RD-109 engine
UDMH is a colorless, hygroscopic liquid with an ammonia odor. In terms of density and melting point, it roughly corresponds to kerosene, at normal temperature and in the absence of air it is stable, but at temperatures above 350 ° C it decomposes with the release of heat and the formation of flammable gaseous products; explodes when overheated in a confined space. It is more stable and less explosive than other hydrazine fuels, stable when stored in sealed containers. It dissolves well in water, alcohols, hydrocarbons, amines and ethers. Low corrosive activity in relation to many construction materials. Among the negative properties of UDMH, one can name, first of all, the high cost of production, a fairly low boiling point (63 ° C) and extremely high toxicity. Thinking to start developing a large family of liquid-propellant rocket engines on the new fuel, V.P. Glushko understood that solid support was needed for a large-scale deployment of work. He hoped to get it from S.P. Korolyov, for whom he proposed to create an engine on the fuel "liquid oxygen - UDMH" for the third stage of the launch vehicle, intended for launching spacecraft to the Moon and for putting a heavy satellite ship into orbit around the Earth (the first two steps - modified R-7). V.P. Glushko guided the customer to an unheard-of large value of the specific impulse of his rocket engine - 350 units! The rocketeers, operating by that time on much smaller quantities, could not but be inspired by this number. According to ballistic calculations, a rocket with an optimal stage with a new liquid-propellant engine made it possible to launch an apparatus with a mass more than twice that of a launch vehicle with a corresponding oxygen-kerosene stage to the Moon. (Initially, SP Korolev proposed to create an oxygen-kerosene rocket engine for the third stage of the carrier based on the steering chamber of the engines of the first stages of the "seven".) When compared with the proposed oxygen-kerosene rocket engine, the calculated advantages of the engine on the new fuel looked very, very clear. S.P. Korolev believed in the new fuel. This option became the main one, but not the only one: preferring to minimize the risk associated with the creation of a product using a poorly studied fuel component, the Chief Designer of OKB-1 instructed the employees of his propulsion department to prepare a draft of an alternative oxygen-kerosene rocket engine. On 02/10/1958, he met with the head of the Voronezh OKB-254 (now KB of Chemical Automatics) S.A. Kosberg and instructed him to create a backup engine for his carrier based on this project using the steering combustion chamber of the "seven" designed by M. Melnikov and the new TNA developed in Voronezh. At the beginning of 1958, the development of a launch vehicle began in Podlipki, which was supposed to provide launches of spacecraft to the Moon in the autumn-winter of the same year. The work on the project of the carrier was supported by the corresponding Decree of the Central Committee and the Council of Ministers of 03/20/1958. The draft design was signed by S.P. Korolev on 07/01/1958.
LV "Vostok" with the satellite "Foton", created on the basis of the satellite-photo reconnaissance aircraft "Zenit-2"
Considering both engines, the designers of OKB-1 realized that the rocket under development would have great prospects as a carrier. In particular, the mass of the heavy satellite, which was originally conceived as a photographic reconnaissance aircraft, became sufficient to design a manned spacecraft (SC) on its basis. Based on the planned characteristics of the third stage liquid-propellant engine, the parameters of the spacecraft and the launch vehicle were selected to put it into orbit. According to their calculations, it turned out that a liquid-propellant rocket engine based on a new synthetic fuel, in comparison with a kerosene engine, made it possible to increase the mass of the ship by 23%. The VPGlushko engine, which had the "proprietary" designation RD-109, was a single-chamber liquid-propellant rocket engine for the upper stages of space rockets. The unprecedented value of the specific impulse was supposed to be achieved not only by using a new high-energy fuel, but also thanks to great pressure in the combustion chamber (over 75 ata) and in a high-altitude nozzle with a high degree of expansion (cutoff pressure - 0.1 ata). The propellant components were fed into the chamber using the THA; after working out on its turbine, the gas was diverted to the steering nozzles, which were used to control the rocket in flight. The rocket engine consisted of a fuel-cooled combustion chamber with a flat nozzle head and a profiled nozzle, a two-shaft TNA with a gas generator, automation units and general assembly units. To drive the TNA turbine, a gas generator (GG) was used, which does not run on steam-gas, as in engines of previous designs, but on the combustion products of the main fuel with a large excess of fuel ("sweet" gas). However, during preliminary tests, due to the excessively low consumption of the oxidizer, serious difficulties with reliable start-up were revealed, so further work with the two-component GG was stopped. The accelerated development and fine-tuning of a one-component gas generator operating on the principle of thermocatalytic decomposition of UDMH began.
Diagram of the RD-109 engine
The combustion chamber with a high-altitude nozzle was the first product of this type developed by OKB-456. At the same time, the possibility of its cooling with dimethylhydrazine was checked and its operational properties were investigated. These results were supposed to be subsequently used for the development of powerful engines using new fuels. Fuel combustion in RD-109 took place at higher temperatures and pressures than in previous liquid-propellant rocket engines, and its chamber operated under more severe thermodynamic conditions. The situation was aggravated by the fact that the efficiency of the chamber cooling system turned out to be lower than the calculated one. SP Korolev met the news of the difficulties faced by the creators of the RD-109 with understanding. He clearly understood that V.P. Glushko was creating an entirely new type of liquid-propellant engine. In the middle of 1958, the attitude of V.P. Glushko to his engine changed markedly. Due to the great difficulties in working out the combustion chamber and gas generator, Valentin Petrovich preferred to retreat and wait. By this time, OKB-456 began to create liquid-propellant rocket engines for new missiles - R-14 and R-16, operating on the components "nitric acid - UDMH". This fuel turned out to be much easier to fine-tune - it did not contain cryogenic components and burned at lower temperatures than oxygen-UDMH, due to which the chambers of the new engines operated under less stressful conditions. In addition, the fuel components spontaneously ignited in contact with each other, which greatly simplified the starting system. All this led to the fact that, despite the large dimension of the new engines, progress with them was much more obvious than with the RD-109. Referring to the great employment of work on new liquid-propellant engines, V.P. Glushko did not pay due attention to his first-born. Active work on it has slowed down. It became obvious that hopes for the creation of a liquid-propellant rocket engine by the fall of 1958 and its participation in the first launches of spacecraft to the Moon were groundless ... satellite ships. The development of RD-109 elements and systems continued, but at a completely different pace. A large volume of tests of the gas generator was carried out, during which it was revealed that at temperatures below 100 ° C, the decomposition process of UDMH stops, and when the wall is heated above 250 ° C, explosions occur in the cooling path of the gas generator. Bench firing tests RD-109 v complete set began only in January 1959, they confirmed the possibility of creating a liquid-propellant rocket engine with a high specific thrust, working on UDMH. Testing of the launch was carried out on a stand equipped with a pressure chamber with a volume of 90 m3, which ensures the operation of the engine at an ambient pressure of about 1 mm Hg. Art. During the firing tests, the sequence of giving commands to start the liquid-propellant engine was selected, the fuel consumption at the preliminary stage was determined, the purge modes were worked out, and the operability of the pyro-ignition device was checked. During the tests, it was found that the zone of stable operation of the engine lies above the previously assumed value, which made it possible to increase the nominal pressure in the combustion chamber from 76 to 79 ata. As a result of hard work, a high-speed, efficient TNA with a cooled gearbox was created. The fine-tuning of the unit was carried out in conditions close to real ones. During bench tests of the first copies of the turbine, it turned out that the power it developed was somewhat lower than the required one. This required special measures to improve it. In the process of fine-tuning tests during 1959, they worked out the start of the engine and checked the joint operation of all its units and assemblies, and some of them had to be significantly modified. So, on the instructions of the design bureau - the customer created and worked out the original design of the mixer for pressurizing the fuel tank. Unfortunately, during the debugging process, it was not possible to get rid of cracks in the welded joints of the blades with the turbine disk. A more complex and heavy version of attaching the blades with the help of a herringbone-type lock was used. Nevertheless, life tests have shown that the RD-109 engine is operational for a given time.
Turbo pump unit of the RD-109 engine
Everything would be fine, but the main results inspired rocket scientists: the specific impulse was much lower than the specified value and barely reached 334 units. Meanwhile, even the first samples were created in record time - in just nine months! - reserve oxygen-kerosene engine RD-0105, which received the “proprietary” name RO-5 in Voronezh, had a specific impulse of over 316 units. Its developers did not see any particular difficulties on the way to increase this indicator by another 10-15 units in the near future. Naturally, such a small difference in the specific impulse of the two competing engines negated the advantages of RD-109 for a three-stage launch vehicle: the maximum calculated mass of the PG (automatic lunar apparatus) of the "main" launch vehicle dropped to 424 kg, and the "duplicate" version increased to 373 kg. The understudy became number one - attractive and promising, and the main variant risked completely disappearing from the stage. Generally speaking, the achieved specific impulse was not a surprise for the employees of OKB-456. The fact is that the influence of a large number of unknown factors in the design has lowered the efficiency, reliability and operability of the combustion chamber and TNA in comparison with the calculated ones. It was required to carry out additional work to improve the already created engine. V.P. Glushko tried to prove to everyone that by making minor changes to the existing structure, the preliminary values of the design parameters could even be surpassed. Having weighed all the pros and cons, S.P. Korolev refused to use an oxygen-UDMG liquid-propellant rocket engine for the carrier of a manned spacecraft, however, V.P. the use of this engine on newly developed products and the results are coordinated with OKB-456 ". An oxygen-fueled rocket engine – UDMG with the parameters specified in the design of a three-stage rocket did not appear either in 1958 or in 1959. At the beginning of 1960, work on the RD-109 was stopped in connection with the beginning of the development of a more advanced RD-119 engine.
