RU2827435C1 - Air launching complex - Google Patents

Abstract

FIELD: aviation.

SUBSTANCE: invention relates to means of air launching of missiles (space, intercontinental, ballistic, geophysical) using carrier aircraft (CA). Air launching complex includes a CA, in the cargo compartment of which the missile is placed on a support structure with guides. Exhaust unit with a constant-thrust solid-propellant rocket engine equipped with wings and empennage is connected to the rocket by cables. Exhaust unit provides stable and safe position of pulling element relative to CA and missile during its removal from cargo compartment of CA.

EFFECT: reduced probability of collision of missile with body of CA during removal of missile from CA, as well as reduction of missile launching time, reduction of disturbing effects on CA and as a result—improvement of overall reliability of the system.

7 cl, 7 dwg

Description

The proposed invention relates to the field of rocket and space technology, and more specifically to systems that provide for the air launch of a rocket by launching it from an aircraft, and can be used to increase the efficiency of the air launch of rockets for various purposes (space, intercontinental, ballistic, geophysical).

A system for air launch of space rockets is known, which has a device for preliminary lifting of the rocket into the troposphere (see patent RU 2268209). The basis of the device is a rigid spatial lattice, on which helicopter and air propellers are located. Electricity is supplied to the spatial lattice with propellers by means of wires, initially laid in pulleys. The main disadvantages of the known device are the large mass-dimensional characteristics of the lattice, the need to supply high electrical power and limitations on the height of the rocket launch, caused by the limited length of the electric cable supplying the propellers. It can also be noted that when the density of the atmosphere drops, the lifting force of the propellers decreases.

Air launch of a space rocket from an aircraft has a number of advantages over other types of launch vehicles:

- there is no need to create huge and expensive launch facilities for the launch of a launch vehicle and to alienate territory; existing airfields can be used: it is enough to build infrastructure on them for pre-launch preparation of the launch vehicle;

- the aircraft plays the role of a reusable first stage of the rocket;

- it is possible to reduce the costs of launching a space rocket by choosing the optimal launch point (including equatorial regions).

Air launch of rockets can be used for the rapid deployment and replenishment of groups of low-orbit satellite navigation and communications systems, including space reconnaissance vehicles, the launch of satellites for environmental monitoring of the Earth's surface and near-Earth space, and the launch of spacecraft for scientific research of the Earth, Moon, and planets of the Solar System onto highly elliptical and departure trajectories.

To carry out an air launch, the missile must be placed on a carrier aircraft. Let us consider three fundamentally different layouts of the missile on the carrier aircraft.

The first scheme is the placement of the launch vehicle underneath (under the fuselage) of the aircraft. In the USA, the aerospace system "Pegasus" (patent US No. 4 901 949) is successfully used, containing a launch aircraft, a launch vehicle with a wing and a payload. The rocket itself is attached to the bottom of the aircraft and, separating at an altitude of 12 km, falls for 5 seconds, after which the engine starts. Similar ideas were proposed in the "Burlak-Diana" project. To launch the two-stage launch vehicle "Burlak" it was proposed to use a converted heavy bomber Tu-160SK, the mass of the payload could be up to 825 kg. [magazine "Aviation Week and Space Technol", 11.01.99, p. 444, USA /

The main disadvantage of this arrangement is the limitation on the rocket diameter, which is determined by the distance between the lower surface of the aircraft and the runway. Also, due to the external arrangement of the rocket, it is more difficult to ensure the temperature regime for the propellant components of the launch vehicle engines: the cryogenic components of the liquid propellant will evaporate while the aircraft is gaining altitude and flying to the launch site. Therefore, it is necessary to have a reserve of propellant components on the launch vehicle, as a result of which the payload weight is reduced.

The second scheme is to place the rocket outside, on top of the aircraft fuselage. With this arrangement, the overall size limitation is removed, but it is extremely difficult to ensure safe separation of the rocket from the aircraft. In addition, as with the "under the belly" arrangement, there are still problems with maintaining the required temperature conditions for the rocket. This arrangement was used to transport the Buran spacecraft on VM-T Atlant aircraft, and then on the An-225 Mriya.

