SpaceX interplanetary program: a detailed analysis of the Raptor LRE

Published on June 28, 2017

SpaceX interplanetary program: a detailed analysis of the Raptor LRE

    So, for 2017, SpaceX is perhaps the closest to sending something to Mars that is different from a probe or a rover. Moreover, the company's plans include quite massive manned expeditions to the Red Planet, which will ensure the long-term human presence on the fourth planet from the Sun. In addition, SpaceX is considering conducting research missions in those parts of the Solar System that even the heads of the most desperate romantics of the rocket industry have not thought about. But what technologies are behind these plans? Let's figure it out. And we begin with a review of the rocket engine, which should ensure the implementation of these very ambitious plans - the Raptor rocket engine.


    Bench tests LRE «Raptor», September 25, 2016. McGregor, TX.

    RRE Raptor: what kind of beast is this?


    So, the Raptor rocket engine is being developed by SpaceX as part of its flight program to distant objects of the solar system.

    The first truly massive engine of SpaceX was Merlin, which runs on an RP-1 / LOX pair. About this engine, it can be said that although it is the most efficient gas-generating engine on this fuel pair in the history of the United States, it has a record-breaking thrust-weight ratio as a whole, it is primarily focused on reliability, reusability and cheapness. It can be said that working on Falcon 9, the task was set first of all to roll up the reusability technology to the routine level, which in the end brought significant results.

    Indeed, saving an entire stage could potentially save a lot more money than reducing the mass of disposable units or increasing their efficiency when switching to new engines or fuel pairs. For example, at pH Union-Y2 as kerosene alternative to "Block A" (second stage) was used syntin , allowing to increase the maximum payload mass 200 kg, compared with the basic version rocket Union-U . Another example is the project of the Soviet lunar rocket UR-700, which proposed the use of completely exotic fuel pairs: it was proposed to replace the already extremely dangerous UDMH for the first stage of the RD-270 engine with an even more dangerous pentabor (B5H9) with an increase in the UD of the RD-270 by 42 s, and on the third step it was generally proposed to install an absolutely fantastic system, based on the complex complexity of operation and preparatory preparation, based on the liquid-fuel liquid / liquid fluoride fuel pair.


    The chemical formulas of fuels that were supposed to be used in RD-270 engines: on the left - NDMH (C2H8N2; blue balls - nitrogen atoms, black balls - carbon atoms, white balls - hydrogen atoms); on the right - pentaboran (B5H9; pink balls - boron atoms, white balls - hydrogen atoms). Both compounds are extremely toxic, pentaboran, in addition to everything, has a tendency to spontaneously ignite on contact with air even with minor pollution. In addition, UDMH and pentaboran are much more expensive than kerosene in production.

    Of course, if you do not have reusable rockets at your disposal, your payloads weigh a lot and the spaceports are far from the equator, then a reasonable conclusion suggests itself - you need to output the maximum possible masses per launch. However, keep in mind that high performance or novelty units can mean higher cost and in this case there is an excellent example: a long time in the third stage of "Soyuz" (so-called "block I") sets the engine RD-0110 ( thrust and turbulence in vacuum - 298 kN and 326 s, respectively). Then, starting with the modification Soyuz-2.1b, a new RD-0124 was put on “Block I”(thrust and turbulence in vacuum - 294.3 kN and an incredible 359 s, respectively). However, despite the fact that RD-0124 is the most highly efficient oxygen-kerosene rocket engine in the world and has a number of other advantages over its predecessor, the transition to the engine created in the 21st century involves a number of financial difficulties: firstly, its operation implies covering the costs of R & D (and the RD-0110 was created already in the 60s); secondly, he received his unique characteristics due to much greater material intensity. Therefore, in the end, it turns out that the RD-0124 is much more expensive than the RD-0110, and from all of this history the following conclusion suggests itself: in modern conditions, the creation of highly efficient disposable missile systems from zero can help in solving current problems but in general, this strategy is not very profitable, and for good it really makes sense to put expensive units on reusable rockets or at least separate reusable stages. And as we will see a little further, the Raptor LRE is designed with the help of an enormous amount of new technologies and modern engineering solutions.


