RU2612228C2 - Launching unit for air-launched missiles - Google Patents

Abstract

FIELD: aircraft; weapon and ammunitions.

SUBSTANCE: invention relates to aircraft weapons and concerns multi-barrel launchers (L). Launchers for air-launched missiles comprises cylindrical casing with suspension assemblies of launcher to aircraft, fairings, set of launch tubes fixed in the end discs of the housing, protector of missiles from aerodynamic heating, electric system for pulsing to missiles and lock for retention. At that the protector of missiles from aerodynamic heating is composed in the form of rotary valves, pivoted with a possibility for opening and smooth shock-free closing of the launch tubes. Launch tubes are equipped with a gas device, configured with a possibility to discharge the part of missile jet stream in the front part of the pipes.

EFFECT: preservation of aircraft controllability after missiles blow-off is achieved by closing launch tubes and reduction of the area swept surface of launcher with air flow.

3 cl, 10 dwg

Description

The invention relates to the field of aviation weapons, namely to multi-barrel launchers (launchers) of the “Block” type for placement in them and launching of aviation missiles (missiles), unguided, adjusted and controlled from aircraft.

From the prior art (see, for example, “Russia's Arms and Technologies. Encyclopedia. XXI Century” / Edited by N. Spassky. - M.: Arms and Technologies Publishing House, Volume 10, 2005), multi-barrel launchers of type B13L, B8M1, differing from each other by the number of launch tubes for missiles, their caliber and dimensions, and including a cylindrical body with PU mounts to the aircraft, fairings, a set of launch tubes fixed in the end disks of the hull, a transverse disposable partition to protect missiles from aerodynamic heating and destructible when their launch, a shutter for their fixation and an electrical system for supplying launch pulses to missiles.

Closest to the claimed invention in terms of technical nature and the technical result achieved by its use is the launcher (see page 266) launcher (B13L gun unit) for aircraft missiles having a typical composition of the incoming elements of known launchers, i.e. including a cylindrical casing with PU mounts for the aircraft, fairings, a set of launch tubes fixed in the end disks of the hull, an electrical system for supplying launch pulses to the rockets and a bolt for holding.

A disadvantage of known aircraft launchers is the deterioration of their aerodynamic properties after firing missiles, as a result of which, due to missile departure and destruction of the transverse, disposable septum, additional stabilizing aerodynamic surfaces appear that exceed the outer surface of launchers by area 2.5 ... 3.0 times, which increases the pilot’s required physical efforts on the airplane’s controls, which become close to maximum and significantly complicate the maneuvering of the airplane after the attack, at the moment enemy air defenses.

The technical task of the present invention is to eliminate the above-mentioned disadvantages and create PU, reducing the deterioration of the aerodynamic properties of the aircraft after firing a set of missiles.

The technical result that can be obtained by implementing the proposed design of the launcher is to maintain the controllability of the aircraft after firing a set of missiles by closing the launch tubes and reducing the area washed by the airflow on the surface of the launcher.

The task underlying the present invention, with the achievement of the claimed technical result, is solved by the fact that in the launcher for aircraft missiles, which includes a cylindrical body with suspension units PU to the aircraft, fairings, a set of launch tubes mounted in the end disks of the body, a means of protecting missiles from aerodynamic heating, an electrical system for supplying launch pulses to the rockets and the shutter to hold, the means of protecting the rockets from aerodynamic heating is made in the form of rotary valves, hinges but mounted for opening and closing smooth stepless launch tubes, the latter being equipped with venting device configured to discharge part reactive rocket jet in the front portion of said pipes.

In addition, each rotary valve is kinematically connected with a one-way pneumatic damper, the retractable rod of which is made with the possibility of different speeds when opening and closing the rotary valve.

In addition, the gas outlet device is made in the form of a deflector and a gas outlet tube, the latter being placed in the housing parallel to the launch tubes, its nose is fixed in the socket of the front end disk and connected to the internal volume of the corresponding launch tube, and the tail part is connected to the internal volume of the corresponding deflector.

The implementation of means of protecting rockets from aerodynamic heating in the form of rotary valves pivotally mounted with the possibility of opening and smooth shockless closing of one-sided pneumatic damper of the launch tubes, and equipping the latter with a gas exhaust device configured to divert part of the rocket jet into the front of these pipes, provides:

- opening of launch tubes at missile launch without touching the valves with their bow parts;

- holding the valve along the longitudinal axis of the launch tube when leaving the missile launcher;

- smooth, shockless closing of the pipe by the valve when leaving the missile launcher.

