What types of multistage rockets are there? The design and principle of operation of the rocket

The invention relates to reusable space transport systems. The proposed rocket contains an axisymmetric body with a payload, a propulsion system and takeoff and landing shock absorbers. Between the struts of these shock absorbers and the nozzle of the main engine there is a heat shield made in the form of a hollow thin-walled compartment made of heat-resistant material. The technical result of the invention is to minimize the gas-dynamic and thermal loads on the shock absorbers from the operating propulsion engine during launches and landings of the launch vehicle and, as a result, ensure the required reliability of the shock absorbers during repeated (up to 50 times) use of the rocket. 1 ill.

Authors of the patent:
Vavilin Alexander Vasilievich (RU)
Usolkin Yuri Yurievich (RU)
Fetisov Vyacheslav Alexandrovich (RU)

Owners of patent RU 2309088:

Federal State Unitary Enterprise "State Missile Center" Design Bureau named after. Academician V.P. Makeeva" (RU)

The invention relates to rocket and space technology, in particular to reusable transport space systems (MTKS) of a new generation of the “Space orbital rocket - single-stage vehicle carrier” (“CORONA”) type with fifty to hundred times of its use without overhaul, which is a possible alternative to winged reusable systems such as the Space Shuttle and Buran.

The CORONA system is designed to launch a payload (spacecraft (SC) and spacecraft with upper stages (UB) into low Earth orbits in the altitude range from 200 to 500 km with an inclination equal to or close to the inclination of the orbit of the launched spacecraft.

It is known that at launch the rocket is located on the launcher, while it is in a vertical position and rests on the four support brackets of the tail compartment, which is subject to the weight of a fully fueled rocket and wind loads that create a capsizing moment, which, when acted simultaneously, are the most dangerous for strength the tail section of the rocket (see, for example, I.N. Pentsak. Flight theory and design of ballistic missiles. - M.: Mashinostroenie, 1974, p. 112, Fig. 5.22, p. 217, Fig. 11.8, p. 219) . The load when parking a fully fueled rocket is distributed across all support brackets.

One of the fundamental issues of the proposed MTKS is the development of takeoff and landing shock absorbers (TSA).

The work carried out at the State Rocket Center (SRC) on the CORONA project showed that the most unfavorable case of loading a rocket launcher is landing a rocket.

The load on the VPA when a fully fueled rocket is parked is distributed over all supports, while during landing, with a high degree of probability, due to the permissible deviation from the vertical position of the rocket body, a case is possible where the load falls on one support. Taking into account the presence of vertical speed, this load turns out to be comparable or even greater than the load in the parking lot.

This circumstance made it possible to decide not to abandon the special launch pad, transferring the power functions of the latter to the rocket’s VPA, which significantly simplifies the launch facilities for systems of the “CORONA” type, and accordingly, the costs of their construction are reduced.

The closest analogue of the proposed invention is a reusable single-stage launch vehicle "CORONA" for vertical take-off and landing, containing an axisymmetric body with a payload, a propulsion system and take-off and landing shock absorbers (see A.V. Vavilin, Yu.Yu. Usolkin "O possible ways of development of reusable transport space systems (MTKS)", RK technology, scientific and technical collection, series XIY, issue 1 (48), part P, calculation, experimental research and design of ballistic missiles with underwater launch, Miass, 2002 ., p.121, fig.1, p.129, fig.2).

The disadvantage of the design of an analogue rocket is that its PPAs are located in the zone of gas-dynamic and thermal influence of the flame emerging from the central nozzle of the main propulsion system (MPU) during repeated launch and landing of the rocket, as a result of which the reliable operation of the design of one PPA is not ensured with the required resource its use (up to one hundred flights with a twenty percent resource reserve).

The technical result when using a single-stage reusable vertical take-off and landing launch vehicle is to ensure the required reliability of the design of one propeller when using the launch vehicle fifty times by minimizing the gas-dynamic and thermal loads on the launch vehicle from the operating MDU during multiple launches and landings of the rocket.

The essence of the invention is that in a well-known single-stage reusable vertical takeoff and landing launch vehicle containing an axisymmetric body with a payload, a propulsion system and takeoff and landing shock absorbers, a heat shield is installed between the struts of the takeoff and landing shock absorbers and the nozzle of the propulsion engine .

Compared to the closest analogue rocket, the proposed single-stage reusable vertical takeoff and landing launch vehicle has better functional and operational capabilities, because it ensures the necessary reliability of the design of one UPA (not lower than 0.9994) for a given service life of one launch vehicle (up to one hundred launches) by isolating (using a heat shield) the UPA struts from the gas-dynamic and thermal loads of the operating MDU for a given resource (up to hundred) flights of the launch vehicle during its multiple launches and landings.

To explain the technical essence of the proposed invention, a diagram of the proposed launch vehicle with an axisymmetric body 1, a nozzle 2 of the propulsion system, struts of the takeoff and landing shock absorber 3 and a heat shield 4 of a hollow thin-walled compartment made of heat-resistant material, which isolates the struts of the takeoff and landing shock absorber from the gas-dynamic and thermal impact of the flame from the central nozzle of the main propulsion system during takeoff and landing of the rocket.

Thus, the proposed reusable vertical takeoff and landing launch vehicle has broader functional and operational capabilities compared to its closest analogue by increasing the reliability of one takeoff and landing shock absorber for a given flight life of the launch vehicle on which this takeoff and landing shock absorber is located.

A single-stage reusable launch vehicle for vertical takeoff and landing, containing an axisymmetric body with a payload, a propulsion system and takeoff and landing shock absorbers, characterized in that a heat shield made in the form of a hollow one is installed between the struts of the takeoff and landing shock absorbers and the nozzle of the propulsion engine thin-walled compartment made of heat-resistant material.

Development of a landing system - the number of supports, their arrangement, while minimizing their mass is a very difficult task...

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The cost of launching spacecraft into orbit is still very high. This is explained by the high cost of rocket engines, an expensive control system, expensive materials used in the stressed structure of rockets and their engines, complex and, as a rule, expensive technology for their manufacture, preparation for launch and, mainly, their one-time use.

The share of the carrier cost in the total cost of launching a spacecraft varies. If the media is serial and the device is unique, then about 10%. If it’s the other way around, it can reach 40% or more. This is very expensive, and therefore the idea arose to create a launch vehicle that, like an airliner, would take off from a cosmodrome, fly into orbit and, leaving a satellite or spacecraft, returned to the cosmodrome.

The first attempt to implement such an idea was the creation of the Space Shuttle system. Based on an analysis of the shortcomings of disposable media and the Space Shuttle system, which was made by Konstantin Feoktistov (K. Feoktistov. Trajectory of life. Moscow: Vagrius, 2000. ISBN 5-264-00383-1. Chapter 8. A rocket like an airplane), one gets an idea of the qualities that a good launch vehicle should have, ensuring the delivery of payload into orbit at minimal cost and with maximum reliability. It should be a reusable system capable of 100–1000 flights. Reusability is needed both to reduce the cost of each flight (development and manufacturing costs are distributed over the number of flights) and to increase the reliability of launching payload into orbit: every car trip and aircraft flight confirms the correctness of its design and high-quality manufacturing. Consequently, it is possible to reduce the cost of insuring the payload and insuring the rocket itself. Only reusable machines - such as a steam locomotive, a car, an airplane - can be truly reliable and inexpensive to operate.

