In his next astronomy video lesson, the professor will talk about a multi-stage rocket, as well as how a place for a spaceport is chosen.
Multistage rocket
The multi-stage rocket is aircraft, consisting of two or more mechanically connected rockets, called stages, that separate in flight. A multi-stage rocket allows you to achieve a speed greater than each of its stages separately. A composite rocket allows more rational use of resources due to the fact that in flight the stage that has exhausted its fuel is separated, and the rest of the rocket fuel is not spent on accelerating the structure of the spent stage, which has become unnecessary for continuing the flight. Structurally, multi-stage rockets are made with transverse or longitudinal separation of stages. With a transverse separation, the stages are placed one above the other and work sequentially one after the other, turning on only after the separation of the previous stage. With longitudinal separation, the first stage consists of several identical rockets (in practice, from 2 to 8), operating simultaneously and located symmetrically around the body of the second stage, so that the resultant of the thrust forces of the first stage engines is directed along the axis of symmetry of the second. Such a 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 weight of the rocket is maximum.
Place for the spaceport
A cosmodrome is a territory on which a complex of structures is located, designed to launch spacecraft into space. The name "cosmodrome" is given by analogy with the airfield for aircraft. Typically, spaceports occupy a large area and are located at a distance from densely populated areas, so that the stages separated during the flight do not harm residential areas or neighboring launch sites. The most advantageous place for a cosmodrome is at the equator, so that the starting carrier can make the most of the energy of the Earth's rotation. A booster rocket, when launched from the equator, can save about 10% of fuel compared to a rocket launched from a cosmodrome located in mid-latitudes. It is also possible to launch into orbit with any inclination from the equator.
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 racks 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 repeated (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-hundredfold use without overhaul, which is a possible alternative to winged reusable systems such as the 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 support 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 on 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 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 technique, scientific and technical collection, series XIY, issue 1 (48), part P, calculation, experimental studies and design of underwater-launched ballistic missiles, 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 take-off 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 analogue rocket, the proposed single-stage reusable vertical take-off and landing launch vehicle has the best functional and operational capabilities, because 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 gas-dynamic and thermal loads of an operating MDU with 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 engine nozzle 2, 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 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 resource 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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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. Home integral part 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. During parallel operation, the sustainer 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. Rocket strategic purpose). 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 creation of a single-stage medium-range missile R-14 and a two-stage intercontinental missile R-16. 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.
While creating modern missiles the choice of the number of stages is determined by many factors, namely, the energy characteristics of the 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, and the creation time is 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.
If the rocket is accelerated for a sufficiently long 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 a rocket to reach the first cosmic velocity and put an artificial satellite into near-Earth orbit, the mass of fuel must be ten 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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Scheme with carrier 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 turn to the oldest and now only historical meaning 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 missiles have already been developed and created. The first samples of liquid-propellant rockets intended for the study of the stratosphere have already been tested. Therefore, the V-2 rocket did not appear out of the blue. But she was the first to go into mass production. 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 shelling of residential areas of London with this rocket. Of course, this step could not affect the general course of military events. Much greater influence was exerted by the famous domestic rocket artillery, the perfect samples of which were tested in the early days Patriotic War directly on the battlefield. But now we are not talking about the military use of missiles. No matter how sad the history of the V-2 missile was, in this case we are only interested in its layout and layout principles. For us, this is a very convenient classroom manual that will help the reader get acquainted with the general structure of all ballistic liquid rockets in general, 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 further developed and shortcomings were eliminated: in what ways was technical progress.
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 the length into several compartments (Fig. 3.1): fuel compartment (T. O), which includes fuel tanks 1 and oxidizer 2; tail compartment (X. O) with the engine and instrument compartment (P. O), to which it is docked warhead(B. Ch). 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 by assembly and subsequent docking. In some types of missiles, there is no instrument compartment as an independent part of the hull, 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 of the cable network.
Like all guided ballistic missiles, the V-2 is equipped with a stabilization machine. Gyro devices and other blocks of the stabilization machine are located in the instrument compartment and mounted on a cross-shaped panel.
The executive bodies of the stabilization machine are gas-jet and air rudders. Gas jet rudders 3 located in the jet flowing from the chamber 4 gases and are attached with their drives - steering gears - on a rigid steering ring 5 . When the rudders deviate, a moment arises that turns the rocket in the right direction. Since gas-jet rudders operate in extremely difficult temperature conditions, they were made from the most heat-resistant material - graphite. Air rudders 6 play an auxiliary role and have an effect only in dense layers of the atmosphere and at a sufficiently high flight speed.