RD-119 engine
The new liquid-propellant rocket engine differed from RD-109 by a significantly increased specific thrust (the height of the nozzle was increased by more than one and a half times, the process of mixture formation in the chamber was improved), as well as by a significantly lower mass and greater reliability. A number of drastic changes aimed at improving its energy and mass characteristics, improved the cooling of the inner wall of the chamber, creating a two-slot belt of additional curtain cooling; a new nozzle head was worked out, which increased the stability of the working process and ensured a greater completeness of combustion of fuel components. These measures made it possible to obtain the highest specific thrust impulse in the void for its time (352 units). At the same time, due to the choice of a rational profile of the supersonic part of the nozzle, as well as due to the widespread use of titanium alloys in the design of the chamber, it was possible, despite a significant increase in the outlet diameter of the nozzle, to somewhat reduce the mass of the combustion chamber.
Combustion chamber of RD-119 engine
TNA engine RD-119 was made according to a single-shaft scheme. Thanks to the simplification of the unit and the improvement of its characteristics, it was possible to significantly reduce the gas consumption for the turbine drive and the mass of the TNA. The gas generator of the engine had an uncooled casing. To increase the efficiency of the flight control system in the first seconds of operation, RD-119, as well as RD-109, provided for the bypass of gas from the gas generator to the steering nozzles, bypassing the turbine. A significant increase in the reliability of the engine was achieved thanks to the nozzle head, which ensured a stable working process in the combustion chamber, as well as due to the introduction of welded joints in the turbine and gas generator instead of flanged ones and the development of the technological process of manufacturing units and assemblies. To control the quality, each RD-119 engine was tested at the stand according to a new method: by control burn-through for 150 seconds and a selective batch test for service life of 260 seconds. The new liquid-propellant rocket engine was developed in the period 1960–1963, in 1963 it underwent final development tests and was put into mass production. However, even before that moment, in 1962, his flying destiny began. As you can understand, at this moment a new stage in the life of the "Khimki motor" began. However, she was no longer associated with the OKB S.P. Korolev. RD-119 was just on its way to the stand, and the Voronezh RO-5 was already successfully tested in flight on a three-stage version of the "seven" when the first "lunar ships" were launched. The next step of this carrier is a rocket for a manned spacecraft. The RD-119 engine already met the requirements set in the design of the carrier for the ship, but it was still out of work. No matter how V.P. Glushko proved his advantages over the oxygen-kerosene counterpart, S.P. Korolyov remained adamant. Perhaps he thought: “Why do we need a new, even promising engine? It's a cat in a poke. We already have reliable motor who has done well. Moreover, it is not yet known how the new component will behave in operation. And we have a long-standing friendship with kerosene. And there is practically no need to modernize the ready-made launch complex ... ”However, the main thing, it seems, is not the point: the RD-0109 (RO-7) engine developed by S.A. Kosberg (an improved version of RO-5) already had a specific impulse of 326 units. The advantages of RD-119 were insignificant. And such disadvantages as the high toxicity of UDMH and its vapors, the high cost of fuel, as well as its low boiling point, outweighed.
Oxygen-kerosene new engine RD-0109
So, must have thought S.P. Korolev, deciding to abandon UDMH in favor of kerosene on his rocket for manned space flight. Did he make the right conclusion? From the heights of today, it is clear that yes. With the exception of the possibility of creating a one-component gas generator, a liquid-propellant rocket engine running on oxygen-UDMH fuel has practically no advantages over an oxygen-kerosene engine of a similar design (at the same pressure in the combustion chamber and the expansion ratio of the nozzle). Its disadvantages are obvious. After S.P. Korolev's refusal from the Khimki engine, V.P. Glushko, of course, did not despair: not all of the rocket engines being developed went into mass production. However, it took too much time and effort to create it. At one of the joint meetings on the industry, Valentin Petrovich proposed to MK Yangel RD-119. Mikhail Kuzmich promised to think it over.
What's the first thing that comes to mind when you say "rocket motors"? Of course, the mysterious space, interplanetary flights, the discovery of new galaxies and the alluring glow of distant stars. At all times, the sky attracted a person to itself, while remaining an unsolved mystery, but the creation of the first space rocket and its launch opened up new horizons for human research.
Rocket engines in essence, these are conventional jet engines with one important feature: they do not use atmospheric oxygen as a fuel oxidizer to create jet thrust. Everything that is needed for its operation is located either directly in its body or in the oxidizer and fuel supply systems. It is this feature that makes it possible to use rocket engines in open space.
There are a lot of types of rocket engines and all of them are strikingly different from each other not only in design features, but also in the principle of operation. That is why each species must be considered separately.
Among the main performance characteristics of rocket engines, special attention is paid to specific impulse - the ratio of the magnitude of the jet thrust to the mass of the working fluid consumed per unit of time. The specific impulse value represents the efficiency and economy of the engine.
Chemical rocket engines (CRM)
This type of engine is currently the only one that is massively used for launching spacecraft into outer space, in addition, it has found application in the military industry. Chemical engines are divided into solid and liquid propellants, depending on the state of aggregation of the propellant.
History of creation
The first rocket engines were solid-propellant, and they appeared several centuries ago in China. At that time, little connected them with space, but with their help it was possible to launch military missiles. As a fuel, a powder was used, resembling gunpowder in composition, only the percentage of its constituents was changed. As a result, during oxidation, the powder did not explode, but gradually burned out, releasing heat and creating jet thrust... Such engines were modified, improved and improved with varying degrees of success, but their specific impulse still remained small, that is, the design was ineffective and uneconomical. Soon, new types of solid fuels appeared, allowing you to get a larger specific impulse and develop more thrust. Scientists from the USSR, the USA and Europe worked on its creation in the first half of the twentieth century. Already in the second half of the 40s, a prototype of a modern fuel was developed, which is still used today.
The rocket engine RD - 170 runs on liquid fuel and an oxidizer.
Liquid propellant rocket engines are the invention of K.E. Tsiolkovsky, who proposed them as a power unit for a space rocket in 1903. In the 20s, work on the creation of liquid-propellant rocket engines began to be carried out in the USA, in the 30s - in the USSR. By the beginning of World War II, the first experimental samples were created, and after its end, liquid-propellant rocket engines began to be mass-produced. They were used in the military industry to equip ballistic missiles. In 1957, for the first time in the history of mankind, a Soviet artificial satellite was launched. To launch it, a rocket equipped with Russian Railways was used.
The device and principle of operation of chemical rocket engines
A solid-fuel engine contains fuel and an oxidizer in a solid state of aggregation in its body, and the container with fuel is also a combustion chamber. The fuel is usually in the form of a rod with a center hole. During the oxidation process, the rod begins to burn from the center to the periphery, and the gases resulting from the combustion exit through the nozzle, forming a thrust. This is the simplest design of any rocket engine.
In liquid jet engines, the fuel and oxidizer are in a liquid aggregate state in two separate tanks. Through the feed channels, they enter the combustion chamber, where they mix and the combustion process takes place. Combustion products exit through the nozzle, forming thrust. Liquid oxygen is usually used as an oxidizer, and the fuel can be different: kerosene, liquid hydrogen, etc.
Pros and cons of chemical RD, their scope
The advantages of solid fuel RDs are:
- simplicity of design;
- comparative safety in terms of ecology;
- low price;
- reliability.
Disadvantages of solid propellants:
- working time limitation: fuel burns out very quickly;
- impossibility of restarting the engine, stopping it and regulating traction;
- small specific gravity in the range of 2000-3000 m / s.
Analyzing the pros and cons of solid propellants, we can conclude that their use is justified only in those cases where you need power unit medium power, fairly cheap and easy to implement. The scope of their use is ballistic, meteorological rockets, MANPADS, as well as side boosters of space rockets (they are equipped with American rockets, they were not used in Soviet and Russian rockets).
Advantages of liquid RD:
- high specific impulse (about 4500 m / s and above);
- the ability to control traction, stop and restart the engine;
- less weight and compactness, which makes it possible to launch even large multi-ton loads into orbit.
LRE disadvantages:
- complex design and commissioning;
- under conditions of weightlessness, liquids in tanks can move randomly. For their deposition, you need to use additional energy sources.
The sphere of application of liquid-propellant rocket engines is mainly astronautics, since these engines are too expensive for military purposes.
Despite the fact that so far chemical RDs are the only ones capable of ensuring the launch of rockets into open space, their further improvement is almost impossible. Scientists and designers are convinced that the limit of their capabilities has already been reached, and other energy sources are needed to obtain more powerful units with a large specific impulse.
Nuclear rocket engines (NRM)
This type of RD, in contrast to chemical ones, generates energy not during the combustion of fuel, but as a result of heating the working fluid with the energy of nuclear reactions. NRE are isotopic, thermonuclear and nuclear.