The rocket must be single-stage. This requirement, like reusability, is related to both minimizing costs and ensuring reliability. Indeed, if the rocket is multi-stage, then even if all its stages return safely to Earth, then before each launch they must be assembled into a single whole, and it is impossible to check the correct assembly and functioning of the stage separation processes after assembly, since with each check the assembled machine must crumble . Untested and unchecked for functionality after assembly, the connections become disposable. And a packet connected by nodes with reduced reliability also becomes, to some extent, disposable. If the rocket is multi-stage, then the costs of its operation are higher than those of a single-stage machine for the following reasons:

The single stage machine does not require any assembly costs.
There is no need to allocate landing areas on the surface of the Earth for landing the first stages, and therefore, there is no need to pay for their rental, for the fact that these areas are not used in the economy.
There is no need to pay for transportation of the first stages to the launch site.
Refueling a multi-stage rocket requires more complex technology and more time. The assembly of the package and delivery of the stages to the launch site cannot be easily automated and, therefore, requires the participation of more specialists in preparing such a rocket for the next flight.

The rocket must use hydrogen and oxygen as fuel, as a result of the combustion of which environmentally friendly combustion products are formed at the exit from the engine with a high specific impulse. Environmental cleanliness is important not only for work carried out at the start, during refueling, in the event of an accident, but also, no less, to avoid the harmful effects of combustion products on ozone layer atmosphere.

Among the most developed projects of single-stage spacecraft abroad, it is worth highlighting Skylon, DC-X, Lockheed Martin X-33 and Roton. If Skylon and X-33 are winged vehicles, then DC-X and Roton are vertical take-off and vertical landing missiles. In addition, both of them got to the point of creating test samples. While Roton only had an atmospheric prototype for testing autorotation landings, the DC-X prototype made several flights to an altitude of several kilometers using a liquid rocket engine (LPRE) powered by liquid oxygen and hydrogen.

Technical description of the Zeya rocket

To radically reduce the cost of launching cargo into space, Lin Industrial proposes to create the Zeya launch vehicle. It is a single-stage, reusable vertical take-off and vertical landing transport system. It uses environmentally friendly and highly efficient fuel components: oxidizer - liquid oxygen, fuel - liquid hydrogen.

The launch vehicle consists of an oxidizer tank (above which there is a heat shield for re-entry and a soft landing system rotor), a payload compartment, an instrument compartment, a fuel tank, a tail compartment with a propulsion system and a landing gear. Fuel and oxidizer tanks are segmental-conical, load-bearing, composite. The fuel tank is pressurized by gasification of liquid hydrogen, and the oxidizer tank is pressurized by compressed helium from high-pressure cylinders. The propulsion system consists of 36 circumferentially located engines and an external expansion nozzle in the form of a central body. During operation of the propulsion engine, pitch and yaw control is carried out by throttling diametrically located engines, and roll control is carried out using eight gaseous propellant engines located under the payload compartment. For control of the orbital flight segment, engines using gaseous fuel components are used.

The Zeya flight pattern is as follows. After entering the reference low-Earth orbit, the rocket, if necessary, performs orbital maneuvers to enter the target orbit, after which, opening the payload compartment (weighing up to 200 kg), separates it.

During one orbit around the Earth's orbit from the moment of launch, having issued a braking impulse, Zeya lands in the area of the launch site. High landing accuracy is achieved by using the lift-to-drag ratio created by the rocket's shape for lateral and range maneuvers. A soft landing is carried out by descending using the principle of autorotation and eight landing shock absorbers.

Economy

Below is an estimate of the time and cost of work before the first launch:

Advance project: 2 months - €2 million
Creation of a propulsion system, development of composite tanks and control systems: 12 months - €100 million
Creation of a bench base, construction of prototypes, preparation and modernization of production, preliminary design: 12 months - €70 million
Testing of components and systems, prototype testing, fire tests flight product, technical design: 12 months - €143 million

Total: 3.2 years, €315 million

According to our estimates, the cost of one launch will be €0.15 million, and the cost of inter-flight maintenance and overhead costs will be about € 0.1 million for the inter-launch period. If you set the launch price to € 35 thousand per 1 kg (at a cost of €1250/kg), which is close to the price of launching on a Dnepr rocket for foreign customers, the entire launch (200 kg payload) will cost the customer € 7 million. Thus, the project will pay for itself in 47 launches.

Zeya variant with a three-fuel engine

Another way to increase the efficiency of a single-stage launch vehicle is to switch to a liquid-propellant rocket engine with three fuel components.

Since the early 1970s, the USSR and the USA have been studying the concept of three-propellant engines that would combine the high specific impulse of using hydrogen as fuel, and a higher average fuel density (and, therefore, smaller volume and weight of fuel tanks), characteristic of hydrocarbon fuel. At startup, such an engine would run on oxygen and kerosene, and at high altitudes it would switch to using liquid oxygen and hydrogen. This approach may make it possible to create a single-stage space launch vehicle.

In our country, three-component engines RD-701, RD-704 and RD0750 were developed, but they were not brought to the stage of creating prototypes. In the 1980s, NPO Molniya developed the Multi-Purpose Aerospace System (MAKS) on the RD-701 liquid-propellant rocket engine with oxygen + kerosene + hydrogen fuel. Calculations and design of three-component liquid propellant engines were also carried out in America (see, for example, Dual-Fuel Propulsion: Why it Works, Possible Engines, and Results of Vehicle Studies, by James A. Martin and Alan W. Wilhite , published in May 1979 in Am erican Institute of Aeronautics and Astronautics (AIAA) Paper No. 79-0878).

We believe that for the three-component Zeya, instead of the kerosene traditionally proposed for such liquid-propellant rocket engines, liquid methane should be used. There are many reasons for this:

Zeya uses liquid oxygen as an oxidizer, boiling at a temperature of -183 degrees Celsius, that is, cryogenic equipment is already used in the design of the rocket and the refueling complex, which means there will be no fundamental difficulties in replacing a kerosene tank with a methane tank at -162 degrees Celsius.
Methane is more efficient than kerosene. The specific impulse (I, a measure of the efficiency of a liquid-propellant rocket engine - the ratio of the impulse created by the engine to the fuel consumption) of the methane + liquid oxygen fuel pair exceeds the I of the kerosene + liquid oxygen pair by about 100 m/s.
Methane is cheaper than kerosene.
Unlike kerosene engines, there is almost no coking in methane engines, that is, in other words, the formation of carbon deposits that are difficult to remove. This means that such engines are more convenient to use in reusable systems.
If necessary, methane can be replaced with liquefied natural gas (LNG) with similar characteristics. LNG consists almost entirely of methane, has similar physical and chemical characteristics and is slightly inferior to pure methane in terms of efficiency. At the same time, LNG is 1.5–2 times cheaper than kerosene and much more affordable. The fact is that Russia is covered by an extensive network of natural gas pipelines. It is enough to take a branch to the cosmodrome and build a small gas liquefaction complex. Russia has also built an LNG production plant on Sakhalin and two small-scale liquefaction complexes in St. Petersburg. It is planned to build five more factories in different parts of the Russian Federation. At the same time, to produce rocket kerosene, special grades of oil are needed, extracted from strictly defined fields, the reserves of which are being depleted in Russia.