The idea of a transport aircraft for transportation and acceleration in the stratosphere of space rockets is known (patent RU 2548829). This is a transport two-body biplane aircraft, on which the rocket is located along the axis of the aircraft between the two bodies and between the two wings. The rocket is launched by its own engines. This idea has not yet found practical implementation. The main disadvantage of such a rocket launch, in addition to the increased complexity of production and operation of such a biplane aircraft, is the intense gas-dynamic and, above all, thermal effect of the jet on the aircraft structure. As estimates have shown, the intensity of the impact on the units operating in such conditions is such that plastic deformations appear in the aircraft structure within 0.1-0.3 seconds, and after 0.5-1.0 seconds it will begin to melt.

An attempt to solve the problem of the thermal effect of the jet from the rocket engines on the aircraft structure are inventions that use a parachute to "pull" the rocket from the aircraft (for example, patent RU 2334190). To do this, the aircraft must perform a hill maneuver with negative overloads. However, performing a hill maneuver with negative overloads tightens the requirements for the aircraft design and control system, due to high overloads. According to information in the same patent, safe separation of the rocket from the aircraft occurs at overloads of -0.4 ... -0.6. The patent considers an option for reducing the modulus of negative overloads due to the use of a parachute wing. The parachute canopy is a multi-section soft wing with air intakes in the front part of the profile, serving to fill the internal cavities of the canopy with air and maintain increased pressure in them due to the velocity pressure.

However, the use of a parachute causes a number of difficulties: the need to create a parachute with a large area, the reliability of its opening and the danger of the parachute being caught in the jets from the aircraft engine. The disadvantages of the parachute scheme will be discussed in more detail below.

The third scheme is the location "inside" the carrier aircraft. An alternative to the previous two schemes is the location of the missile inside the aircraft, for example in the cargo compartment. This method facilitates maintaining the temperature regime and is more convenient from the point of view of the layout (in comparison with installing the missile under the fuselage), but complicates the separation from the aircraft (removal from the aircraft) of the missile.

There is a patent RU 2160214 "Method of controlling an aerospace system for launching a payload", in which the carrier aircraft performs a maneuver with high overloads, and the rocket is parachuted ("falls out") from the cargo compartment of the carrier aircraft under the action of gravity. The disadvantage of the scheme is low reliability, and significant limitations that are imposed on the rocket and carrier aircraft due to high overloads during flight and launch.

The literature discusses three fundamentally different schemes for landing a missile located inside an aircraft: mortar launch, catapult and parachute schemes. Let us consider each scheme in turn.

Mortar launch scheme. In this air launch scheme, the missile is in a transport and launch container (TLC) on board the carrier aircraft and is ejected from it using a powder pressure accumulator. For example, a method of vertical air launch of a missile is known (patent RU 2722633), using a mortar scheme.

A technical solution is known for launching a payload into space using an AN-124 transport aircraft and a Polet launch vehicle. [http://eurasian-defence.ru/?q=node/2644&vsclid=liqvtzimyl211262681]. In this technical solution, the rocket is located inside the launch vehicle; to launch the rocket, the aircraft performs a “hill” maneuver; at the moment the maximum angle of inclination of the trajectory to the local horizon is reached, the launch vehicle is ejected from the launch vehicle using a special launch container using a pneumatic ejection system with a powder pressure accumulator.

The advantage of such schemes is that the space rocket is placed in the TPK at the manufacturing plant. Thus, it is protected from external influences and the required temperature and humidity conditions are maintained in it.

However, the scheme has a number of disadvantages:

- when a rocket is launched from a TPK, large dynamic loads are transferred to the aircraft structure and the rocket, not only in the vertical direction, but also in the longitudinal direction if the TPK axis is parallel or inclined at a small angle to the carrier aircraft (the longitudinal reaction acting on the aircraft is equal to the rocket launch force) - they must be transferred to the aircraft structure in such a way as to eliminate large local loads;

- this reaction will affect the trajectory of the aircraft (in particular, the moment from the force of the launch relative to the center of mass of the SN will cause disturbances in pitch);

- after the last obturating element of the rocket leaves the container, gases from the container come out (the unsealing process), and it is difficult to exclude the shock-wave and thermal impact on the aircraft structure - ensuring its strength and heat resistance under such impact requires the use of special measures that complicate and weigh down the structure;

- during the launch of the rocket, the propellant gases generated by the PAD have a thermal effect on the bottom of the rocket;

- of all the air launch schemes, the launch system implementing the mortar launch has the largest dimensions due to the large length and diameter of the TPK - in connection with this, problems with the placement of the launch system in the cargo compartment of the aircraft are possible.