History of creation
The design and principle of operation of the nuclear rocket engine were developed back in the 50s. Already in the 70s, experimental samples were ready in the USSR and the USA, which were successfully tested. The solid-phase Soviet engine RD-0410 with a thrust of 3.6 tons was tested at a bench base, and the American reactor "NERVA" was to be installed on the Saturn V rocket before the sponsorship of the lunar program was stopped. In parallel, work was carried out on the creation of gas-phase NRE. Now there are scientific programs for the development of nuclear taxiways, experiments are being carried out at space stations.
Thus, there are already operating models of nuclear rocket engines, but so far none of them has been used outside laboratories or scientific bases. The potential of such engines is quite high, but the risk associated with their use is also considerable, so so far they exist only in projects.
Device and principle of operation
Nuclear rocket engines are gas-, liquid- and solid-phase, depending on the state of aggregation of nuclear fuel. The fuel in solid-phase NRE are fuel rods, the same as in nuclear reactors. They are located in the engine housing and release thermal energy during the decay of fissile material. The working fluid - gaseous hydrogen or ammonia - in contact with the fuel element, absorbs energy and heats up, expanding and contracting, after which it exits through the nozzle under high pressure.
The principle of operation of a liquid-phase NRE and its structure is similar to a solid-phase one, only the fuel is in a liquid state, which makes it possible to increase the temperature, and hence the thrust.
Gas-phase nuclear propulsion engines operate on fuel in a gaseous state. They usually use uranium. The gaseous fuel can be kept in the housing by an electric field, or it can be kept in a sealed transparent flask - a nuclear lamp. In the first case, there is a contact of the working fluid with the fuel, as well as a partial leakage of the latter, therefore, in addition to the main mass of fuel, the engine must provide for its reserve for periodic replenishment. In the case of a nuclear lamp, no leakage occurs, and the fuel is completely isolated from the flow of the working fluid.
Advantages and disadvantages of NRE
Nuclear rocket engines have a huge advantage over chemical ones - they have a high specific impulse. For solid-phase models, its value is 8000-9000 m / s, for liquid-phase models - 14000 m / s, for gas-phase models - 30,000 m / s. At the same time, their use entails contamination of the atmosphere with radioactive emissions. Work is underway to create a safe, environmentally friendly and efficient nuclear engine, and the main "contender" for this role is a gas-phase NRE with a nuclear lamp, where the radioactive substance is in a sealed flask and does not come out with a jet flame.
Electric rocket motors (EPM)
Another potential competitor to chemical RDs is an electric RD, powered by electrical energy... ERE can be electrothermal, electrostatic, electromagnetic or pulsed.
History of creation
The first ERE was designed in the 30s by the Soviet designer V.P. Glushko, although the idea of creating such an engine appeared at the beginning of the twentieth century. In the 60s, scientists from the USSR and the United States actively worked on the creation of an electric propulsion engine, and already in the 70s, the first samples began to be used in spacecraft as control engines.
Device and principle of operation
The electric rocket propulsion system consists of the ERE itself, the structure of which depends on its type, the systems for supplying the working fluid, control and power supply. Electrothermal RD heats the flow of the working fluid due to the heat generated by the heating element or in an electric arc. Helium, ammonia, hydrazine, nitrogen and other inert gases are used as a working fluid, less often hydrogen.
Electrostatic RDs are divided into colloidal, ionic, and plasma. In them, charged particles of the working fluid are accelerated by the electric field. In colloidal or ionic RDs, gas ionization is provided by an ionizer, a high-frequency electric field, or a gas-discharge chamber. In plasma RDs, the working fluid — the inert gas xenon — passes through the annular anode and enters the gas-discharge chamber with a compensator cathode. At high voltage, a spark ignites between the anode and cathode, ionizing the gas, resulting in a plasma. The positively charged ions exit through the nozzle at a high speed, acquired due to acceleration by an electric field, and the electrons are removed outside by the compensator cathode.
Electromagnetic RDs have their own magnetic field - external or internal, which accelerates the charged particles of the working fluid.
Pulse RDs work due to the evaporation of solid fuel under the influence of electrical discharges.
Advantages and disadvantages of ERE, scope of use
Among the advantages of an ERE:
- high specific impulse, the upper limit of which is practically unlimited;
- low fuel consumption (working fluid).
Disadvantages:
- high level of electricity consumption;
- the complexity of the design;
- little thrust.
Today, the use of electric propulsion engines is limited to their installation on space satellites, and solar batteries are used as sources of electricity for them. At the same time, it is these engines that can become those power plants that will make it possible to explore space, therefore, work on the creation of their new models is being actively carried out in many countries. It was these power plants that were most often mentioned by science fiction writers in their works devoted to the conquest of space, and they can also be found in science fiction films. So far, it is the ERD that is the hope that people will still be able to travel to the stars.
In early 1996, NPO Energomash's RD-180 engine project was declared the winner of the tender for the development and supply of the first stage engine for the modernized Atlas launch vehicle by the American company Lockheed Martin. This is a two-chamber engine with afterburning of an oxidizing generator gas, with thrust vector control due to the rocking of each chamber in two planes, with the possibility of providing deep throttling of the engine thrust in flight. This design is based on well-proven designs of assemblies and elements of the RD-170/171 engines. The creation of a powerful engine of the first stage was carried out in a short time, and testing was carried out on a small amount of material. Having signed a contract for the development of the engine in the summer of 1996, already in November 1996, the first firing test of the prototype engine was carried out, and in April 1997, a firing test. standard engine... In 1997-1998, a series of firing tests of the engine as part of the LV stage was successfully carried out in the USA. In the spring of 1999, the engine was certified for use in the Atlas 3 launch vehicle. In the summer of 2001, the certification of the engine for use in the Atlas 5 LV was completed.
The engine is made in a closed circuit with the afterburning of the oxidizing generator gas after the turbine.
Fuel components: oxidizer - liquid oxygen, fuel - kerosene.
The engine consists of two chambers, a turbo pump unit (TNA), a booster fuel pump unit (BNAG), an oxidizer booster pump unit (BNAO), a gas generator, an automation control unit, a cylinder block, an automation drive system (SPA), a steering drive system (SRP) , a fuel flow regulator in the gas generator, an oxidizer throttle, a fuel throttle, start-off valves for the oxidizer and fuel, two ampoules with starting fuel, a starting tank, an engine frame, a bottom screen, sensors for an emergency protection system, a heat exchanger for heating helium for pressurizing the oxidizer tank.
When creating the RD-180 engine, due to the halving of the consumption of fuel components in comparison with the RD-170 prototype, it was necessary to redesign the THA and a number of automation units. According to the initial estimate, the unification of the RD-180 and RD-170 engines was 70 ... 75%. However, in the process of working out the RD-180 engine according to the Lockheed Martin technical assignment, more advanced design solutions were found than those used in the RD-170 engine for a number of units, including the design of the pump guide vanes, improved working conditions for the TNA bearings, increased efficiency supply units, a new sub-tank separating valve has been developed. In addition, the flange structure of the gas generator was replaced by a welded one, and the engine circuit was simplified. In connection with these works, the degree of unification of the RD-180 and RD-170 engines has significantly decreased. Essentially, the RD-180 engine is new development using the RD-170 engine as a basic version.
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Table 1. Technical parameters of the engine
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Fig. 1. RD-180 engine (enlarged image) |
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The engine contains two combustion chambers 1, a turbopump unit 2, consisting of a turbine 3, a two-stage fuel pump 4 and a single-stage oxidizer pump 5, a gas generator 6, a fuel booster pump 7 driven by a hydraulic turbine 8, and an oxidizer booster pump 9, which is driven is a gas turbine 10.
The booster pump of the oxidizer (BLLW) 9 is connected through the pipeline 11 to the inlet of the oxidizer pump 5, the outlet of which is connected through the cut-off valve 12 to the collector cavity 13 of the mixing head 14 of the gas generator 6. An oxidizer filter is installed at the inlet of the BLLW.
The fuel booster pump (BNAG) 7 is connected through the pipeline 15 to the inlet of the first stage 16 of the fuel pump 4. The first stage of the fuel pump 16 is connected to the inlet of the second stage 17 of the fuel pump and through the pipeline 18, in which the throttle 19 with the electric drive 20 is installed, is connected to the manifold 21 of the combustion chamber 1, from which the fuel is distributed through the channels 22 of the regenerative cooling of the combustion chamber 1. A fuel filter is installed at the inlet of the BNAG.
The channels 22 of the regenerative cooling of the nozzle 23 through the manifold 24 are connected to the start-cutoff valve 25. The outlet of this valve is connected to the manifold 26 located on the cylindrical part of the combustion chamber. The outlet of the collector 26 through the regenerative channels 27 for cooling the cylindrical part of the combustion chamber is connected to the fuel cavity 28 of the mixing head 29 of the combustion chamber 1.