The operation scheme of a three-component launch vehicle is as follows. First, methane is burned - a fuel with high density, but a relatively low specific impulse in a vacuum. Hydrogen is then burned, a low-density fuel with the highest possible specific impulse. Both types of fuel are burned in a single propulsion system. The higher the proportion of fuel of the first type, the smaller the mass of the structure, but the greater the mass of fuel. Accordingly, the higher the share of fuel of the second type, the lower the required fuel supply, but the greater the mass of the structure. Consequently, it is possible to find the optimal ratio between the masses of liquid methane and hydrogen.

We carried out the corresponding calculations, taking the coefficient of fuel compartments for hydrogen equal to 0.1, and for methane - 0.05. The fuel compartment ratio is the ratio of the final mass of the fuel compartment to the mass of the available fuel supply. The final mass of the fuel compartment includes the mass of the guaranteed fuel supply and unused component residues rocket fuel and the mass of charge gases.

Calculations have shown that the three-component Zeya will launch 200 kg of payload into low Earth orbit with a mass of its structure of 2.1 tons and a launch mass of 19.2 tons. The two-component Zeya on liquid hydrogen is greatly inferior: the mass of the structure is 4. 8 tons, and the launch weight is 37.8 tons.

APUSK was produced using a multi-stage rocket,” we have read these words many times in reports about the launch of the world’s first artificial Earth satellites, about the creation of a solar satellite, about the launch of space rockets to the Moon. Just one short phrase, but how much inspired work of scientists, engineers and workers of our Motherland is hidden behind these six words!

What are modern multistage rockets? Why did it become necessary to use rockets for space flights consisting of large quantity steps? What technical effect does increasing the number of rocket stages give?

Let's try to briefly answer these questions. Flights into space require huge reserves of fuel. They are so large that they cannot be placed in the tanks of a single-stage rocket. With the modern level of engineering science, it is possible to build a rocket in which the fuel would account for up to 80-90% of its total weight. And for flights to other planets, the required fuel reserves must be hundreds and even thousands of times greater than the own weight of the rocket and the payload in it. With the fuel reserves that can be placed in the tanks of a single-stage rocket, it is possible to achieve flight speeds of up to 3-4 km/sec. Improving rocket engines, finding the most advantageous types of fuel, using better structural materials and further improving the design of rockets will certainly make it possible to slightly increase the speed of single-stage rockets. But it will still be very far from cosmic speeds.

To achieve cosmic speeds, K. E. Tsiolkovsky proposed the use of multi-stage rockets. The scientist himself figuratively called them “rocket trains.” According to Tsiolkovsky, a rocket train, or, as we say now, a multi-stage rocket, should consist of several rockets mounted on one another. The bottom rocket is usually the largest. She carries the entire “train” on herself. Subsequent steps are made of smaller and smaller sizes.

When taking off from the surface of the Earth, the engines of the lower rocket operate. They operate until all the fuel in its tanks is used up. When the tanks of the first stage are empty, it is separated from the upper rockets so as not to burden their further flight with dead weight. The separated first stage with empty tanks continues to fly upward for some time by inertia, and then falls to the ground. To preserve the first stage for reuse, it can be lowered by parachute.

After separation of the first stage, the second stage engines are switched on. They begin to operate when the rocket has already risen to a certain altitude and has a significant flight speed. The second stage engines accelerate the rocket further, increasing its speed by several kilometers per second. After all the fuel contained in the tanks of the second stage is consumed, it is also dumped. The further flight of the composite rocket is ensured by the operation of the third stage engines. Then the third stage is reset. The line is approaching the fourth stage engines. Having completed the work assigned to them, they increase the speed of the rocket by a certain amount, and then give way to the fifth stage engines. After the fifth stage is reset, the engines of the sixth begin to operate.

Thus, each stage of the rocket successively increases its flight speed, and the last, upper stage reaches the required cosmic speed in vacuum. If the task is to land on another planet and return back to Earth, then the rocket launched into space, in turn, must consist of several stages, sequentially switched on when descending to the planet and when taking off from it.

It is interesting to see the effect of using a large number of stages on rockets.

Let's take a single-stage rocket with a launch weight of 500 tons. Let us assume that this weight is distributed as follows: payload - 1 ton, stage dry weight - 99.8 tons and fuel - 399.2 tons. Consequently, the structural perfection of this rocket is such that the weight fuel is 4 times the dry weight of the stage, that is, the weight of the rocket itself without fuel and payload. The Tsiolkovsky number, that is, the ratio of the launch weight of the rocket to its weight after all the fuel has been consumed, for this rocket will be equal to 4.96. This number and the speed at which gas flows out of the engine nozzle determine the speed that the rocket can achieve. Let's now try to replace the single-stage rocket with a two-stage one. Let us again take a payload of 1 ton and assume that the design perfection of the stages and the gas flow rate will remain the same as in a single-stage rocket. Then, as calculations show, to achieve the same flight speed as in the first case, a two-stage rocket with a total weight of only 10.32 tons will be required, that is, almost 50 times lighter than a single-stage one. The dry weight of a two-stage rocket will be 1.86 tons, and the weight of the fuel placed in both stages will be 7.46 tons. As we can see, in the example under consideration, replacing a single-stage rocket with a two-stage one makes it possible to reduce metal and fuel consumption by 54 times when launching the same payload .

Let's take, for example, a space rocket with a payload of 1 ton. Let this rocket must penetrate the dense layers of the atmosphere and, flying into airless space, develop a second escape velocity of 11.2 km/sec. Our charts show weight changes like this space rocket depending on the weight fraction of fuel in each stage and the number of stages (see page 22).

It is easy to calculate that if you build a rocket whose engines eject gases at a speed of 2,400 m/sec and in each stage the fuel accounts for only 75% of the weight, then even with six stages, the take-off weight of the rocket will be very large - almost 5.5 thousand tons. By improving the design characteristics of rocket stages, it is possible to achieve a significant reduction in launch weight. So, for example, if fuel accounts for 90% of the stage's weight, then a six-stage rocket can weigh 400 tons.

An exceptionally great effect comes from using high-calorie fuel in rockets and increasing the efficiency of their engines. If in this way we increase the speed of gas flow from the engine nozzle by only 300 m/sec, bringing it to the value indicated on the graph - 2,700 m/sec, then the launch weight of the rocket can be reduced several times. A six-stage rocket, in which the weight of the fuel is only 3 times greater than the weight of the stage structure, will have a launch weight of approximately 1.5 thousand tons. And by reducing the weight of the structure to 10% of the total weight of each stage, we can reduce the launch weight of the rocket with the same number of stages up to 200 t.