    Third-stage engines of the Soyuz family of missiles: RD-0110 (left) and RD-0124 (right). Despite the external similarity in size and geometry, the RD-0124 is much more technological and young unit, which positively affects its characteristics and negatively affects the final cost of its operation.

    In general, SpaceX is famous for its thoughtful approach in matters of expenditure of funds, for that it is a young private company, and not a fat cumbersome corporation like Boeing or Lockheed and the like, who like to suck money from the state feederor state monopoly structure. Therefore, every step of SpaceX on the way to the goal has long been discussed and studied for a possible reduction in the cost of development, production and repeated operation, and it would be senseless to expect from this company projects to develop exotic units like the RD-301 LRE on fuel vapor “liquid ammonia / liquid fluorine” ”Which created a whole mountain of technological and medical-ecological problems. Just as it would be senseless to expect from SpaceX parallel large-scale work on the development of several missiles at once (as was the case during the Soviet lunar program - super-heavy carriers N-1 and UR-700 were developed in parallel) or engines on an extremely toxic pair UDMH / AT .


    LRE RD-301 (liquid ammonia / liquid fluorine) in the Museum of the Gas-Dynamic Laboratory (GDL) in St. Petersburg. By the way, a very interesting excerpt from the first volume of the three-volume book entitled “Selected Works of Academician V.P.Glushko” is posted on the Internet , in which the motives and prospects for creating engines with liquid fluorine as an oxidizer are discussed .

    I propose to begin the discussion of the Raptor LRE with the consideration of the main drawbacks of the RP-1 / LOX and LH2 / LOX fuel pairs, which should be considered when choosing fuel for a rocket engine:
    • For example, in some sense, a significant disadvantage of kerosene rocket is relatively low as compared with a cryogenic fuel specific pulse (337 in vacuo at RD-180 on a pair of RP-1 / LOX against working on a pair of LH2 / LOX RD-0120 with its 455 in vacuum (4 of these engines were installed at the second stage of the Energy rocket, the technology / technological chain of production of this unit was lost, according to some industry representatives ). In this case, the specific impulse can be crucial when it comes to the rocket, the starting weight Oui is thousands of tons;


      LRE RD-0120 (RSC Energia Museum), which was installed at the second stage of the Energia launch vehicle. LH2 / LOX was used as a fuel pair for this engine. The ability to produce the engine in the form in which it was installed on the Energia launch vehicle is currently lost.

    • Also, the use of kerosene implies the accumulation in engines of a greater amount of soot, which can increase the cost of maintaining a reusable engine or simply simply reduce its reliability or service life;

    • Another disadvantage of oxygen-kerosene engines is the fact that kerosene is prone to coking, which results in the need to supply excess oxygen oxygen to the combustion chambers to avoid formation of solid petroleum coke on the internal parts of the engines . This creates two difficulties at once if the developer’s goal is a reusable rocket: first, there is a need to clean the engines of petroleum coke before re-launches; secondly, the excess supply of oxygen into the combustion chamber accelerates corrosion processes and leads to wear of pumping systems.

    • Another disadvantage of kerosene is that it is impossible to find it anywhere except the Earth, therefore, in fact, the only way to fill interplanetary ships in the case of a kerosene engine is to send kerosene from the Earth. At the same time, kerosene by itself, although it has a high density (especially compared to hydrogen), would still be better to somehow learn how to deliver the most indispensable components of fuel synthesis to other planets from Earth, and to produce and extract the missing reagents on site landing an interplanetary ship. In addition, in the case of a long flight with kerosene on board, it may trivially lose its properties;