The second stage 17 of the fuel pump 4 (through which 20% of the total fuel consumption passes) through the pipeline 30 is connected to the main inlet 31 of the draft regulator 32, controlled by an electric drive 33 and having a check valve 34 at the inlet. The outlet 35 of the draft regulator 32 is connected to 36 filled starting fuel triethylaluminum Al (C 2 H 5) h. Outlets from these ampoules through start-off valves 37 are connected to the fuel cavity 38 of the mixing head 39 of the gas generator 6. The outlet of the gas generators 40 is connected to the turbine 3, the outlet of which is connected through pipelines 41 to the cavity 42 of the mixing heads 29 of the combustion chambers 1.
In addition, the outlet from the turbine 3 through the pipeline 43, in which the heat exchanger 44 and the pressure valve 45 are installed, is connected to the manifold of the turbine 46 of the drive of the booster pump 9 of the oxidizer.
The pneumohydraulic circuit of the liquid-propellant engine also contains a starting system, which includes 47 with a separating membrane 48, a high-pressure gas supply pipe 49 and an outlet pipe 50. The outlet pipe 50 of the starting tank 47 is connected through the filling valve 51 to the fuel supply pipeline 15 from the fuel booster pump 7. In addition In addition, the outlet pipe 50 on one side through the pipeline 52, in which the check valve 53 is installed, is connected to the second inlet 54 of the draft regulator 32, through which the engine is started, and on the other hand, through the check valve 55, it is connected to 56 filled with a starting combustible triethylaluminum Al (C 2 H 5) z, the output of which through the valve 57 is connected to the line 58 for supplying the starting fuel to the ignition nozzles 59 of the combustion chamber. A nozzle 60 is installed in the line 58, which provides a metered supply of starting fuel to the ignition nozzles.
To reduce the aftereffect impulse start-off fuel valves are installed between the cooling ducts of the nozzle and the combustion chamber (valves 25), as well as in front of the collector of the second and third curtain belts.
The pneumatic valves are driven by helium from the high pressure cylinder bank by means of solenoid valves. Engine operation
The engine is started according to the "self-starting" scheme. Pre-drives 20 and 33 are set to positions that provide the initial installation of the thrust regulator 32 and throttle 19. Then open the rocket tank valves (not shown in the diagram) and, under the influence of hydrostatic head and boost pressure, the fuel components fill the cavities of the oxidizer and fuel pumps up to the cut-off valves 12 and 25 and check valve 34 of the draft regulator 32, respectively. The engine cavities are filled with fuel up to the starting ampoules 36 and 56 through the filling valve 51, check valves 53 and 55. 47 is also filled with the main fuel. This condition is considered the initial condition for starting the engine.
When the engine is started, the fuel is pressurized and displaced from it, the pressure of which breaks through the membranes (not shown) of the starting ampoules 36 and 56. At the same time, the starting-cutoff valves 12 and 37 and 25, respectively, are opened. As a result, the starting fuel from 36 and 56, under the action of the pressure created by the starting tank, enters the gas generator (through the open valve 37) and the chambers (through the check valves 57). The starting fuel entering the gas generator is ignited with oxygen, which also enters the gas generator due to the pre-launch pressurization of the rocket tanks and the hydrostatic head in them. The fuel, passing through the cooled path of the combustion chambers, after a fixed time enters the mixing heads of the combustion chambers 1. During this delay time, the combustion process begins in the gas generator and the generated generator gas spins the turbine 3 THA 2. After the turbine, the oxidizing gas flows through two cooled gas ducts 41 to the mixing heads 29 of the two combustion chambers, where it is ignited with the starting fuel coming from the ignition nozzles 59 and is subsequently burned up with the fuel entering the chambers. The time of entry of both components into the combustion chambers is selected so that THA 2 has time to enter the operating mode, while the back pressure has not yet been established in chambers 1.
As the pressure behind the fuel pump 17 rises, the starting tank 47 is automatically shut off from operation by closing the check valves 53 and 55, and the fuel supply to the gas generator 6 is switched to the pump 17 due to the programmed opening of the throttle of the draft regulator 32.
Part of the oxidizing gas from the outlet of the turbine is taken to the drive of the two-stage gas turbine 10 of the booster pre-pump 9. This gas, passing through the heat exchanger 44, heats the gas that goes to pressurize the tanks of the rocket. After the turbine 10, the gas is discharged into the outlet manifold 11, where it is mixed with the main stream of the oxidizer and condensed. The use of gas taken from the outlet of the TNA turbine as a working medium for the drive of the turbine of the booster pump of the oxidizer makes it possible to reduce the temperature in the gas generator and, accordingly, to reduce the power of the TNA turbine.
Part of the fuel from the output of the pump 4 goes to the drive of the single-stage hydraulic turbine 8 of the fuel booster pump 7.
A small part of liquid oxygen is taken from the gas generator manifolds and enters the cooling path of the turbine housing and gas ducts.
At the entire stage of starting the engine, program control of the opening of the throttle of the thrust regulator 32 and the throttle of fuel 19 from the positions of the initial setting to the positions corresponding to the nominal mode of the engine is carried out using the corresponding drives 33 and 20.
Thus, a smooth engine start is carried out with an exit to the main mode after 3 seconds.
Before switching off, the motors are transferred to the final stage mode, which is 50% of the nominal.
Figure 2.3. Simplified cyclogram of the RD-180 engine as part of the Atlas 3 and Atlas 5 launch vehicles
(see also; image is enlarged)
The chamber is a brazed-welded one-piece unit and consists of a mixing head, a combustion chamber and a nozzle. The chamber is attached to the gas path using a flange connection.
Table 2. Technical parameters of the camera
 | Fig. 4. Diagram of the fuel supply to the chamber cooling path:
|
The chamber body consists of a combustion chamber and a nozzle. The chamber body includes an outer load-bearing shell 11 and an inner fire wall 13 with milled channels forming a path for external regenerative cooling of the chamber with three coolant inlets. The first inlet is in communication with the nozzle throat cooling path, the second inlet is in communication with the nozzle outlet part cooling path, and the third in communication with the combustion chamber cooling path. In this case, the first outlet is in communication with the third inlet, and the first inlet, the second inlet and the supply to the two lower belts of slotted curtains are united by a common branch pipe, branched and located outside the chamber.
Internal cooling is provided by three belts of slotted curtains in the subcritical part of the combustion chamber. Through them, about 2% of the fuel is supplied to the wall in the form of films that evaporate and protect it from heat fluxes, which reach values of the order of 50 MW / m2 in the nozzle throat.
The means of ignition are made of four equally spaced around the circumference of the jet nozzles 6, installed behind the front (fire) bottom 3 in the power housing of the chamber 11. The axes of the flow openings of the jet nozzles are located at an acute angle to the outlet of the power housing and are deflected in a circle in the transverse plane from the longitudinal axis power housing in the same direction, and the axis of the flow hole of each jet nozzle is crossed with respect to the axes of the flow holes of the adjacent nozzles. The injectors are hydraulically united by a common manifold.
All nozzles are two-component with axial supply of oxidizing gas and tangential supply of fuel. The nozzles located near the fire (inner) wall of the chamber are made with increased hydraulic resistance along the fuel line in comparison with other nozzles due to a decrease in the diameters of the fuel supply holes, i.e. providing reduced fuel consumption compared to other injectors.
To suppress pressure pulsations, the initial zone of mixture formation and combustion, in which, as a rule, high-frequency oscillations arise, is divided into seven approximately equal volumes using antipulsation partitions consisting of nozzles protruding beyond the fire bottom, which are loosely adjacent to each other along their cylindrical generatrices. Due to this, the natural vibration frequencies in the volumes between the partitions sharply increase, shifting far from the resonant frequencies of the combustion chamber structure. In addition, the protruding nozzles stretch the combustion zone, which also reduces the possibility of high frequency phenomena. Gaps between protruding nozzles that do not fit tightly together provide additional damping effect.
The part of the nozzle protruding beyond the fire bottom is cooled by the fuel passing through the spiral channels (screw swirler) 6 of the inner sleeve.
The rest of the nozzles are buried in the fire bottom (their outlet cavities 4 go out into conical bores 5 in the fire bottom 7) and are made with different hydraulic resistance when fuel is supplied with a division according to the mass flow rate of fuel into three groups with the possibility of providing a difference in fuel consumption between each group from 3% up to 10% at nominal mode. In this case, the nozzles (except for those located near the fire wall of the chamber) are fixed in the fire bottom and the middle bottom so that nozzles from different groups are adjacent to each other by cyclical sequential spiral repetition of the arrangement of the nozzles from the first to the last group.
The introduction of injectors with different flow rates is necessary in order to reduce the effects of high-frequency vibrations at engine operating conditions.
Fig. 6.2 Arrangement of nozzles on the mixing head (images are enlarged),
Each of the two chambers is equipped with a swing unit. The thrust force is transmitted from the camera to the power frame through the gimbal. The generator gas triggered by the turbine is supplied to the compressor station through a 12-layer composite bellows placed inside the gimbal. The bellows is armored with special rings and is cooled by a small amount of cold oxygen flowing between the inner surface of the bellows and the thin inner wall.
