Among the technological achievements of mankind, rocket engines occupy a special place. Devices created by the mind of man and his hands are not only the pinnacle of scientific and technological progress. Thanks to these most complex machines, humanity has managed to escape from the embrace of our planet and enter the expanses of space.
It is today at the disposal of man the most powerful rocket engines in the world, capable of developing thrust of hundreds of tons of forces. The rocket race began thousands of years ago, when craftsmen in ancient China managed to create the first powder charges for fireworks. It will take a huge period of time before the first jet engine in the truest sense of the word is created.
Throwing aside gunpowder and getting jet thrust on liquid fuel, man moved on to the construction of jet aircraft and got the opportunity to create more powerful models of rocket technology.
The first steps of man into the world of rocket technology
Mankind has long been familiar with jet propulsion. Even the ancient Greeks tried to use mechanical devices driven by compressed air. Later, devices and mechanisms began to appear that fly due to the combustion of a powder charge. Created in China, and then appeared in Western Europe, the first primitive rockets were far from perfect. However, already in those distant years, the theory of a rocket engine began to take on its first outlines. Inventors and scientists tried to find an explanation for the processes that occurred during the combustion of gunpowder, ensuring the rapid flight of the physical, material body. Jet propulsion was more and more interested in man, opening up new horizons in the development of technology.
The story of the invention of gunpowder gave a new impetus to the development of rocket technology. The first ideas about what the thrust of a jet engine was formed in the process of lengthy experiments and experiments. Work and research were carried out using black powder. It turned out that the process of burning gunpowder causes a large amount of gases that have a huge working potential. Firearms gave scientists the idea to use the energy of powder gases with greater efficiency.
It was not possible to use other fuel to create jet propulsion due to the imperfection of the technical base. It was the powder rocket engine that became the first solid-propellant device, the prototype of modern rocket engines in the service of man.
Until the beginning of the 20th century, rocket technology was in its primitive state, based on the most primitive ideas about jet propulsion. It was only at the end of the 19th century that the first attempts were made to explain from a scientific point of view the processes that contribute to the emergence of jet propulsion. It turned out that with an increase in charge, the traction force increased, which was the main factor in a running engine. This ratio explained how the rocket engine worked and in what direction it was necessary to go in order to achieve greater efficiency of the launched device.
The leadership in this field belongs to Russian scientists. Nikolai Tikhomirov already in 1894 tried to mathematically explain the theory of jet propulsion and create a mathematical model of a rocket (jet) engine. An outstanding scientist of the 20th century, Konstantin Tsiolkovsky, made a huge contribution to the development of rocket technology. The result of his work was the foundations of the theory of rocket engines, which were subsequently used by any designer of rocket engines. All subsequent developments, the creation of rocket technology, went with the use of the theoretical part created by Russian scientists.
Tsiolkovsky, absorbed in the theory of space flight, first voiced the idea of using liquid components, hydrogen and oxygen, instead of solid fuels. With his filing, a liquid jet engine appeared, which today is the most efficient and efficient type of engine. All subsequent developments of the main models of rocket engines that were used to launch rockets, for the most part, worked on liquid fuel, where oxygen could be an oxidizer, and other chemical elements were used.
Types of rocket engines: design, scheme and device
Looking at the schematics of the rocket engine and the industrial finished products, it's hard to call this the pinnacle of technical genius. Even such a perfect device as the Russian Rd-180 rocket engine, at first glance, looks quite prosaic. However, the main thing in this device is the technology used and the parameters that this miracle of technology possesses. The essence of a rocket engine is a conventional jet engine, in which, due to the combustion of fuel, a working fluid is created that provides the necessary traction force. The only difference is in the type of fuel and in the conditions under which the fuel is burned and the working fluid is formed. In order for the engine to develop maximum thrust in the first seconds of its operation, a lot of fuel is required.
In jet engines, the combustion of fuel components is carried out with the participation of atmospheric air. A ramjet engine is today the main workhorse, where aviation kerosene in the combustion chamber burns together with oxygen, forming a powerful jet gas stream at the outlet. A rocket engine is a completely autonomous system where jet thrust is generated by the combustion of solid or liquid fuel without the participation of atmospheric oxygen. For example, a liquid rocket engine runs on fuel, where the oxidizer is one of the chemical elements supplied to the combustion chamber. Solid rockets run on solid fuels that are in the same tank. When they are burned, a huge amount of energy is released, which, under high pressure, comes out of the combustion chamber.
Before starting work, the mass of fuel is 90% of the mass of the rocket engine. As fuel is consumed, its initial weight decreases. Accordingly, the thrust of the rocket engine increases, which ensures the performance of useful work on the transfer of cargo.
Combustion processes occurring inside the combustion chamber of a rocket engine without the participation of air make the use of rocket engines ideal devices for flights to high altitudes and into outer space. Among all rocket engines with which modern rocket technology works, the following types should be distinguished:
- solid rocket engines (TRD);
- liquid (LRE);
- chemical rocket engines (CRD);
- ion rocket engine;
- electric rocket engine;
- hybrid rocket engine (GRD).
A separate type includes a detonation rocket engine (impulse), which is mainly installed on spacecraft traveling in outer space.
Depending on the operation and technical capabilities, the devices are divided into starting rocket engines and steering ones. The first type includes the most powerful rocket engines, which have enormous thrust and are able to overcome the force of gravity. The most famous representatives of this type are the Soviet engine, the liquid-propellant RD-170/171, which develops thrust during a rocket launch of 700 tf. The pressure created in the combustion chamber has a colossal value of 250 kgf / cm2. This type of engine was created for the Energia launch vehicle. A mixture of kerosene and oxygen is used as fuel for the operation of the plant.
Soviet technology turned out to be more powerful than the famous American F-1 device, which ensures the flight of rockets of the American Apollo lunar program.
Starting rocket engines or marching engines can be used as a propulsion system for the first and second stages. It is they that provide a given speed and stable flight of a rocket along a given trajectory and can be represented by all types of rocket engines that exist today. The last type - steering engines - is used to maneuver rocket technology both during a march in the atmospheric layers and during adjustment of spacecraft in space.
To date, only a few states have the technical capabilities to manufacture high-power sustainer rocket engines capable of launching large volumes of cargo into space. Such devices are produced in Russia, the USA, Ukraine and the countries of the European Union. The Russian rocket engine RD -180, Ukrainian engines ZhRD 120 and ZhRD 170 are today the main propulsion systems for rocket technology used for the development of space programs. Today, Russian rocket engines are used to equip American Saturn and Antares launch vehicles.
