid propellant rocket, and the design features of the solid propellant rocket motor largely determine the design of the rocket and its main technical characteristics. In turn, the design features of the solid propellant rocket engine are mainly determined (Table 1.16):

the shape and basic design of the body;

the form of the solid fuel charge, the method of its fastening in the housing;

number and arrangement of nozzles;

type and layout of devices for creating control forces;

draft cut-off device.

1 .3.1 ROCKET CASE AND NOZZLE

The body and nozzle is a hollow multi-block (see Fig. 1.1) or prefabricated single-section (multi-section) cylindrical shell, closed at the ends with front and rear bottoms. The housings can have other shapes, for example, spherical or elliptical. The bottoms are made monolithically with a cylindrical part and separately. The internal structure of the housing is determined by the design of the solid fuel charge.

Table 1.16

Characteristics of various solid propellant rocket motor schemes

Power shells"cocoon" type are made of a composite material by spiral winding on a mandrel with the bottoms made together with the cylindrical part of the shell.

The thickness of the hull shell at the junction of the bottom and the cylindrical part is determined by the formula

Where R- maximum pressure in the engine; D – internal diameter of the cylindrical part of the shell; d- diameter of the pole hole; - tensile strength of glass tape.

An equal-strength cylindrical shell is obtained at = 2…3( d, where is the thickness of the annular layers; - thickness of spiral layers.

Rear bottom thickness

where is the winding angle.

The connecting skirts (see Fig. 1.1) are made by winding integral with the body, and flange fittings are wound into them. Docking skirts are part of the rocket structure and must withstand combined loads: axial (compression and bending), shear and torsion.

The cylindrical part of the power shell can be manufactured by longitudinal-transverse winding on a mandrel.

The wall thickness of the housing shell is determined by the formula

D/(2), where [ is the tensile strength of fiberglass (0.1...1.1 GPa); n– safety factor (1.35...1.5). This formula is valid when one layer of longitudinal tapes is applied to two layers of circumferential tapes.

The load-bearing shells are made without units with thickening at both ends, followed by their mechanical processing to prepare the joints with metal bottoms.

Metal body shells

They are divided according to their shape into cylindrical, conical and spherical, and according to manufacturing technology - into welded (with circular, spiral and longitudinal seams) and seamless (rolled and seamless).

Combined shell shells are metal shells reinforced with an outer braid made of glass fibers or other high-strength reinforcing materials, which are made with a certain tension that creates stress in the braid layer before loading the shell. If the braid takes on half the circumferential load acting on the entire cylindrical shell, then the ratio of the thicknesses of the metal shell and braid is optimal. In this case, the thickness of the metal shell is determined from the condition of ensuring strength in the axial direction D/4, and the lack of strength in the circumferential direction is compensated by a braid with a thickness equal to D/4. In these formulas and are the permissible stresses in the metal shell and reinforcing braid, respectively.

Connections of structural elements are provided using special units, the main requirements for which are to ensure the strength and tightness of the connections with minimal weight and overall dimensions in relation to each specific case, taking into account the materials of the connected elements and types of loading.

With the same type of detachable connection, a huge number of modifications of the ring seals at the joint are possible. The main element of the seal is the rubber ring. The dimensions of rubber rings and grooves for them, as well as recommendations for the use of rubber sealing rings are given in the relevant national and industry standards (GOST 9833-73).

IN nozzle block Solid propellant rocket engines may contain a different number of nozzles: one (coaxial with the engine or rotated relative to the engine axis at an angle of 90), two (rotary) or four, as well as 10...20, inclined to the plane of the nozzle cover, for example, in turbojet shells (see Fig. 1.2).

The nozzle can be round or annular (the latter have not yet found application in solid propellant rocket engines).

The solid propellant rocket engine design with one central nozzle is characterized by the best energy and mass characteristics. To reduce the length of the engine, the nozzle can be sealed into the housing (see Fig. 1.1). In rocket engines in which the solid propellant rocket engine is located near the center of the rocket, the entrance to the nozzle is made in the form of an elongated pipe. The overall dimensions of the variable geometry nozzle in the working position exceed the original ones, such as the sliding nozzle (Fig. 1.3).