Fig. 7. Appearance swing unit
 Fig. 8. Swing unit diagram | The swing unit consists of support rings 9 and 10, which are respectively hermetically connected to the combustion chamber and the gas duct (outlet from the turbine), in which there are consumption elements of external flow cooling 11 and 12, also shown in the view A... The bellows 13 is located inside the cardan ring 14. The cardan ring 14 through the hinges 15, forming two pivot axes, is connected by the power brackets 16 and 17 with the support rings 9 and 10. Inside the bellows 13 there are two shells 18 and 19, each of which is a body of revolution and cantilevered, respectively, to one of the said support rings, and the free end of the shell 18 is made in the form of a nipple with a spherical end 20 and is installed with a gap a in the shell 19. The center of the sphere of the nipple with a spherical end 20 is located on the rocking axis of the chamber. The size of the specified gap is chosen so as to ensure the flow rate of the cooling working fluid (oxidizer) required for reliable cooling of the bellows 13. |
The bellows 13 is made multilayer and is equipped with protective rings 21 inserted between the corrugations 22 of the bellows 13. Outside the protective rings 21 there is a tightly adhering casing 23 made of layers of cylindrical spirals 24 connected at the ends with the support rings 9 and 10 of the bellows assembly. Adjacent layers of spirals are adjacent to each other, and their turns are wound in opposite directions.
Installation of a metal power casing in the form of a metal cylindrical spiral outside the protective rings 21 of the bellows 13 increases its strength properties and at the same time limits the spontaneous bending of the bellows 13 when the engine chamber is rotated at relatively large angles (10-12 °), thereby increasing its stability.
The turbopump unit is made according to a single-shaft scheme and consists of an axial single-stage reactive turbine, a single-stage screw centrifugal oxidizer pump and a two-stage screw centrifugal fuel pump (the second stage is used to supply part of the fuel to the gas generators).
 |
| Figure 10.2. TNA rotor configuration |
 |
| Figure 10.3. Sectional diagram of the THA rotor |
On the main shaft with the turbine there is an oxidizer pump, coaxial with which two stages of the fuel pump are located on the other shaft. The shafts of the oxidizer and fuel pumps are connected by a toothed spring to relieve the shaft from thermal deformations arising from the large temperature difference between the working bodies of the pumps, as well as to prevent the fuel from freezing.
To protect the angular contact shaft bearings from excessive loads, effective auto-unloading devices are used.
The turbine is an axial single-stage reactive turbine.
To prevent ignition due to breakdowns of structural elements or friction of rotating parts against stationary ones (due to the selection of gaps from deformations or work hardening on the mating surfaces from vibration), the gap between the blades of the nozzle apparatus and the rotor is made relatively large, and the edges of the blades are relatively thick.
To exclude fire and destruction of parts of the gas path of the turbine, nickel alloys are used in the design, including heat-resistant ones for hot gas lines. The stator and exhaust tract of the turbine are forcibly cooled with cold oxygen. In places of small radial or end clearances, various kinds of heat-shielding coatings are used (nickel for the rotor and stator blades, sintered for the rotor), as well as silver or bronze elements that exclude ignition even with possible contact with rotating and stationary parts of the turbo pump unit.
To reduce the size and mass of foreign particles that can lead to a fire in the gas path of the turbine, a filter with a cell of 0.16x0.16 mm is installed at the engine inlet.
The high pressure of liquid oxygen and, as a consequence, increased combustion rate caused the design features of the oxidizer pump.
So, instead of floating O-rings on the impeller collars (usually used on less powerful HPA), stationary gap seals with a silver lining are used, since the process of "floating" of the rings is accompanied by friction at the contact points of the impeller with the casing and can lead to a pump fire.
The auger, impeller and torus outlet require particularly careful profiling, and the rotor as a whole needs special measures to ensure dynamic balance during operation. In the opposite case, due to large pulsations and vibrations, pipelines are destroyed, fires in joints due to mutual movement of parts, friction and work hardening.
To prevent ignition due to breakdowns of structural elements (auger, impeller and guide vane blades) under dynamic loading with subsequent ignition due to rubbing of debris, such means were used as an increase in design perfection and strength due to geometry, materials and cleanliness of mining, and also the introduction of new technologies: isostatic pressing of cast blanks, the use of granular technology and other types.
Fig. 11. Oxidizer pump impeller made of granules
nickel alloy EP741NP with mechanically untreated
hydrodynamic path.
The oxidizer booster pump consists of a high-pressure screw and a two-stage gas turbine, which is driven by an oxidizing gas taken after the main turbine with its subsequent bypass to the inlet to the main pump.
 Fig. 12. Simplified diagram of an oxidizer booster pumping unit (the image is enlarged). | The composite casing, consisting of flanged housings 1 and 2, has a bushing 4 fixed on the load-bearing ribs 3, the inner cavity of which is closed by a fairing 5. Inside the bushing 4 there is a ball bearing 6, seated on the pump impeller, made in the form of an auger 7. Fairing 5 the liner 8 installed in the bushing 4 is pressed in. The liner 8 has holes 9 that communicate the cavity of the liner 8 with the high pressure channel 10. The body 2 contains a fairing 11, fixed in it with the help of straightening blades 12. In this fairing, a ball bearing 13 is installed, fixed with a nut 14 on the auger 7. The auger has blades 15. Along these blades, the auger is inserted into the turbine impeller 16 (which actually consists from two stages, and not from one, as shown in the simplified diagram) and welded with it, i.e. the turbine impeller is fixed on the peripheral part of the pump impeller. The turbine impeller has profiled blades 17, the inter-blade spaces of which are communicated by nozzles in the nozzle apparatus with the inlet manifold. The supply of combustion products with an excess of oxygen is carried out through the inlet branch pipe 18. The outlet cavity of the turbine, made in the housing 2 in the form of an annular cylindrical cavity, is communicated by channels 19 with a conical annular branch pipe 20, which is connected with the cylindrical outlet 22 by openings 21. |
During the operation of the LLLW, liquid oxygen is supplied to the pump inlet (shown by an arrow), and the combustion products with an excess of oxygen taken from the gas duct after the turbine of the main HPA (see ASG in Fig. 2) are fed to the turbine inlet (shown by the arrow). The combustion products then fall on the profiled turbine blades 17, providing liquid oxygen supply by the screw 7. Behind the turbine, the combustion products through the holes 19 enter the cavity of the branch pipe 20, and then through the holes 21 to the pump outlet, where they are mixed with liquid oxygen and condensed. To solve the problem of the occurrence of low-frequency pulsations during gas condensation, the splitting of the gas dumping stream was used.
The unloading of the auger 7 from the action of axial forces is provided by the supply of high pressure liquid oxygen (see Fig. 2.2) through the high pressure channel 10 into the high pressure cavity of the auto unloading device. In the place of a small gap between the impeller and the body in the high-pressure cavity of the auto-unloading device, a silver lining is used, which prevents ignition from possible contact.
A "hot gas" valve (45 in Fig. 2.1) is installed in the line for supplying combustion products to the BNAO turbine, which operates under conditions of oxygen generator gas with a high temperature and at high pressure.
The fuel booster pump consists of a high-pressure auger and a single-stage hydraulic turbine powered by kerosene taken after the main pump.
Structurally, the fuel booster pump is similar to the oxidizer booster pump with the following differences:
- a single-stage hydraulic turbine runs on fuel taken from the fuel pump outlet of the main HPA;
- The discharge of high-pressure fuel for unloading the auger from the axial actions is made from the inlet manifold of the BNAG turbine.
A single-zone gas generator producing gas with an excess of oxidizer for driving a turbine consists of a body of a brazed-welded structure with a spherical outer shell and an outlet pipe rigidly connected to it, a cylindrical fire chamber with a diameter of 300 mm and a mixing head equipped with two-component and two-stage oxidizer nozzles, design which is made with a combustion zone and a gas ballasting zone inside the nozzles. In fact, each nozzle forms, together with the channel of the thick-walled fire bottom, in which it is located, an individual two-zone gas generator. As a result, the uniformity of the temperature field along the cross-section of the total gas flow formed by such nozzles is ensured at a high flow rate.
Fig. 13. Gas generator diagram, (the image is enlarged):
1 - spherical force shell; 2 - outlet branch pipe; 3 - cover; 4 - bushing; 5 - fire bottom; 6 - through chambers in the fire bottom; 7 - oxidizer cavity; 8 - spacer (outer wall of the fire chamber); 9 - annular cavity; 10 - shell (inner wall) of the fire chamber; 11 - fire chamber; 12 - mixing module (nozzle); 13 - housing of the mixing module; 14 - fuel channel; 15 - annular oxidizer channel; 16 - mixing chamber; 17 - fuel supply pipe; 18 - fuel cavity; 19 - oxidizer inlet branch pipe; 20 - windows in bushing 4; 21 - tangential holes for the oxidant supply; 22 - grooves on the outer surface of the nozzle body; 23 - calibrated fuel supply channels; 25 - tangential fuel supply holes; 26 - tapered bores; 27 - cooling cavity; 28 - channels forming a cooling cavity; 29 - holes for supplying the oxidizer to the cooling cavity; 30 - annular slit of the exit of the oxidant from the cooling cavity.
