«No engine - and the most perfect design any rocket with all her stuffed dead» i> Glushko
In recent years, a private space company Space X, headed by Elon Musk, never ceases to amaze the world with its fantastic success. The main highlight of space transportation systems of the company are considered RN series Falcon, and in particular the Merlin engine 1, already dubbed "the most efficient in the world».
It creates the false impression that Space X in a relatively short time was able to create the engine, eclipsed developments in this area of such giants as « ENERGOMASH » and « Rocketdyne ». Under the cut in a popular form, we get acquainted with the modern world of rocket engines and try to understand this is not an unambiguous characterization as their effectiveness.
In 2012, the company Space X, spent the last test firing at the moment engine modifications Merlin 1 - D. During these tests, the engine thrust-weight ratio was increased to 150 units, which allowed him to name Space X «самым effective in history ».
In the field of engine, thrust-weight ratio is called the ratio of engine thrust (CU) to its dry weight. In the case of missile LRE usually this ratio is engine thrust in vacuum (in TC) divided by its dry weight (in tonnes).
Merlin 1D is capable of thrust from the earth in the 67t and 82T in vacuum (Melin 1D Vacuum), with a mass of about 600kg. At pH 1.1 using the Falcon 9 9 overall thrust of these engines in 600ts. The engine does not have its own system of thrust vector control (SWT), and stage of the launch control is effected by varying thrust on opposite axes, as well as at pH 1 H (throttling to reduce thrust and / or afterburner to increase the rated thrust).
Merlin 1D engines on the Falcon 9 rocket 1.1 i>
The difference between the engine thrust in the ground and in a vacuum is typical, and is connected with the deterioration of motor performance in the dense layers of the atmosphere of the planet. Atmospheric drag aerodynamics of the engine is increased in proportion to the size of the exhaust nozzle of the engine (the combined resistance of the atmospheric pressure only increases with the square of jet exhaust). It would seem that only need to minimize the size of the nozzle and improve engine performance. However, with increasing altitude atmospheric drag is reduced, and with it the characteristics of the engine on the contrary increases with the size of the nozzle.
The key here is not the absolute size of the nozzle, and expansion ratio - the ratio of the greatest cross sectional area of the cone of the nozzle to its smallest cross-sectional area. The higher the value, the higher the efficiency of the engine in a given environment. Therefore, the nozzle size the vast majority of engines a first stage launch vehicles represent a compromise for optimal engine performance, both in the atmosphere and in vacuum.
Thrust indirectly depends on the fuel used, and in contrast to the characteristics such as thrust and specific impulse, is not applicable to all types of engines - solid rocket systems are themselves engines.
The most high thrust-weight ratio differ LRE runs on a mixture of heptyl and dinitrogen tetroxide. So RD 275M ("Energomash") carrier rocket Proton M have the highest thrust-weight ratio in the world - more than 170 (187t thrust in vacuum, the mass of the engine 1, 1t). Their "big" brother RD 270 , developed in time for the summer is not ur RN 700, had a thrust-200 units! All the more surprising that it is a closed-cycle engines (see. Below). These characteristics are achieved because of the hypergolic fuel by which greatly simplified the design (and weight) of rocket engines. At the same time these motors have a pretty high parameters of specific impulse (285s for RD 275M).
RD 275M, Russian modification of the base of the Soviet RD 253 for RN series Proton (payload mass is increased to 750kg) i>
Specific impulse (sometimes referred to as specific thrust) expresses the time during which the engine develops thrust in one newton (1N = 1kgs / 0, 102), using 1 kg of fuel. The higher the UI the lower the engine requires fuel message payload certain amount of movement. In contrast, Thrust, this value and engineers often taken as an index of efficiency of the engine.
Mr H h5> Modern hydrogen engines are the most efficient among all types of used engines. The greatest value UI possessed Soviet RD 0120 (455s in vacuum and traction in 200ts). The highest values of UI Sea level has RS 68 (365s and thrust 295ts) from Rocketdyne, used on the world's only fully hydrogen rocket Delta 4. hydrogen rocket engine at the same time have the lowest values of thrust-weight ratio (in the range of 50-75 units), which is ignored because of the high energy opportunities these LRE. This allows more than offset the "extra" several tons of engine, compared with the rest of the LRE high thrust-weight ratio.