Rice. 1.3 Rotary sliding nozzle:

1 – drive termination; 2 – drive; 3 – sliding parts.

The multi-nozzle design makes it possible to organize control of the rocket in two planes and in roll. However, in this case, the conditions for the entry of combustion products into the nozzle worsen, and the carryover of heat-protective coatings at the entrance to the nozzle and in the socket increases.

Also considered are the design diagrams of a solid propellant rocket engine with an annular nozzle, the movable central body of which allows you to adjust the thrust, and with a disc nozzle (metal-free fuel), the outer section of the expanding part of which is formed by the rear bottom of the engine (the same nozzle with a muffled minimum cross-section also serves as the front bottom of the lower stage) .

For features of solid propellant rocket motor thrust cut-off nozzles, see paragraph 1.3.5.

Materials thermal protection Solid propellant rocket motors are artificial isotropic and anisotropic compositions that provide thermal insulation of the supporting structure and predictable carryover of the surface layer.

With some degree of convention, thermal protection materials can be divided into linings, thermal insulation layers and nozzles (Fig. 1.4). The linings provide the specified resistance of the first layer of thermal protection of the path from destruction when interacting with a two-phase working fluid; In this case, material can be carried away at a predictable rate.

Thermal insulation layers have low heat conductivity, but are subject to significant entrainment even at a low level of convection of the working fluid.

Rice. 1.4 Thermal protection:

CCCM – carbon-carbon composite materials; USP – carbon and fiberglass; TZM – heat-protective materials; NO – non-oriented materials; O – oriented materials.

The attachments at the end parts of the nozzles simultaneously perform the functions of both thermal protection and supporting structure. Depending on the level of influence of the flow around the same material can serve as both a cladding and an insulator. For example, the charge geometry of a modern solid propellant rocket engine with a central recessed nozzle eliminates the occurrence of high flow velocities around the body elements; thermal protection materials are mainly subject to heating by radiation. Then the thermal protection of the housing is made of lightweight, elastic, low-thermal conductivity materials based on rubber and rubber without reinforcement with fillers. And for the four-nozzle design of the solid propellant rocket engine, the thermal protection of the nozzle cover, exposed to the influence of a high-speed multiphase jet from the charge channel, is a material made of materials reinforced with asbestos or silica fabric on phenol-formaldehyde binders, which have sufficient erosion resistance and a high density value (up to 1800 kg/m ).

In multilayer structures, thermal insulation layers are placed between the erosion-resistant layer and the protected element in order to minimize the total mass of this unit (Fig. 1.5). Depending on the level of stress-strain state and temperature of the elements, the insulator can be a heat-protective material based on rubber, as well as low-thermal conductivity carbon and fiberglass. The materials of the sealing and diffusion layers of the engine housing are at the same time insulators when the structure warms up.

Rice. 1.5 Elements of the nozzle path:

1 – carbon fiber used as cladding; 2 – fiberglass used as an insulator; 3 – heat insulator made of TMZ.

Non-metallic cladding materials are isotropic and anisotropic compositions consisting of a binder (matrix) and filler. Carbon and fiberglass plastics have an organic binder and fillers made of carbon or silica tissue. Details of the thermal protection of the nozzle path are obtained by pressing and winding. By pressing it is possible to obtain layered (anisotropic) composites.

Large-sized elements of the duct (nozzle bells) are obtained by marking filler tapes impregnated with a binder onto mandrels, followed by hardening under pressure and mechanical processing.

Graphites are produced by pressing a mixture of coal tar pitch (binder) with oil sand (filler), followed by graphitization at T>2400K.

Pyrographites are produced by the deposition of carbon during the decomposition of methane on the surface of graphite in the temperature range of 2373...2673 K, and pyrographite in its properties approaches the properties of a single crystal; it is characterized by sharp anisotropy and extreme values ​​of thermal conductivity and other characteristics.