Ejection system. This system uses a catapult mounted on board an aircraft. The rocket arms rest on guides secured to the support truss. After the launch command is received, gas is fed from the gas generators into the two working cylinders of the catapult. The force from the pistons is transmitted through the rods and the crossbar connecting them to the rocket and pushes it out of the aircraft [Rocket Ejection Systems / Yu.A. Kruglov [et al.]; Baltic State Tech. Univ., 2010. 184 p].

The advantage of the scheme (in comparison with the mortar scheme) is the smaller size of the launch device and the absence of thermal impact on the bottom of the missile. The disadvantage is that during launch, higher axial loads are transferred to the aircraft and missile than with a mortar launch, since the acceleration path with the catapult scheme is shorter and to achieve the same launch speed, the axial acceleration of the missile, and therefore the force acting on the aircraft, must be significantly greater.

Parachute scheme. In this scheme, a parachute is used for landing, which is thrown out of the SN, opens and pulls out the rocket. Before landing the rocket, the pressure in the cargo compartment of the aircraft is released and the cargo hatch is opened.

A method for launching a rocket from an aircraft is known - patent RU 2068169. In this method, the rocket is placed on a platform and is pulled out of the carrier aircraft by the force of an exhaust parachute.

The advantages of the scheme are the simplicity of the design of the rocket-pulling element and the absence of longitudinal loads on the aircraft, except for the friction force. However, there are significant disadvantages, which will be discussed in detail in a comparison of the proposed device with the closest analogue.

The closest in technical essence to the claimed invention is a device for launching a rocket from an aircraft and a method for doing so (RU 2422329 C1), adopted as the closest analogue. The device contains a launch device with a rocket, connected to an exhaust parachute by means of a connecting link, a main parachute system having a canopy and slings, and an automatic launch system for the rocket power plant.

The main disadvantage of the known device described in patent RU 2422329 C1 is its insufficient reliability, due to the fact that due to the design features of the parachute it is more difficult to ensure the required mode of launching the rocket from the carrier aircraft. To create a parachute drag force sufficient for the shock-free launch of a rocket (for example, a space rocket) weighing 100 tons from the cargo compartment, according to our estimates, a dome with a diameter of at least 10 m will be required, which will entail difficulties associated with its opening and possible entry into a turbulent wake or jet from the aircraft engine.

The parachute is unstable in the turbulent flow of the aircraft in the direction perpendicular to its movement, and can be burned by jets of aircraft engines if it gets into the jet. If it gets into the wake of the carrier aircraft, it can lose its shape, there will be pulsations of tension force and vibrations in the flow, which will cause vibrations of the cables and will be transmitted to the rocket. Several cables in this case can get tangled, since the parachute will have uncontrolled jerks.

As a result, the time of the rocket launch is unpredictable and the "hill" maneuver may not be enough for safe launch. The limitations of the dome size do not allow creating an effort that will pull the rocket in the optimal mode, that is, with the maximum allowable overloads for the rocket, in the minimum possible time.

In the parachute scheme, the rocket begins to turn around the front yokes relative to the aircraft when it is already stretched by 2/3 of its length or more, since due to the parachute's features, it will be strictly at the same level with the aircraft and the jets from the engine, which means that the force it creates will not pull the rocket "up", compensating for the increase in weight falling on the parachute due to its "removal" from the yokes on the aircraft. As a result, the rocket may collide with the upper or lower edge of the cargo compartment (the problem of a shock-free exit arises).

All this reduces the efficiency of using the rocket's air launch and increases vulnerability to random factors.