However, the high price of hydrogen rocket engine (about $ 20 million for RS 68) still makes use of engineers in the early stages of trade-offs, often with kerosene LRE.
Transparent table influence the degree of expansion of the efficiency of cryogenic engines in different environments i>
From left to right: RS 68, Vulcain, RS 25, RD 0120 i>
The champion in a narrow sense. H5> The highest specific impulse at sea level (311 seconds) of kerosene LRE have RD 171 RD 180 (a stripped-down version with ½ RD 171 with a thrust of 384ts) and RD 191 (stripped down to ¼ version of RD 171 with a thrust of 196ts) NGO "Energomash". Thrust-weight ratio of these engines does not exceed 90 units. Against the background of these masterpieces of technical thought, efficiency Merlin 1D looks quite modest (285 sec), although dominated by thrust-weight ratio of kerosene LRE.
RD 171/180/191 i>
The list of the most famous LRE Energomash i>
Such a difference in the characteristics connected with different design approach in the design of engines:
- LRE "family" RD 170/171 are designed using a closed cycle - to initialize the engine operating pressure is supplied to the gasifier gases of which are driven by the turbine, rotating pumps the fuel and oxidizer. Incoming fuel takes part in cooling the nozzle and further into the combustion chamber, the other part goes to the maintenance operation of the gasifier with an oxidant, and then after the gas outlet of the turbine mixture enters the combustion chamber. The whole vicious cycle is repeated until all the fuel in the tanks. In this combustion chamber at Staged combustion cycle substantially less than Gas-generator cycle. As you probably guessed it provides high values of pressure in the combustion chamber (usually 200 atmospheres or more) and a greater degree of expansion of the engine nozzle, allowing LRE provide high efficiency (specific impulse) in the planet's atmosphere.
Disadvantages - high load on the turbine engine, the relatively high cost and complexity of such engines.
The approximate scheme Staged combustion cycle on the example of Russian RD 191 and NC 33 1 - The gas generator; 2 - turbine; 3 - Supply of fuel (kerosene); 4 - Submission of an oxidant (oxygen); 5 - The fuel pump; 6 - Pump oxidant; 7 - Withdrawal of the fuel nozzles for cooling; 8 - Retraction gasifier fuel / oxidizer turbine combustor; 9 - Transfer of oxidant gas generator i>
- LRE family Merlin and RD 107/108 (RN Union) are typical representatives of the open-cycle engines. The working fluid of the turbine engine (coming from the gasifier), not closed to the combustion chamber, and is output to the external environment, together with a portion of the fuel is only partially participating in the creation of additional thrust. To compensate for the loss of efficiency is possible to increase the pressure in the gasifier, increasing the efficiency of the turbine and consequently the pressure in the combustion chamber (which is about 100 atmospheres). Engines of this scheme is easier, safer, easier and cheaper Staged combustion cycle.
Among the shortcomings should be noted low expansion nozzle of the engine and correspondingly smaller values of specific impulse at work in the atmosphere of the planet (263 / 257s for RD 107/108 and 255s to RS 27A).
The scheme open cycle engines on the example of the LRE F 1 and Merlin 1D. 1 - Supply of fuel and oxidizer tanks; 2 - The gas generator; 3 - Pump oxidant (oxygen); 4 - Fuel pump (kerosene); 5 - Turbine; 6 - high pressure oxidant conduit; 7 - conduit high pressure fuel; 8a - Fuel System Cooling nozzle; 8b - gas gasifier heat conductor; 9 - Exhaust gasifier mixture (Merlin 1D) / venting gas generator for cooling the nozzle head (F 1); 10 - Nozzle extension i>
Comparing the first-stage engine, it should be noted that the thrust-weight ratio of the engine is not directly related to the thrust-weight ratio of all stages. At equal thrust rocket engines will be of decisive importance not their comparative thrust-weight ratio, namely, the specific impulse. As we said, the higher the value, the less fuel is used to disperse the LRE specific weight and, therefore, the higher thrust-stage rocket.