    • Finally, with all the advantages of liquid hydrogen (as mentioned above, the specific impulse of the LH2 / LOX pair in vacuum is about 35% higher than that of the RP-1 / LOX pair, in addition, the low molecular weight of molecular hydrogen reduces the engine wear rate, and the burning process virtually eliminates the accumulation of soot); its use is fraught with a number of difficulties:

      • The extremely low temperature of liquid hydrogen (about -253 degrees Celsius) makes it not the most convenient fuel;

      • Contact of hydrogen with metals leads to hydrogen embrittlement . High-strength steels and alloys of titanium and nickel are most susceptible to hydrogen embrittlement, which poses a danger to rockets, while the mechanism of hydrogen embrittlement has not yet been established, so it is not clear how to deal with it;

      • Despite the fact that hydrogen shows excellent specific impulses in vacuum, the LH2 / LOX vapor does not have similar high levels at sea level. For example, the specific impulse of the first stage hydrogen propulsion engine of the “Delta IV” RS-68A LV at sea level is 360 seconds, which is less than 12% higher than the similar indicator for a kerosene RD-180 - 311.3 seconds (I recall that in vacuum for hydrogen engines, superiority over kerosene engines in a specific impulse of 35% was achieved);

      • Finally, the LH2 / LOX pair has a catastrophically low density compared to the same kerosene: 0.29 g / cm ^ 3 for LH2 / LOX versus 1.03 g / cm ^ 3 for RP-1 / LOX, that is, it is more than three times less ! Of course, a higher specific impulse allows the use of less fuel and oxidizer in the case of LH2 / LOX, but it is not so great, so the use of LH2 / LOX inevitably leads to a very significant increase in the volume of fuel tanks. In the case of ITS LV, this would mean a transition from the already gigantic dimensions to completely unimaginable.


        Comparison of the size of some missile systems. It can be noted that in spite of approximately the same indicators of the load on the LEO at the Proton M (23 tons) and Delta IV Heavy (26 tons) and almost equal starting masses (705 tons for the Proton M and 723 tons in Delta IV Heavy, the use of the LH2 / LOX fuel pair on the Delta makes Proton M seem to be a midget compared to the American flying hydrogen monster.

        A simple example: the fully hydrogen "Delta IV Heavy" and the best friend of the Kazakh environmentalist RN Proton M working at UDMH / AT are able to take approximately the same cargo to the NOU (slightly less than 26 tons for Delta and about 23 tons for Proton) . At the same time, the fuel tanks of the “Delta IV Heavy” are so large that its structure essentially consists of as many as three first steps, each of which has a height of 40.8 meters. The height of the fully assembled Proton M LV is 58.2 meters. By the way, the “Delta IV Heavy” is also heavier than the “Proton M”: its starting mass is 732 tons, which is 27 tons more than the starting mass of the “Proton”. In general, as an intermediate result, it can be said that the existence of the benefits of using the LH2 / LOX pair in the first steps is a rather individual and debatable question.

    Such drawbacks of the LH2 / LOX pair have resulted in the fact that mainly hydrogen or stages fly overclocking blocks whose engines are turned on exclusively in a vacuum, an example being the one currently being developed at the State Space Research and Production Center. MV Khrunicheva hydrogen accelerating unit "KVTK" , which means "heavy oxygen-hydrogen class" (in the framework of the project to create an accelerating unit "KVTK" on the Voronezh Design Bureau of Chemical Automation , a gas-free RD-0146 engine was created using the gas-generator circuit ) as well as the project of the hydrogen upper stage of the Angara-A5 PH. At the same time, it is expected that the use of a cryogenic upper stage will allow increasing the mass output by the Angaroy-A5 to the LEO from 24.5 tons to 34-38 tons during launches from the Vostochny cosmodrome. Therefore, theoretically, SpaceX engineers could follow a similar path: the first stage is on kerosene or other fuels, and the upper ones are on hydrogen. However, such a concept in the case of ITS LV is not without significant drawbacks, the main of which is the need to build a launch complex filling a giant rocket with large volumes of several types of rocket fuels, and SpaceX always seeks to reduce costs in everything. In addition, if SpaceX wants to return the upper stages, then liquid hydrogen is again not the best choice. In general, the engineers of a small but very proud company had a difficult choice.

