Cluster of two pairs of tanks and flying launcher equipped with such cluster. RU patent 2509039.
 Method and device for filling onboard cylinders of missile carriers with helium / 2440505Helium filling method of onboard cylinders of missile carriers, which consists in cooling of gaseous helium in cryostat to freezing-out temperature of hazardous impurities, filtration of frozen-out particles, further heating to intermediate pressure with further draining for cleaning of internal surfaces of cylinders and their final filling to the working pressure; at that, there provided is triple filling of onboard cylinders to intermediate pressure and their filling to working pressure is performed from different receivers pre-filled to various initial pressures, and helium draining from onboard cylinders is performed in the section of the third additional receiver by pressure balancing in onboard cylinders and sections of the third receiver, from which, after missile carrier is launched, helium is supplied to compressor for compression and further filling of helium receivers. Device for the filling method's implementation includes helium and nitrogen receivers, cryostats, eclectic heaters, liquid nitrogen tank, compressor, shutoff valves, electric pneumatic valves, reduction gears and throttles. From helium receivers there performed is alternate filling of cylinders to intermediate pressure from the first receiver; then pressure is brought to working parameter from the other helium receiver. At that, in case of failed launching with the third receiver there performed is helium drain from onboard cylinders. Cooling of gaseous helium to the required freezing-out temperature of hazardous impurities is performed in cryostats, and further heating to the required temperature is performed in electric heaters. Regeneration of cryostats at contamination of their filter elements by supplying gaseous nitrogen from nitrogen receiver heated to the required temperature in electric heater. |
 Liquid-propellant rocket engine / 2406857Proposed liquid-propellant rocket engine comprises combustion chamber, gas generator, turbine, fuel pump, oxidiser pump, fuel and oxidiser pipelines communicating pumps outlets with gas generator and combustion chamber, starting tank for fuel component to be displaced therefrom on starting the engine. In compliance with this invention, working space of said starting tank is tightly divided by piston or flexible membrane into two parts, one being communicated with fuel pipeline and another communicates with oxidiszer pipeline. |
 Mixing system of hydrogen-oxygen steam generator / 2379590Mixing system of hydrogen-oxygen steam generator with spray nozzles includes ignition device, housing and flame bottom; mixing elements are made with the ratio of fuel and oxidiser nozzles of 2.2 to 12, at that, the main fuel and oxidiser nozzles located in one and the same radial direction are installed so that intersection of their axes is at one point at the distance from flame bottom of not less than 8 gauges of oxidiser nozzles; and additional fuel nozzles are installed in intermediate radial directions. |
 Liquid propellant jet engine / 2352804Invention relates to rocket engine production. The proposed rocket liquid propellant engine comprises a combustion chamber with a cooling circuit and jet head, gas generator, fuel pump, oxidiser pump and the turbine with its intake communicating the gas generator and its outlet communicating with the jet head. Note here that it comprises additional turbine with its intake communicating with the cooling circuit outlet and its outlet communicating with the jet head. |
 Oxygen-hydrogen-fuelled engine installation / 2304727The oxygen-hydrogen-fuelled engine installation has a body, combustion chamber coupled to a nozzle for exit of gases, a microporous structure is positioned in the combustion chamber, the structure is made of hollow microspheres and/or sealed hollow capillaries, a portion of which is filled with oxygen, and the other-with hydrogen. The microspheres and the capillaries may by made of combustible polymeric material. The capillaries may wound on an extension positioned along the installation axis. The microporous structure may be made of capillaries wound on an extension and located in the portion of the combustion chamber adjoining the engine body, and of microsphere filling the central portion of the chamber. The microporous structure may be made of capillaries located along the installation axis in the portion of the combustion chamber adjoining the engine body, and of microspheres filling the central portion of the chamber. Besides, the microporous structure may be made of microspheres perforated by capillaries. |
 Device (modifications) and method for combustion of rocket propellant / 2303154The device for combustion of rocket propellant has a source of rocket propellant based on ionic salt, for example, hydroxylammoniumonitrate, combustion chamber, and a reactor. The reactor serves for dissociation of the main part of ionic salt contained in the rocket propellant and is installed for supply of the products of dissociation together with and additional propellant component contained in the rocket propellant to the combustion chamber. The method for propellant combustion in the device for combustion of rocket propellant baaed an hydroxylammoniumnitrate consists in dissociation of hydroxylammoniumnitrate and combustion of the dissociation products of the mentioned hydroxylammoniumnitrate together with not reacting propellant components contained in the rocket propellant. According to the mentioned method, the rocket propellant is introduced in the reactor, and decomposed in the reactor with dissociation at least of the greatest part of hydroxylammoniumnitrate contained in the rocket propellant. Then the output products of the reactor are directed to the combustion chamber. |