If we increase the gas flow rate by another 300 m/sec, that is, take it equal to 3 thousand m/sec, then an even greater reduction in weight will occur. For example, a six-stage rocket with a fuel weight fraction of 75% will have a launch weight of 600 tons. By increasing the fuel weight fraction to 90%, it is possible to create a space rocket with only two stages. Its weight will be about 850 tons. By doubling the number of stages, you can reduce the weight of the rocket to 140 tons. And with six stages, the take-off weight will drop to 116 tons.

This is how the number of stages, their design perfection and gas flow rate affect the weight of the rocket.

Why, as the number of stages increases, the required fuel reserves decrease, and with them the total weight of the rocket? This happens because than larger number stages, the more often empty tanks will be discarded, the faster the rocket will be freed from useless cargo. In this case, as the number of stages increases, first the take-off weight of the rocket decreases very much, and then the effect of increasing the number of stages becomes less significant. It can also be noted, as can be clearly seen in the graphs above, that for rockets with relatively poor design characteristics, increasing the number of stages has a greater effect than for rockets with a high percentage of fuel in each stage. This is quite understandable. If the bodies of each stage are very heavy, then they must be dropped as quickly as possible. And if the hull is very light in weight, then it does not burden the missiles too much and frequent drops of empty hulls no longer have such a great effect.

When rockets fly to other planets, the required fuel consumption is not limited to the amount required for acceleration when taking off from Earth. Approaching another planet, the spacecraft falls into its sphere of gravity and begins to approach its surface with increasing speed. If the planet is deprived of an atmosphere capable of extinguishing at least part of the speed, then the rocket, when falling on the surface of the planet, will develop the same speed as is necessary to depart from this planet, that is, the second escape velocity. The value of the second escape velocity, as is known, is different for each planet. For example, for Mars it is 5.1 km/sec, for Venus - 10.4 km/sec, for the Moon - 2.4 km/sec. In the case when the rocket approaches the sphere of gravity of the planet, having a certain speed relative to the latter, the speed of the rocket's fall will be even greater. For example, the second Soviet space rocket reached the surface of the Moon at a speed of 3.3 km/sec. If the task is to ensure a smooth landing of the rocket on the surface of the Moon, then additional fuel reserves must be on board the rocket. To extinguish any speed, it is necessary to consume the same amount of fuel as is necessary for the rocket to develop the same speed. Consequently, a space rocket designed to safely deliver some cargo to the lunar surface must carry significant reserves of fuel. Single stage rocket with a payload of 1 ton should have a weight of 3-4.5 tons, depending on its design perfection.

Previously, we showed what enormous weight rockets must have in order to carry outer space a load of 1 ton. And now we see that of this load only a third or even a fourth can be safely lowered to the surface of the Moon. The rest must be fuel, fuel storage tanks, engine and control system.

What should ultimately be the starting weight of a space rocket designed to safely deliver scientific equipment or other payload weighing 1 ton to the lunar surface?

In order to give an idea of ships of this type, our figure conventionally shows a sectional view of a five-stage rocket designed to deliver a container with scientific equipment weighing 1 ton to the lunar surface. The calculation of this rocket was based on technical data given in a large number of books (for example, in the books by V. Feodosyev and G. Sinyarev “Introduction to Rocketry” and Sutton “Rocket Engines”).

Rocket engines running on liquid fuel were taken. To supply fuel to the combustion chambers, turbopump units driven by hydrogen peroxide decomposition products are provided. The average gas outflow velocity for the first stage engines is assumed to be 2,400 m/sec. Upper stage engines operate in highly rarefied layers of the atmosphere and in airless space, so their efficiency turns out to be somewhat greater and for them the gas outflow velocity is assumed to be 2,700 m/sec. For the design characteristics of the stages, the following values were adopted that are found in rockets described in the technical literature.

With the selected initial data, the following weight characteristics of the space rocket were obtained: take-off weight - 3,348 tons, including 2,892 tons - fuel, 455 tons - structure and 1 t - payload. The weight of the individual stages was distributed as follows: the first stage - 2,760 tons, the second - 495 tons, the third - 75.5 tons, the fourth - 13.78 tons, the fifth - 2.72 tons. The height of the rocket reached 60 m, the diameter of the lower stage - 10 m.

The first stage contains 19 engines with a thrust of 350 tons each. On the second - 3 of the same engines, on the third - 3 engines with a thrust of 60 tons. On the fourth - one with a thrust of 35 tons and on the last stage - an engine with a thrust of 10 tons.

When taking off from the surface of the Earth, the first stage engines accelerate the rocket to a speed of 2 km/sec. After the empty casing of the first stage is released, the engines of the next three stages are turned on, and the rocket acquires a second escape velocity.

Then the rocket flies by inertia towards the Moon. Approaching its surface, the rocket turns its nozzle down. The fifth stage engine turns on. It dampens the speed of fall, and the rocket smoothly descends to the lunar surface.

The above figure and the calculations related to it, of course, do not represent a real project for a lunar rocket. They are given only to give a first idea of the scale of multi-stage space rockets. It is absolutely clear that the design of a rocket, its dimensions and weight depend on the level of development of science and technology, on the materials available to the designers, on the fuel used and the quality of the rocket engines, on the skill of its builders. The creation of space rockets provides endless scope for the creativity of scientists, engineers, and technologists. There are still many discoveries and inventions to be made in this area. And with each new achievement, the characteristics of missiles will change.

Just as modern airships such as IL-18, TU-104, TU-114 are not similar to the airplanes that flew at the beginning of this century, so space rockets will be continuously improved. Over time, rocket engines will use more than just energy to fly into space. chemical reactions, but also other energy sources, for example the energy of nuclear processes. As the types of rocket engines change, the design of the rockets themselves will also change. But K. E. Tsiolkovsky’s wonderful idea to create “rocket trains” will always play an honorable role in the exploration of the vast expanses of space.

Scheme with load-bearing tanks

Transition circuit

Scheme with hanging tanks

SINGLE-STAGE LIQUID ROCKETS.

A lot of long-range liquid ballistic missiles and launch vehicles have been created to date. But we must start with the simplest and most obvious. Therefore, we will turn to the oldest one, which now only has historical significance German V-2 rocket. It is considered the first liquid-propellant ballistic missile.

The word "first", however, needs clarification. Already in the pre-war, thirties, the principles of the design of a ballistic liquid rocket were well known to specialists. Quite advanced liquid-propellant rocket engines already existed (primarily in the Soviet Union). Gyroscopic systems for stabilizing rockets have already been developed and created. The first samples of liquid-propellant rockets intended for exploration of the stratosphere have already been tested. Therefore, the V-2 rocket did not appear out of the blue. But it went into mass production first. It was also the first to find military use, when, in a paroxysm of despair, in 1943 the German command

gave the order for the senseless firing of this missile into residential areas of London. Of course, this step could not in any way affect the general course of military events. Much greater influence was exerted by the famous domestic rocket artillery, perfect samples of which were tested in the early days Patriotic War directly on the battlefields. But now we are not talking about the military use of missiles. No matter how sad the history of the V-2 missile is, in this case we are only interested in its design and layout principles. For us, this is a very convenient classroom aid that will help the reader get acquainted with common device in general, all ballistic liquid missiles, and not only with the device. From the heights of the experience accumulated to date, it is easy to evaluate this design and show how its advantages were subsequently developed and disadvantages were eliminated: in what ways technical progress took place.