    The first messages about the engine being prepared for flights to other planets began to appear in the summer of 2010, when the then director of the SpaceX rocket development and testing complex (SpaceX Rocket Development and Test Facility), MacGregor, Texas (probably this small city with a population of about 5,000 people known to many readers for the video takeoff and landing of experimental test benches for testing the landing of the first stage - Grasshoppers) Tom Markusic announced the start of work on the Merlin 2 gas engine. It was assumed that he would use the RP-1 / LOX fuel pair and have a thrust of 7.6 MN at sea level and 8.5 MN in a vacuum, which exceeded the same-size single-chamber “monster” F-1, which is five in number. used at the first stage of the “Saturn V” LV for launching lunar missions. Also in the statement it was said that the engine will have unprecedented efficiency, although it is rather difficult to say what these statements were based on, and the project to develop “Merlin 2” very quickly came to naught.


    Returned first stage PH "Falcon 9" - the result of the test tests of Grasshoper'ov.

    The second announcement by Tom Markazika was the announcement of a project to develop an LH2 / LOX LRE Raptor, which was supposed to be brought to the level of thrust ~ 0.67 MN with a specific impulse of 470 seconds. This iteration assumed that the Merlin 2 engines would be on the first degree, and the Raptor LRE would be installed on the top. As a result, the story of the kerosene-hydrogen superracket ended with the statement by Ilon Mask that the plans voiced earlier should be understood not as an approved development program, but as a result of brainstorming and subject for further discussion. Soon, SpaceX left and
    Tom Markazik.

    The first hints that SpaceX was preparing something on the exotic liquid vapor "methane / liquid oxygen" (CH4 / LOX) were news in May 2011 that SpaceX was in contact with the US Air Force for possible participation in the state program on the development of high-thrust engines for reusable accelerators. And there really was something to discuss. The fact is that this application by the United States Air Force implied very high requirements for engine efficiency, moreover, it clearly stated that engines needed on RP-1 / LOX steam. At that time, only two units met the requirements of the United States Air Force: the AJ-26-500 Aerojet and RD-191 engine developed on the basis of the Soviet Lunar Heritage NK-33production NPO "Energomash". In turn, SpaceX just consulted with customers from the Air Force for an opportunity to squeeze into this Soviet-Russian “sweet couple” with certain own engines running on other fuels. And since the application for the program was about high-propulsion engines, it became clear that this was not about the upgraded Merlin 1 LRE, but about something completely new. Time passed and the new engine, which eventually received the name "Raptor", acquired all new and new parts and details. Initially, in 2011, the desired level of thrust at 2.2 MN was announced, in the second quarter of 2013, the project thrust was increased from the initial 2.2 MN to 2.9 MN, and in 2014 there was information about the thrust of 4.5 MN.


    Oxygen-kerosene LRE RD-191 produced by NPO Energomash, built according to a closed scheme with oxidizing generator gas (about what this means will be written a little lower), MAKS-2013. Extremely efficient, reusable, record holder for throttlingthrust at sea level. The largest recorded throttling is 27% of the maximum value, which was confirmed in actual operating conditions during the launch of the Angara-A5 LV: the RD-191 installed on the central unit was chocked to a level of 30%. One problem: in Russia there are not yet reusable rocket stages, therefore this rather expensive, in fact, reusable engine is lost after the first launch. On August 25, 2015, NPO Energomash set about creating an upgraded version of the RD-191M engine, which should be 10-15% more powerful than the basic version.