The most common engines with which modern technology works today are solid propellant and liquid rocket engines. The first type is the easiest to use. The second type - liquid rocket engines are powerful and complex closed-cycle devices in which chemical elements are the main fuel components. These two types of propulsion systems include chemical rocket engines, which differ only in the state of aggregation of the fuel components. However, the operation of this type of equipment takes place in extreme conditions, in compliance with high security measures. The main fuel for this type of engine is hydrogen and carbon, which interact with oxygen, which acts as an oxidizer.
For chemical jet engines, kerosene, alcohol and other flammable substances are used as fuel components. Fluorine, chlorine or oxygen serve as an oxidizing agent for such a mixture. The fuel mass for the operation of chemical engines is very toxic and dangerous to humans.
Unlike their solid fuel counterparts, whose duty cycle is too fast and uncontrollable, liquid fuel engines allow you to regulate their work. The oxidizer is located in a separate container and is fed into the combustion chamber in a limited amount, where, together with other components, a working fluid is formed, leaving through the nozzle, creating thrust. This feature of the propulsion systems allows not only to regulate the thrust of the engine, but also, accordingly, to monitor the speed of the rocket. The best rocket engine currently used to launch space rockets is the Russian RD-180. This device is high performance and economical, making it cost effective to operate.
Both types of engines have their advantages and disadvantages, which are offset by the scope of their use and the technical challenges facing the creators of rocket technology. The latest in a cohort of chemical engines is the SpaceX Raptor cryogenic methane rocket engine, which is being built for a rocket capable of interplanetary flight.
Modern types of rocket engines
The main operating characteristic of rocket engines is the specific impulse. This value is determined by the ratio of the generated thrust to the amount of fuel consumed per unit of time. It is this parameter that today determines the effectiveness of rocket technology, its economic feasibility. Modern technologies are aimed at achieving high values of this parameter in order to obtain a high specific impulse. It may be necessary to use other types of fuel to achieve fast and infinite movement of the spacecraft.
Chemical rocket engines, both solid and liquid, have reached the peak of their development. Despite the fact that these types of engines are the main ones for ballistic and space rockets, their subsequent improvement is problematic. Today, work is underway to use other energy sources.
There are two priority areas:
- nuclear rocket engines (ion);
- electric rocket engines (pulse).
Both types seem to be a priority in the field of spacecraft construction. Despite the shortcomings that the first prototypes of these propulsion systems have today, launching them into space will be much cheaper and more efficient.
Unlike chemical propulsion, on which humanity entered the space age, nuclear propulsion does not provide the necessary momentum by burning liquid or solid fuel. Hydrogen or ammonia heated to a gaseous state act as a working fluid. High-pressure gases heated by contact with nuclear fuel leave the combustion chamber. The specific impulse of these types of engines is quite high. Such installations are also called nuclear and isotope. Their power is estimated quite highly. The work of the NRE from the start on Earth is considered impossible due to the high risk of radioactive contamination of the area and the personnel of the launch complex. Such engines can only be used during a cruising flight in space.
It is believed that the potential of nuclear rocket engines is quite high, but the lack of effective methods for controlling a thermonuclear reaction makes their use under current conditions rather problematic and dangerous.
The next type, electric propulsion engines, are experimental from start to finish. Four types of this propulsion system are considered at once: electromagnetic, electrostatic, electrothermal and pulsed. Of the greatest interest from this group is electrostatic devices, which are also called ionic or colloidal. In this installation, the working fluid (as a rule, it is an inert gas) is heated by an electric field to a plasma state. Ion rocket engines among all others have the highest specific impulse, but it is too early to talk about the practical implementation of the project.
Despite the high momentum, this development has significant drawbacks. The engine requires constant sources of electricity to operate, capable of providing an uninterrupted supply of electricity in large volumes. Accordingly, such an engine cannot have a large thrust, which reduces the efforts of designers to create efficient and economical spacecraft to poor results.
The rocket engine, which humanity has today, has ensured the exit of man into space, made it possible to conduct space exploration at great distances. However, the technical limits that the devices used have reached create the preconditions for intensifying work in other directions. Perhaps in the foreseeable future, ships with nuclear power plants will plow space, or we will plunge into the world of plasma rocket engines flying at speeds close to the speed of light.
A liquid propellant rocket engine is an engine that is fueled by liquefied gases and chemical liquids. Depending on the number of components, liquid-propellant rocket engines are divided into one-, two- and three-component ones.
Brief history of development
For the first time, the use of liquefied hydrogen and oxygen as fuel for rockets was proposed by K.E. Tsiolkovsky in 1903. The first prototype of the rocket engine was created by the American Robert Howard in 1926. Subsequently, similar developments were carried out in the USSR, USA, Germany. The greatest successes were achieved by German scientists: Thiel, Walter, von Braun. During World War II, they created a whole line of rocket engines for military purposes. There is an opinion that if they had created the V-2 Reich earlier, they would have won the war. Subsequently, the Cold War and the arms race became the catalyst for accelerating the development of liquid propellant rocket engines with a view to applying them to the space program. With the help of RD-108, the first artificial Earth satellites were put into orbit.
Today, LRE is used in space programs and heavy rocket weapons.
Scope of application
As mentioned above, LRE is used mainly as an engine for spacecraft and launch vehicles. The main advantages of LRE are:
- the highest specific impulse in the class;
- the ability to perform a full stop and restart paired with traction control gives increased maneuverability;
- significantly less weight of the fuel compartment in comparison with solid fuel engines.
Among the disadvantages of LRE:
- more complex device and high cost;
- increased requirements for safe transportation;
- in a state of weightlessness, it is necessary to use additional engines to deposit fuel.
However, the main disadvantage of liquid-propellant rocket engines is the limit of the energy capabilities of the fuel, which limits space exploration with their help to the distance of Venus and Mars.
Device and principle of operation
The principle of operation of the LRE is the same, but it is achieved using different device schemes. Fuel and oxidizer are pumped from different tanks to the nozzle head, injected into the combustion chamber and mixed. After ignition under pressure, the internal energy of the fuel is converted into kinetic energy and flows out through the nozzle, creating jet thrust.
The fuel system consists of fuel tanks, pipelines and pumps with a turbine for pumping fuel from the tank into the pipeline and a control valve.
Pumping fuel supply creates a high pressure in the chamber and, as a result, a greater expansion of the working fluid, due to which the maximum value of the specific impulse is achieved.