The claimed invention is tasked with reducing the probability of a missile colliding with the SN body when the missile is launched from the cargo compartment, as well as reducing the missile launch time, reducing the impact on the SN, ensuring a stable position of the pulling element during the missile launch in order to increase the overall reliability of the system and ensure greater reliability of the air launch of the missile from the SN.

The essence of the complex is that the air launch complex consists of a rocket, a carrier aircraft with a supporting structure with guides placed in it, on which the rocket is located, connected to the extension element by a connecting link equipped with cables fixed on the extension block (EB) between the nozzles. The rocket is rigidly locked in the transport position. The extension element is made in the form of an extension block EB with a solid propellant rocket engine of constant thrust (SRRE PT) with several nozzles, equipped with wings and empennage, and the EB itself is rigidly fixed in the transport position by a locking device.

The supporting structure can be made in the form of a spatial supporting truss equipped with a deformation compensation system SN, consisting of hydraulic compensators located between the spatial supporting truss and the support points SN and combined into 6 groups according to the number of degrees of freedom of the supporting structure relative to SN, the hydraulic cavities of the hydraulic compensators within one group are connected. The spatial supporting truss can be equipped with a pneumatic pusher VB with an air pressure accumulator.

The connecting link may consist of a load-bearing frame connected to the rocket bottom with explosive bolts, three drums secured to the load-bearing frame, and three cables wound onto the drums, wherein the end of each cable facing the VB may be covered with a heat-protective coating for a length of 2 to 4 lengths of the nozzles of the solid propellant rocket engine of constant thrust of the exhaust unit. The drums may be equipped with locking and locking devices controlled by a hydraulic brake with a drum revolution counter and power stoppers with a design force of at least 2/3 of the launched rocket each. The cables on the drums may be sealed with the possibility of disconnection. The nozzles of the solid propellant rocket engine PT may be equipped with plugs.

The technical result of using the proposed complex will be a decrease in the probability of a missile collision with the SN body when launching the missile from the cargo compartment of the carrier aircraft, as well as a decrease in the missile launch time, a decrease in the impact on the SN and a stable position of the pulling element during the missile launch, which leads to an increase in the overall reliability of the system. As a result, an air launch of a missile using a VB has all the advantages of an air launch and is devoid of the disadvantages of the schemes discussed above.

Fig. 1, 2 and 3 show the flight sequence of the carrier aircraft, insertion and launch of the rocket. Fig. 4, 5 and 6 show a variant of the implementation of the exhaust unit with a solid fuel rocket engine of constant thrust (implementation of the optimal variant, with three nozzles). Fig. 7 shows the exhaust unit and the rocket in the transport position (in the cargo compartment of the carrier aircraft).

In Fig. 1, 2, 3, 4, 5, 6, 7 the following designations are used:

1. rocket

2. carrier aircraft (CA)

3. supporting structure

4. cables

5. exhaust unit

6. nozzles

7. wing

8. plumage

9. drums

10. power frame

11. pneumatic pusher

12. hydraulic compensators.

The air launch complex of the rocket 1 (see Fig. 7) with the carrier aircraft (CA) 2 includes a support structure 3 with guides on which the rocket 1 is located, rigidly fixed in the transport position. The rocket 1 is connected by means of a connecting link and cables 4 with an exhaust unit (EU) 5, equipped with a solid propellant rocket engine of constant thrust (SPRE) with several nozzles 6, which can be equipped with plugs. The nozzles are set at an angle (see Fig. 4 and Fig. 5), at which the thermal effect of the jets on the cables 4 of the connecting link is reduced (the penetration of cables 4 inside the jets flowing out of the nozzles of the SPRE is excluded). The seal of the connecting link on the EU 5 is located between the nozzles 6

The solid propellant rocket motor PT. VB 5 is equipped with a wing 7 with an asymmetrical profile, which is characterized by a large angle of stall, and a tail unit 8 (see Fig. 4), while the VB 5 itself is rigidly fixed in the transport position by a locking device (not shown in the Fig.) on the guides of the support structure 3.