Carbon-carbon composite materials (CCCM) have fillers made of carbon and graphite fabrics and fibers (including three-dimensional weaving) and a matrix of pyrolytic carbon. A number of parts are obtained by impregnating a carbon-graphite filler with a binder of organic resins during carbonization of the workpiece, and in an inert environment at a temperature of 1273...1373 K and compacting the carbonized workpiece with pyrolytic carbon - deposition of films of organic substances at a temperature of 1373...1473 K.

Other parts are produced by basting or laying carbon-graphite strips or fibers not impregnated with a binder onto a mandrel, followed by compaction with pyrolytic carbon in an oven.

Nozzles - the end parts of nozzles with radiation cooling - are made of alloys based on molybdenum or niobium, which have a high melting point and sufficient strong properties at the equilibrium temperature of the nozzle, and they can also be made of CCCM.

The condition for operability can be taken to be the condition of non-destruction of structural elements, and this extremely complex problem is divided into two simpler ones and, in some cases, independent of each other:

determination of temperature fields in power elements;

determination of stresses and strains in elements under force loading and comparison with permissible values ​​at known temperature fields.

For the liner and elements of solid propellant rocket motor thrust vector control devices exposed to the influence of the working fluid, the restrictions are the conditions of the permissible drift value. In some cases, a limitation is imposed on the permissible spread in the thickness of the entrained layer of materials.

1.3.2 CHARGE OF SOLID FUEL ROCKET

In rocketry, various forms of solid fuel charges are used (Fig. 1.6, Table 1.17): burning mainly on internal surfaces (surfaces whose combustion must be prevented are covered with an armoring compound or a protective-fastening layer to secure the charge to the body); burning on almost all side surfaces, for example, unarmored tubular bombs (Fig. 1.7); burning from the end.

Solid propellant charges are manufactured using injection molding technology, free vacuum casting and the through-pressing method.

A charge made by casting is formed either directly in the solid propellant rocket motor body, or in a special frame, or separately in a special mold. The geometry of the internal surface of the charge is formed by a technological needle placed inside the housing.

Technological process Manufacturing a charge includes preparing a mixture of powdery components, preparing a binder (vacuuming, mixing liquid elements, preparing a mixture of binder with aluminum), preparing a fuel mass and forming a charge, and polymerizing a charge.

When producing charges by injection molding, continuous mixers are used. The fuel mass prepared in the mixer is transported using screws into the mold or into the engine housing. The pressure of the fuel mass at the beginning of filling, equal to 0.5...1.0 MPa, increases when bleeding at the end of filling to 2...4 MPa.

Rice. 1.6 Forms of solid fuel charges

A– multi-checker; b– telescopic; V– with a star-shaped channel; G– with a wheel-shaped channel; d– end combustion; e– cylindrical; and– slotted.

With free casting, the preparation of liquid components and the displacement of the fuel mass are carried out in separate mixers, then the mass is poured into a mold or housing with the preliminary creation of a vacuum in it.

The polymerization process is carried out under a pressure of 3...8 MPa, depending on the design of the charge and engine at a temperature of 40...80 C for 15...25 days. After polymerization, the technological needle, which determines the internal configuration of the charge, is removed. Molding technology allows you to create a charge design from several fuels (different combustion rates, combustion temperatures, etc.).

The charges are made by through-compression using a screw that forces the fuel mass through the mold, which forms the outer and inner cross-sectional shapes of the charge, after which the charge hardens.

A charge formed by pouring directly into the housing and bonded to the inner surface of the housing is called a bonded solid propellant charge (see Fig. 1.1).

The bonded charge is pre-made and then glued into the engine housing. The manufacture of the glued-in charge is carried out in a thick-walled mold with an internal diameter slightly smaller than that of the body.

Rice. 1.7 Cross-sectional shapes of all-combustion charges

A– single-channel checkers; b– multi-channel; V– ductless.

Table 1.17

Characteristics of charges various forms

Table 1.18

Charging parameters with a star-shaped channel

A charge made separately and loosely inserted into the engine body is called an insert charge (Fig. 1.8). Before the advent of mixed fuels, the only way to load them was to loosely pack the charges into the engine housing. Part of the charge surface is armored.

The basic requirements for the armor coating are as follows:

chemical and physical compatibility with TPT and stability under operating conditions;

good adhesion to the charge surface;

high erosion resistance;

low thermal conductivity;

low level of smoke formation (in the case of ballistic fuel).