Injector head - a block of injectors for injecting fuel components into the combustion chamber. The main requirement for the nozzle is high-quality mixing and the speed of fuel supply to the combustion chamber.
Cooling system
Although the proportion of heat transfer from the structure during the combustion process is insignificant, the problem of cooling is relevant due to the high combustion temperature (>3000 K) and threatens with thermal destruction of the engine. There are several types of chamber wall cooling:
Regenerative cooling is based on creating a cavity in the chamber walls, through which fuel passes without an oxidizer, cooling the chamber wall, and the heat, together with the coolant (fuel), returns to the chamber.
The near-wall layer is a layer of gas created from combustible vapors near the walls of the chamber. This effect is achieved by installing injectors on the periphery of the head that supply only fuel. Thus, the combustible mixture lacks an oxidizing agent, and combustion near the wall is not as intense as in the center of the chamber. The temperature of the near-wall layer isolates the high temperatures in the center of the chamber from the walls of the combustion chamber.
The ablative method of cooling a liquid-propellant rocket engine is carried out by applying a special heat-shielding coating to the walls of the chamber and nozzles. The coating at high temperatures changes from a solid to a gaseous state, absorbing a large proportion of heat. This method of cooling a liquid rocket engine was used in the Apollo lunar program.
The launch of a rocket engine is a very responsible operation in terms of explosiveness in case of failures in its implementation. There are self-igniting components with which there are no difficulties, however, when using an external initiator for ignition, ideal coordination of its supply with the fuel components is necessary. The accumulation of unburned fuel in the chamber has a destructive explosive force and promises dire consequences.
The launch of large liquid rocket engines takes place in several stages, followed by reaching maximum power, while small engines are launched with an instant exit to one hundred percent power.
The automatic control system of liquid-propellant rocket engines is characterized by the implementation of a safe engine start and exit to the main mode, control of stable operation, thrust adjustment according to the flight plan, adjustment of consumables, shutdown when reaching a given trajectory. Due to the moments that cannot be calculated, the liquid-propellant rocket engine is equipped with a guaranteed supply of fuel so that the rocket can enter the desired orbit in case of deviations in the program.
The propellant components and their choice during the design process are decisive in the design of a liquid rocket engine. On this basis, the conditions of storage, transportation and production technology are determined. The most important indicator of the combination of components is the specific impulse, on which the distribution of the percentage of the mass of fuel and cargo depends. The dimensions and mass of the rocket are calculated using the Tsiolkovsky formula. In addition to specific impulse, density affects the size of tanks with fuel components, boiling point can limit the operating conditions of missiles, chemical aggressiveness is characteristic of all oxidizers and, if the rules for operating tanks are not followed, can cause a tank fire, the toxicity of some fuel compounds can cause serious harm to the atmosphere and the environment . Therefore, although fluorine is a better oxidizing agent than oxygen, it is not used due to its toxicity.
Single-component liquid-propellant rocket engines use liquid as fuel, which, interacting with the catalyst, decomposes with the release of hot gas. The main advantage of single-component rocket engines is their simplicity of design, and although the specific impulse of such engines is small, they are ideally suited as low-thrust engines for orientation and stabilization of spacecraft. These engines use a displacement fuel supply system and, due to the low process temperature, do not need a cooling system. Single-component engines also include gas-jet engines, which are used in conditions where thermal and chemical emissions are unacceptable.
In the early 1970s, the United States and the USSR were developing three-component liquid-propellant rocket engines that would use hydrogen and hydrocarbon fuels as fuel. This way the engine would run on kerosene and oxygen at startup and switch to liquid hydrogen and oxygen at high altitude. An example of a three-component rocket engine in Russia is the RD-701.
Rocket control was first used in V-2 rockets using graphite gas-dynamic rudders, but this reduced engine thrust, and modern rockets use rotary chambers attached to the body with hinges that create maneuverability in one or two planes. In addition to rotary chambers, control motors are also used, which are fixed with nozzles in the opposite direction and are turned on if it is necessary to control the apparatus in space.
A closed-cycle liquid-propellant rocket engine is an engine, one of the components of which is gasified by combustion at a low temperature with a small part of the other component, the resulting gas acts as a working fluid of the turbine, and then is fed into the combustion chamber, where it burns with the remains of fuel components and creates jet thrust. The main disadvantage of this scheme is the complexity of the design, but the specific impulse increases.
The prospect of increasing the power of liquid rocket engines
In the Russian school of LRE creators, headed by Academician Glushko for a long time, they strive for the maximum use of fuel energy and, as a result, the maximum possible specific impulse. Since the maximum specific impulse can be obtained only by increasing the expansion of the combustion products in the nozzle, all developments are carried out in search of the ideal fuel mixture.
1) Study of the scheme and principle of operation of a liquid-propellant rocket engine (LRE).
2) Determination of the change in the parameters of the working fluid along the path of the LRE chamber.
- GENERAL INFORMATION ABOUT LRE
2.1. The composition of the rocket engine
A jet engine is a technical device that creates thrust as a result of the outflow of a working fluid from it. Jet engines provide acceleration of moving vehicles of various types.
A rocket engine is a jet engine that uses only the substances and energy sources that are stored on board a moving vehicle.
A liquid-propellant rocket engine (LRE) is a rocket engine that uses fuel (primary energy source and working fluid) that is in a liquid state of aggregation for operation.
LRE generally consists of:
2- turbopump units (TPU);
3- gas generators;
4 pipelines;
5- automation units;
6- auxiliary devices
One or more liquid-propellant rocket engines, together with a pneumohydraulic system (PGS) for supplying fuel to the engine chambers and auxiliary units of the rocket stage, constitute a liquid-propellant rocket propulsion system (LPRM).
As a liquid propellant (LFR), a substance or several substances (oxidizer, fuel) are used, which are capable of forming high-temperature combustion (decomposition) products as a result of exothermic chemical reactions. These products are the working body of the engine.
Each LRE chamber consists of a combustion chamber and a nozzle. In the LRE chamber, the primary chemical energy of liquid fuel is converted into the final kinetic energy of the gaseous working fluid, as a result of which the reactive force of the chamber is created.
A separate LRE turbopump unit consists of pumps and a turbine that drives them. TNA provides the supply of liquid fuel components to the chambers and gas generators of the LRE.
The LRE gas generator is a unit in which the main or auxiliary fuel is converted into gas generation products used as the working fluid of the turbine and the working fluids of the pressurization system for tanks with LRE components.