The connecting link is not shown separately in all three figures, since it can consist of several elements, depending on the embodiment of the air launch complex. The connecting link can be made in the form of one or several cables 4 (see Fig. 7), which can be wound onto drums 9, secured to the load-bearing frame 10, fastened to the bottom of the rocket 1 with pyrobolts. The drums 9, in turn, can be equipped with a controlled hydraulic brake and load-bearing stoppers. In the case of the embodiment with three cables 4 and three drums 9, they must provide a calculated force of at least 2/3 of the weight of the rocket 1 being launched each. The ends of each cable 4 facing the VB 5 can be treated with a heat-protective coating for a length of 2 to 4 lengths of the nozzles 6 of the PT solid-propellant rocket motor.

The preferred embodiment of the complex is the embodiment when the supporting structure 3 is equipped with a pneumatic pusher 11, made in the form of an air pressure accumulator with a piston device, the connecting link consists of a power frame 10, connected by means of pyrobolts to the bottom of the rocket 1, with three drums 9 fixed to it (according to the number of cables 9 of the connecting link) with controlled hydraulic brakes and a revolution counter, and the VB 5 with the solid propellant rocket motor PT has three nozzles 6, which are directed in the opposite trajectory of the SN 2 direction. The supporting structure 3 itself can be optimally designed as a spatial supporting truss (not shown in the diagram) equipped with a deformation compensation system SN 2, consisting of hydraulic compensators 12, combined into 6 groups according to the number of degrees of freedom of the supporting structure relative to SN 2. The hydraulic cavities of the hydraulic compensators 12 within one group are connected by tubes (not shown in the diagram). The spatial supporting truss is a support for the guides on which the rocket 1 and VB 5 lie, and rests on the SN 2 structure at nodal points that are better able to withstand a concentrated load.

It is necessary to take into account the limitations on dimensions, wing span 7 and tail dimensions 8 (see Fig. 6), associated with the layout of the cargo compartment 5 in the cargo compartment SN 2 (for example, the width of the cargo compartment of the An-124 Ruslan aircraft is 6.4 meters, and the height is 4.4, and the length is 36.5 m).

In case of an accident with VB 5, provision may be made for the possibility of “dropping” VB 5 while preserving rocket 1 - for this purpose, provision may be made for the possibility of disconnecting the seal of cables 4 on drums 9.

The proposed air launch complex for rocket 1 with SN 2 functions as follows. Rocket 1 is located in the cargo compartment of SN 2, on support structure 3 (see Fig. 1).

When using a spatial support truss with a deformation compensation system SN 2 as a supporting structure 3, deformations of the SN 2 structure during flight, as well as force loads acting on the SN 2 from the side of the supporting structure 3 during the insertion of the VB 5, do not cause the appearance of significant stresses near the support points of the SN 2 structure. This is achieved, on the one hand, due to the high rigidity of the spatial support truss and, secondly, due to the use of a deformation compensation system based on hydraulic compensators 12. The hydraulic compensators 12 are divided into 6 groups, and the hydraulic cavities of the hydraulic compensators 12 within one group are connected by tubes in such a way that the liquid flows from one group to another, as a result of which equal pressure is maintained in them. In this way, uniform distribution of the force from the spatial support truss is ensured to a plurality of nodal points of the SN 2 structure, which eliminates the concentration of stresses in the SN 2 structure in the area of the support points.

During flight, the SN 2 structure can be freely deformed in accordance with flight loads, and the rigid support structure 3 will not prevent this, due to the flow of fluid between the hydraulic compensators 12 of one group. During the launch, the rigid structure of the spatial support truss 3 and the system of hydraulic compensators 12 ensure uniform distribution of forces acting from the launch device on a large number of nodal points of the SN 2 structure.

At a given altitude and at a given point in the flight of SN 2, preparations are made for the landing of missile 1: the pressure in the cargo compartment of SN 2 is released and the hatch is opened. Then the locking device VB 5 is unlocked and it can be removed from SN 2.