During operation of the gas generator, the fuel from the nozzle 17 fills the cavity 18 and is fed through the calibrated channels 23 and tangential holes 25 into the channels 14 and further into the mixing chambers 16. The oxidizer is fed through the nozzle 19 into the annular cavity 9, through the windows 20 it fills the cavity 7. Part of the oxidizer through the tangential holes 21 enter the mixing chamber 16, where, mixing with the fuel, it causes it to ignite. Through the slots 22, the oxidant is also fed into the chamber 6, providing mixing of the high-temperature combustion products. Further, in the fire chamber 11, the high-temperature combustion products are cooled with simultaneous evaporation of the liquid and heating of the gaseous oxidizer. At the outlet of the gas generator, an oxidizer is added to the gas generation products supplied through the annular slot 30.
The gas generator provides an oxidizing gas at the outlet in a wide temperature range (from 190 to 600 ° C), which makes it possible to regulate the engine thrust from 40 to 105% of the nominal value.
Unlike the prototype (RD-170), in which the connection of the body and the mixing head is carried out using a split flange, in the RD-180 welded joints of the body and the mixing head are used. However, at the stage of development, serial units from the RD-171 were widely used, which can be seen in some of the published photographs.
To ensure an acceptable level of temperature stresses in the bearing body parts, the gas ducts between the gas generators, the turbine and the chambers are cooled with oxygen.
To prevent ignition in gas ducts, rocking units of the mixing head of the chamber, oxidizer valve, increased (compared to less powerful engines) requirements for the cleanliness of gas paths and prevention of the presence of organic substances are set.
 |
 | The ampoule contains a body 1 with inlet 2 and outlet 3 nozzles of the membrane assemblies 4 and 5 installed inside the body 1, and means for refueling the body with starting fuel 6. Each membrane assembly 4, 5 contains a piston 7, which can be made in one piece with the membrane 8 or in which the membrane 8 is sealed to its outer surface. The piston 7 is installed in the housing guide 9 along a sliding fit. The peripheral portion of the membrane 8 is hermetically welded to the body 1 under the guide 9. The piston 7 is connected to the shank 10, which can be cylindrical or of any other shape and is located in the sleeve 11. The sleeve 11 is attached to the ampoule body 1 on the brackets 12. The sleeve 11 has a spring clip 13, for example made in the form of a spring ring, and the shank 10 is made with an annular groove 14. When the diaphragm assembly is triggered, the spring lock 13 restricts the movement of the shank 10. The shank 10 is made with holes 15 for venting gas from the stagnant zone when filling the ampoule. The membrane 8 from the side of the inlet 2 is made thin in the form of an annular bridge 16, which is ruptured when interacting with the working medium at diameter D. Dimension D is slightly smaller than the diameter of the piston 7. At the junction of the membrane 8 with the piston 7, it is made with a smaller thickness in order to exclude seizure marks during the movement of the piston 7 in the guide 9 of the housing 1. |
| Fig. 14. Diagram of an ampoule with starting fuel (the image is enlarged). |
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The design includes a means for filling the housing with starting fuel 6, which is installed in the partition 17 of the housing 1 and consists of two plugs - a filling plug 18 and a drain plug 19, which are installed in the filling 20 and drain 21 channels, respectively. Each of the plugs has a screw plug 22, a sealing plug 23, a gasket 24, and a nut 25. The screw plug 22 has a flow hole 26.
The ampoule is filled with starting fuel as follows. On the assembled ampoule, before installing the nuts 25 and sealing plugs 23, screw plugs 22 are not completely screwed in so that the opening of the flow section of the filling 20 and drain 21 channels through the hole 26 is provided. the cavity of the body 1 between the membrane assemblies 4 and 5, and then through the drain channel to the drain. After filling the ampoules, screw in the threaded plugs 22 until they stop, after which the starting fuel is drained before the threaded plug 22 of the filling plug 18 and after the threaded plug 22 of the drain plug 19. After that, the sealing plugs 23, the sealing gaskets 24 and the nuts 25 are installed. After that, the ampoule is ready. for installation on a rocket engine. A gas cushion is formed in the inner cavity of the ampoule in the housing 1 between the membranes 8 as a result of the assembly and filling of the ampoule. The presence of a gas cushion helps to ensure the reliability of the ampoule during storage and effective movement with the acceleration of the piston 8 when the medium pressure is applied to the ampoule inlet.
The device works as follows. When the high-pressure component acts on the inlet side of the diaphragm assembly 4, the membrane 8 is deformed, and then destruction along the circumference D. With uneven destruction of the membrane 8, with the appearance of a leak, the pressure in front of the piston 7 does not drop, due to the operation of the throttling gap formed by the housing guide 9 and piston 7, piston 7 continues to move, and after complete destruction of membrane 8, it accelerates. The movement of the piston 7 with acceleration is provided due to the presence of an effort from a differential pressure acting on the surface area determined by the diameter D.
The length "A", at which the piston moves with acceleration and the gap between the piston 7 and the guide 9 are chosen such as to ensure a guaranteed shearing of the membrane 8 along the entire perimeter, the required delay in opening the flow section of the line after the membrane 8 is cut, the acceleration of the piston 7, which is necessary for operation spring retainer 13. The dimensions of the diaphragm bridges 8 are determined on the basis of a given pressure, which ensures the destruction of the bridging.
Further, the moving shank 10 along the flow is fixed by means of a spring lock 13, while the hydraulic characteristics of the open diaphragm assembly 4 are reproduced with high accuracy, since there are no structural elements with an undefined position in the flow of the component.
After opening the diaphragm unit 4 due to the increased pressure of the starting fuel, the diaphragm unit 5 opens in the same way.
The starting tank is designed to create the pressure required to break through the membranes of ampoules with starting fuel.
Fig. 15. Starting tank diagram
The starting tank contains a power shell 1, made in the form of a hemisphere, and a tubular flange 2, mated at its end with the end of the power shell 1. The tubular flange 2 is located along the longitudinal axis of the said hemisphere of the power shell 1 and an annular groove 3 is made on its inner surface. 4 for filling and dispensing liquid is installed in the load-bearing shell 1. The pressure ring 5 is located coaxially with the longitudinal axis of the load-bearing shell 1. The elastic diaphragm 6 is fixed between the tubular flange 2 and the pressure ring 5 and is made in the form of a hemisphere mated with the cylinder on the outer surface at the base of which the end protrusion 7 is made, located in the annular groove 3 of the tubular flange 2. The outer surface of the clamping ring 5 and the inner surface of the tubular flange 2 at the location of the end protrusion 7 in the annular groove 3 are cylindrical. The device has a bottom 8, made in the form of a part of a sphere, with the possibility of its end impact on the end of the clamping ring 5 and a hermetic connection with the tubular flange 2 of the power shell 1. The connection 9 for supplying the control gas is installed in the bottom 8. A thin-walled ring 10 is introduced into the structure. on which the collar 11 is made and which is installed between the clamping ring 5 and the elastic diaphragm 6 at the location of its annular protrusion 7.
The divider 16 is made in the form of a plate perforated with holes 21, the edges of which are attached to the inner surface of the bottom 8 in the cavity 14 connected to the connecting piece 9 for supplying the control gas. The divider 16 with holes 21 serves to uniformly influence the gas flow on the elastic diaphragm 6.
The device works as follows (see also section). Through the nozzle 4, the tank is filled with the main fuel, while the elastic diaphragm 6 is shifted to the bottom 8. Then the control gas is supplied through the nozzle 9, under the action of which the diaphragm 6 is shifted to its original position, displacing the main fuel through the nozzle 4.
Thanks to the adopted design of the attachment point for the end section of the elastic diaphragm at high pressure, tightness is ensured with reusable transfers (more than 450), and it is possible to bend the elastic shell practically without stretching it.
 |
 | The container is intended for transportation of the engine, while the container includes a frame, a transverse power rack fixed to it, and attachment points mounted on it with a transportable rocket engine, which in the container is cantilevered on a transverse power rack. The transverse power stand is made in the form of a transport ring, and the container is equipped with a means of installing and fixing this ring on the frame in a vertical or deviated from the vertical at an angle of not more than 10 °, and the fastening of this ring on the frame is carried out using lanyards, and the frame and the transport ring equipped with fastening elements to the end sections of lanyards. The overall dimensions of the container are 4.6 x 3.67 x 3.0 m, the weight with the engine is about 9 tons. |
 |
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| Fig. 16. Shipping container (enlarged image). | ||
- Katorgin B.I.Prospects for the creation of powerful liquid-propellant rocket engines
- George P. Sutton "History of Liquid Propellant Rocket Engines"
- Prospect NPO Energomash
- Description of the invention to the patent of the Russian Federation RU 2159351. Gas generator (US Patent 6244040. Video (size 46 MB, duration 6 min. 52 sec.)
- Description of the invention to the patent of the Russian Federation RU 2106534. Booster turbopump unit.
- Description of the invention to the patent of the Russian Federation RU 2159353. Ampoule with starting fuel for ignition of LPRE fuel components.
- Description of the invention to the patent of the Russian Federation RU 2158699. Tank for storage and displacement of liquid.
In the US, it is badly about regaining the status of a "great space power" by abandoning the Russian RD-180 rocket engines.
Many are worried that the withdrawal of American military satellites depends on the goodwill of the Russians.
On this occasion, an interesting dualism arose in the states:
The US Air Force and ULA are asking Congress to authorize the delivery of the RD-180 to the United States, and Senator John McCain is categorically forbidding Congress to allow it.