 Method of and device for feeding working medium of spacecraft engine plant / 2293200Invention relates to designing and operation of spacecraft rocket engine plants. Proposed method includes separated feeding of working medium into rocket engine from high-pressure bottle with output main line, and energy from its source, for instance, plasma source for interaction in jet engine during its operation. Novelty in proposed method is provision of interaction of working medium with additional energy source before getting into jet engine, electric heater being used as such energy source. System of working medium feeding of spacecraft engine plant includes high-pressure bottle filled with working medium and provided with output main line with restrictor, main line being connected to pressure reducing main line including service-connected starting-and-cutoff valve functionally coupled with control unit, pressure reducer and receiver installed before jet engine furnished with energy source, for instance, plasma one, separately feeding energy into engine, and temperature pickup, restrictor being made in form of electric heater coaxially installed in output main line and provided with threaded throttling joint of its outer surface with inner surface, and functionally coupled with control unit, output main line being furnished with external thermal insulation; receiver is made with ribbed through pipe from side of its space in which electric heater is installed functionally coupled through control unit with temperature pickup installed on receiver provided with external thermal insulation; output of additional pressure reducing main line is connected to receiver, said pressure reducing main line includes series connected starting-and-cutoff valve and pressure reducer made similar to those of main pressure reducing main line, and its input is connected between restrictor and starting-and-cutoff valve of main pressure reducing main line. |
 Pipe line articulated joint / 2282094Proposed articulated joint has body which is connected with fixed elbow-shaped branch pipe. Second pipe line is mounted inside articulated joint coaxially relative to it. One end of this pipe line is rigidly connected with one of elbow-shaped branch pipes. Mating elbow-shaped branch pipe has blind hole which is coaxial relative to joint. End-piece of second pipe line is inserted into this blind hole by sliding fit. Laid in between engageable surfaces of blind hole and end-piece is seal. End sections of second pipe line are connected with supply and discharge pipe unions through passages made in elbow-shaped branch pipes. |
 Method of and system for charging on-road helium bottles of launch vehicles and spacecraft / 2267023Proposed method includes cooling of gaseous helium in cryostat to harmful freeze-out temperature. Then helium is filter and heated to temperature tolerable for serviceability of sealing members of on-board systems. Cryostat inlet and outlet pressure is monitored. Then triple filling of bottles with helium, thus obtained, is carried out, and their venting is done after which bottles are charged to operating pressure. If pressure differential at cryostat inlet and outlet exceeds 1 MPa, cryostat is disconnected and second cryostat is connected in its place. Proposed method is implemented by helium charging system consisting of helium receiver connected by main line with electric heater, pipeline, tank, nitrogen and additional helium receiver and at least two cryostats connected with electric heater. Each helium receiver is connected with one of cryostats and, through pipelines with fitted-in shutoff valves, with tank. Electric heater including at least three coils is connected by manifold with fitted-in shutoff-and-control valves, with on-board bottles. One coil of electric heater communicates nitrogen receiver with each cryostat. Each of remaining coils is located between corresponding cryostat manifold. Each manifold is furnished with check valve, manifolds being connected into common output manifold with fitted-in shutoff valves. Shutoff-and-control valves of manifold are essentially electropneumatic valves, one of which is normally open and the other, normally closed. Check valves are installed between manifolds and common output manifold. |
 Method of hitting dangerous space objects and device to this end / 2504503Set of invention release to space engineering, particularly, to displacement in space using space resources and can be used for hitting the dangerous space objects (DSO) Proposed method comprises selecting the comet nucleus of one of mini-comet in orbits 6 approaching the Earth 3 as hitting space body (HSB). Rocket engine unit exploiting comet nucleus evaporant as a working medium is launched in space from the Earth and placed in orbit 8 to approach said HSB. Landing thereto is carried out at point 9. Said rocket engine unit is used to change said HSB from initial orbit 7 to trajectory 10 that ensures its collision with DSO 3 at point 11. This imparts to DSB the impulse that changes it from initial orbit 4 that implies the collision of the Earth at point 5 to safe orbit. Device to this end (not shown) comprises said rocket engine unit, snapper with comet matter evaporator, power plant (with solar concentrator and astronavigation device. With snapper penetrated into comet nucleus, evaporator sublimates nucleus volatile matters. Evaporated vapors are heated by solar concentrator effuse from rocket engine nozzle generates the thrust. Astronavigator sets required direction of thrust vector. |