In a multi-checker charge (see Fig. 1.6, a) the number of checkers providing the highest loading density is equal to n= 1 + 3(i+i), Where i 0.714

The design of the charge of the last stages of ballistic missiles must ensure the possibility of stopping the engine at any time during the flight within a given range. It is necessary that by the time the speed corresponding to the minimum range is reached, the openings of the thrust cut-off system communicate with the free volume of the solid propellant rocket engine combustion chamber. For this purpose, special channels may be provided in the charge.

Depending on the operational requirements for the solid propellant rocket engine, the form of the charge and mechanical properties solid fuel, the method of securing the charge in the solid propellant housing is selected.

The advantage of a bonded charge is that there is no thermal protection coating on most of the internal surface, and this helps to increase the filling density. The walls of the housing are partially loaded from internal pressure with a charge on initial stage operation of the solid propellant rocket motor. The engine does not have special charge mounting motors.

When the charge is placed freely in the housing, a device is inserted for securing the charge in the form of diaphragms (Fig. 1.9), radial supports and ring seals located in the gap between the thermally insulated wall of the engine housing and the armored surface of the charge (see Fig. 1.8). The charge fastening system must provide strong and reliable fixation when the charge is exposed to longitudinal and transverse overloads and vibrations. The fastening design should not cause high local stresses in the charge, which can violate its integrity, cause local destruction of the charge, leading to a distortion of the pressure diagram and a decrease in the completeness of fuel combustion.

Rice. 1.8 Free charge and its attachment points in the housing:

A– front unit; B- rear node.

The diaphragms are designed to reliably fix the solid fuel charge in the housing and at the same time serve as a grate, ensuring better combustion of the charge and complete combustion of its particles in the combustion chamber without ejecting them from the engine.

A radial support for a solid fuel charge may consist of a number of thin-walled support elements or strips that are located circumferentially between the charge and the housing wall; the supporting elements elastically rest against the wall of the housing and the charge, supporting the latter along its entire length. The radial support can also be made in the form of flat elastic strips, which are inserted into the gap with prestress.

Rice. 1.9 Diaphragms:

A– for attaching multi-shot charges; b– for attaching a single-checker charge.