In the end, all the same Congress lifted the ban- apparently, while the US Air Force turned out to be more convincing than the downed American pilot who threatened to vote against the budget (because of the RD-180).
:)
At the same time, the report of a special Pentagon commission under the leadership of retired US Air Force Major General Howard Mitchell came to the disposal of the American media, in which he noted that without the RD-180 space launches of military satellites after 2016 would be disrupted. Transferring launches from Atlas V rockets powered by Russian engines to Delta IV rockets (powered by RS-68 liquid propellant rocket engines) will still result in significant delays and potential losses that could amount to $ 5 billion.
Have you forgotten about the astronauts who may not get a return ticket from the ISS?
They also fly by Soviet Russian "Soyuz".
Reference:
RD-180 is produced by NPO Energomash named after academician V.P. Glushko since 1999.
Why can't the Americans make the RD-180?
P.P.S.
United Launch Alliance will purchase 20 more RD-180 engines
A joint venture between Lockheed Martin Corp and Boeing Co - United Launch Alliance, has ordered 20 additional Russian RD-180 rocket engines.
Customer representative Jessica Roj clarified that The new batch will start shipping immediately after the previous order for 29 engines is completed, - reports Reuters.
Russian engines will be used on US Atlas-5 missiles until the US develops and certifies its own new engine. RD-180s are used in the first stage of American missiles.
The House of Representatives of the US Congress in December 2014, as an anti-Russian measure on events in Ukraine, adopted an amendment by Senator John McCain, which provides for complete abandonment of the RD-180 rocket engines by the United States until 2019... An exception is made for the contract concluded by the Boeing and Lockheed Martin (ULA) consortium with the Russian NPO Energomash until 2019. At the same time, it was reported that Congress allocated $ 220 million for the development of new American engines.
220 million "cut" is clearly not enough, as we have already seen above.
Academician Boris Katorgin, the creator of the world's best liquid-propellant rocket engines, explains why the Americans still cannot repeat our achievements in this area and how to keep the Soviet head start in the future.
On June 21, at the St. Petersburg Economic Forum, the Global Energy Prize winners were awarded. An authoritative commission of industry experts from different countries selected three applications out of 639 submitted and named the winners of the 2012 prize, which is already commonly called the “Nobel Prize for Power Engineers”. As a result, 33 million premium rubles this year were shared by the famous inventor from Great Britain, Professor Rodney John Allam, and two of our outstanding scientists - Academicians of the Russian Academy of Sciences Boris Katorgin and Valery Kostyuk.
All three are related to the creation of cryogenic technology, the study of the properties of cryogenic products and their application in various power plants. Academician Boris Katorgin was awarded "for the development of highly efficient liquid-propellant rocket engines on cryogenic fuels, which provide reliable operation of space systems with high energy parameters for the peaceful use of space." With the direct participation of Katorgin, who devoted more than fifty years to the OKB-456 enterprise, now known as NPO Energomash, liquid-propellant rocket engines (LRE) were created, the performance of which is still considered the best in the world. Katorgin himself was engaged in the development of schemes for organizing the working process in engines, mixture formation of fuel components and elimination of pulsation in the combustion chamber. Also known are his fundamental work on nuclear rocket engines (NRE) with a high specific impulse and developments in the field of creating powerful continuous chemical lasers.
In the most difficult times for Russian science-intensive organizations, from 1991 to 2009, Boris Katorgin headed NPO Energomash, combining the positions of General Director and General Designer, and managed not only to keep the company, but also to create a number of new engines. The lack of an internal order for engines forced Katorgin to look for a customer in the external market. One of the new engines was the RD-180, developed in 1995 specifically for participation in a tender organized by the American corporation Lockheed Martin, which chose a liquid-propellant engine for the Atlas carrier rocket being upgraded at that time. As a result, NPO Energomash signed a contract for the supply of 101 engines, and by the beginning of 2012 had already delivered more than 60 rocket engines to the United States, 35 of which had been successfully operated on Atlas when launching satellites for various purposes.
Before the awarding of the award, "Expert" talked with academician Boris Katorgin about the state and prospects of development of liquid-propellant rocket engines and found out why engines based on developments of forty years ago are still considered innovative, and RD-180 could not be recreated at American factories.
Boris Ivanovich, what exactly is your merit in the creation of domestic liquid-propellant jet engines, which are now considered the best in the world?
To explain this to a layman, you probably need a special skill. For liquid-propellant rocket engines, I developed combustion chambers, gas generators; in general, he supervised the creation of the engines themselves for the peaceful exploration of outer space. (In the combustion chambers, the fuel and oxidizer are mixed and burned, and a volume of hot gases is formed, which, then ejected through the nozzles, create the actual jet thrust; gas generators also burn the fuel mixture, but already for the operation of turbo pumps, which pump fuel and oxidizer under enormous pressure into the same combustion chamber. - "Expert".)
You are talking about peaceful space exploration, although it is obvious that all engines with thrust from several tens to 800 tons, which were created at NPO Energomash, were intended primarily for military needs.
We did not have to drop a single atomic bomb, we did not deliver a single nuclear charge on our missiles to the target, and thank God. All military developments went into peaceful space. We can be proud of the huge contribution of our rocket and space technology to the development of human civilization. Thanks to astronautics, whole technological clusters were born: space navigation, telecommunications, satellite television, and sensing systems.
The engine for the R-9 intercontinental ballistic missile, on which you worked, then formed the basis of almost all of our manned program.
Back in the late 1950s, I carried out computational and experimental work to improve mixture formation in the combustion chambers of the RD-111 engine, which was intended for that very rocket. The results of the work are still used in the modified RD-107 and RD-108 engines for the same Soyuz rocket; about two thousand space flights were performed on them, including all manned programs.
Two years ago, I interviewed your colleague, Global Energy Laureate Academician Alexander Leontyev. In a conversation about specialists closed to the general public, which Leontyev himself once was, he mentioned Vitaly Ievlev, who also did a lot for our space industry.
Many academics who worked for the defense industry were classified - this is a fact. Now a lot has been declassified - this is also a fact. I know Alexander Ivanovich very well: he worked on the creation of calculation methods and methods for cooling the combustion chambers of various rocket engines. It was not easy to solve this technological problem, especially when we began to squeeze out the chemical energy of the fuel mixture as much as possible to obtain the maximum specific impulse, increasing, among other measures, the pressure in the combustion chambers to 250 atmospheres. Let's take our most powerful engine - RD-170. Fuel consumption with an oxidizing agent - kerosene with liquid oxygen flowing through the engine - 2.5 tons per second. Heat flows in it reach 50 megawatts per square meter - this is a huge energy. The temperature in the combustion chamber is 3.5 thousand degrees Celsius. It was necessary to come up with a special cooling for the combustion chamber so that it could work calculated and withstand the thermal head. Alexander Ivanovich did just that, and, I must say, he did an excellent job. Vitaly Mikhailovich Ievlev - Corresponding Member of the Russian Academy of Sciences, Doctor of Technical Sciences, professor, who, unfortunately, died quite early, - was a scientist of the broadest profile, possessed an encyclopedic erudition. Like Leontiev, he worked a lot on the methodology for calculating high-stress thermal structures. Their work somewhere intersected, somewhere they were integrated, and as a result, an excellent method was obtained by which it is possible to calculate the heat intensity of any combustion chambers; now, perhaps, using it, any student can do it. In addition, Vitaly Mikhailovich took an active part in the development of nuclear, plasma rocket engines. Here our interests intersected in the years when Energomash was doing the same.
In our conversation with Leontyev, we touched upon the sale of the RD-180 energomashevsky engines in the USA, and Alexander Ivanovich said that in many ways this engine is the result of developments that were made just during the creation of the RD-170, and in a sense, its half ... Is this really the result of backscaling?
Any engine in a new dimension is, of course, a new apparatus. RD-180 with a thrust of 400 tons is actually half the size of the RD-170 with a thrust of 800 tons. The RD-191, designed for our new Angara rocket, has a thrust of 200 tons. What do these engines have in common? All of them have one turbo pump, but the RD-170 has four combustion chambers, the "American" RD-180 has two, and the RD-191 has one. Each engine needs its own turbo pump unit - after all, if the four-chamber RD-170 consumes about 2.5 tons of fuel per second, for which a turbo pump with a capacity of 180 thousand kilowatts was developed, which is more than two times higher than, for example, the reactor power of the atomic icebreaker "Arktika" , then the two-chamber RD-180 - only half, 1.2 tons. In the development of turbo pumps for the RD-180 and RD-191, I participated directly and at the same time led the creation of these engines as a whole.
The combustion chamber, then, is the same on all these engines, only their number is different?
Yes, and this is our main achievement. In one such chamber with a diameter of only 380 millimeters, a little more than 0.6 tons of fuel per second is burned. Without exaggeration, this camera is a unique high-heat-stress equipment with special belts to protect against powerful heat fluxes. Protection is carried out not only due to external cooling of the chamber walls, but also due to an ingenious method of "lining" a fuel film on them, which evaporates and cools the wall. On the basis of this outstanding camera, which has no equal in the world, we manufacture our best engines: RD-170 and RD-171 for Energia and Zenit, RD-180 for the American Atlas and RD-191 for the new Russian missile "Angara".
