What are multistage rockets? Multistage missile: Ministry of Defense of the Russian Federation. Single stage liquid propellant rockets

On fig. 22 shows that the trajectory of a ballistic missile, and hence the range of its flight, depends on the initial velocity V 0 and the angle Θ 0 between this velocity and the horizon. This angle is called the throw angle.

Let, for example, the throwing angle is equal to Θ 0 = 30°. In this case, the rocket, which started its ballistic flight at point 0 with a speed V 0 = 5 km/sec, will fly along the elliptic curve II. At V 0 = 8 km/sec, the rocket will fly along an elliptical curve III, at V 0 = 9 km/sec, along curve IV. When the speed is increased to 11.2 km/s, the trajectory from a closed elliptic curve will turn into an open parabolic one and the rocket will leave the sphere of gravity of the earth (curve V). At an even higher speed, the rocket will escape along a hyperbole (VI). This is how the rocket trajectory changes with a change in the initial speed, although the angle of throw remains unchanged.

If you keep the initial speed constant, and change only the angle of throw, then the trajectory of the rocket will undergo no less significant changes.

Let, for example, the initial "speed is equal to V 0 = 8 km / h. If the rocket is launched vertically upwards (throwing angle Θ 0 = 90 °), then theoretically it will rise to a height equal to the radius of the Earth and return to Earth not far from the start ( VII) At Θ 0 = 30°, the rocket will fly along the elliptical trajectory we have already considered (curve III).Finally, at Θ 0 = 0° (launch parallel to the horizon), the rocket will turn into a satellite of the Earth with a circular orbit (curve I).

These examples show that only by changing the angle of throw, the range of missiles at the same initial speed of 8 km / s can have a range from zero to infinity.

At what angle will the rocket start its ballistic flight? It depends on the control program that is given to the rocket. It is possible, for example, for each initial speed to choose the most advantageous (optimal) throwing angle at which the flight range will be the greatest. As the initial speed increases, this angle decreases. The resulting approximate values ​​​​of the range, altitude and flight time are shown in Table. four.

Table 4

If the throwing angle can be changed arbitrarily, then the change in the initial speed is limited, and its increase by every 1 km / s is associated with great technical problems.

K. E. Tsiolkovsky gave a formula that makes it possible to determine the ideal speed of a rocket at the end of its acceleration by engines:

V id \u003d V ist ln G start / G end,

where V id - the ideal speed of the rocket at the end of the active section;

V ist - the speed of the outflow of gases from the jet nozzle of the engine;

G beg - the initial weight of the rocket;

G con - the final weight of the rocket;

ln is the sign of the natural logarithm.

We got acquainted with the value of the speed of the outflow of gases from the nozzle of a rocket engine in the previous section. For liquid fuels given in table. 3, these speeds are limited to 2200 - 2600 m / s (or 2.2 - 2.6 km / s), and for solid fuels - to 1.6 - 2.0 km / s.

G start denotes the initial weight, i.e., the total weight of the rocket before launch, and G end is its final weight at the end of acceleration (after running out of fuel or turning off the engines). The ratio of these weights G beg /G con, included in the formula, is called the Tsiolkovsky number and indirectly characterizes the weight of the fuel used to accelerate the rocket. Obviously, the larger the Tsiolkovsky number, the greater the speed the rocket will develop and, consequently, the farther it will fly (ceteris paribus). However, the Tsiolkovsky number, as well as the speed of the outflow of gases from the nozzle, has its limitations.

On fig. 23 shows a section of a typical single-stage rocket and its weight diagram. In addition to fuel tanks, the rocket has engines, controls and control systems, skin, payload, and various structural elements and auxiliary equipment. Therefore, the final weight of the rocket cannot be many times less than its initial weight. For example, the German V-2 rocket weighed 3.9 tons without fuel, and 12.9 tons with fuel. This means that the Tsiolkovsky number of this rocket was: 12.9 / 3.9 = 3.31. At the current level of development of foreign rocket science, this ratio for foreign rockets reaches 5–7.

Let's calculate the ideal speed of a single-stage rocket, taking V 0 = 2.6 km/sec. and G start / G end = 7,

V id \u003d 2.6 ln 7 \u003d 2.6 1.946 ≈ 5 km / s.