The LRE automation system is a set of devices (valves, regulators, sensors, etc.) of various types: electrical, mechanical, hydraulic, pneumatic, pyrotechnic, etc. Automation units provide starting, control, regulation and shutdown of the LRE.
LRE parameters
The main traction parameters of the LRE are:
The reactive force of the rocket engine - R - the resultant gas and hydrodynamic forces acting on the internal surfaces of the rocket engine during the outflow of matter from it;
LRE thrust - R - the resultant of the reactive force of the LRE (R) and all environmental pressure forces that act on the outer surfaces of the engine, with the exception of the forces of external aerodynamic resistance;
LRE thrust impulse - I - integral of the LRE thrust over the time of its operation;
The specific thrust impulse of the LRE - I y - the ratio of thrust (P) to the mass fuel consumption () of the LRE.
The main parameters that characterize the processes occurring in the LRE chamber are pressure (p), temperature (T) and flow rate (W) of the products of combustion (decomposition) of liquid rocket fuel. In this case, the values of the parameters at the nozzle inlet (section index “c”), as well as in the critical (“*”) and outlet (“a”) nozzle sections are highlighted.
Calculation of parameter values in various sections of the LRE nozzle tract and determination of the thrust parameters of the engine is carried out according to the corresponding equations of thermogasdynamics. An approximate methodology for such a calculation is discussed in Section 4 of this manual.
- SCHEME AND PRINCIPLE OF OPERATION LRE "RD-214"
3.1. General characteristics of LRE "RD-214"
The RD-214 liquid-propellant rocket engine has been used in domestic practice since 1957. Since 1962, it has been installed on the 1st stage of the Kosmos multi-stage launch vehicles, with the help of which many satellites of the Kosmos and Interkomos series have been launched into near-Earth orbits.
LRE "RD-214" has a pumping fuel supply system. The engine runs on a high-boiling nitric acid oxidizer (a solution of nitrogen oxides in nitric acid) and hydrocarbon fuel (kerosene processing products). A special component is used for the gas generator - liquid hydrogen peroxide.
The main parameters of the engine have the following meanings:
Thrust in the void R p = 726 kN;
The specific impulse of thrust in the void I yn = 2590 N×s/kg;
Gas pressure in the combustion chamber p k = 4.4 MPa;
Degree of gas expansion in the nozzle e = 64
LRE "RD-214", (Fig. 1) consists of:
Four chambers (pos. 6);
One turbopump unit (TPU) (pos. 1, 2, 3, 4);
Gas generator (pos. 5);
pipeline;
Automation units (pos. 7, 8)
The THA of the engine consists of an oxidizer pump (pos. 2), a fuel pump (pos. 3), a hydrogen peroxide pump (pos. 4) and a turbine (pos. 1). The rotors (rotating parts) of the pumps and the turbine are connected by a single shaft.
Units and units that provide the supply of components to the engine chamber, gas generator and turbine are combined into three separate systems - lines:
Oxidizer supply system
fuel supply system
Hydrogen peroxide steam and gas generation system.
Fig.1. Schematic of a liquid propellant rocket engine
1 - turbine; 2 – oxidizer pump; 3 - fuel pump;
4 – hydrogen peroxide pump; 5 – gas generator (reactor);
6 – engine chamber; 7, 8 - elements of automation.
3.2. Characteristics of the LRE units "RD-214"
3.2.1. LRE chamber
Four LRE chambers are connected into a single block along two sections with the help of bolts.
Each LRE chamber (pos. 6) consists of a mixing head and a housing. The mixing head includes top, middle and bottom (firing) bottoms. A cavity for the oxidizer is formed between the upper and middle bottoms, and a cavity for fuel is formed between the middle and fire bottoms. Each of the cavities is connected with the internal volume of the engine housing by means of the corresponding injectors.
In the process of LRE operation, liquid fuel components are supplied, sprayed and mixed through the mixing head and its nozzles.
LRE chamber housing includes part of combustion chamber and nozzle. The liquid-propellant rocket engine nozzle is supersonic, has converging and diverging parts.
The housing of the LRE chamber is double-walled. The inner (fire) and outer (power) walls of the body are interconnected by spacers. At the same time, with the help of spacers, channels of the body liquid cooling path are formed between the walls. Fuel is used as a coolant.
During engine operation, fuel is supplied to the cooling path through special manifold pipes located at the end of the nozzle. Having passed the cooling path, the fuel enters the corresponding cavity of the mixing head and is injected through the nozzles into the combustion chamber. At the same time, through another cavity of the mixing head and the corresponding nozzles, an oxidizer enters the combustion chamber.
In the volume of the combustion chamber, spraying, mixing and combustion of liquid fuel components takes place. As a result, a high-temperature gaseous working fluid of the engine is formed.
Then, in the supersonic nozzle, the thermal energy of the working fluid is converted into the kinetic energy of its jet, upon expiration of which the LRE thrust is created.
3.2.2. Gas generator and turbopump unit
The gas generator (Fig. 1, item 5) is a unit in which liquid hydrogen peroxide is converted into a high-temperature vaporous working fluid of the turbine as a result of exothermic decomposition.
The turbopump unit provides pressure supply of liquid fuel components to the chamber and engine gas generator.
THA consists of (Fig. 1):
Screw-centrifugal oxidizer pump (pos. 2);
Screw-centrifugal fuel pump (pos. 3);
Hydrogen peroxide centrifugal pump (item 4);
Gas turbine (pos. 1).
Each pump and turbine has a fixed stator and a rotating rotor. The rotors of pumps and turbines have a common shaft, which consists of two parts, which are connected by a spring.
The turbine (pos. 1) serves as a pump drive. The main elements of the turbine stator are the housing and the nozzle apparatus, and the main elements of the rotor are the shaft and the impeller with blades. During operation, peroxide vapor gas is supplied to the turbine from the gas generator. When the steam gas passes through the nozzle apparatus and the blades of the turbine impeller, its thermal energy is converted into mechanical energy of rotation of the wheel and the turbine rotor shaft. The exhaust steam gas is collected in the outlet manifold of the turbine housing and discharged into the atmosphere through special waste nozzles. This creates some additional thrust LRE.