 Liquid-propellant rocket engine thrust unit / 2502645Invention relates to rocketry. Liquid-propellant rocket engine thrust unit comprises frame, combustion chamber with nozzle and thrust unit protection device with bottom shields. Said thrust unit protection device incorporates extra frame thermal protection device composed of combustion chamber wall cooler with channels communicating with channels feed one of fuel components to het chamber. |
 Spacecraft power plant (versions) and method of its operation / 2497730Invention relates to rocketry. Proposed power plant comprises cryogenic tank with shield-vacuum insulation, flow control valve, feed pipeline, engine combustion chamber and cryogenic tank intake including cryogenic tank lower bottom, capillary accumulator with heat exchanger under screen separator and throttling device. Engine combustion chamber has two-way start-shutoff valve to allow release of evaporated cryogenic fluid before engine start outside spacecraft and to feed said fluid in operation and after engine start into combustion chamber. In compliance with first version, engine comprises the channel to communicated heat exchanger outlet with chamber between flow control valve and booster pump that freezes constantly the engine before start and between starts. In compliance with second version, engine comprises pipeline with displacement compensator communicating heat exchanger outlet with feed pipeline downstream of booster pump. Method of engine operation comprises cryogenic fluid fee from accumulator into heat exchanger via throttling device and cooling said fluid in accumulator with the help of heat exchanger. Before regular engine start, feed pipeline is communicated with booster pipeline. Then, at regular start and at standard engine operation, engine feed pipeline is communicated with combustion chamber to communicated said pipeline after engine operation termination to booster pump tool engine down to the next start. |
 Coat for jet accelerator booster to be used at tenth-generation aircraft, spacecraft, submarines and torpedoes / 2495790Invention relates to jet machinery. Accelerator booster coat consists of flat plates with rounded angles made of light strong and refractory metal alloy 3-5 mm-thick 30-70 mm-long plates. Said coat is fixed to the surface or articulated therewith to change angle of attack of air flow at turning of coat plate brackets around hinge pins by hydraulically driven push levers with solenoids fitted under said coat plates. Said solenoids are made from high-performance materials and supplied by onboard power source. Solenoid conductors carry electric current to induce magnetic fields around coat plate interacting with charged particles, i.e. ions filling said air flow to impart extra kinetic energy pulses so that ionised airflow speed is increased. Airflow pass through ionisers arranged upstream of booster coat surface. |
 Aircraft / 2494020Invention relates to aerospace engineering. Aircraft consists of control unit, two jets and two ignited fuel exhaust pipes bent and turned in opposite directions, both being rigidly secured to airframe. Note here that said fuel escapes via exhaust nozzle arranged behind combustion chamber with front taper ledge. Control unit increases rate of fuel preset-amount portions feed. |
 Method for space burial of radioactive wastes in far space and spacecraft for its realisation / 2492537Inventions are related to methods of extraterrestrial burial of radioactive wastes and to spacecrafts (SC) with an electrojet propulsion system (EPS) for transportation of these wastes to areas of burial. The method includes placement of ampoules with specified wastes into heat-conductive matrices inside a container. Heat released in them is removed by coolant of the first closed circuit to a heat accumulator and further to a hot junction of a thermal electric generator. The cold junction is cooled with coolant of the second closed circuit, removing heat to the emitting radiator surface. Electric energy produced in the generator is supplied to the EPS. The coolant of the first circuit (for instance, lithium, potassium. caesium, mercury, bismuth, or gases) is partially used as a working body of the EPS. The flow of the latter is increased in accordance with the exponential law, selecting excessive coolant above the necessary one for removal of the current thermal power. This power drops as a result of radioactive decomposition of transported wastes. The first circuit of the coolant is equipped with a thermal throttle and an electric valve and has a branch at the outlet from the container. The first outlet of the branch is connected to the inlet of the thermal accumulator, and the outlet of the first circuit is connected with the inlet of the electric valve. The outlet of the electric valve is connected with the inlet of the thermal throttle, the outlet of which is connected with the outlet of the EPS working body reservoir. |