When using drums 9 and cables 4 as elements of the connecting link, after opening the hatch cover of the SN 2, drums 9 are unlocked and with the help of pneumatic pusher 11 VB 5 is thrown out of the cargo compartment of the SN 2. Cables 4 begin to unwind, while the rotation speed of drums 9 is limited by the hydraulic brake. Using drums 9 makes it possible to limit the speed of VB 5 withdrawal, make the withdrawal smooth and avoid an impact at the end of VB 5 withdrawal when cables 4 are fully pulled out. This is achieved by using a controlled hydraulic brake on drums 9: it limits the rotation speed of drums 9 by selecting the law of changing the flow section of the throttle of the adjustable section, which smoothly closes. After a certain number of revolutions (turns), due to the operation of the revolution counter of drums 9, the throttle section is finally closed and the locking device latches the drums. In this case, smooth braking of VB 5 due to the use of a controlled hydraulic brake eliminates the effect of shock loads on rocket 1 when cables 4 are fully extended. Only after cables 4 are unwound to the required length, drums 9 are stopped, and VB 5 takes the required position and the PT solid propellant rocket motor is triggered, the stoppers of rocket 1 are removed and rocket 1 begins to move along the guides of support structure 3. The signal for removing the stoppers of rocket 1 will be the failure of the nozzle plugs, which are designed for a certain pressure in the combustion chamber of the PT solid propellant rocket motor.

After VB 5 has left the carrier aircraft 2 and the cables 4 have unwound to the required length according to the revolution counter, VB 5 takes the required position, which depends on the design of SN 2, or more precisely VB 5 should be located above the aerodynamic wake of SN 2, so as not to fall into the wake of SN 2. Compliance with this requirement allows achieving a stable position of the pulling element relative to SN 2. The lifting force of the wing 7 of VB 5 should be greater than the weight of VB 5 by the value of the vertical projection of the tension force of the connecting link during preparation for the landing of rocket 1.

In comparison with the closest analogue, in which the parachute system is used as the pulling element, the proposed complex has advantages, in particular, VB 5 occupies a more fixed and stable position in the air before the rocket is launched than a parachute. The parachute force is limited by the dome area, and the force of VB 5, due to the use of the PT solid propellant rocket motor, can be easily changed by selecting the parameters of the charge and the critical section. In this way, it is possible to achieve the thrust value that ensures the optimal launch mode: taking into account the maximum permissible overloads of the rocket 1, in the minimum possible time (the shorter the time, the safer). The lifting force of the wing 7 and the parameters of the empennage 8 on VB 5 are selected in such a way that they ensure its stable position before the rocket 1 is launched and located above the trace of the SN 2 and the jets from its engines. This allows avoiding vibrations during the launch of rocket 1 associated with the pull-out element getting into the wake or jet stream from the SN 2 engines, and also reduces the vertical force of rocket 1 during launch due to the fact that cables 4 are located at an angle to the axis of rocket 1 and the pull-out force during the launch has a vertical projection, partially compensating for the weight of rocket 1 and significantly reducing the angular acceleration after leaving the guides of the penultimate support elements, which simplifies ensuring the shock-free exit of rocket 1.

There is a limitation on the angle of inclination of the cables 4 of the connecting link so that during the process of launching the rocket 1 the body of the SN 2 is not touched. This is achieved by combining the lifting force created by the wing 7 and stabilization provided on the one hand by the empennage 8, on the other hand by the tension of the cables 4 of the connecting link, as well as by selecting the length of the cables 4 and, accordingly, the distance of the VB 5 from the SN 2. According to our estimates, the safe angle of the cables 4 of the connecting link can be from 8 to 12 degrees. For example, for a space rocket weighing 100 tons, launched from the SN 2 of the AN-124 type, the safe angle of inclination of the cables 4 is 9 degrees (the oscillations of the VB 5 during the launch are taken into account); the length of the cables 4 is about 56 m.