Pumps for oxidizer (pos. 2) and fuel (pos. 3) are screw-centrifugal type. The main elements of each of the pumps are the housing and the rotor. The rotor has a shaft, an auger and a centrifugal wheel with blades. During operation, mechanical energy is supplied from the turbine to the pump through a common shaft, which ensures the rotation of the pump rotor. As a result of the action of the screw blades and the centrifugal wheel on the liquid (fuel component) pumped by the pumps, the mechanical energy of rotation of the pump rotor is converted into potential energy of the liquid pressure, which ensures the supply of the component to the engine chamber. The screw in front of the centrifugal impeller of the pump is installed to preliminarily increase the pressure of the liquid at the inlet to the interblade channels of the impeller in order to prevent cold boiling of the liquid (cavitation) and disruption of its continuity. Disturbances in the continuity of the flow of the component can cause instability of the fuel combustion process in the engine chamber, and, consequently, the instability of the LRE as a whole.
A centrifugal pump (pos. 4) is used to supply hydrogen peroxide to the gas generator. The relatively low consumption of the component creates conditions for non-cavitational operation of a centrifugal pump without installing a screw prepump in front of it.
3.3. The principle of the engine
Start, control and stop of the engine is carried out automatically by electrical commands from the rocket board to the corresponding automation elements.
For the initial ignition of the fuel components, a special starting fuel is used, self-igniting with an oxidizer. Starting fuel initially fills a small section of the pipeline in front of the fuel pump. At the moment of launching the LRE, starting fuel and oxidizer enter the chamber, they spontaneously ignite, and only then do the main components of the fuel begin to enter the chamber.
During engine operation, the oxidizer sequentially passes through the elements and assemblies of the line (system), including:
Dividing valve;
Oxidizer pump;
Oxidizer valve;
Mixing head chamber motor.
The flow of fuel flows through the line, including:
Dividing valves;
fuel pump;
Collector and path for cooling the engine chamber;
mixing head chamber.
Hydrogen peroxide and the resulting vapor gas sequentially pass through the elements and units of the steam and gas generation system, including:
Dividing valve;
Hydrogen peroxide pump;
Hydraulic reducer;
gas generator;
Turbine nozzle apparatus;
Turbine impeller blades;
turbine manifold;
Waste nozzles.
As a result of the continuous supply of fuel components by the turbopump unit to the engine chamber, their combustion with the formation of a high-temperature working fluid and the expiration of the working fluid from the chamber, an LRE thrust is created.
Variation of the thrust value of the engine during its operation is provided by changing the flow rate of hydrogen peroxide supplied to the gas generator. This changes the power of the turbine and pumps, and, consequently, the supply of fuel components to the engine chamber.
Shutdown of the LRE is carried out in two stages with the help of automation elements. From the main mode, the engine is first switched to the final mode of operation with less thrust and only then is completely switched off.
- WORK METHODOLOGY
4.1. Scope and order of work
In the course of the work, the following actions are sequentially performed.
1) The scheme of the RD-214 rocket engine is being studied. The purpose and composition of the LRE, the design of the units, the principle of operation of the engine are considered.
2) The geometrical parameters of the LRE nozzle are measured. The diameter of the inlet ("c"), critical ("*") and outlet ("a") sections of the nozzle (D c, D * , D a) is found.
3) The value of the parameters of the LRE working fluid in the inlet, critical and outlet sections of the LRE nozzle is calculated.
Based on the results of the calculations, a generalized graph of the change in temperature (T), pressure (p) and velocity (W) of the working fluid along the nozzle path (L) of the LRE is constructed.
4) The thrust parameters of the liquid-propellant rocket engine are determined at the design mode of operation of the nozzle ().
4.2. Initial data for calculating the parameters of the rocket engine "RD-214"
Gas pressure in the chamber (see option)
Temperature of gases in the chamber
Gas constant
Isentropic exponent
Function 
It is assumed that the processes in the chamber proceed without energy losses. In this case, the energy loss coefficients in the combustion chamber and nozzle, respectively, are 
The nozzle operation mode is calculated (index " r»).
The measurement determines:
Nozzle throat diameter ;
Nozzle outlet diameter .
4.3. Sequence of calculation of LRE parameters
BUT) The parameters in the outlet section of the nozzle ("a") are determined in the following sequence.
1) Nozzle exit area
2) Nozzle throat area
3) Geometric degree of gas expansion
From everyday practice it is known that in the internal combustion engine, the furnace of a steam boiler - wherever combustion takes place, atmospheric oxygen takes the most active part. Without it, there is no combustion. There is no air in outer space, therefore, for the operation of rocket engines, it is necessary to have fuel containing two components - fuel and an oxidizer.
Liquid thermochemical rocket engines use alcohol, kerosene, gasoline, aniline, hydrazine, dimstylhydrazine, liquid hydrogen as a fuel, and liquid oxygen, hydrogen peroxide, nitric acid, liquid fluorine as an oxidizer. Fuel and oxidizer for LRE are stored separately, in special tanks and under pressure or with the help of pumps are fed into the combustion chamber, where, when they are combined, a temperature of 3000 - 4500 ° C develops.
Combustion products, expanding, acquire a speed of 2500-4500 m / s, creating jet thrust. The greater the mass and velocity of the outflow of gases, the greater the thrust of the engine. The pumps supply fuel to the engine head, in which a large number of injectors are mounted. Through some of them, an oxidizing agent is injected into the chamber, through others - fuel. In any car, during the combustion of fuel, large heat flows are formed that heat the walls of the engine. If you do not cool the walls of the chamber, then it will quickly burn out, no matter what material it is made of. LRE, as a rule, is cooled by one of the propellant components. For this, the chamber is made double-walled. The fuel component flows in the gap between the walls.
A large specific thrust impulse is created by an engine running on liquid oxygen and liquid hydrogen. In the jet stream of this engine, gases rush at a speed of a little more than 4 km / s. 2
The temperature of the jet is about 3000°C, and it consists of superheated water vapor, which is formed during the combustion of hydrogen in oxygen. The main data of typical fuels for LRE (on Earth) are given in the table.
Oxidizer Fuel Density, kg/m3 Specific thrust impulse, m/s Specific calorific value, kJ/kg
Nitric acid Kerosene 1400 2900 6100
Liquid oxygen Kerosene 1036 3283 9200
Liquid oxygen Liquid hydrogen 345 4164 13400
Liquid oxygen Dimethylhydrazine 1000 3381 9200
Liquid Fluorine Hydrazine 1312 4275 9350
Main characteristics of liquid rocket propellants
But oxygen, along with a number of advantages, has one drawback - at normal temperatures it is a gas. It is clear that it is impossible to use gaseous oxygen in a rocket, because in this case it would have to be stored under high pressure in massive cylinders. Therefore, already Tsiolkovsky, who was the first to propose oxygen as a component of rocket fuel, spoke about liquid oxygen. To turn oxygen into a liquid, it must be cooled to a temperature of -183 ° C. However, liquefied oxygen evaporates easily and quickly, even if it is stored in special heat-insulated vessels. Therefore, it is impossible, for example, to keep a rocket equipped for a long time, the engine of which runs on liquid oxygen. You have to fill the oxygen tank of such a rocket just before launch.