From Table. 4 shows that such a missile is capable of reaching a range of about 3,200 km. However, its actual speed will be less than 5 km/sec. since the engine expends its energy not only on rocket acceleration, but also on overcoming air resistance, on overcoming the force of gravity. The actual speed of the rocket will be only 75 - 80% of the ideal. Consequently, it will have an initial speed of about 4 km/sec and a range of no more than 1800 km*.

* (The range given in table. 4 is given approximately, since a number of factors were not taken into account when calculating it. For example, sections of the trajectory lying in dense layers of the atmosphere and the influence of the Earth's rotation were not taken into account. When firing in an easterly direction, the flight range of ballistic missiles is greater, since the speed of rotation of the Earth itself is added to their speed relative to the Earth.)

To create an intercontinental ballistic missile, launch artificial Earth satellites and spacecraft, and even more so to send space rockets to the Moon and planets, it is necessary to impart a significantly higher speed to the carrier rocket. So, for a missile with a range of 9000 - 13000 km, an initial speed of about 7 km / s is required. The first cosmic velocity that must be given to a rocket so that it can become a satellite of the Earth with a low orbital altitude is, as is known, 8 km/sec.

To exit the Earth's sphere of gravity, the rocket must be accelerated to the second cosmic velocity - 11.2 km / s, to fly around the Moon (without returning to Earth) a speed of more than 12 km / s is required. A flyby of Mars without returning to Earth can be carried out at an initial speed of about 14 km/s, and with a return to orbit around the Earth - about 27 km/s. A speed of 48 km/s is required to reduce the duration of a flight to Mars and back to three months. Increasing the speed of the rocket, in turn, requires the expenditure of an ever-increasing amount of fuel for acceleration.

Suppose, for example, we have built a rocket weighing 1 kg without fuel. If we want to tell her the speed of 3, 6, 9 and 12 km / s, then how much fuel will need to be filled into the rocket and burned during acceleration? The required amount of fuel * is shown in table. 5.

* (With an outflow velocity of 3 km/sec.)

Table 5

There is no doubt that in the body of a rocket, the "dry" weight of which is only 1 kg, we will be able to accommodate 1.7 kg of fuel. But it is very doubtful that it can accommodate his 6.4 kg. And, obviously, it is absolutely impossible to fill it with 19 or 54 kg of fuel. A simple but strong enough tank that can hold such an amount of fuel already weighs much more than a kilogram. For example, a twenty-liter canister known to motorists weighs about 3 kg. The "dry" weight of the rocket, in addition to the tank, should include the weight of the engines, structure, payload, etc.

Our great compatriot K. E. Tsiolkovsky found another (and so far the only) way to solve such a difficult task as achieving the rocket speeds that are required by practice today. This path consists in the creation of multi-stage rockets.

A typical multi-stage rocket is shown in Fig. 24. It consists of a payload AND several detachable stages with a power plant and a supply of fuel in each. The engine of the first stage informs the payload, as well as the second and third stages (the second sub-rocket) with the speed ν 1 . After the fuel is used up, the first stage separates from the rest of the rocket and falls to the ground, and the second stage engine is turned on on the rocket. Under the action of its thrust, the remaining part of the rocket (the third sub-rocket) acquires an additional speed ν 2 . Then the second stage, after running out of fuel, also separates from the rest of the rocket and falls to the ground. At this time, the third stage engine turns on and informs the payload of the additional speed ν 3 .

Thus, in a multi-stage rocket, the payload accelerates many times. The total ideal speed of a three-stage rocket will be equal to the sum of the three ideal speeds obtained from each stage:

V id 3 \u003d ν 1 + ν 2 + ν 3.

If the speed of the outflow of gases from the engines of all stages is the same and after the separation of each of them the ratio of the initial weight of the remaining part of the rocket to the final one does not change, then the speed increments ν 1 , ν 2 and ν 3 will be equal to each other. Then we can assume that the speed of a rocket consisting of three (or even n) stages will be equal to triple (or increased by n times) the speed of a single-stage rocket.