 Solid-fuel missile / 2492417Solid-fuel missile comprises a start engine of I stage with a powder pressure accumulator of stages separation, a forward-flow rocket engine of II stage with a nozzle and an air duct, in the afterburning chamber of which there is a start engine installed with loose fit in series with a solid fuel gas generator. The solid fuel missile is equipped with a motor of III stage, with its nozzle bottom sunk into the afterburning chamber of the forward-flow engine and joined connection compartment. Around the front bottom of the start engine and the nozzle bottom of the engine of III stage there are power structures installed, which are rigidly fixed in flanges of the connection compartment and the front bottom of the start engine. Between the flanges there is a transition cylindrical spacer, which is rigidly connected with the power structure installed on the start engine, and resting with a free end to the power structure of the engine of the III stage. On the inner surface of the afterburning chamber there i an afterburning chamber nozzle installed with loose fit and equipped with a fixator of final position on the end cut of the afterburning chamber, and is fixed with a rupture joint on the front bottom or the appropriate power structure of the start engine. The gas generator body is made in the form of a hollow cylinder covering the cylindrical insert and installed with a circular gap between the inner surface of the afterburning chamber. The air ducts are arranged along the perimeter of the afterburning chamber between the nozzle bottom of the engine of III stage and the gas generator. Between flanges of the nozzle bottom of the engine of III stage and the flange of the connection compartment there is a rupture link. |
 Carrier rocket first stage fly-back booster built around unified rocket unit / 2492123Invention relates to aerospace engineering. Fly-back booster comprises housing, oxidiser and fuel tanks, nose compartment with nose cone and air intakes stopped by rotary plug, intertank and tail compartments, rocket and air-jet engines. Liquid oxygen oxidiser tank is arranged between nose and intertank compartments, liquid methane fuel tank is arranged between nose and intertank compartments and adjoins directly to the latter and shares bottom with oxidiser tank while liquid methane fuel tank is arranged between intertank and tail compartments. Total volume of tanks arranged between nose and intertank compartments equals that of oxidiser tank while that of the tank located between intertank and tail compartments equals that of fuel tank in case kerosene is used as fuel. Four-engine rocket propulsor allows shutting down of emergent engine and boosting the trust of serviceable engines to 133%. Fly-back booster is equipped with three air-jet engines. Two of the latter are arranged on sides of nose compartment while the third one is arranged above it. Air intakes of said engines are stopped by their appropriate plugs furnished with drives. |
 Carrier rocket stage / 2486114Invention relates to space engineering. Carrier rocket stage comprises central sustainer and several solid-propellant engines arranged around said central sustainer. Every solid-propellant engine features conical generatrix of solid propellant charge attached to conical capsule-type all-wound airframe made of composite material. Said conical generatrix of solid propellant engine is parallel about carrier rocket cylindrical surface generatrix. Angle of solid-propellant engine inclination to carrier rocket axis is selected to make the former pass through center of gravity of equipped carrier rocket. Every solid-propellant engine comprises hinged control nozzle turning through minor angle not exceeding inclination of solid-propellant engine axis to carrier rocket axis. |
 Space object cryogenic liquid-propellant engine starting system / 2486113Invention relates to space engineering. Starting system relates to liquid-propellant rocket engine including cryogenic fuel tank 1, turbo pump units 2, 6, gas generator 7 communicated with turbine 18 of turbo pump unit 6, combustion chamber 19 and nozzle 20. System is equipped with high-pressure cylinders 3 with charging valve 17. Additionally, it comprises service pipeline 14 its inlet being communicated with tank 1 and outlet being connected to combustion chamber nozzle jackets 8, 9. Besides, it includes pipeline 5 to feed gaseous cryogenic fuel from nozzle jacket 9 to turbine 9 of the unit feeding fuel to gas generator 7. Pipeline 5 is connected via starting valve 16 with cylinders 3. Besides, system is equipped with makeup fuel pipeline 21 with valve 15 to communicate combustion chamber jacket 8 with cylinders 3. Systems serves to start rocket engine and to charge cylinder 3 via pipeline 21 with cryogenic fuel vapor from jacket 8. There is no need in special means of nozzle inner wall cooling. |
 Aircraft skin panel with ultrasound transducer / 2507128Invention relates to aircraft engineering, particularly, to system intended for control over fuel feed to aircraft engine. in compliance with this invention, aircraft skin represents a fluid tank wall with skin layer with outer surface that makes an aircraft aerofoil. Fuel tank skin layer accommodates the panel with fuel level transducer built in skin layer inner surface nearby hole in said skin layer and access panel to close said hole which can be detached to allow access to said transducer. |
 Aircraft fuel tank / 2507127Invention relates to aircraft engineering, particularly, to aircraft fuel tank. Fuel tank comprises body two gas supercharge and fuel feed devices. Extra shell is fitted with clearance relative to the body there inside. Cutouts are made in said shell opposite said gas supercharge and fuel feed devices. Note here that said clearance is selected from the relationship: 0,002·D≤δ≤0,145·D, where δ is clearance between tank body inner surface and shell; D is tank body ID. |
FIELD: aircraft engineering.