Liquid oxygen and ethyl alcohol are used as fuel components in the V-2 rocket. Since the acute problem of engine cooling could not be properly solved at that time, the designers went to the loss of 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 section is the chamber 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 (steam + oxygen), which are formed in the so-called steam-gas generator (SGG)(not visible in the picture). Hydrogen peroxide is supplied to the PGG reactor from the tank 3 and decomposes in the presence of a catalyst - an aqueous solution of sodium permanganate supplied from a tank 9. These components are forced out of the tanks by the compressed air contained in the cylinders. 10. Thus, the operation of the propulsion system is provided by a total of four components - two main and two auxiliary for steam and gas generation. Of course, one should not forget about compressed air, the supply of which is necessary for the supply of auxiliary components and for the operation of pneumatic automation.
The items listed are the camera, TNA, tanks of auxiliary components, cylinders with compressed air - together with supply pipelines, valves and other fittings are mounted on the power frame 11 and form a common energy block, which is called a liquid rocket engine (LPRE).
When assembling a 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 compartment, equipped with four stabilizers.
The engine thrust of the V-2 rocket on Earth is 25 ts, and in the void - about 30 ts. If this thrust is divided by the total weight consumption, consisting of 50 kgf/s alcohol, 75 kgf/s oxygen and 1.7 kgf/s hydrogen peroxide and permanganate, we get a specific thrust of 198 and 237 units on Earth and in the void, respectively. According to modern concepts, such a specific thrust for liquid engines is, of course, considered very low.
Let us turn to the so-called power scheme. It is difficult to find a short and clear definition for this rather clear concept. The power circuit is a constructive solution, which is based on considerations of strength and rigidity of the entire structure, its ability to withstand the loads acting on the rocket as a whole.
You can draw an analogy. In higher animals, the power circuit is skeletal. The bones of the skeleton are the main load-bearing elements that support the body and close all muscle efforts. But the skeletal scheme is not the only one. The shell of cancer, crab and other similar creatures can be considered not only as a means of protection, but also as an element of the general power scheme. Such a scheme should be called a shell scheme. With a deeper knowledge in the field of biology, one could apparently find examples of other power circuits in nature. But now we are talking about the power circuit of the rocket design.
At the launch site of the V-2 rocket, the engine thrust is transferred to the rear power frame 12. The rocket moves with acceleration, and in all cross sections of the hull, located above the power frame, there is an axial compressive force. The question is what elements of the hull should take it - tanks, longitudinal reinforcements, a special frame, or maybe enough in
tanks create high blood pressure, and then the structure will acquire a load-bearing capacity like a well-inflated car tire. The solution of this issue is the subject of the choice of the power circuit.
In the V-2 rocket, the scheme of the external power body and external tanks is adopted. Power Corps 13 is a steel shell with a longitudinal-transverse set of reinforcing elements. Longitudinal reinforcing elements are called stringers, and the most powerful of them - spars. Transverse ring elements are called frames. For ease of installation, the rocket body has a longitudinal bolted connector.
Lower oxygen tank 2 relies on the same power frame 12, to which, as already mentioned, the engine frame with the tail fairing is attached. The alcohol tank is suspended on the front power frame 14, with which the instrument compartment is joined.
Thus, in the V-2 rocket, the fuel tanks play only the role of containers and are not included in the power circuit, and the rocket body is the main power element. But it is calculated not only on 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 turning off the engine, the gas-jet rudders cannot perform their functions, and since the shutdown is already carried out at a high altitude, where there is practically no atmosphere, the air rudders and tail stabilizer also completely lose their effectiveness. Therefore, after turning off the engine, the rocket becomes non-orientable. The flight takes place in the mode of indefinite rotation relative to the center of mass. When entering relatively dense layers of the atmosphere, the tail stabilizer orients the rocket along the flight, and in the final section of the trajectory it moves with its 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 hull and tail unit. This is an uncontrolled flight with angles of attack varying within ±180°. The skin heats up, and significant bending moments arise in the cross sections of the body, for which the strength is mainly calculated.
At first glance, it seems unclear whether it is really necessary to care about the strength of the rocket in the final section of the trajectory. The rocket almost flew, and the job, as it were, is done. Even if the body is destroyed, the warhead will still reach the target, the fuses will work, and the destructive effect of the rocket will be ensured.