A LITTLE THEORY From a school physics course (the law of conservation of momentum) it is known that if a mass m is separated from a body at rest with a mass M with a speed V, then the remaining part of the body mass M-m will move with speed m/(M-m) x V in the opposite direction. This means that the greater the ejected mass and its speed, the greater the speed the remaining part of the mass will acquire, i.e. the greater will be the force driving it. To operate a rocket engine (RE), like any jet engine, you need an energy source (fuel), a working fluid (RM) that ensures the accumulation of source energy, its transfer and transformation), a device in which energy is transferred to the RT, and a device in which internal energy The RT is converted into the kinetic energy of the gas stream and transferred to the rocket in the form of thrust. Chemical and non-chemical fuels are known: in the former (liquid rocket engines - liquid rocket engines - liquid rocket engines 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 gaseous products formed in this case serve as the working fluid; in the latter, for heating the working fluid bodies use other energy sources (for example nuclear energy). The efficiency of the thruster, like 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 liquid propellant engines and solid propellant rocket engines, the flow rate of the working fluid coincides with the fuel flow rate and the specific impulse is the reciprocal value of the specific fuel flow rate. The specific impulse characterizes the efficiency of the thruster - the larger it is, the less fuel (in the general case, the working fluid) is consumed to create a unit of thrust. In the SI system, the specific impulse is measured in m/sec and practically coincides in value with the speed of the jet stream. IN technical system 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 second. 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 was used instead of density, etc. In rocket technology, when calculating specific impulse, they also used not mass but weight fuel consumption. As a result, the specific impulse (in the MKGSS system) was measured in seconds (in magnitude it is 9.81 times less than the specific “mass” impulse). The magnitude of the specific impulse of the thruster is inversely proportional to the square root of the molecular mass of the working fluid and directly proportional to the square root of the temperature of the working fluid 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 maximum specific impulse of the rocket engine to no more than 500 sec. The magnitude of the specific impulse depends on the thermal efficiency of the thruster - the ratio of the kinetic energy imparted to the working fluid in the engine to the total calorific value of the fuel. The conversion of the calorific value of the fuel into the kinetic energy of the outflowing jet in the engine occurs with losses, since part of the heat is carried away with the outflowing working fluid, and some is not released at all due to incomplete combustion of the fuel. Electric reaction engines have the highest specific impulse. For a plasma electric propulsion engine it reaches 29,000 sec. The maximum impulse of serial Russian RD-107 engines is 314 sec. The characteristics of the RD are 90% determined by the fuel used. Rocket fuel is a substance (one or more) that represents a source of energy and RT for rocket engines. It must satisfy the following basic requirements: have a high specific impulse, high density, the required state of aggregation of components under operating conditions, must be stable, safe to handle, non-toxic, compatible with structural materials, have raw materials, etc. Most existing The current RD runs on chemical fuel. The main energy characteristic (specific impulse) is determined by the amount of heat released (calorific value of the fuel) and the chemical composition of the reaction products, on which the completeness of the conversion of thermal energy into kinetic energy of the flow depends (the lower the molecular weight, the higher the specific impulse). According to the number of separately stored components, chemical rocket fuels are divided into one-(unitary), two-, three- and multi-component; according to the aggregate state of the components - into liquid, solid, hybrid, pseudo-liquid, jelly-like. Single-component fuels - compounds such as hydrazine N 2 H 4, hydrogen peroxide H 2 O 2 in the RD chamber disintegrate and release large quantity heat and gaseous products, have low energy properties. For example, 100% hydrogen peroxide has a shock pulse of 145s. and is used as auxiliary fuel for control and orientation systems, drives of RD turbopumps. Gel-like fuels are usually fuel thickened with salts of high-molecular organic acids or special additives (less often an oxidizer). 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 its flight. aluminum powder. Two-component liquid and solid fuels are most widely used. LIQUID FUEL Two-component liquid fuel consists of an oxidizer and a fuel. The following specific requirements are imposed on liquid fuels: a possibly wider temperature range of the liquid state, the suitability of at least one of the components for cooling a liquid propellant (thermal stability, high boiling points and heat capacity), the possibility of obtaining generator gas from the main components high performance, minimal viscosity of the components and its low dependence on temperature. To improve characteristics, various additives are introduced into the fuel composition (metals, for example Be and Al to increase the shock pulse, corrosion inhibitors, stabilizers, ignition activators, substances that lower the freezing point). The fuel used is 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); hydrazine (N 2 H 4), and its derivatives (dimethylhydrazine), liquid ammonia (NH 3), aniline, methyl-, dimethyl- and