In fact, in each stage of multi-stage rockets there may be engines that give different exhaust velocities; a constant weight ratio may not be maintained; air resistance as the flight speed changes and the attraction of the Earth as you move away from it change. Therefore, the final speed of a multi-stage rocket cannot be determined by simply multiplying the speed of a single-stage rocket by the number of stages*. But it remains true that by increasing the number of stages, the speed of the rocket can be increased many times over.

* (It should also be borne in mind that between turning off one stage and turning on another, there may be a time interval during which the rocket flies by inertia.)

In addition, a multi-stage rocket can provide a given range of the same payload at a much lower total fuel consumption and launch weight than a single-stage rocket. Has the human mind managed to circumvent the laws of nature? No. Just a person, having learned these laws, can save on fuel and weight of the structure, performing the task. In a single-stage rocket, from the very start to the end of the active section, we accelerate all of its "dry" weight. In a multi-stage rocket, we don't do that. So, in a three-stage rocket, the second stage no longer spends fuel to accelerate the "dry" weight of the first stage, because the latter is discarded. The third stage also does not waste fuel for acceleration of the "dry" weight of the first and second stages. It accelerates only itself and the payload. The third (and in general the last) stage could no longer be disconnected from the head of the rocket, because further acceleration is not required. But in many cases, it still separates. Thus, the separation of the last stages is practiced in carrier rockets of satellites, space rockets and such combat missiles as Atlas, Titan, Minuteman, Jupiter, Polaris, etc.

When scientific equipment placed in the head part of the rocket is launched into space, separation of the last stage is envisaged. This is necessary for the correct functioning of the equipment. When a satellite is launched, it is also provided for its separation from the last stage. Due to this, resistance is reduced and it can exist for a long time. When launching a combat ballistic missile, the separation of the last stage from the combat head is provided, as a result of which it becomes more difficult to detect the combat head and hit it with an anti-missile. Moreover, the last stage separated during the descent of the rocket becomes a decoy. If during reentry into the atmosphere it is planned to control the warhead or stabilize its flight, then without the last stage it is easier to control it, since it has a smaller mass. Finally, if the last stage is not separated from the combat head, then it will be necessary to protect both from heating and combustion, which is unprofitable.

Of course, the problem of obtaining high speeds will be solved not only by the creation of multi-stage rockets. This method also has its drawbacks. The fact is that with an increase in the number of stages, the design of rockets becomes much more complicated. There is a need for complex mechanisms for separating steps. Therefore, scientists will always strive for the minimum number of steps, and for this, first of all, it is necessary to learn how to get more and more speeds of the outflow of combustion products or products of some other reaction.

The project was developed at the request of a venture investor from the EU.

The cost of launching spacecraft into orbit is still very high. This is due to the high cost of rocket engines, an expensive control system, expensive materials used in the stressed design 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 cost of the carrier in the total cost of launching a spacecraft varies. If the media is serial, and the device is unique, then about 10%. On the contrary, it can reach 40% or more. It is very expensive, and therefore the idea arose to create a launch vehicle that, like an air liner, would take off from the cosmodrome, fly into orbit and, leaving a satellite or spacecraft there, would return to the cosmodrome.

The first attempt to implement such an idea was the creation of the Space Shuttle system. Based on the analysis of the shortcomings of disposable carriers and the Space Shuttle system, which was made by Konstantin Feoktistov (K. Feoktistov. The trajectory of life. Moscow: Vagrius, 2000. ISBN 5-264-00383-1. Chapter 8. Rocket as an airplane), there is an idea of ​​the qualities that a good launch vehicle should have to ensure the delivery of a 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 a payload into orbit: every trip by car and flight of an aircraft 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 can be truly reliable and inexpensive to operate - such as a steam locomotive, a car, an airplane.