This approach, however, is unacceptable. There are no guarantees that the warhead itself will not be damaged during the destruction of the hull, and such damage, combined with local overheating, is fraught with a premature trajectory explosion. In addition, under conditions of structural failure, the process of subsequent movement has an obvious unpredictability. Even a serviceable, non-destructive rocket even gets some indefinite change in the velocity vector in the atmospheric part of the free flight. Aerodynamic forces can and do lead the rocket away from the calculated trajectory. In addition to the inevitable errors for the launch site, new unaccounted errors appear. The missile falls short, overshoots, lies to the right or left of the target. Dissipation occurs, which, due to uncertain re-entry conditions, increases noticeably. If, however, we accept the destruction of the hull and, accordingly, the loss of stabilization and speed, then the protracted uncertainty of movement will lead to an unacceptable increase in dispersion. Something similar happens to what we see when we follow the trajectory of crumbling leaves: the same uncertainty of the trajectory and the same loss of speed. By the way, a decrease in speed at the target for a combat missile of the type "V-2" also undesirable. The kinetic energy of the mass of the rocket and the energy of the explosion of the remnants of the fuel components for this type of weapon gave a quite tangible increase in the combat action of a ton explosive located in the head of the rocket.
So, the body of the rocket must be strong enough in all parts of the trajectory. And if now, without delving into the details, we take a critical look at the V-2 rocket as a whole, then we can conclude that it is the power circuit that is the weakest point of this design, since the need for excessive strengthening of the hull significantly reduces the weight characteristics of the rocket. Therefore, it is necessary to look for another constructive solution.
When analyzing the power circuit, naturally, the idea arises to abandon the supporting body and assign power functions to the walls of the tanks, additionally, perhaps, strengthening them and supporting them with moderate internal pressure. But such a solution is only suitable for the active site. As for the stabilization of the ranet when returning to the atmospheric part of the trajectory, this will have to be abandoned and the warhead should be made detachable.
Thus, a power circuit with carrier tanks is born. Fuel tanks must satisfy the strength conditions only under regulated, predetermined loads and thermal regimes of the core. After the engine is turned off, the head part is separated, equipped with its own aerodynamic stabilizer. From this moment on, the rocket body with the propulsion system already turned off and the warhead fly practically along a common trajectory, separately and without a certain angular orientation. When entering the dense layers of the atmosphere, the body, which has a large aerodynamic resistance, begins to lag behind, collapses, and its parts fall, not reaching the target. The warhead stabilizes, maintains a relatively high speed and brings the warhead to a given point. With such a scheme, it is clear that the kinetic energy of the mass of the rocket is not included in the effect combat action. However, reducing the overall weight of the structure allows you to compensate for this loss by increasing the payload. In the case of a transition to a nuclear warhead, the kinetic energy of the rocket's mass does not matter at all.
Now let's see what we gain and what we lose; what is the asset and liability in the transition to the scheme of the carrier tanks and the separating warhead. Obviously, the absence of a power hull and the absence of a tail stabilizer, the need for which is now no longer necessary, should be recorded as an asset. The asset should include the possibility of switching from steel to lighter aluminum-magnesium alloys: the atmospheric launch site of the rocket passes at a relatively low speed, and the heating of the hull is low. And finally, there is another important circumstance. The design loads on the core have a fairly high degree of reliability; they are regulated by precisely maintained conditions of withdrawal. As for reentry into the atmosphere, the load trajectories for this section are determined with less accuracy. Confidence in the design loads of the core allows you to reduce the assigned safety factor, which for a rocket with a separating warhead gives an additional weight reduction.
Some increase in the weight of the tanks will have to be made into the liability; they need to be strengthened. You may have to write down the additional weight of compressed air and fuel tank pressurization systems here. The weight of the new head stabilizer will also be recorded in the 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 be preserved from the old stabilizer. They have two tasks. The pylons provide some stabilizing effect, which makes it possible to somewhat simplify the conditions for the operation of the stabilization machine. In addition, the pylons allow you to move the air control surfaces, if any, away from the hull into a free and "unobscured" aerodynamic flow.
Naturally, in such arguments for and against one cannot be satisfied with only speculative statements. A detailed design analysis, numerical estimates and calculation are needed. And such a calculation indicates the undoubted weight advantages of the new power circuit.
The above considerations apply only to rockets having a turbopump delivery system. If the supply of components is carried out 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 supply, 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. Therefore, it is destined for them to be carriers. Suspended tanks with displacement flow would be an obvious nonsense.
A tank designed for a high internal pressure of displacement supply, as a rule, also satisfies the condition of the strength of the hull upon entry into the atmosphere. Therefore, the separation of the head part for such a missile 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 a little later. The design of the B2A rocket is a curious and instructive hybrid of old and new nascent propulsion schemes and deserves discussion as an example of the development of design thought.