Nitric acid does not have this disadvantage and is therefore a "permanent" oxidizing agent. This explains its strong position in rocket technology, despite the significantly lower specific thrust impulse that it provides.
Left - Solid Rocket Engine (TPRD)
Right - Hybrid rocket engine
The use of fluorine, the most powerful oxidizing agent known to chemistry, will make it possible to significantly increase the efficiency of a liquid propellant rocket engine. True, liquid fluorine is inconvenient to use due to its toxicity and low boiling point (-188 °C). But this does not stop rocket scientists: experimental fluorine engines already exist. F. A. Zander suggested using light metals as fuel - lithium, beryllium, etc., especially as an additive to conventional fuel, for example, hydrogen-oxygen. Such "triple compositions" are capable of providing the highest possible outflow velocity for chemical fuels up to 5 km/s. But this is probably the limit of chemistry resources. She can't really do more than that.
The efficiency of a propulsion system (PS) with an LRE increases with an increase in the specific thrust impulse and fuel density. Moreover, in recent years more and more requirements have been imposed on the environmental cleanliness of both the fuel components themselves and their combustion products. At present, liquid oxygen and liquid hydrogen are the best highly efficient, environmentally friendly fuels. However, the extremely low density of liquid hydrogen (only 70 kg/m3) significantly limits the possibility of its application. The best fuel components for the first stage PS are liquid oxygen and hydrocarbon fuel. Until now, kerosene is most often used as a hydrocarbon fuel (HCF). However, kerosene has a number of disadvantages, and therefore the use of methane (CH4), propane (C3H8) and liquefied natural gas is being considered.
1 - Combustion chamber
3 - Turbine
4 - Oxidizer pump
5 - Fuel pump
7 - Gas generator
SCHEME OF LRE WITHOUT AFTERBURNING OF GAS GENERATOR GAS
Increasing the pressure in the combustion chamber is the second most important way to improve the energy characteristics of a rocket engine. An increase in pressure in the LRE chambers also contributes to a reduction in the overall dimensions of the power plant. It should be noted that an increase in the specific thrust impulse of a liquid-propellant rocket engine, a reduction in the overall dimensions of engines and a carrier as a whole can be achieved by using a retractable nozzle nozzle (two-position nozzle), i.e., using a nozzle with height compensation
1 - Combustion chamber
2 - Gas pipeline
3 - Turbine
4 - Oxidizer pump
5 - Fuel pump
6 - Generator fuel pump
7 - Gas generator
SCHEME OF LRE WITH AFTERBURNING OF GAS GENERATOR GAS
Although we started the story with a liquid-propellant rocket engine, it must be said that the thermochemical solid-propellant rocket engine, the TTRD, was the first to be created. Fuel - special gunpowder - is located here directly in the combustion chamber. A chamber with a jet nozzle - that's the whole design. Solid propellant rocket engines have many advantages over liquid fuel engines: they are easy to manufacture, can be stored for a long time, are always ready for action, and are explosion-proof. But in terms of specific impulse, solid propellant rocket engines are 10 - 30% inferior to liquid ones.
For many years, the development of domestic fuels was carried out by scientists from the State Institute of Applied Chemistry under the leadership of V.S. Shpak in the city of Leningrad. Foreign launch vehicles use:
Mixed solid fuel based on polybutadiene rubber (NTRV);
Mixed solid fuel based on polybutadiene acrylonitrile rubber (PBAN).
LIQUID ROCKET ENGINE (LRE) - rocket engine, powered by liquid rocket fuel. The transformation of the fuel into a thrust-producing jet of gas occurs in the chamber. In modern LRE Both bipropellant propellants, consisting of an oxidizer and fuel stored in separate tanks, and monopropellant propellants, which are liquids capable of catalytic decomposition, are used. According to the type of oxidizing agent used LRE there are nitric acid, nitrogen tetroxide (oxidizing agent - nitrogen tetroxide), oxygen, hydrogen peroxide, fluorine, etc. Depending on the thrust value, they distinguish LRE small, medium and high thrust. The conditional boundaries between them are 10 kN and 250 kN (aircraft were equipped with LRE with thrust from tenths of N to 8 MN). LRE are also characterized by specific thrust impulse, operating mode, dimensions, specific gravity, pressure in the combustion chamber, general arrangement and design of the main units. LRE is the main type of space engines and is also widely used in high-altitude research rockets, long-range combat ballistic missiles, anti-aircraft guided missiles; limited - in combat missiles of other classes, on experimental aircraft, etc.
The main problems when creating LRE: rational choice of fuel that meets energy requirements and operating conditions; organization of the workflow to achieve the calculated specific impulse; ensuring stable operation at specified modes, without developed low-frequency and high-frequency pressure fluctuations that cause destructive engine vibrations; cooling a rocket engine exposed to aggressive combustion products at very high temperatures (up to 5000 K) and pressures up to many tens of MPa (this effect is exacerbated in some cases by the presence of a condensed phase in the nozzle); supply of fuel (cryogenic, aggressive, etc.) at pressures reaching for powerful engines up to many tens of MPa and at flow rates up to several tons per second; ensuring the minimum mass of units and the engine as a whole, operating in very stressful conditions; achieving high reliability.
LRE was proposed by K. E. Tsiolkovsky in 1903 as an engine for space flight. Scientist developed a concept LRE, pointed out the most profitable rocket fuels, investigated the issues of the design of the main units. Practical work on creating LRE were started in 1921 in the USA by R. Goddard. In 1922, he first registered thrust when testing an experimental LRE, and in 1926 launched a small liquid rocket. In the late 20's - early 30's. to the development LRE started in Germany, the USSR and other countries. In 1931, the first Soviet LRE ORM and ORM-1, created by V.P. Glushko in the Gas Dynamics Laboratory. In 1933, the OR-2 propulsion system designed by F. A. Zander was tested, and engine 10, created by the Jet Propulsion Study Group, ensured the flight of a liquid rocket.