The launch weight of the V-2 rocket was approximately 13 ts, and its range was close to 300 km. A cross-section of the rocket is shown on the poster.

The body of a liquid-propellant ballistic missile is divided along its length into several compartments (Fig. 3.1): fuel compartment (F.O), which includes fuel tanks 1 and oxidizer 2; the tail compartment (X. O) with the engine and the instrument compartment (P. O), to which the warhead (B. Ch) is docked. The very concept of “compartment” is associated not only with the functional purpose of some part of the rocket, but, first of all, with the presence of transverse connectors that allow separate assembly and subsequent docking. In some types of rockets, the instrument compartment is like independent part there is no housing, and control devices are placed block by block in free space, taking into account the convenience of approaches and maintenance at the start and the minimum length cable network.

Like all guided ballistic missiles, the V-2 is equipped with an automatic stabilization system. Gyro devices and other automatic stabilization units are located in the instrument compartment and mounted on a cross-shaped panel.

The executive bodies of the automatic stabilization system are gas-jet and air rudders. Gas jet rudders 3 are located in the stream flowing out of the chamber 4 gases and are mounted with their drives - steering gears - on a rigid steering ring 5 . When the rudders are deflected, a moment arises that turns the rocket in the desired direction. Since gas-jet rudders operate in extremely heavy temperature conditions, they were made from the most heat-resistant material - graphite. Air rudders 6 play a supporting role and produce an effect only in dense layers of the atmosphere and at a sufficiently high flight speed.

The V-2 rocket uses liquid oxygen and ethyl alcohol as fuel components. Since the acute problem of engine cooling could not be properly solved at that time, the designers decided to lose specific thrust by ballasting ethyl alcohol with water and reducing its concentration to 75%. The total supply of alcohol on board the rocket is 3.5 g, and liquid oxygen - 5 g.

The main elements of the engine located in the tail compartment are the camera 4 and turbopump unit (TNA) 7, designed to supply fuel components to the combustion chamber.

The turbopump unit consists of two centrifugal pumps - alcohol and oxygen, installed on a common shaft with a gas turbine. The turbine is driven by the decomposition products of hydrogen peroxide (water vapor + oxygen), which are formed in the so-called steam and gas generator (PGG)(not visible in the picture). Hydrogen peroxide is supplied to the GHG reactor from the tank 3 and decomposes in the presence of a catalyst - an aqueous solution of sodium permanganate supplied from the tank 9. These components are forced out of the tanks by compressed air contained in the cylinders 10. Thus, the operation of the propulsion system is ensured by a total of four components - two main and two auxiliary for steam and gas generation. We should not, of course, forget about compressed air, the supply of which is necessary for supplying auxiliary components and for the operation of pneumatic automation.

The items listed are the camera, TNA, tanks of auxiliary components, compressed air cylinders - together with supply pipelines, valves and other fittings are mounted on a load-bearing frame 11 and form a common energy block, which is called liquid rocket engine (LPRE).

When assembling the rocket, the engine frame is docked to the rear frame 12 and is closed by a thin-walled reinforced shell - the body of the tail section, equipped with four stabilizers.

The thrust of the V-2 rocket engine on Earth is 25 ts, and in emptiness - about 30 ts. If this thrust is divided by the total weight flow, consisting of 50 kgf/sec alcohol, 75 kgf/sec oxygen and 1.7 kgf/sec hydrogen peroxide and permanganate, we get a specific thrust of 198 and 237 units on Earth and in vacuum, respectively. According to modern concepts, such a specific thrust for liquid engines is, of course, considered very low.

Let's turn to the so-called power circuit. It is difficult to find a short and clear definition for this concept, which is quite clear in meaning. The power circuit is a design solution based on considerations of the strength and rigidity of the entire structure, its ability to withstand the loads acting on the rocket as a whole.

An analogy can be drawn. In higher animals the power circuit is skeletal. The bones of the skeleton are the main load-bearing elements that support the body and absorb all muscle efforts. But the skeletal diagram is not the only one. The shell of a crayfish, crab and other similar creatures can be considered not only as a means of protection, but also as an element of the overall power scheme. Such a scheme should be called shell. With a deeper understanding of biology, one could presumably find examples of other force circuits in nature. But now we are talking about the power circuit of the rocket structure.

At the launch site of the V-2 rocket, engine thrust is transferred to the rear power frame 12. The rocket moves with acceleration, and an axial compressive force arises in all cross sections of the body located above the power frame. The question is which elements of the hull should receive it - tanks, longitudinal reinforcements, a special frame, or perhaps enough

create increased pressure in the tanks, and then the structure will acquire load-bearing capacity like a well-inflated car tire. The solution to this issue is the subject of choosing a power circuit.

The V-2 rocket adopts the design of an external power body and external tanks. Power Corps 13 It is a steel shell with a longitudinal-transverse set of reinforcing elements. Longitudinal reinforcing elements are called stringers, and the most powerful of them are spars. The transverse ring elements are called frames. For ease of installation, the rocket body has a longitudinal bolt connector.

Lower oxygen tank 2 rests on the same power frame 12, to which, as already mentioned, the engine frame with tail fairing is attached. The alcohol tank is suspended on the front power frame 14, with which the instrument compartment is also connected.

Thus, in the V-2 rocket, fuel tanks play only the role of containers and are not included in the power circuit, and the main power element is the rocket body. But it is calculated not only for the load of the launch site. It is also important to ensure the strength of the rocket when approaching the target, and this circumstance deserves special discussion.

After the engine is turned off, the gas-jet rudders cannot perform their functions, and since the shutdown is performed at a high altitude, where there is practically no atmosphere, the air rudders and tail stabilizer also completely lose their effectiveness. Therefore, after the engine is turned off, the rocket becomes unorientable. The flight occurs in a mode of indefinite rotation relative to the center of mass. Upon entering the relatively dense layers of the atmosphere, the tail stabilizer orients the missile along the flight, and at the final part of the trajectory it moves with the head part forward, slowing down somewhat in the air, but maintaining a speed of 650-750 by the time it meets the target m/sec.

The stabilization process is associated with the occurrence of large aerodynamic loads on the body and tail. This is an uncontrolled flight with angles of attack varying within ±180°. The casing heats up, and significant bending moments arise in the cross sections of the body, for which strength calculations are mainly carried out.

At first impression, it seems unclear whether it is really necessary to care about the strength of the rocket in the final part of the trajectory. The rocket has almost reached, and the job seems to be done. Even if the body is destroyed, the warhead will still reach the target, the fuses will go off, and the destructive effect of the rocket will be ensured.