The rocket has only one carrier tank - the front, alcohol, and the oxygen tank is placed in a lightweight power case, designed only for the load of the active site. Detachable head 2 equipped with its own tail stabilizer 3, representing a reinforced shell in the form of a truncated cone. In the geophysical version, the stabilizer 3 salvage head has a mechanism for opening the brake flaps 4, which reduce the fall speed of the head to 100-150 m/s, after which the parachute opens. Figure 2 shows the reentry vehicle after landing. The crumpled nose tip is visible 1 and open shields 4, partially melted during braking in the atmosphere.
The end frame of the head part 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 It has freely unlockable sealed hatches and is located not in the upper, but in the lower part of the rocket, which provides certain convenience for pre-launch operations.
Considering the B2A rocket in more detail, one could note its other features. But that's not the point. A striking and at the same time very instructive feature of this design is the logical discrepancy between the principle of a detachable head and the presence of a tail stabilizer. At the launch site, the orientation of the rocket is provided by a stabilization machine. As for aerodynamic stabilization when entering the dense layers of the atmosphere, the tail unit cannot help here, since the hull 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 found in engineering practice. technical compromise- concession to temporary circumstances. Experience has already been accumulated in the creation of missiles 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 did not require serious readjustment, which would be inevitable when switching to new aerodynamic forms. Therefore, in an environment where theoretical discussions were still underway, what threatens the transition to a non-stabilizer aerodynamically unstable scheme, it was easier, without waiting for the creation of new proven control systems, to stop at the old one. Having lost something in terms of weight, it was easier to establish itself in certain already won positions. On the way to the real implementation of the scheme with carrier tanks, it was necessary to find something between the desire to achieve the goal as soon as possible and the danger of lengthy experimental refinement, between the inevitable readjustment of production and the use of existing workshop equipment, between the risk of failure and reasonable foresight. Otherwise, a series of failures during launches, which is not at all excluded, could compromise the idea at its very core and give food to persistent distrust of the new scheme, no matter how promising and logically justified it may be.
And one more, not so important, but curious psychological aspect. The design of the B2A rocket did not seem unusual at that time. The force of the habit of seeing tails on all the small and large rockets that existed before kept the illusion of everyday life for an outside observer, and appearance missiles 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 at that time was the focus dissenting opinions 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 carrying alcohol tank regularly performs power functions, that the warhead successfully separates and safely reaches the target, and 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 naturally associated 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, mounted on the V2A rocket, provided 37 and 41.3 ts terrestrial and void thrust or 214 and 242 units of specific thrust at the Earth's surface and in the void, respectively. This was achieved by increasing the concentration of alcohol to 92%, increasing the pressure in the chamber and further expanding the outlet section of the nozzle.
The creators of the engine abandoned the liquid catalyst for the decomposition of hydrogen peroxide. It was replaced by a solid catalyst, which was placed in advance in the working cavity of the steam and gas generator. Thus, the number of liquid components decreased from four, as was the case with the V-2, to three. There was also a new, which soon became traditional, torus cylinder for hydrogen peroxide, which fits comfortably into the layout of the rocket. Some other innovations were also initiated, listing which does not make sense here.
Naturally, the B2A missile, as a transitional option 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 carrying tanks and a detachable warhead, which was done by S.P. Korolev in subsequent developments.
The first samples of missiles with carrier tanks were tested and tested in the early 50s. After that, some modifications were worked out. So, in particular, the B5V meteorological missile (combat missile R-5) also appeared. Today, a mock-up sample of a ballistic missile with carrier tanks takes pride of place as a historical exhibit in front of the museum entrance. Soviet army in Moscow.
When switching to a new upgraded scheme, in order to increase the range, the starting weight was increased and the engine operation mode was forced. The transition to the scheme of carrier tanks, of course, is more high level technology and careful study of the design made it possible to bring the weight quality factor 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 processing in the B5V rocket. After all, it was the first aerodynamically unstable rocket equipped with a very small tail and air rudders. On the same rocket, a gyroplatform and a new principle of functional engine shutdown were later used for the first time.
The B5B rocket continued to use 92% ethyl alcohol and liquid oxygen as fuel. Rocket testing showed that the lack of thermal insulation on the side surface of the oxygen tank does not entail unpleasant consequences. A somewhat increased evaporation of oxygen during prelaunch preparation is easily compensated by replenishment, i.e., automated refueling of oxygen immediately before the start. This operation is necessary in general for all rockets on low-boiling fuel components.
Thus, after the B5V rocket, the scheme of the carrier tanks and the detachable warhead became a reality. All modern long-range liquid-propellant ballistic missiles and more high step- launch vehicles are now created only on the basis of this power scheme. It was its development on the basis of modern technology and countless design improvements that gave rise to a generalized image of the machine that rightly symbolizes the pinnacle of technological progress of our time.