trimethylamines, etc. The following oxidants are used: liquid oxygen, concentrated nitric acid (HNO 3), nitrogen tetroxide (N 2 O 4), tetranitromethane; liquid fluorine, chlorine and their compounds with oxygen, etc. When supplied to the combustion chamber, fuel components can spontaneously ignite (concentrated nitric acid with aniline, nitrogen tetroxide with hydrazine, etc.) or not. The use of self-igniting fuels simplifies the design of the RD and makes it easier to carry out reusable launches. The maximum specific impulse is for hydrogen-fluorine (412s), hydrogen-oxygen (391s) pairs. From a chemical point of view, the ideal oxidizing agent is liquid oxygen. It was used in the first FAA ballistic missiles and 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 liquid propellant engines, was more suitable for the military. It has high density, low cost, and is produced in large quantities , quite stable, including at high temperatures, fire and explosion-proof. Its main advantage over liquid oxygen is its high boiling point, and therefore the ability to be stored indefinitely without any thermal insulation. But nitric acid is such an aggressive substance that it continuously reacts with itself - hydrogen atoms are split off from one acid molecule and join neighboring ones, forming fragile but extremely chemically active aggregates. Even the most resistant grades of stainless steel are slowly destroyed by concentrated nitric acid (as a result, a thick greenish “jelly”, a mixture of metal salts, is formed at the bottom of the tank). To reduce the corrosive activity, various substances began to be added to nitric acid; just 0.5% hydrofluoric acid reduces the corrosion rate of stainless steel tenfold. To increase the shock pulse, nitrogen dioxide (NO 2) is added to the acid. 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 a temperature of +21 0 C, and freezes at –11 0 C. The gas consists mainly of NO 2 molecules, the liquid is 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 water entering the oxidizer, which reduces the corrosive activity of the acid, increases the density of the solution, reaching a maximum at 14% dissolved AT. The Americans used this concentration for their military missiles. Ours to obtain the maximum beat. impulse, a 27% AT solution was used. This oxidizer was designated 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), used on the first Soviet missiles R-1, R-2, R-5 (“legacy” of the FAU-2). In addition to the low energy performance, the military was obviously not satisfied with the low resistance of personnel to “poisoning” with such fuels. The military was most satisfied with the product of petroleum distillation, but the problem was that such fuel does not spontaneously ignite upon contact with nitric acid. This drawback was circumvented by using starting fuel. Its composition was discovered by German rocket scientists during World War II, and it was called “Tonka-250” (in the USSR it was called TG-02). Substances that contain, in addition to carbon and hydrogen, nitrogen in their composition are best ignited with nitric acid. Such a substance with high energetic characteristics was hydrazine (N 2 H 4). By physical properties it is very similar to water (density is several percent higher, freezing point +1.5 0 C, boiling point +113 0 C, viscosity and everything else - like 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), and the Americans used a simpler process for producing monomethylhydrazine. Both of these liquids were extremely toxic but less explosive, absorbed less water vapor, and were thermally more stable than hydrazine. But the boiling point and density decreased compared to hydrazine. Despite some shortcomings, the new fuel suited both designers and military personnel 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 a consequence of the invention of a technological process in which they were produced simultaneously. After ballistic missiles began to be placed in silos, in a sealed container with a temperature control system, the requirements for the operating temperature range of rocket fuel were reduced. As a result from nitric acid refused, switching to pure AT, which also received an unclassified name - “amil”. The boost pressure in the tanks increased the boiling point to an acceptable value. The corrosion of tanks and pipelines with the use of AT has decreased so much that it has become possible to keep the missile fueled throughout the entire period of combat duty. The first missiles using AT as an oxidizer were the UR-100 and the heavy R-36. They could stand fueled for up to 10 years in a row. Main characteristics of two-component liquid fuels at an optimal ratio of components (pressure in the combustion chamber, 100 kgf/cm2, at the nozzle exit 1 kgf/cm2) Oxidizer Fuel Calorific value Density Temperature Specific impulse of fuel*, g/cm2 * in the chamber in the void , kcal/kg combustion, K sec Nitrogen Kerosene 1460 1.36 2980 313 quantity (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 * ratio of the total mass of the oxidizer and fuel to their volume. SOLID FUEL Solid fuel is divided into ballistic pressed - nitroglycerin powder) which is a homogeneous mixture of components (not used in modern powerful rocket engines) and mixed fuel which is a heterogeneous mixture of an oxidizer, a fuel-binder (promoting 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 that burn on the outer or inner surface. The main specific requirements for solid fuels: uniform distribution of components and, consequently, constancy of physicochemical and energy properties in the block, stability and regularity of combustion in the propellant chamber, as well as a set of physical and mechanical properties that ensure engine performance in conditions of overload, variable temperature, vibration. In terms of 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 