This approach, however, is unacceptable. There is no guarantee that if the case is destroyed, the combat charge itself will not be damaged, and such damage, combined with local overheating, is fraught with a premature trajectory explosion. In addition, in conditions of structural destruction, the process of subsequent movement is obviously unpredictable. Even a serviceable, non-destructive rocket even receives some indefinite change in the velocity vector during the atmospheric phase of free flight. Aerodynamic forces can and do lead the rocket away from its intended trajectory. In addition to the inevitable errors for the launch site, new unaccounted errors appear. The missile falls undershot, overshot, or falls to the right or left of the target. Dispersion occurs, which, due to the uncertain conditions of entry into the atmosphere, increases noticeably. If we accept the destruction of the hull and, accordingly, the loss of stabilization and speed, then the prolonged uncertainty of movement will lead to an unacceptable increase in dispersion. Something similar happens to what we see when we follow the trajectory of falling leaves: the same uncertainty of the trajectory and the same loss of speed. By the way, reducing the speed at the target for a combat missile like "V-2" also undesirable. The kinetic energy of the rocket's mass and the energy of the explosion of the remaining fuel components for this type of weapon gave a quite noticeable increase in the combat effect of the tons of explosive located in the head of the rocket.

So, the rocket body must be strong enough in all parts of the trajectory. And if now, without going into details, take a critical look at the V-2 rocket as a whole, then we can conclude that it is the power circuit that is most weak point this design, since the need to excessively strengthen the body significantly reduces the weight characteristics of the rocket. Therefore, it is necessary to look for another constructive solution.

When analyzing the power circuit, the idea naturally arises of abandoning the load-bearing body and assigning power functions to the walls of the tanks, perhaps additionally strengthening them and maintaining moderate internal pressure. But this solution is only suitable for the active section. As for stabilizing the rocket when returning to the atmospheric part of the trajectory, this will have to be abandoned and the warhead will have to be made detachable.

Thus, a power circuit with load-bearing tanks is born. Fuel tanks must satisfy strength conditions only under regulated, predetermined loads and thermal conditions of the active section. After turning off the engine, the head section, equipped with its own aerodynamic stabilizer, separates. From this moment, the rocket body with the propulsion system already turned off and the warhead fly almost along a common trajectory, separately and without a specific angular orientation. Upon entering the dense layers of the atmosphere, the body, which has high aerodynamic resistance, begins to lag behind, collapses, and its parts fall without reaching the target. The warhead stabilizes, maintains a relatively high speed and delivers the warhead to a given point. With this scheme, it is clear that the kinetic energy of the rocket mass is not included in the effect combat action. However, reducing the overall weight of the structure makes it possible to compensate for this loss by increasing the payload. In the case of a transition to a nuclear warhead, the kinetic energy of the missile mass does not matter at all.

Now let's see what we gain and what we lose; what are the assets and liabilities when moving to the scheme of supporting tanks and a detachable head section. Obviously, the absence of a power body and the absence of a tail stabilizer, the need for which is now eliminated, should be noted as an asset. An asset should include the possibility of switching from steel to lighter aluminum-magnesium alloys: the rocket passes through the atmospheric launch phase at a relatively low speed, and the heating of the body is small. And finally, there is one more important circumstance. The calculated loads on the active section have a fairly high degree of reliability; they are regulated by precisely maintained breeding conditions. As for re-entry into the atmosphere, for this section the load trajectories are determined with less accuracy. Reliance on the calculated loads of the active section makes it possible to reduce the assigned safety factor, which for a rocket with a separating warhead results in additional weight reduction.

The liability will have to include some increase in the weight of the tanks; they need to be strengthened. You may have to add the additional weight of compressed air and fuel tank pressurization systems here. The weight of the new head stabilizer will also be recorded as a liability. But, of course, such a stabilizer weighs much less than the old one, intended for the rocket as a whole. And finally, some rudiments in the form of so-called pylons may remain from the old stabilizer. They have two tasks. The pylons provide some stabilizing effect, which makes it possible to somewhat simplify the operating conditions of the stabilization machine. In addition, the pylons allow the air rudders, if any, to be moved away from the hull into a free and “unshaded” aerodynamic flow.

Naturally, in such arguments for and against one cannot be content with only speculative statements. Detailed design analysis, numerical estimates and calculations are required. And such a calculation indicates the undoubted weight advantages of the new power scheme.

The above considerations apply only to rockets that have a turbopump feed system. If the components are supplied by high pressure created in the fuel tanks (such a supply is called displacement), then the logic of the power circuit changes somewhat.

In the case of displacement feed, fuel tanks are designed primarily for internal pressure, and, satisfying the pressure strength condition, such tanks, as a rule, automatically satisfy both strength and temperature requirements in all flight modes. Consequently, they were destined to be carriers. Suspended tanks with displacement feeding would be an obvious absurdity.

A tank designed for high internal pressure of displacement supply, as a rule, also satisfies the condition of the strength of the hull upon re-entry into the atmosphere. Consequently, separation of the warhead for such a rocket is not necessary, but then the body must be equipped with a tail stabilizer.

The idea of a detachable warhead was first implemented in 1949 on one of the earliest domestic ballistic missiles, the R-2. On its basis, a geophysical modification of the rocket, B2A, was created somewhat later. The design of the B2A rocket is an interesting and instructive hybrid version of the old and new emerging power schemes and deserves discussion as an example of the development of design thought.

The rocket has only one load-bearing tank - the front, alcohol tank, and the oxygen tank is placed in a lightweight power housing, designed only for the loads of the active section. Detachable head 2 equipped with its own tail stabilizer 3, representing a reinforced shell in the shape of a truncated cone. In the geophysical version, the stabilizer 3 the salvageable head part has a mechanism for opening the brake flaps 4, which reduce the rate of fall of the head part to 100-150 m/sec, after which the parachute opens. Figure 2 shows the head section after landing. The crumpled nasal shock-absorbing tip is visible 1 and open shields 4, partially melted during braking in the atmosphere.

The end frame of the head stabilizer is attached with special locks to the support frame located in the upper part of the alcohol tank. After the command to separate, the locks open, and the head part receives a small impulse from the spring pusher.

Instrument compartment 8 has freely unlocked locking hatches with sealing and is located not in the upper, but in the lower part of the rocket, which provides certain convenience for pre-launch operations.

Looking at the B2A rocket in more detail, one could note its other features. But that's not the main point. A striking and at the same time very instructive feature of this design is the logical discrepancy between the principle of a detachable nose section and the presence of a tail stabilizer. At the launch site, the missile's orientation is ensured by a stabilization machine. As for aerodynamic stabilization when entering dense layers of the atmosphere, the tail unit cannot help here, since the body does not have the necessary strength for this.

Of course, it would be naive to believe that the designers did not see or understand this. The design, simply put, was common, often encountered in engineering practice technical compromise- a concession to temporary circumstances. Experience has already been gained in creating rockets with a stabilizer circuit and with external tanks. The proven system of gas-jet and air rudders was reliable and did not cause concern, and the automatic stabilization system did not require serious readjustment, which would be inevitable when moving to new aerodynamic forms. Therefore, in a situation when there were still theoretical discussions about the dangers of switching to a non-stabilized aerodynamically unstable scheme, it was easier, without waiting for the creation of new proven control systems, to stay with the old one. Having lost something in weight, it was easier to establish a position in certain already won positions. On the way to the real implementation of the scheme with load-bearing tanks, it was necessary to find something between the desire to quickly achieve the goal and the danger of lengthy experimental development, between the inevitable readjustment of production and the use of existing workshop equipment, between the risk of failure and reasonable forethought. Otherwise, a series of failures during launches, which is not at all impossible, could compromise the idea at its very core and give food to persistent distrust in new scheme, no matter how promising and logical it may be.