The rocket must be single-stage. This requirement, like reusability, is associated with 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 processes of stage separation after assembly, since with each check the assembled machine must crumble . Not tested, not tested for functioning after assembly, the connections become, as it were, disposable. And a packet connected by nodes with reduced reliability also becomes disposable to some extent. If the rocket is multi-stage, then the cost of its operation is greater than the cost of operating a single-stage machine for the following reasons:

• For a single stage machine, no assembly costs are required.
• It is not necessary to allocate landing areas on the Earth's surface for the landing of the first stages, and therefore, it is not necessary to pay for their rent, for the fact that these areas are not used in the economy.
• There is no need to pay for the transportation of the first steps to the launch site.
• Refueling a multi-stage rocket requires more complex technology, more time. The assembly of the package and the delivery of the stages to the launch site are not amenable to simple automation and, therefore, require 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 combustion of which, at the exit from the engine, environmentally friendly combustion products are formed at 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 to avoid the harmful effects of combustion products on the ozone layer of the atmosphere.

Skylon, DC-X, Lockheed Martin X-33 and Roton are among the most developed projects of single-stage spacecraft abroad. If Skylon and X-33 are winged vehicles, then DC-X and Roton are vertical takeoff and vertical landing missiles. In addition, both of them went as far as creating test samples. If Roton had only an atmospheric prototype for practicing landing in autorotation, then the DC-X prototype made several flights to a height of several kilometers on a liquid-propellant rocket engine (LRE) on 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 a Zeya launch vehicle (LV). 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 is a heat shield for atmospheric entry and a soft landing 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 - segmental-conical, load-bearing, composite. The fuel tank is pressurized by liquid hydrogen gasification, and the oxidizer tank is pressurized by compressed helium from high-pressure cylinders. The marching propulsion system consists of 36 engines located around the circumference and an external expansion nozzle in the form of a central body. Control during operation of the main engine in pitch and yaw is carried out by throttling diametrically located engines, in roll - with the help of eight engines on gaseous fuel components located under the payload compartment. Engines on gaseous propellant components are used for control in the orbital flight segment.

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

During one revolution in near-Earth orbit from the moment of launch, having given out a braking impulse, the Zeya lands in the area of ​​the launch cosmodrome. High landing accuracy is ensured by using the lift-to-drag ratio created by the shape of the missile 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 start-up:

• Pilot project: 2 months - €2 million
• Creation of the propulsion system, development of composite tanks and control system: 12 months - €100 million
• Creation of a bench base, construction of prototypes, preparation and modernization of production, draft design: 12 months - €70 million
• Development of components and systems, prototype testing, fire testing of a 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 interlaunch period. If you set the launch price in € 35 thousand per 1 kg (at a cost of €1250/kg), which is close to the launch price on the Dnepr rocket for foreign customers, the entire launch (200 kg payload) will cost the customer € 7 million. Thus, the project will pay off in 47 launches.

Zeya variant with a three-component engine

Another way to increase the efficiency of a single-stage launch vehicle is to switch to an LRE with three fuel components.

Since the beginning of the 1970s, the concept of three-component engines has been studied in the USSR and the USA, which would combine a high specific impulse when using hydrogen as a fuel, and a higher average fuel density (and, consequently, a smaller volume and weight of fuel tanks), characteristic of hydrocarbon fuels. At start-up, such an engine would run on oxygen and kerosene, and at high altitudes it would switch to using liquid oxygen and hydrogen. Such an approach may make it possible to create a single-stage space carrier.

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 Multipurpose Aerospace System (MAKS) based on the RD-701 liquid-propellant rocket engine with oxygen + kerosene + hydrogen fuel. Calculations and design of three-component rocket 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, liquid methane should be used instead of the kerosene traditionally offered for such liquid-propellant rocket engines. 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 that 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 (SI, a measure of LRE efficiency - the ratio of the impulse created by the engine to fuel consumption) of the methane + liquid oxygen fuel pair exceeds the SI 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 hard-to-remove soot. And, therefore, such engines are more convenient to use in reusable systems.
• If necessary, methane can be replaced with a similar liquefied natural gas (LNG). LNG consists almost entirely of methane, has similar physical and chemical characteristics, and is slightly less efficient than pure methane. 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. Also in Russia, an LNG plant was built on Sakhalin and two small-scale liquefaction complexes in St. Petersburg. It is planned to build five more plants in different parts of the Russian Federation. At the same time, the production of rocket kerosene requires special grades of oil extracted from strictly defined fields, the reserves of which are depleted in Russia.