SUBSTANCE: flying launcher comprises cluster of tanks, fasteners, wing, engine and payload. Said cluster of tanks comprises two pairs of identical-volume cylindrical tanks with rocket liquid propellant of equal density and equal volume flow. Said four tanks are secured to each other by reinforcing belts to make parts of tanks with invariable center of gravity at efflux of propellant. Fasteners are secured to two tanks to allow attachment of the wing thereto. Linkage of tanks is arranged at top stage of square section and rounded angles.
EFFECT: decreased length of launcher.
4 cl, 5 dwg
The technical field
The invention relates to a bunch of tanks and aircraft launchers, containing this link.
The level of equipment
Modern level of equipment in the aircraft launchers used tanks circular cross-section, which are located in a line along the axis of the aircraft launchers. This configuration imposes certain restrictions on the size on carrying aircraft, in particular in relation systems for fixing and release launchers, and relating to the location of the chassis.
Disclosure of the invention
The present invention allows to mitigate these shortcomings through the creation bundles containing two pairs of identical in capacity cylindrical tanks, and each pair contains two tanks with capacity of rocket fuel of the same density, with the ability of expiration, the same volumetric rate, with four tank attached directly to each other through reinforcing zones so that the center of gravity of each of these pairs is axis ligament at the expiration of rocket fuel.
In the example, run reinforcing belt form part of these tanks are made integral with them.
The concept of combined into a bunch of tanks never been used in relation to a flying start-up facilities.
For professionals in this field it is clear that the engine is fed with fuel from these tanks shall be managed in such a way that four tank was oporozhnenija at the same time that the center of gravity of the stage continuously remained on the axis of the ligaments.
As not imposing restrictions example, one pair of tanks can hold dioxide nitrogen (chemical formula N 2 O 4 ), and another pair of tanks, the monomethylhydrazine (MMH).
When using aerial mapping in the launcher it is best way is to significantly reduce the length of the launcher, because tanks already are not in sequence, one after the other.
In practice, for the same mass ratio between the length and width of the stage, which is the conjunction of the invention (upper stage), this is equivalent to a normal level, in which tanks are located one after another.
Best way increase the main cross-section in the result of use of the invention is not significant because of the size of the nose cone of the aircraft launchers.
In particular example, the invention, the cross-section of the upper stage performed a square with rounded corners. This feature best way is to keep under control of the main cross-section.
Besides, in particular example, the implementation of some of reinforcing zones contain fixing solutions for an engine or payload. In other words, the best way themselves tanks form the design launchers. Thus, the mechanical force, in particular with the development of thrust, perceived tanks and their own fixing means. Therefore, thanks to the invention, there is no need to use resistant cone, as is the case in conventional constructions launchers that provides favorable saving of weight.
For the second aspect of the invention, created aircraft missile launcher, containing the above-described bunch of tanks together with fastening tools, attached to two of these tanks in combination, and provides for direct mounting to them the wing of the aircraft launchers.
This aspect of the invention wing aircraft launchers can be attached directly to a bunch of tanks, for example, with attachments attached to tanks.
A brief list of drawings
Further, with reference to the accompanying drawings will be described in detail with examples of the implementation of inventions that do not have a restrictive nature, its other features and benefits. Drawings:
figure 1 depicts a future aircraft missile launcher for the invention in the specific example of the implementation,
figure 2 depicts the aircraft missile launcher in figure 1 on the form, in the future, partially cut,
figure 3 depicts the aircraft missile launcher of figure 1 in the form of a longitudinal section and its nodes are enlarged, illustrating mount payload and engine to a bunch of tanks, missile launchers,
figure 4 depicts the view in the section on IV-IV, figure 3,
figure 5 depicts site design 3 is enlarged.
The implementation of the invention
Figure 1 is a very schematically shown in perspective aircraft launcher 100 in accordance with the invention.