And one more, not so important, but interesting psychological aspect. The design of the B2A rocket did not seem unusual at that time. The force of the habit of seeing the tail fin on all the small and large rockets that existed before preserved the illusion of routine for an outside observer, and the appearance of the rocket did not provoke premature and unqualified criticism of the design as a whole. The same can be said about the design of the oxygen tank. The use of liquid oxygen was the focus of dissenting opinions at the time, based on concerns about the low boiling point of this fuel component. The presence of thermal insulation of the oxygen tank on the B2A rocket reassured many and did not overload the already sufficient range of concerns facing the chief designer. It was necessary to show that the supporting alcohol tank regularly performs power functions, that the head part is successfully separated and safely reaches the target, and that the automation and control devices located near the engine, despite the increased level of vibration, are able to work as well as they worked when they were in the head compartment.

The transition to a new power scheme was associated, naturally, with the simultaneous solution of a number of other fundamental issues. This concerned, first of all, the design of the engine. The RD-101 engine installed on the V2A rocket provided 37 and 41.3 ts earth and void thrust or 214 and 242 units of specific thrust at the Earth's surface and in void, respectively. This was achieved by increasing the alcohol concentration to 92%, increasing the pressure in the chamber and additionally expanding the nozzle exit section.

The engine creators abandoned the liquid catalyst for the decomposition of hydrogen peroxide. It was replaced by a solid catalyst, which was placed in advance into the working cavity of the steam and gas generator. Thus, the number of liquid components was reduced from four, as was the case with the V-2, to three. A new, soon to become traditional, torus cylinder for hydrogen peroxide also appeared, conveniently fitting into the rocket’s layout. The beginning was also made of some other innovations, which it makes no sense to list here.

Naturally, the B2A rocket, as a transitional version from one power scheme to another, could not and should not have been reproduced in subsequent modernized forms. It was necessary to fully implement the idea of load-bearing tanks and a detachable warhead, which was done by S.P. Korolev in subsequent developments.

The first samples of missiles with load-bearing tanks were tested and developed in the early 50s. After that, some modifications were worked out. Thus, in particular, the B5B meteorological missile (R-5 combat missile) appeared. Nowadays, a prototype of a ballistic missile with load-bearing tanks occupies a place of honor as a historical exhibit in front of the museum entrance. Soviet Army in Moscow.

When switching to a new modernized design, in order to increase the range, the starting weight was increased and the engine operating mode was forced. The transition to a load-bearing tank scheme is, of course, more high level technology and careful design work made it possible to increase the weight quality coefficient α k to 0.127 (instead of 0.25 for the V-2) with a relative final weight µ k ~ 0.16.

The control system was subjected to the most serious changes in the B5B rocket. After all, it was the first aerodynamically unstable rocket equipped with a very small tail unit and air rudders. Later on the same rocket was used for the first time to use a gyroplatform and new principle functional engine shutdown.

The B5B rocket still used 92% ethyl alcohol and liquid oxygen as fuel. Testing of the rocket showed that the lack of thermal insulation on the side surface of the oxygen tank does not entail unpleasant consequences. The slightly increased evaporation of oxygen during pre-launch preparation is easily compensated for by make-up, i.e., automated refueling of oxygen immediately before the start. This operation is generally necessary for all rockets using low-boiling fuel components.

Thus, after the B5B rocket, the design of the load-bearing tanks and detachable warhead became a reality. All modern long-range liquid-propellant ballistic missiles and their higher stage - launch vehicles - are now created only on the basis of this power scheme. It is its development based on modern technology and countless design improvements gave rise to a generalized image of that machine, which rightly symbolizes the heights of technical progress of our time.

Now the B5B rocket can be viewed as critically as the V-2 rocket was viewed at the time of its creation. While maintaining the general layout and the basic principles of the power circuit, it is possible to further reduce weight and increase the main characteristics, and ways to solve this problem are easily visible and understood using examples of later designs.

In Fig. 3.3 shows a single-stage version of the American Thor ballistic missile; it is also made according to the typical design of load-bearing tanks and has a detachable head part. The total weight of fuel components (oxygen + kerosene) is 45 ts with a net weight of the structure (without the head part) of 3.6 ts. This means the following. If we conditionally take the total weight of fuel residues to be 0.4 ts, then for the familiar weight quality coefficient α k we get a value of 0.082. Taking the weight of the head part approximately 2 ts, we obtain the parameter µ K = 0.12. It can also be established that with the specific void thrust of oxygen-kerosene fuel assumed to be 300 units, the range of this rocket is 3000 km.

The high weight indicators of modern missiles, in particular this one, are based on a careful study of many elements, which would be very difficult to list, but some, quite general and typical, can be indicated.

Fuel tank walls 1 And 2 have a waffle design. This is a thin-walled shell made of high-strength aluminum alloy with frequently located longitudinal-transverse reinforcements, playing the same role as the power set in the body of the V-2 rocket, but with greater weight quality. The now widespread wafer structure is usually produced by mechanical milling. In some cases, however, chemical milling is also used. Shell blank of original thickness h 0 undergoes carefully controlled etching in acid along the part of the surface where it is necessary to remove excess metal (the rest of the surface is first coated with varnish). Thickness remaining after etching h should ensure the tightness and strength of the resulting panel at a given internal pressure, and the longitudinal and transverse ribs provide the shell with increased bending rigidity, which determines the stability of the structure under axial compression. The regularity of the distribution of longitudinal and transverse ribs is deliberately disrupted in the area welds, which, as is known, have a slightly reduced strength compared to rolled sheets, as well as at the ends of the shell, where the bottoms have yet to be welded. In these places, the thickness of the workpiece remains unchanged.

There are other ways to make waffle structures. However, we deliberately focused on chemical milling in order to show at what cost, literally and figuratively, the design weight indicators that are characteristic of modern rocketry are achieved.

The Thor rocket has a shortened and lightweight tail section Z, at the end of which two control motors are mounted. The rejection of gas-jet rudders is naturally associated with their high gas-dynamic resistance in the stream of escaping gases. The use of control motors somewhat complicates the design, but provides a significant gain in specific thrust.

From the above, one should not get the impression that control chambers appeared for the first time on this ballistic missile. This system of power control elements has been used in various versions before, in particular, on the Vostok or Soyuz launch vehicle, which will be discussed later. The single-stage version of the Thor missile is considered here solely as an example of the next generation of ballistic missiles after the B5B missile.

Almost all ballistic missiles Braking solid fuel engines are also installed 6. This is also not one of the latest innovations. The task of the braking engines is to, by braking the rocket body, move it away from the head part when it separates; namely, the body, without imparting additional speed to the head.