    The use of methane has a number of important advantages compared with LH2 / LOX and RP-1 / LOX:

    • The CH4 / LOX pair is characterized by a fairly high density, which is 0.82 g / cm ^ 3 (remember, this is 0.23 g / cm ^ 3 for LH2 / LOX, 1.03 g / cm ^ 3 for RP-1 / LOX). Thus, it will be sufficient to increase the size of the tanks by only 25-30% relative to the “kerosene design” of the same flight quality;
    • Although methane is a cryogenic fuel, its temperature in the liquid state is far from that of liquid hydrogen (about -161 degrees Celsius for liquid methane versus -253 degrees Celsius in liquid hydrogen). In addition, compared with liquid hydrogen, liquid methane is much less aggressive with respect to materials used in rocket production;
    • The use of liquid methane as a fuel significantly reduces the amount of soot generated in engines in comparison with RP-1 / LOX, which allows to reduce the cost of prelaunch preparation of reusable stages and, in general, increase the reliability of a reusable engine;
    • Finally, methane is an affordable and cheap fuel.

    But SpaceX decided not to limit themselves to the “native” advantages of the methane system and went even further: “Raptor” is the first in the world to be launched into full-scale production of LREs with the most effective closed cycle - the so-called “full-flow closed cycle” (that is, afterburning previously gasified and oxidative and fuel components).

    In general, both in our media and in foreign documentaries, you can hear words like “The first engine of the closed cycle was the NK-33, then everyone forgot about this technology, and then on its basis made the RD-180. And all other countries envy us (Russia) (c) ”. For example, the story is described in the British film “The Hot Engines of a Cold Country” (“The Engines That Came In From The Cold.” Channel 4, London). In fact, there are a lot of engines with one form or another of the closed cycle (they will be discussed below).


    Documentary film “Hot Engines of the Cold Country” (“The Engines That Came In From The Cold”. Channel 4, London). In school, this film greatly strengthened the desire of the author of this article to go and study as a rocket engineer or physicist.

    Liquid rocket engine of the closed circuit (liquid rocket engine of the closed cycle) is a liquid-propellant rocket engine, made according to the scheme with afterburning of the generator gas. In a rocket engine of a closed circuit, one of the components is gasified in the gas generator by burning at a relatively low temperature with a small part of the other component, and the resulting hot gas is used as the working fluid of the turbine turbine unit (THA). The generator gas, which has been triggered on the turbine, is then fed into the engine's combustion chamber, where the rest of the unused component of the fuel is also fed. The combustion chamber completes the combustion of components with the creation of thrust. The following types of closed-end LREs are distinguished:

    • With oxidizing gas generator. Examples: RD-253 (“Proton M”), RD-170/171 (“Energy”, “Zenith”, in the future, perhaps, “Soyuz-5”), RD-180 (Atlas-V), RD-191 / 193 ("Angara", "Naro-1" (aka KSLV-1), Soyuz-2.1v, may also be installed on "Antares" instead of NK-33) RD-120 (second stage of "Zenith"), NK-33 (N-1, Soyuz-2.1v, “Antares”, possibly, Soyuz-2-3);
    • With regenerative gas. Examples: RD-0120 (second stage of the PH "Energy", SSME (Space Shuttle Main Engine), RD-857 (Soviet ICBM RT-20P ), LE-7 / LE-7A (Japanese engines for H-II missiles )
    • With full gasification of components. Examples: RD-270 (UR-700 and UR-900), "Raptor" SpaceX.

    Quote from the article "Liquid Rocket Engine Closed Cycle" , Wikipedia, with minor additions of the author.

    An example of an engine working on such a scheme was developed in OKB-456 (now NPO Energomash named after academician V.P. Glushko ) in the OKB-456 ) RD-270 LRE (used UDMH / AT) for the Soviet lunar project / Martian missile UR-700 / UR-900 (all the same it’s good that the choice fell on the kerosene N-1: if in the Kazakh steppes a minute after the launch a super-heavy rocket exploded on UDMH / AT, then with ecology at Baikonur it would be quite bad) .