The scheme of operation of a three-component launch vehicle is as follows. First, methane is burned - a fuel with a high density, but a relatively small specific impulse in a vacuum. Then hydrogen is burned - a fuel with a low density and 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 the fuel. Accordingly, the higher the proportion of fuel of the second type, the lower the required fuel supply, but the greater the mass of the structure. Therefore, 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 masses of the guaranteed fuel supply, the unusable residues of propellant components, and the mass of pressurization gases.

Calculations showed 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 loses a lot: the mass of the structure is 4, 8 tons, and the starting weight is 37.8 tons.

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

Patent Authors:
Vavilin Alexander Vasilievich (RU)
Usolkin Yury Yuryevich (RU)
Fetisov Vyacheslav Aleksandrovich (RU)

The owners of the patent RU 2309088:

Federal State Unitary Enterprise "State Missile Center" KB im. Academician V.P. Makeev" (RU)

The invention relates to rocket and space technology, in particular to reusable transport space systems (MTKS) of a new generation of the type "Space orbital rocket - a single-stage vehicle carrier" ("CROWN") with its fifty-fold use without major repairs, which is a possible alternative to cruise reusable systems like "Space Shuttle" and "Buran".

The KORONA system is designed to launch a payload (spacecraft (SC) and SC with upper stages (US) 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 SC.

It is known that at launch, the rocket is located on the launcher, while it is in a vertical position and rests on four supporting brackets of the tail compartment, which is affected by the weight of a fully fueled rocket and wind loads that create a capsizing moment, which, at the same time, are the most dangerous for strength missile tail section (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 to all support brackets.

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

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

The load on the VPA during the parking of a fully fueled rocket is distributed to all supports, while during landing, with a high degree of probability, due to the permissible deviation from the vertical position of the rocket body, the case may occur when the load falls on one support. Given the presence of vertical speed, this load is comparable or even exceeds the load in the parking lot.

This circumstance made it possible to make a decision not to use a special launch pad, transferring the power functions of the latter to the VPA of the rocket, which greatly simplifies the launch facilities for KORONA-type systems, and, accordingly, reduces the cost of their construction.

The closest analogue of the present invention is a reusable single-stage launch vehicle "CROWN" of vertical takeoff and landing, containing an axisymmetric body with a payload, a sustainer propulsion system and takeoff and landing shock absorbers (see A.V. Vavilin, Yu.Yu. Usolkin "O possible ways of development of reusable transport space systems (MTKS), RK technics, 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 analog rocket design is that its VPA is located in the zone of gas-dynamic and thermal effects of the flame coming out of the central nozzle of the sustainer propulsion system (MDU) during multiple launch and landing of the rocket, as a result of which reliable operation of the design of one VPA with the required resource is not ensured. its use (up to one hundred flights with a twenty percent reserve for the resource).

The technical result when using a single-stage reusable vertical takeoff and landing launch vehicle is to ensure the required reliability of the design of one VPA with a fifty-fold use of the launch vehicle by minimizing gas-dynamic and thermal loads on the VPA from a working MDU during multiple rocket launches and landings.

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

Compared with the closest analogous rocket, the proposed single-stage reusable vertical takeoff and landing launch vehicle has the best functional and operational capabilities, tk. it provides the necessary reliability of the design of one VPA (not lower than 0.9994) for a given period of operation of one launch vehicle (up to one hundred launches) by isolating (using a heat shield) the RPA racks from the gas-dynamic and thermal loads of the operating MDU for a given resource (up to one hundred) flights of the launch vehicle during its multiple launches and landings.

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

Thus, the proposed reusable vertical takeoff and landing launch vehicle has wider functional and operational capabilities compared to the 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 vertical takeoff and landing launch vehicle, containing an axisymmetric body with a payload, a sustainer propulsion system and takeoff and landing shock absorbers, characterized in that a heat shield made in the form of a hollow thin-walled compartment made of heat-resistant material.

The development of a landing system - the number of supports, their device, provided that their mass is minimized, is a very difficult task ...