Shutting down a liquid engine is not instantaneous. After closing the valves of the fuel lines, combustion and evaporation of the remaining components still continue in the chamber for the next fraction of a second. As a result, the rocket receives a small additional impulse, called aftereffect impulse. When calculating the range, a correction is introduced for it. However, this is definitely impossible to do, since the aftereffect impulse does not have stability and changes from case to case, which is one of the significant reasons for range dispersion. In order to reduce this dispersion, braking motors are used. The moment of their activation is coordinated with the command to turn off the liquid engine in such a way that the aftereffect impulse is basically compensated.

It will be instructive to compare the geometric proportions of the B5B and Thor missiles. The B5B rocket is more elongated. The ratio of length to diameter (the so-called rocket extension) for it is significantly more than that of the Thor rocket; approximately 14 versus 8. The difference in elongations also raises different concerns. With increasing elongation, the frequency of the rocket's own transverse oscillations, like an elastic beam, decreases, and this forces one to take into account the disturbances that arrive at the input of the stabilization system as a result of angular movements when the body is bent. In other words, stabilization of a bending rocket, rather than a rigid one, must be ensured. In some cases this causes serious difficulties,

With a small elongation of the rocket, this issue naturally disappears, but another nuisance arises - the role of disturbances from transverse vibrations of the liquid in the tanks increases, and if the proper selection of the parameters of the stabilization machine fails to fend them off, it is necessary to install tanks partitions that limit fluid mobility. The figure partially shows units 7 for attaching vibration dampers in the fuel tank. Naturally, such a solution leads to a deterioration in the weight characteristics of the rocket.

The Thor rocket should not be viewed as a model of perfection. At the same time, the designers could probably counter any critical remarks about its layout with their own counter-arguments. Using the example of the B2A rocket, we have already seen that justified criticism of a design solution can only be carried out taking into account the specific conditions of design and production, and most importantly, the long-term tasks that the creators of the new machine set for themselves. And the Thor rocket is one of those on the basis of which it is possible to create rocket and space systems.

2. Operating principle of a multi-stage rocket

The rocket is very “costly” vehicle. Spacecraft launch vehicles “transport” mainly the fuel necessary to operate their engines and their own structure, consisting mainly of fuel containers and a propulsion system. The payload accounts for only a small portion of the rocket's launch mass.

A composite rocket allows for a more efficient use of resources due to the fact that during flight a stage that has exhausted its fuel is separated, and the rest of the rocket fuel is not wasted on accelerating the design of the spent stage, which has become unnecessary to continue the flight. An example of a calculation confirming these considerations is given in the article Tsiolkovsky Formula.

Missile configuration options. From left to right:
1. single-stage rocket;
2. two-stage rocket with transverse separation;
3. two-stage rocket with longitudinal separation.
4. A rocket with external fuel tanks that are separated after the fuel in them is exhausted.

Three-stage transversely separated Saturn V rocket without adapters

Structurally, multistage rockets are made with transverse or longitudinal separation of stages.
With transverse separation, the stages are placed one above the other and work sequentially one after another, turning on only after the separation of the previous stage. This scheme makes it possible to create systems, in principle, with any number of stages. Its disadvantage is that the resources of subsequent stages cannot be used in the work of the previous one, being a passive load for it.

Three-stage launch vehicle with longitudinal-transverse separation Soyuz-2.

With longitudinal separation, the first stage consists of several identical rockets operating simultaneously and located symmetrically around the body of the second stage, so that the resultant thrust forces of the first stage engines are directed along the axis of symmetry of the second. This scheme allows the engine of the second stage to operate simultaneously with the engines of the first, thus increasing the total thrust, which is especially necessary during the operation of the first stage, when the mass of the rocket is maximum. But a rocket with longitudinal separation of stages can only be two-stage.
There is also a combined separation scheme - longitudinal-transverse, which allows you to combine the advantages of both schemes, in which the first stage is divided from the second longitudinally, and the separation of all subsequent stages occurs transversely. An example of this approach is the domestic carrier Soyuz.

Space Shuttle layout.
The first stage is side solid propellant boosters.
The second stage is an orbiter with a detachable external fuel tank. At start, the engines of both stages are started.

Launch of the Space Shuttle.

The Space Shuttle has a unique design of a two-stage longitudinally separated rocket, the first stage of which consists of two side-mounted solid rocket boosters, and the second stage contains part of the fuel in the orbiter tanks, and most of it in a detachable external fuel tank. First, the orbiter propulsion system consumes fuel from the external tank, and when it is depleted, the external tank is reset and the engines continue to operate on the fuel contained in the orbiter tanks. This scheme makes it possible to make maximum use of the orbiter’s propulsion system, which operates throughout the entire launch of the spacecraft into orbit.

When transversely separated, the stages are connected to each other by special sections - adapters - load-bearing structures of cylindrical or conical shape, each of which must withstand the total weight of all subsequent stages, multiplied by the maximum value of the overload experienced by the rocket in all flight segments in which this adapter is included. rockets.
With longitudinal separation, power bands are created on the body of the second stage, to which the blocks of the first stage are attached.
The elements connecting the parts of a composite rocket give it the rigidity of a solid body, and when the stages are separated, they should almost instantly release the upper stage. Typically, the steps are connected using pyrobolts. A pyrobolt is a fastening bolt, in the rod of which a cavity is created next to the head, filled with a high explosive with an electric detonator. When a current pulse is applied to the electric detonator, an explosion occurs, destroying the bolt rod, causing its head to come off. The amount of explosives in the pyrobolt is carefully dosed in order, on the one hand, to ensure that the head comes off, and, on the other, not to damage the rocket. When the stages are separated into electric detonators of all pyrobolts connecting the separated parts, a current pulse is simultaneously applied and the connection is released.
Next, the steps should be spaced a safe distance from each other. When separating stages in the atmosphere, the aerodynamic force of the oncoming air flow can be used to separate them, and when separating in the void, auxiliary small solid rocket engines are sometimes used.
On liquid-propellant rockets, these same engines also serve to “sediment” the fuel in the tanks of the upper stage: when the engine of the lower stage is turned off, the rocket flies by inertia, in a state of free fall, while liquid fuel in tanks is suspended, which can lead to failure when starting the engine. Auxiliary engines provide the stage with a slight acceleration, under the influence of which the fuel “settles” on the bottom of the tanks.
In the above photo of the Saturn 5 rocket, on the body of the third stage, the black body of one of the auxiliary solid propellant propulsion engines of the 3rd and 2nd stages is visible.

Increasing the number of steps gives a positive effect only up to a certain limit. The more stages, the greater the total mass of adapters, as well as engines operating only on one part of the flight, and, at some point, a further increase in the number of stages becomes counterproductive. In modern rocket science practice, more than four stages, as a rule, are not made.

When choosing the number of stages, reliability issues are also important. Pyrobolts and auxiliary solid propellant rocket motors are single-use elements, the functioning of which cannot be checked before the launch of the rocket. Meanwhile, the failure of just one pyrobolt can lead to an emergency termination of the rocket's flight. An increase in the number of disposable elements that are not subject to functional testing reduces the reliability of the entire rocket as a whole. This also forces designers to refrain from using too many steps.

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