    Created in 1962-1967 in OKB-456 (now NPO Energomash), the RD-270 LRE-270 (NDMG / AT) for the Soviet lunar / Marsian program UR-700 / UR-900. It is the first in the world and one of two (sometimes also referred to as the third engine - Integrated Powerhead Demonstrator produced by Rocketdyne and Aerojet ) created for all time engines with a full-flow closed loop scheme. The second such engine had to wait more than 50 years.

    A few words about the RD-270. Its development began in 1962 and was completed in 1967, that is, after 5 years. In total from October 1967 until the close of the program for the creation of the UR-700 / UR-900 in July 1969, 27 fire tests of this unit were carried out and a total of 22 pieces of this engine were assembled. Three engines were tested again, and one - three times. Then the project UR-700 / UR-900 was closed.

    In addition to the increase in the specific impulse, a closed circuit with full gasification of components implies an engine design with a reduced number of potential points of failure as compared with a partial gasification of liquid propellant rocket engines. Also, a scheme with full gasification implies no need for forcing and burning liquid components in the combustion chamber, which negates the risk of cavitation.components of liquid fuel and thereby increases the reliability of the system. However, this design concealed some difficulties: because of the simultaneous operation of four deeply integrated important engines - two gas generators and two turbopumps and their essentially interrelated work on supplying full gasification products to the main combustion chamber in the RD-270, low-frequency pulsations in gas generators and in the main combustion chamber. The main reason for this dangerous mode of operation of the engine was the difficulty of synchronizing the joint operation of two turbopumps, which tried to overpower each other. In the framework of the RD-270 project, this engineering task could not be solved, and for the first time, it was possible for the American engineers to cope with it only 10 years later when the RS-25 LRE was created.(the main engine of the Space Shuttle shuttle) only through the use of an onboard digital computer , which simply did not exist in the USSR at the time of the development of the RD-270.


    Scheme of a rocket engine with full gasification. This architecture can significantly improve the reliability (for example, by reducing the number of required pumps and pipelines) and engine performance while reducing its mass. Preburner - gas generator; Pump - turbo pumps; Combustion Chamber - the main combustion chamber. For comparison, in the spoiler below is a diagram of a closed-cycle engine with regenerative generating gas, in which fuel is supplied only through a gas generator, and the oxidizer is also directly from the tanks.

    Scheme of a closed cycle engine with regenerative gas

    There is, however, a scheme with a full gasification of the pitfall - the main combustion chambers of engines produced by this technology is very difficult to test. The fact is that most modern engines can be tested in parts: pumps separately, combustion chambers separately, and so on. When using full gasification, this is not possible due to the fact that all engine parts are very dependent on each other. A closed circuit with full gasification of fuel components implies gasification in two separate gas generators (a gas generator is a device for converting solid or liquid fuel into a gaseous form): in one small part of the fuel is burned with a huge amount of oxidant consumption (in fact, it is a kind of oxidizing gas generator) and in the other, a surplus of fuel is burned with a small amount of oxidizer (in fact, it is a kind of fuel gas generator). The supply of the oxidizer and fuel to the gas generators is carried out using turbopumps and these same turbopumps immediately after starting the engine operate at the expense of the energy of the gasification products obtained in the gas generators. Finally, unlike all other schemes, a full-flow closed cycle implies that the oxidant fuel enters the combustion chamber exclusively in gaseous form, that is, it (the combustion chamber) is connected exclusively to the gas generators, but not to the tanks, it would test the combustion chamber without gas generators and corresponding Turbopumps are fundamentally impossible. In general, for testing you need to assemble the engine completely. The supply of the oxidizer and fuel to the gas generators is carried out using turbopumps and these same turbopumps immediately after starting the engine operate at the expense of the energy of the gasification products obtained in the gas generators. Finally, unlike all other schemes, a full-flow closed cycle implies that the oxidant fuel enters the combustion chamber exclusively in gaseous form, that is, it (the combustion chamber) is connected exclusively to the gas generators, but not to the tanks, it would test the combustion chamber without gas generators and corresponding Turbopumps are fundamentally impossible. In general, for testing you need to assemble the engine completely. The supply of the oxidizer and fuel to the gas generators is carried out using turbopumps and these same turbopumps immediately after starting the engine operate at the expense of the energy of the gasification products obtained in the gas generators. Finally, unlike all other schemes, a full-flow closed cycle implies that the oxidant fuel enters the combustion chamber exclusively in gaseous form, that is, it (the combustion chamber) is connected exclusively to the gas generators, but not to the tanks, it would test the combustion chamber without gas generators and corresponding Turbopumps are fundamentally impossible. In general, for testing you need to assemble the engine completely. Finally, unlike all other schemes, a full-flow closed cycle implies that the oxidant fuel enters the combustion chamber exclusively in gaseous form, that is, it (the combustion chamber) is connected exclusively to the gas generators, but not to the tanks, it would test the combustion chamber without gas generators and corresponding Turbopumps are fundamentally impossible. In general, for testing you need to assemble the engine completely. Finally, unlike all other schemes, a full-flow closed cycle implies that the oxidant fuel enters the combustion chamber exclusively in gaseous form, that is, it (the combustion chamber) is connected exclusively to the gas generators, but not to the tanks, it would test the combustion chamber without gas generators and corresponding Turbopumps are fundamentally impossible. In general, for testing you need to assemble the engine completely.