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If the rocket accelerates for a long enough time - so that the astronauts do not experience excessive overloads - the gas emitted from the nozzle transfers momentum not only to the shell, but also to the huge supply of fuel that the rocket continues to "carry with it." Since the mass of the propellant is much greater than the mass of the shell, the acceleration of the rocket is much slower than if all the propellant were ejected at once. Calculations show that in order for the rocket to be able to reach the first cosmic velocity and put an artificial satellite into near-Earth orbit, the mass of fuel must be tens of times greater than the mass of the payload. To reduce the mass of the “accelerated” part of the rocket, the rocket is made multistage .

The first and second stages are containers with fuel, combustion chambers and nozzles. As soon as the fuel contained in the first stage burns out, this stage separates from the rocket, as a result of which the mass of the rocket is significantly reduced. The engines of the second stage immediately turn on and work until the fuel contained in the second stage runs out. Finally, this stage is also discarded, and then the engines of the third stage are turned on, completing the acceleration of the rocket to the design speed.

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Multistage rocket

A rocket whose launch vehicle includes more than one stage. A stage is a part of the rocket that is separated during the flight, including units and systems that have completed their operation by the time of separation. The main component of the stage is the propulsion system (see Rocket engine) of the stage, the operation time of which determines the operation time of other elements of the stage.

Propulsion systems belonging to different stages can operate both in series and in parallel. During sequential operation, the marching propulsion system of the next stage is switched on after the operation of the marching propulsion system of the previous stage is completed. When operating in parallel, the marching propulsion systems of adjacent stages work together, but the propulsion system of the previous stage completes its operation and is separated before the completion of the next stage. The stage numbers are determined by the order in which they are separated from the rocket.

The prototype of multi-stage rockets are composite rockets, in which it was not supposed to sequentially separate the spent parts. For the first time, composite rockets were mentioned in the 16th century in the work “On Pyrotechnics” (Venice, 1540) by the Italian scientist and engineer Vannoccio Biringuccio (1480-1539).

In the 17th century, the Polish-Belarusian-Lithuanian scientist Kazimir Seminovich (Seminavichus) (1600-1651) in his book "The Great Art of Artillery" (Amsterdam, 1650), which for 150 years was the fundamental scientific work on artillery and pyrotechnics, cites drawings of multi-stage missiles. It is Semenovich, according to many experts, who is the first inventor of a multi-stage rocket.

The first patent in 1911 for a multi-stage rocket was received by the Belgian engineer Andre Bing. Bing's rocket moved due to the successive detonation of powder bombs. In 1913, the American scientist Robert Goddard became the owner of the patent. The design of Godard's rocket provides for a sequential separation of stages.

At the beginning of the 20th century, a number of well-known scientists were engaged in the study of multistage rockets. The most significant contribution to the idea of ​​creating and practical use of multistage rockets was made by K.E. Tsiolkovsky (1857-1935), who expressed his views in the works "Rocket space trains" (1927) and "The highest speed of the rocket" (1935). Ideas of Tsiolkovsky K.E. have been widely adopted and implemented.

In the Strategic Missile Forces, the first multi-stage missile, put into service in 1960, was the R-7 missile (see Strategic Missile). The propulsion systems of two stages of the rocket, placed in parallel, using liquid oxygen and kerosene as fuel components, ensured the delivery of 5400 kg. payload at a range of up to 8000 km. It was impossible to achieve the same results with a single-stage rocket. In addition, it was found in practice that when switching from a single-stage to a two-stage rocket design, it is possible to achieve a multiple increase in range with a less significant increase in the launch mass.

This advantage was clearly manifested in the development of the R-14 single-stage medium-range missile and the R-16 two-stage intercontinental missile. With the similarity of the main energy characteristics, the flight range of the R-16 rocket is 2.5 times greater than the R-14 rocket, while its launch mass is only 1.6 times greater.

When creating modern rockets, the choice of the number of stages is determined by many factors, namely, the energy characteristics of fuels, the properties of structural materials, the perfection of the design of rocket units and systems, etc. It is also taken into account that the design of a rocket with a smaller number of stages is simpler, its cost is lower, the creation time shorter. An analysis of the design of modern rockets makes it possible to reveal the dependence of the number of stages on the type of fuel and flight range.