fuel is its 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 had been carried out in the United States since the mid-40s, which made it possible already in 1962. adopt the first solid-fuel ICBM, Minuteman-1. In our country, large-scale research began with a significant delay. By resolution of November 20, 1959 it was envisaged to create a three-stage RT-1 rocket with solid propellant rocket engines (solid propellant rocket motors) and a range of 2500 km. Since by that time there was practically no scientific, technological and production base for mixed charges, there was no alternative to the use of ballistic solid fuels. The maximum permissible diameter of powder bombs produced by the through-pressing method did not exceed 800 mm. Therefore, the engines of each stage had a package arrangement of 4 and 2 blocks for the first and second stages, respectively. The inserted powder charge burned along the internal cylindrical channel, the ends and the surface of 4 longitudinal slits located in the front part of the charge. This shape of the combustion surface provided the necessary pressure diagram in the engine. The rocket had unsatisfactory characteristics, such as with a launch weight of 29.5 tons. "Minuteman-1" had a maximum range of 9300 km, and for RT-1 these characteristics were, respectively, 34 tons. and 2400 km. The main reason for the lag of the RT-1 missile was the use of ballista powder. To create a solid-fuel ICBM with characteristics approaching those of Minuteman-1, it was necessary to use mixed fuels that would provide higher energy and better mass characteristics of the engines and the rocket as a whole. In April 1961 A government decree was issued on the development of solid fuel ICBMs - RT-2, an orientation meeting was held and the Nylon-S program was prepared for the development of mixed fuels with a specific impulse of 235s. These fuels were supposed to provide the ability to manufacture charges weighing up to 40 tons. by casting into the engine housing. At the end of 1968 the missile was accepted for service, but required further improvement. Thus, the mixed fuel was molded in separate molds, then the charge was inserted 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 PAL-17/7 solid fuel based on butyl rubber, which has high plasticity and does not have noticeable aging and cracking during storage, while the fuel was poured directly into the engine housing, then it was polymerized and molded required charge burning surfaces. In terms of its flight performance characteristics, the RT-2P was close to the Minuteman-3 missile. Mixed fuels based on potassium perchlorate and polysulfide were the first to find widespread use in solid propellant rocket engines. Significant increase in beat. The pulse of the solid propellant rocket engine occurred after ammonium perchlorate began to be used instead of potassium perchlorate, and instead of polysulfide rubbers, polyurstane, and then polybutadiene and other rubbers, and additional fuel was introduced into the fuel composition - powdered aluminum. Almost all modern solid propellant rocket engines contain charges made from ammonium perchlorate, aluminum and butadiene polymers (CH 2 =CH-CH=CH2). The finished charge looks like hard rubber or plastic. It is subjected to careful control for the continuity and homogeneity of the mass, strong adhesion of the fuel to the body, etc. Cracks and pores in the charge, as well as detachments from the body, are unacceptable as they can lead to an unreasonable increase in the thrust of the solid propellant rocket engine (due to an increase in the burning surface), burnout 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 (Al, Be, Mg powders and their hydrides) 10-20%, hardener 0.5-2.0% and combustion catalyst 0.1-1.0%. (iron oxide) In modern space solid propellant rocket engines Modified dibasic or mixed dibasic fuel is also used relatively rarely. In composition, it is intermediate between conventional ballistic dibasic (dual-base powders are smokeless powders in which there are two main components: nitrocellulose - most often in the form of pyroxylin, and a non-volatile solvent - most often nitroglycerin) fuel and mixed. Dibasic mixed fuel usually contains crystalline ammonium perchlorate (oxidizer) and powdered aluminum (fuel), bound using 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 polybutadiene blend fuel, the modified dual-base fuel has a slightly higher specific impulse. Its disadvantages are a higher combustion temperature, higher cost, and increased explosion hazard (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 dibasic fuels. HYBRID FUEL In hybrid fuel, the components are in different states of aggregation. Fuel 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 . By specific impulse These fuels occupy an intermediate position between liquid and solid. The following fuels have the maximum specific impulse: BeH 2 -F 2 (395 s), BeH 2 -H 2 O 2 (375 s), BeH 2 -O 2 (371 s). The hybrid fuel developed by Stanford University and NASA is based on paraffin. It is non-toxic and environmentally friendly (when burned, it produces only carbon dioxide and water), its thrust is adjustable within a wide range, and restarting is possible. The engine has a fairly simple design: an oxidizer (oxygen gas) is pumped through a paraffin tube located in the combustion chamber; upon ignition and further heating, the surface layer of the fuel evaporates, supporting combustion. The developers managed to achieve high speed combustion and thus solve the main problem that previously hampered the use of such engines in space rockets. The use of metal fuel may have good prospects. 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. Patents have been published in the United States for solid rocket fuel containing 51-68% lithium metal.

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