Before the start of the 2nd World War 1939-45. prototypes appeared in the USSR and the USA LRE with thrust up to several kN, designed for experimental aircraft. Intensive work in the field of rocket technology, carried out in Germany during the war, caused the appearance of various types LRE military purpose, many of which were mass-produced. The best were LRE designs by H. Walter (including HVK 109-509A (HWK 109-509A)) and H. Zborowski, LRE anti-aircraft guided missile "Wasserfall" (Wasserfall) and ballistic missile V-2 (V-2). Until the 2nd half of the 40s. the largest Soviet LRE were D-1-A-1100 and RD-1, developed by the Jet Research Institute. The first serial Soviet LRE steel engines RD-1 and RD-1KhZ, created by the end of the war in the GDL-OKB. In the same place in 1947-53. developed the first in the USSR powerful LRE: RD-100, RD-101, RD-103. During the same period, the United States produced LRE with a thrust of ~ 350 kN for the Redstone ballistic missile.
Further development LRE and their current state was determined by the beginning in the mid-50s. in the USSR and the USA, the development of ICBMs and launch vehicles. To implement them, it was necessary to create powerful, economical and compact LRE. The first among them were the RD-107 and RD-108, with the advent of which the thrust LRE doubled, thrust control - 10 times. Specific impulse LRE increased by almost 30%, the specific gravity decreased by more than 1.5 times. These results were made possible thanks to the development of a fundamentally new design. LRE, which made it possible to switch from oxygen-ethyl alcohol fuel to oxygen-kerosene fuel while simultaneously increasing the pressure in the combustion chamber by 2–2.5 times.
From the beginning of the 60s. on launch vehicles (LV) also began to be applied LRE operating on high-boiling fuels. The first of these was the RD-214. Of great importance for the development of astronautics was the creation in the mid-60s. oxygen-hydrogen LRE(designed for the upper stages of the launch vehicle), which in terms of specific impulse exceed the oxygen-kerosene ones by 30%. Because oxygen-hydrogen fuel, compared with oxygen-kerosene fuel, requires three times more volume for its placement with the same mass, and hydrogen tanks have to be provided with thermal insulation, then the Tsiolkovsky number is 40% larger for oxygen-hydrogen fuel. This drawback is more than offset by the high efficiency of oxygen-hydrogen LRE. With an equal launch mass of the launch vehicle, they are able to launch three times more payload into low Earth orbit than oxygen-kerosene LRE.
Mastering ever more efficient fuels, designers LRE sought at the same time to convert the chemical energy of the fuels into the kinetic energy of the jet stream with the greatest possible efficiency. For this purpose, a scheme was developed LRE with afterburning of generator gas in the chamber. To implement this scheme, it was necessary to create chambers operating under conditions of high mechanical and thermal loads, as well as compact high-power power units. LRE with afterburning since the mid-60s. are widely used on launch vehicles, in particular, they are used at all stages of the Proton launch vehicle.
Along with powerful space LRE numerous LRE medium and low thrust. The trouble-free operation of spacecraft (SC) engines is ensured to a large extent by the use of high-boiling single-component and self-igniting propellants, which are not difficult to store on board the SC. remote control with LRE single-component fuels are simpler in design, but have a significantly lower specific impulse. By the mid 60s. in subsidiary LRE Hydrogen peroxide has received the greatest use, which then began to be replaced by hydrazine and two-component fuels. The use of hydrazine made it possible to increase the specific impulse LRE on single-component fuel by about 40%.
Most of the Soviet space LRE created in GDL-OKB V.P. Glushko, OKB A.M. Isaev and OKB S.A. Kosberg. Engines RD-107, RD-108, RD-214, RD-216, RD-253 and other GDL-OKB designs ensured the launch of all Soviet launch vehicles; on the second stages of a number of launch vehicles are also installed LRE designs GDL-OKB: RD-119, RD-219, etc. The engines of the Kosberg Design Bureau are installed on the upper stages of the Vostok, Voskhod (Soyuz) and Proton launch vehicles. Isaev Design Bureau engines are mainly used on artificial Earth satellites (AES), interplanetary spacecraft and spacecraft (KK) (KRD-61, KDU-414, TDU-1, KTDU-5A, etc.).
The largest of the foreign organizations involved in the development LRE are located in the USA. The leading company is Rocketdyne, which created LRE Jay-2 (J-2), LR-79-NA (LR-79-NA), LR-89-NA (LR-89-NA), LR-105-NA (LR-105-NA), RS- 2701 (RS-2701), H-1 (H-1), F-1 (F-1), SSME (SSME), numerous LRE medium and low thrust on high-boiling two-component fuel. Most of the powerful LRE created under the leadership of S. Hoffman. Aerojet General Corporation has created a series of LRE on high-boiling two-component fuel, incl. LRE LR-87-ADJ-5 (LR-87-AJ-5) and LR-91-ADJ-5 (LR-91-AJ-5), series LRE medium thrust ADJ-10 (AJ-10), including ADJ-10-137 (AJ-10-137) and ADJ-10-138 (AJ-10-138). Pratt & Whitney created the world's first oxygen-hydrogen LRE RL-10 (RL-10), Bell Aerospace Textron (Bell Aerospace Textron) - numerous auxiliary LRE, as well as LRE medium thrust LR-81-BA-9 (LR-81-BA-9), TRV company - LRE medium thrust LMDE (LMDE), Marquardt company (Marquardt) - series LRE on high-boiling two-component propellant for spacecraft and interplanetary spacecraft. In the United States, several dozen types of hydrazine LRE(tested in flight LRE with thrust from 0.4 N to 2.7 kN). Among the developers LRE for interplanetary spacecraft - Reaction Motors, which also created a powerful LRE LR-99-RM-1 (LR-99-RM-1). The most famous of Western European LRE- AshM-7 (HM-7), Valois (Valois), Vexen (Vexen), Viking (Viking, France), Gamma-2 (Gamma), Gamma-8, RZet- 2 (RZ-2, UK). Western Europe is also developing LRE low thrust on two- and one-component propellants for satellites. Japan produces under license American LRE LR-79-NA for its own version of the Delta launch vehicle (Delta). For one of the stages of this launch vehicle, Mitsubishi has developed a liquid-propellant rocket engine running on high-boiling fuel with a thrust of 53 kN with displacement feed. The stands tested oxygen-hydrogen LRE thrust up to 0.1 MN with pumping. Chinese launch vehicles use LRE thrust 0.7 MN with pumping high-boiling fuel.