So thrust-RN Falcon 9 1.1 is 1, 2 (Rod 600ts / 503t weight rockets) and Zenit rocket with RD 171 2 1 5 (Rod 720t / 470t mass of the rocket) with a similar payload into LEO 13T.
For understanding this approach domestic designers should consider the geographical location of the specifics of the Russian and American space centers. The latter is to the south, have a 15% greater energy advantage thanks to the contribution of the Earth rotation (additional ~ 200 m / s). Therefore, a high thrust-weight ratio is common for Russian rockets (1, 5-1, 7 for LV Energy and H1, 1 against, 1 for the Saturn 5). And as we have learned, the rocket engine thrust-right has no value for this key.
However, in the Soviet Union was still created kerosene engine combines high thrust-weight ratio and high specific impulse. LRE NC 33 from OKB Kuznetsov, created on the basis of engine NK 15 moon rocket H 1, with thrust-weight ratio of 136 (171ts / 1, 25t), had a specific impulse in the 297s (at sea level). Modern modification of the engine used on RN Anateres, a private company Orbital Sciences (AJ26). Russian modification of NK 33-1 PH is used on Soyuz 2, 1B, at the start of thrust in the 185ts at specific impulse at 305c! The basic version of the Tax Code of 33, the engine is characterized primarily by the presence of the thrust vector control (UHT).
In the future, NC 33-1 plan to use at high altitudes retractable Nozzle extension, significantly improves engine performance.
NK 33-1 with the nozzle head. On the right, the graph of increasing performance with LRE nozzle head i>
Price of the issue. H5> is no doubt that one of the main "characteristics" of any type of technology is its cost. Because of the large difference in the specifications of engines, it would be preferable to compare their relative price variable. In this case, this value is approximately the ratio of prices to the LRE his rod ($ / ton).
Poster "Energomash" showing the pricing scheme LRE and their share of the price of the entire booster i> 15,261,141
Obviously that price increases in proportion to the complexity of the engine and its efficiency.
So RS 68 standing on the Delta 4 rocket, NASA cost of 60 000 $ / ton of traction (20 million $).
Kerosene LRE with a big draw, but lower specific Pulses RD 180 (PH Atlas 5) nominally manages NASA's half the price - 30 000 $ / ton ($ 11 million).
For comparison, the cost of RD 171 on the basis of which RD 180/191, is within 22,000 $ / ton ($ 13-15 million). This variation is partly due to the fact that the last two engine designed for the US domestic market, in particular for the Atlas 5 rocket (RD 180 as the main engine of the central unit, and RD 191 engine for the side blocks). However, the RD 191 was left unclaimed in the US, even after the creation of a budget RD 193 (the version without UHT).
Most "cheap" closed-cycle engine can be considered the LRE NC 33-1. Subject to restore production price modification NK 33-1 for the new "Soyuz 2-3" may amount to 25 000 $ / ton (4, $ 5 million). Officially NC 33-1 will be used until depletion of stocks of old NC 33 and replaced on RD 193 .
Merlin 1D with an price of 15 000 $ / ton (~ $ 1 million), very well, "joined" in the domestic market of the US missile engines. After the closure of the Apollo program, the United States half a century have focused on the development of cryogenic (hydrogen), toxic (heptyl) and solid rocket engines. The implications of this approach and we are seeing today - ahead of Russia on the part of the development and operation of the cryogenic liquid fuel rocket engine and TRD, USA is far behind in the development of the already kerosene LRE.
Even with the development in the US own kerosene LRE, it is doubtful that they will be able to compete on price and the degree of perfection with Russian engines and even more so with the "budget" brainchild Space X. Therefore, Elon Musk and Co. have every reason to look optimistically to the future of their developments. Development is extremely successful, reliable and promising, which is not necessarily "try on" controversial epithets long deserved other talented developers < / a>.