    Simulation of physicochemical processes in the Raptor LRE. It is highly recommended for viewing to people who love bright beautiful diagrams and models, obtained on the basis of mathematical calculations.

    Another “challenge” on the way to the creation of the finished product is the fact that in engines with full gasification only gaseous components of the fuel and the gaseous products of its combustion already enter the combustion chamber, and the physicochemical aspects of this process have not been widely studied before. the fact that no one in the United States, and in the world, has ever used a full-flow closed cycle before. And even if we take into account the existence of the RD-270, then, firstly, SpaceX could hardly get detailed documentation on this product, secondly, it is unlikely that at the end of the 60s of the last century the computing power allowed to get results, which it would not make sense to clarify or even recheck in 2017.


    Ilon Musk presents to the public a review of the characteristics of the Raptor engine at the International Astronautical Congress, September 27, 2016, Guadalajara, Mexico.

    It is also known that in order to optimize launches, the fuel and oxidizer for the Raptor LRE will be in tanks at temperatures close to the freezing temperature, and not to the boiling point, which is not typical of existing cryogenic missile systems. Overcooling of methane and oxygen should increase their density, which will lead to a decrease in the volume of fuel tanks and the rocket as a whole. In addition, supercooled fuel and oxidizer are less prone to cavitation processes in turbopump units, which also affects the reliability of the system in a very positive way.

    In addition, the possibility of transferring the production of individual Raptor units to 3D printing technology is being studied. Thus, in 2016, an experimental reduced engine sample with a load of about 1 MN was tested, 40% of the details of which (by weight) were printed.


    Summary table with characteristics of some widely used single-chamber engines. Engines created in the USA are marked in blue, red - created in the USSR / Russia. The signs (***) of the indicators for the Raptor and Merlin 1D engines indicate that these figures refer not to the basic modifications of these engines that are on the first steps, but to the special Raptor Vacuum versions installed on the upper stages. and Merlin 1D Vacuum, respectively.

    Thus, as a conclusion, it can be said that the “Merlin” open-cycle was a very successful one, and its upgraded version “Merlin 1D” has the highest ratio of thrust / mass and thrust / cost, and is also the most effective oxygen kerosene engine ever produced in the United States, but still in many matters "Merlin" remains far from the most advanced unit. In turn, the new generation Raptor, developed by SpaceX, has absorbed, if not all, then very many of the most advanced technologies that exist in the rocket engine building industry today. And the planned reusable use of this unit will more than compensate for the high cost of such advanced solutions.

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