Space LRE varied in design and features. The biggest difference exists between powerful LRE, providing acceleration of the launch vehicle, and LRE reactive spacecraft control systems. The first ones run on two-component fuel. The thrust of these LRE reaches 8 MN (with a total thrust of up to 40 MN), dimensions - several meters, and weight - several tons. They are usually designed for a single inclusion (except for some LRE upper stages of the pH) and work for 2-10 minutes when changing parameters within narrow limits. To these LRE a requirement is made to provide a high specific impulse with small dimensions and weight. Therefore, they use pumping fuel into the chamber (the exception is LRE"Vexin" and "Valois"). To this end, in LRE envisaged turbopump unit(THA) and gas generator(GG). HPP contains high-pressure fuel pumps (usually axial-centrifugal) and a turbine that drives them, which is rotated by the gas produced in the GG. IN LRE without afterburning, the generator gas exhausted in the turbine is discharged into the exhaust pipe, steering nozzle or chamber nozzle. IN LRE with afterburning, this gas enters the afterburner chamber with the rest of the fuel.
IN LRE without afterburning, 2-3% of the total fuel can be consumed through the GG, and the reasonable pressure limit in the combustion chamber is limited to ~ 10 MPa, which is associated with the loss of specific impulse for the HPP drive: for LRE in general, this parameter is lower than for the camera, because the additional thrust created by the expiration of the spent generator gas is small. The reason for this is the low pressure and temperature of this gas. For LRE RD-216 they are, for example, 0.12 MPa and 870 K, respectively; in this case, the specific impulse losses reach 1.5% (over 40 m/s). With an increase in pressure in the combustion chamber, an increase in its specific impulse is observed, but for this it is necessary to increase the flow of generator gas (to ensure the required power of the fuel pumps). From a certain moment, the ever-increasing losses of the specific impulse to the drive of the TPU balance, and then exceed the increase in the specific impulse of the chamber. IN LRE with afterburning through the GG, it means that part of the entire fuel (20-80%) is consumed, however, the HPP drive is carried out without compromising efficiency LRE(values of the specific impulse of the chamber and LRE match). In the combustion chambers of these LRE it is possible to realize a pressure of 15-25 MPa (the pressure in the GG is approximately twice as high). For powerful LRE with pumped fuel supply, the specific impulse reaches 3430 m/s when using oxygen-kerosene fuel and 4500 m/s when using oxygen-hydrogen; specific gravity LRE can be as low as 0.75-0.85 g/N.
In addition to the camera, TNA and GG, powerful LRE contain fuel lines with bellows hoses and angular and linear motion compensators, facilitating assembly and installation LRE, as well as providing unloading from thermal stresses and allowing the chamber to be deflected in order to control the movement of the launch vehicle; generator gas pipelines and fuel drainage; devices and systems rocket engine start; automation units with electric drives, pneumatic, pyro- and hydraulic systems and devices for controlling work LRE(including for his throttling); units of the emergency protection system; telemetric measurement system sensors; electric cable trunks for sending signals to automation units and receiving signals from telemetry sensors; heat-insulating covers and screens that ensure the proper temperature in the engine compartment and exclude overheating or hypothermia of individual elements; elements of the tank pressurization system (heat exchangers, mixers, etc.); articulated suspension or frame for mounting LRE to the launch vehicle (the frame that perceives the thrust is at the same time the element on which the engine is assembled); often - steering chambers and nozzles with systems that ensure their operation; elements of the general assembly (brackets, fasteners, seals). According to the device, they distinguish block liquid rocket engines, single- and multi-chamber (with power supply of several chambers from one TNA).
LRE jet control systems belong to small thrust engines, their mass usually does not reach 10 kg, and their height is 0.5 m; mass of many LRE does not exceed 0.5 kg, and they fit in the palm of your hand. A characteristic feature of these LRE is the operation in a pulsed mode (for several years of the SC operation, the total number of switching on LRE can reach several hundred thousand, and the operating time is several hours). These LRE are single-walled chambers equipped with start-up and shut-off fuel valves, and are designed for displacement flow high-boiling fuel (two-component self-igniting or one-component). The pressure in the combustion chambers indicated LRE, determined mainly by the boost pressure of the PS tanks and the hydraulic resistance of the supply lines, is in the range of 0.7-2.3 MPa. In the case when the gas for boosting fuel tanks is located in the tanks themselves, its pressure decreases as the fuel is consumed, which leads to deterioration in performance. LRE. Relatively high specific impulse LRE(up to 3050 m/s for two-component fuel and up to 2350 m/s for hydrazine) is achieved due to the relatively large size of the jet nozzle, which ensures the expansion of combustion products to a very low pressure. Despite the small absolute mass LRE reactive control systems, their specific gravity is large (with a decrease in thrust from 500 to 1 N, it increases from approximately 5 to 150 g/N).
LRE spacecraft occupy an intermediate position in their characteristics between powerful LRE launch vehicles and LRE reactive control systems. Their thrust ranges from hundreds of N to tens of kN and can be either unregulated or adjustable; they can continuously operate for tenths of seconds and several thousand seconds with the number of switchings from 1 to several tens. In the indicated LRE the same types of fuels are used as in LRE jet control systems (single-component fuel is used only in LRE low thrust).
Plans for further space exploration LRE plays a big role. Powerful LRE, designed for the economical use of efficient fuels, are still in the spotlight. By 1981, an oxygen-hydrogen LRE with a thrust of more than 2 MN, designed to accelerate the aircraft from launch to launch into low Earth orbit. Thanks to advances in the field of cryogenic technology and thermal insulation materials, it becomes expedient to create LRE on low-boiling fuels developing a high specific impulse for use in spacecraft operating in space. Development progress LRE with a thrust of up to several tens of kN, operating on fuels containing fluorine and its derivatives (see, for example, RD-301), makes it possible to use fluorine LRE in upper stages of launch vehicles and in automatic spacecraft that will fly to the planets. During bench tests in 1977 of an experimental oxygen-hydrogen LRE(thrust 0.1 MN), developed for these purposes, achieved a specific impulse of 4690 m/s. Experimental studies of various problems of creating LRE on the metal-containing fuel.
Along with development for LRE new fuels are being searched for technical principles that provide a further increase in efficiency and a reduction in size and weight LRE. The improvement in parameters achieved by increasing the pressure in the chamber becomes less and less noticeable with increasing pressure, and the difficulties of creating LRE are increasing more and more. An increase in this parameter above 25-30 MPa is ineffective and difficult to implement. There is interest in LRE equipped with nozzles with a central body. In order to reduce the cost of launching payloads, LRE(for reusable spacecraft), designed for several dozen flights and a resource of several hours with a small amount of inter-flight routine maintenance.
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