After VB 5 has taken the required position, SN 2 is ready to perform the "hill" maneuver and carry out the air launch of rocket 1 (see Fig. 2). The moment of giving the command to launch the PT solid propellant rocket motor is selected in such a way that the pitch angle of SN 2 is close to the maximum in the middle of the process of launching rocket 1. The PT solid propellant rocket motor is launched and creates the required thrust during the calculated time, which is transmitted to rocket 1 by means of cables 4 of the connecting link. Rocket 1 is unlocked and begins to move along the guides, while the required launch mode and limitations on the loads transmitted from rocket 1 to the SN 2 structure are ensured due to the selected value of the PT solid propellant rocket motor thrust and the direction of the PT solid propellant rocket motor thrust. It is assumed that during the launch of rocket 1, SN 2 maintains a stable position by its own means.

To increase the reliability of the air launch of rocket 1 with VB 5, a variant of the complex can be used that provides for the possibility of disconnecting the termination of cables 4 on drums 9. In this variant, in the event of a failure of the PT solid propellant rocket motor or other abnormal modes, VB 5 with cables 4 and drums 9 are separated from the load frame 10, while rocket 1 with the load frame 10 will remain in the cargo compartment of SN 2. This will happen due to the unlocking of the hydraulic brake - the throttle hole opens. VB 5 continues to pull cables 4 and unwind drums 9, while after unwinding the spare turns, an instantaneous tension of cables 4 will occur due to the inertia of VB 5 and the load will pass to the termination of cables 4, which is specially designed so that it is torn off by the appropriate force.

It should be noted that the combustion products flow out of the nozzles 6 in the direction opposite to the movement of the CH 2. Cables 4 may be protected, for example, by a heat-protective coating in the form of paint or mastic. It must be heat-resistant on the one hand (withstand the effects of a given high temperature), and also be fire-resistant, i.e., retain its characteristics under a given thermal effect from the solid propellant rocket motor.

After rocket 1 has completely left the yokes in the cargo compartment of SN 2 and has finally exited the cargo compartment, after a specified time (to prevent the rocket from colliding with SN 2), the explosive bolts connecting the load frame 10 of the connecting link to the bottom of rocket 1 are detonated (see Fig. 3). Then the engine of rocket 1 is started and it flies along a specified trajectory.

Thus, a technical result has been achieved, namely, the probability of a missile colliding with the SN body during the missile's launch from the SN cargo compartment has been reduced, and the missile launch time has been reduced, the impact on the SN has been reduced, and a stable position of the pulling element has been ensured during launch, which leads to an increase in the overall reliability of the system.

Claims (7)

1. An air launch complex that includes a rocket, a carrier aircraft (CA) and a support structure with guides placed in it, on which the rocket is located, connected to the extension element by a connecting link and rigidly locked in the transport position, characterized in that the extension element is made in the form of an extraction unit (EU) with a solid propellant rocket engine with constant thrust (SPRE) with several nozzles, equipped with wings and empennage, the connecting link is equipped with cables embedded on the EU between the nozzles, and the EU itself is rigidly fixed in the transport position by a locking device. 2. An air launch complex according to paragraph 1, characterized in that the support structure is made in the form of a spatial support truss equipped with a deformation compensation system SN, consisting of hydraulic compensators located between the spatial support truss and the support points SN and combined into 6 groups according to the number of degrees of freedom of the support structure relative to the SN, wherein the hydraulic cavities of the hydraulic compensators within one group are connected. 3. An air launch complex according to item 2, characterized in that the spatial support truss is equipped with a pneumatic pusher of the exhaust unit with an air pressure accumulator. 4. An air launch complex according to paragraph 1, characterized in that the connecting link consists of a load-bearing frame connected to the bottom of the rocket by explosive bolts, three drums secured to the load-bearing frame, and three cables wound onto the drums, wherein the end of each cable facing the VB is covered with a heat-protective coating for a length of 2 to 4 times the length of the solid propellant rocket motor nozzles of the PT exhaust unit. 5. An air launch complex according to paragraph 4, characterized in that the drums are equipped with locking and stopping devices controlled by _{a hydraulic brake with a drum revolution counter and power stoppers with a calculated force of at least 2/3} ^of the weight of the rocket being launched each. 6. An air launch complex according to item 4, characterized in that the cables on the drums are secured with the possibility of being disconnected. 7. An air launch complex according to paragraph 1, characterized in that the solid rocket motor PT uses three nozzles equipped with plugs.