On this drawing position 15 marked the nose cone bullet-shaped form, designed to protect the payload (in this drawing is not shown) while giving the aerodynamic form of the aircraft launchers 100.
For the nose cone is skirt 51 intermediate stage, followed by the stage 91 solid fuel. It placed the block 80 solid fuel.
The diagram mounting between the levels indicated positions 90.
A bunch of the invention is in the top step, closed nose cone 15 and skirt 51 intermediate steps.
In the example of the execution of figure 1 aircraft launcher 100 according to the invention, contains 70 wing and tail fins at the rear.
Figure 2 depicts the aircraft missile launcher on the form to partially cut, representing total internal design launchers.
Inside the nose cone is 15 payload of 40, for example, satellite.
Item 10 designated in General, a bunch of four tanks, which are directly bound to each other reinforcing belts 20. Reinforcing zone 20 executed at the same time, with tanks, as seen in figure 3. As shown in figure 2, the axis ligament coincides with the axis of the aircraft launchers.
According to the invention payload 60 attached directly to the bunch, like the engine of 50. According to the invention 70 wing also attached to a bunch 10.
In the described example run-aircraft missile launcher 100 is equipped with four spheres 30 with helium under high pressure for the purpose of ousting the rocket fuel.
Fig.3. aircraft missile launcher on 2 shows a longitudinal section, with its two node shown in the increased kind. First, it is shown how useful load and 60 sphere 30 with helium attached to a bunch 10 and, secondly, as an engine of 50 attached to a bunch 10.
In this example, run skirt 51 executed with double essentially a fork of the form for education:
outer ring attached to nasal cone 15 mounting 90, which is done is cut off with the first subversive pyrotechnic cord to release the nose cone 15, and
the inner ring attached to strengthening the zones 20, with the help of connecting plates 80, which is shown in more detail in figure 5 and made cutting through the second subversive pyrotechnic cord 81 to free up skirt 51 intermediate steps.
Figure 4 shows the section on IV-IV, figure 3. The diagram shows that in this example, the implementation of the shell upper stage 75 executed in the cross-section is essentially a square with rounded corners. This form serves to reduce to a minimum the main cross-section.
Bunch of 10 in this example, the implementation consists of the first pair of tanks, designated as 11, and the second pair of tanks, designated as 12. Both tank each pair contain the same rocket fuel and served with the same volumetric rate.
Tanks are bound to each other directly reinforcing their belts 20 with attachments 25. These mounts are 25 thus, in order to minimize the distance between the two tanks. Sphere 30 with helium attached to the reinforcement belts 20 links 31.
In this example, run the payload 60 attached to the base plate 61, which is in turn connected with the tanks 11 and 12 spatial kozhuhovsky sectors 62, which are welded to the reinforcement zones 20, as shown in figure 4.
According to the invention of the injection tools 52 aircraft launchers 100 are designed to control fuel supply to the engine 50 so that all four tank 11, 12 emptied simultaneously.
In this example, the implementation engine 50 executed managed with the possibility to change the direction of his thrust. For this purpose the engine 50 installed on the universal block 52, United with the reinforcing zones 20 skid plate 53, with the orientation of the engine can be adjusted steering actuators 54. In this example, the implementation of support for actuators 54 attached to the reinforcement belts 20 fastenings 55.
In this example, the implementation of the wing 70 attached directly to tanks 11, 12 aircraft launchers mounting tools 71, which is attached to tanks.
More specifically, in this example, the implementation of the wing equipped with 70 eyes 71, snapped his fingers 74 mounting devices or plugs 72, which are welded to the reinforcement belts 20. Those fingers 74 represent cut off by the explosion fingers that enables you to reset 70 wing in flight.
1. Aircraft missile launcher, containing a bunch of tanks, with the specified bunch of tanks contains two pairs of identical in capacity cylindrical tanks, and each pair contains two tanks with capacity of rocket fuel of the same density, with the ability of expiration, the same volumetric rate, with four tank attached directly to each other through reinforcing zones so that the center of gravity of each of these pairs is axis ligament at the expiration of rocket fuel, while aircraft launcher additional fastening the funds attached to two of these tanks, providing the opportunity for direct mounting to them the wing of the aircraft launchers.
2. Installation according to claim 1, wherein reinforcing belt form part of these tanks are made integral with them.
3. Installation according to claim 1 or 2, wherein at least some of reinforcing zones contain fixing solutions for the engine and the payload.
4. Installation according to claim 1, characterized in that section of the upper stage that hosts the specified bundle, made a square with rounded corners.