- "Angara" was supposed to replace "Proton-M" several years ago, but the creators of the rocket faced serious problems, the first flight tests were repeatedly postponed, and the project seems to continue to stall.
There were indeed problems. A decision has now been made to launch the rocket in 2013. The peculiarity of the Angara is that, on the basis of its universal rocket modules, it is possible to create a whole family of launch vehicles with a payload capacity of 2.5 to 25 tons to launch cargo into low-earth orbit on the basis of the RD-191 universal oxygen-kerosene engine. Angara-1 has one engine, Angara-3 - three with a total thrust of 600 tons, Angara-5 will have 1000 tons of thrust, that is, it will be able to put more cargo into orbit than Proton. In addition, instead of the very toxic heptyl, which is burned in the Proton engines, we use environmentally friendly fuel, after which only water and carbon dioxide remain.
How did it happen that the same RD-170, which was created back in the mid-1970s, still remains, in fact, an innovative product, and its technologies are used as the basis for new rocket engines?
A similar thing happened with an aircraft created after World War II by Vladimir Mikhailovich Myasishchev (a long-range strategic bomber of the M series, developed by the Moscow OKB-23 of the 1950s - "Expert"). In many respects, the aircraft was thirty years ahead of its time, and the elements of its design were then borrowed by other aircraft manufacturers. So it is here: in the RD-170 there are a lot of new elements, materials, design solutions. According to my estimates, they will not become obsolete for several more decades. This is the merit, first of all, of the founder of NPO Energomash and its general designer Valentin Petrovich Glushko and Corresponding Member of the Russian Academy of Sciences Vitaly Petrovich Radovsky, who headed the company after Glushko's death. (Note that the world's best energy and performance characteristics RD-170 is largely provided due to Katorgin's solution to the problem of suppressing high-frequency combustion instability through the development of anti-pulsation baffles in the same combustion chamber. - "Expert.") And the first stage RD-253 engine for the Proton launch vehicle? Introduced back in 1965, it is so perfect that it has not yet been surpassed by anyone. This is how Glushko taught to design - at the limit of the possible and always above the world average. It is also important to remember another thing: the country has invested in its technological future. How was it in the Soviet Union? The Ministry of General Machine Building, which, in particular, was in charge of space and rockets, spent 22 percent of its huge budget on R&D alone - in all areas, including propulsion. Today, research funding is much less, and that says a lot.
Doesn't the achievement of some perfect qualities by these liquid-propellant rocket engines, and this happened half a century ago, that a rocket engine with a chemical energy source is in some sense outdated: the main discoveries have been made in new generations of rocket engines, now we are talking more about the so-called supporting innovations?
Certainly not. Liquid-propellant rocket engines are in demand and will be in demand for a very long time, because no other technology is able to more reliably and economically lift a load from Earth and put it into low-Earth orbit. They are safe from an environmental point of view, especially those that run on liquid oxygen and kerosene. But for flights to stars and other galaxies, liquid-propellant rocket engines, of course, are completely unsuitable. The mass of the entire metagalaxy is 10 to 56 degrees of grams. In order to accelerate on a rocket engine to at least a quarter of the speed of light, an absolutely incredible amount of fuel will be required - 10 to 3200 grams, so even thinking about it is stupid. The liquid-propellant engine has its own niche - sustainer engines. On liquid engines, you can accelerate the carrier to the second cosmic speed, fly to Mars, and that's it.
The next stage - nuclear rocket engines?
Of course. It is not known whether we will live to see some of the stages, but much has been done for the development of nuclear-powered rocket engines already in Soviet times. Now, under the leadership of the Keldysh Center, headed by Academician Anatoly Sazonovich Koroteev, the so-called transport and energy module is being developed. The designers came to the conclusion that it is possible to create a gas-cooled nuclear reactor that is less stressful than it was in the USSR, which will work both as a power plant and as a source of energy for plasma engines when traveling in space. Such a reactor is currently being designed at the NIKIET named after N. A. Dollezhal under the leadership of Corresponding Member of the Russian Academy of Sciences Yuri Dragunov. The Kaliningrad design bureau Fakel also participates in the project, where electric propulsion engines are being created. As in Soviet times, it will not do without the Voronezh Design Bureau of Chemical Automatics, where gas turbines, compressors in order to drive a coolant - a gas mixture in a closed loop.
In the meantime, are we going to the rocket engine?
Of course, we also clearly see the prospects for the further development of these engines. There are tactical, long-term tasks, there is no limit here: the introduction of new, more heat-resistant coatings, new composite materials, a decrease in the mass of engines, an increase in their reliability, and a simplification of the control scheme. A number of elements can be introduced to better control the wear of parts and other processes occurring in the engine. There are strategic tasks: for example, the development of liquefied methane and acetylene as fuel together with ammonia or three-component fuel. NPO Energomash is developing a three-component engine. Such a liquid-propellant rocket engine could be used as an engine for both the first and second stages. At the first stage, it uses well-developed components: oxygen, liquid kerosene, and if you add about five percent more hydrogen, then the specific impulse will significantly increase - one of the main energy characteristics of the engine, which means that more payload can be sent into space. At the first stage, all the kerosene with the addition of hydrogen is produced, and at the second, the same engine switches from operation on three-component fuel to two-component fuel - hydrogen and oxygen.
We have already created an experimental engine, albeit of a small dimension and a thrust of only about 7 tons, carried out 44 tests, made full-scale mixing elements in the nozzles, in the gas generator, in the combustion chamber and found out that you can first work on three components, and then smoothly switch to two. Everything is working out, a high combustion efficiency is achieved, but in order to go further, a larger sample is needed, the benches need to be refined in order to launch the components that we are going to use in a real engine into the combustion chamber: liquid hydrogen and oxygen, as well as kerosene. I think this is a very promising direction and a big step forward. And I hope to have time to do something during my lifetime.
Why the Americans, having received the right to reproduce the RD-180, have not been able to make it for many years?
Americans are very pragmatic. In the 1990s, at the very beginning of their work with us, they realized that in the energy field we were far ahead of them and we had to adopt these technologies from us. For example, our RD-170 engine in one start, due to a higher specific impulse, could take out a payload two tons more than their most powerful F-1, which meant at that time $ 20 million in gain. They announced a competition for a 400-ton engine for their Atlases, which was won by our RD-180. Then the Americans thought that they would start working with us, and in four years they would take our technologies and reproduce them themselves. I told them at once: you will spend more than a billion dollars and ten years. Four years have passed, and they say: yes, six years are needed. More years have passed, they say: no, we need another eight years. Seventeen years have passed, and they have not reproduced a single engine. They now need billions of dollars for bench equipment alone. At Energomash we have stands where the same RD-170 engine can be tested in a pressure chamber, the jet power of which reaches 27 million kilowatts.
- I heard right - 27 gigawatts? This is more than the installed capacity of all Rosatom NPPs.
Twenty-seven gigawatts is jet power that develops in a relatively short time. During tests on the stand, the energy of the jet is first extinguished in a special pool, then in a diffusion pipe with a diameter of 16 meters and a height of 100 meters. It takes a lot of money to build a stand like this, which can house an engine that generates such power. The Americans have now given up on this and are taking the finished product. As a result, we are not selling raw materials, but a product with a huge added value, in which highly intellectual labor is invested. Unfortunately, in Russia this is a rare example of high-tech sales abroad in such a large volume. But this proves that for correct setting question we are capable of a lot.
- Boris Ivanovich, what should be done in order not to lose the head start gained by the Soviet rocket engine building? Probably, in addition to the lack of funding for R&D, another problem is also very painful - personnel?
To stay on the world market, you have to go forward all the time, create new products. Apparently, until the end of us was pressed down and the thunder struck. But the state needs to realize that without new developments it will find itself on the margins of the world market, and today, in this transitional period, while we have not yet grown to normal capitalism, it must first of all invest in the new - the state. Then you can transfer the development for the release of a series to a private company on terms that are beneficial to both the state and business. I do not believe that it is impossible to come up with reasonable methods of creating something new, without them it is useless to talk about development and innovations.
There are personnel. I am the head of a department at the Moscow Aviation Institute, where we train both engine specialists and laser specialists. The guys are smart, they want to do the business they are learning, but you need to give them a normal initial impulse so that they do not leave, as many people do now, to write programs for distributing goods in stores. For this it is necessary to create an appropriate laboratory environment, to give a decent salary. Build the correct structure of interaction between science and the Ministry of Education. The same Academy of Sciences solves many issues related to personnel training. Indeed, among the current members of the academy, corresponding members, there are many specialists who manage high-tech enterprises and research institutes, powerful design bureaus. They are directly interested in the departments assigned to their organizations to bring up the necessary specialists in the field of technology, physics, chemistry, so that they immediately receive not just a specialized university graduate, but a ready-made specialist with some life and scientific and technical experience. It has always been this way: the best specialists were born in institutes and enterprises where educational departments existed. At Energomash and at NPO Lavochkin we have departments of the branch of the Moscow Aviation Institute "Kometa", which I am in charge. There are old cadres who can pass the experience on to the young. But there is very little time left, and the losses will be irrecoverable: in order to simply return to the current level, you will have to spend much more effort than is needed today to maintain it.
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