Now the B5B rocket can be considered as critically as the V-2 rocket was considered at the time of its creation. While maintaining the overall layout and the basic principles of the power circuit, further weight reduction and an increase in the main characteristics are possible, and the ways to solve this problem are easily seen and understood using examples of later designs.
On fig. 3.3 shows a single-stage version of the American ballistic missile "Thor"; it is also made according to the typical scheme of carrier tanks and has a detachable head. The total weight of the fuel components (oxygen + kerosene) is 45 ts with a net weight of the structure (without head part) 3.6 ts. This means the following. If we conditionally accept the total weight of fuel residues 0.4 ts, then for the familiar weight quality factor α to we get the value 0.082. Bearing the weight of the head about 2 ts, we obtain the parameter µ K = 0.12. It can also be established that with a specific void thrust of oxygen-kerosene fuel taken equal to 300 units, the range of this rocket is 3000 km.
The basis of the high weight indicators of modern missiles, in particular this one, is the 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 often located longitudinal-transverse reinforcements, which play the same role as the power pack in the V-2 rocket body, but with a greater weight quality. The waffle structure, which is currently widespread, is usually manufactured by mechanical milling. In some cases, however, chemical milling is also used. Shell blank of initial thickness h 0 subjected to carefully controlled etching in acid on that part of the surface where excess metal must be removed (the rest of the surface is pre-coated with varnish). Remaining thickness after pickling h should ensure the tightness and strength of the resulting panel at a given internal pressure, and the longitudinal and transverse ribs give the shell 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 disturbed in the zone of welds, which, as is known, have a slightly reduced strength compared to the rolled sheet, and also 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 stopped at chemical milling in order to show at what cost, in the literal and figurative sense, those weight indicators of the structure that are characteristic of modern rocket technology are achieved.
Rocket "Thor" has a shortened and lightweight tail section Z, on the end of which two control motors are mounted. The rejection of gas-jet rudders is associated, of course, with their high gas-dynamic resistance in the jet of outflowing gases. The use of control motors somewhat complicates the design, but gives a significant gain in specific thrust.
From what has been said, one should not get the impression that the control cameras appeared for the first time on this particular ballistic missile. Such a system of power controls has been used in various versions before, in particular, on the carrier rocket of the Vostok or Soyuz systems, which will be discussed below. The single-stage version of the Thor missile is considered here solely as an example of the next generation of ballistic missiles following the B5B missile.
Almost all ballistic missiles brake solid propellant engines are also installed 6. This is also not the latest news. The task of the brake engines is to, having slowed down the body of the rocket, take it away from the warhead during its separation; namely, the hull, without imparting additional speed to the warhead.
Shutdown of the liquid engine is not instantaneous. After the valves of the fuel lines are closed, combustion and evaporation of the remaining components still continue in the chamber for the next fractions of a second. As a result, the rocket receives a small additional impulse, called aftereffect impulse. When calculating the range, an amendment is introduced to it. However, it is definitely impossible to do this, since the aftereffect impulse does not possess stability and varies from case to case, which is one of the significant reasons for range dispersion. In order to reduce this dispersion, brake motors are used. The moment of their inclusion 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 B5V and Thor missiles. Rocket B5B is more elongated. The ratio of length to diameter (called rocket extension) for her significantly more than the missile "Tor"; about 14 versus 8. The difference in elongations causes various concerns. With an increase in elongation, the frequency of natural transverse oscillations of the rocket, as an elastic beam, decreases, and this forces us to take into account the perturbations that enter the stabilization system as a result of angular displacements during body bending. In other words, stabilization should be ensured not of a rigid, but of a curving rocket. In some cases, this causes serious difficulties,
With a small elongation of the rocket, this issue is naturally removed, but another nuisance arises - the role of perturbations from transverse oscillations of the liquid in the tanks increases, and if it is not possible to fend them off by proper selection of the parameters of the stabilization automaton, it is necessary to set them in tanks baffles that restrict fluid flow. The figure partially shows the nodes 7 for mounting vibration dampers in the fuel tank. Naturally, such a decision leads to a deterioration in the weight characteristics of the rocket.
The Thor missile should not be regarded as a model of perfection. At the same time, the designers could probably oppose their own counterarguments to any critical remarks about its layout. On the example of the B2A rocket, we have already seen that a reasonable criticism of a constructive solution can only be carried out taking into account the specific design and production conditions, and most importantly, the long-term tasks that the creators of the new machine set for themselves. And the Thor rocket is just one of those on the basis of which it is possible to create rocket and space systems.
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