After shooting a set of missiles, closing the launch tubes with valves and eliminating the penetration of airflow into them, the carrier aircraft receives a regular opportunity to maneuver after an attack.

The presence of essential features distinctive from the prototype allows us to recognize the claimed technical solution as new.

A comparative analysis of the proposed technical solution with identified analogues of the prior art, from which the invention does not explicitly follow for an aviation weapons specialist, showed that it is unknown and the influence of its distinctive features on the technical result indicated by the applicant is not confirmed, therefore, the claimed invention meets the condition "Inventive step", and taking into account the possibility of industrial manufacturing launcher, we can conclude that its subject to patentability criteria.

The preferred embodiment of the proposed technical solution is described further on the basis of the drawings, where:

- in FIG. 1 shows a General view of PU for aircraft missiles;

- in FIG. 2 shows the extension element A in FIG. 1, supplemented by a local cut in the damper zone;

- in FIG. 3 shows the remote control B in FIG. one;

- in FIG. 4 is a view B of FIG. one;

- in FIG. 5 shows a simplified view of the PU, along section G-D in FIG. 4, showing the principle of operation of the valve system, in statics;

- in FIG. 6 is the same in dynamics along section DD in FIG. four;

- in FIG. 7 shows a frame of a video recording - the moment the rocket starts;

- in FIG. 8 - the same, knocking out the plug due to overpressure in the pipe;

- in FIG. 9 - the same, rocket exit from the pipe;

- in FIG. 10 - the same, the movement of the plug under the action of a rocket,

and in which the front fairing is conventionally not shown, the arrows indicate the direction of gas movement, in order to increase the visibility in FIG. 1 and FIG. A 2 valve system is shown for one launch pipe.

In graphic materials, the corresponding structural elements of PU are indicated by the following positions:

1 - housing;

2 - a rear fairing;

3 - nodes of the suspension;

4 - electrical connector;

5 - launch tubes;

6 - rocket;

7 - gas outlet pipe;

8 - front end disk;

9 - nests;

10 - rotary valve;

11 - pneumatic damper;

12 - shutter;

13 - exhaust holes;

14 - a deflector;

15 - a piston;

16 - flap valve;

17 - a stock;

18 - spring;

19 - drive fork;

20 - lever;

21 - drainage hole.

PU includes a housing 1, a front fairing (not shown conventionally), a rear fairing 2, suspension units 3 for mounting a PU on a holder of an aircraft, an electric connector 4 for docking with the electrical system of the aircraft, launch tubes 5 for accommodating missiles 6, mounted in end disks housing, a shutter 12 for holding rockets, a means of protecting the latter from aerodynamic heating, made in the form of rotary valves 10, pivotally mounted with the possibility of opening and smooth shockless closing of the launch tubes. Each rotary valve 10 is kinematically connected with a one-way pneumatic damper 11 and gasdynamically with the exhaust of the rocket, and the retractable rod 17 of the damper 11 is made with the possibility of different speeds when opening and closing the rotary valve. Launcher tubes 5 are equipped with a gas outlet device (one for each launcher tube) configured to divert part of the rocket jet into the front of these tubes in the form of a deflector 14 and a gas outlet tube 7, the latter being placed in the housing 1 parallel to the launcher tubes 5, its nose the part is fixed in the socket of the front end disk 8 and is connected to the internal volume of the corresponding launch tube 5, and the tail part is connected to the internal volume of the corresponding deflector 14, which directs the pressure part of the exhaust to the cancer you're in the bow of the PU. In the rear part of the PU on the housing 1, a removable shutter 12 is installed, having exhaust ports 13.

PU works as follows.

When an electric pulse passes to launch a rocket, a solid propellant bomb of its engine is ignited with a jet gas jet ejected through openings 13 in the shutter 12 at a speed of ≈2000 m / s, while the rocket does not move until the engine has set a given thrust (≈600 kg), the valve 10 is closed and protects the missiles 6 from the oncoming flow. Part of the jet is captured by the deflector 14 and rotated through it by 180 °, due to the braking pressure, which is ≈12 atm, which is much greater than the counter pressure of the incoming air flow, and is supplied through the gas exhaust pipe 7 to the front of the launch tube 5, to the volume limited by the valve 10 and the head part of the rocket 6, after which the valve 10 by rotation opens the launch tube 5, putting into operation the mechanism of the pneumatic damper 11, which contains: a piston 15 with a flap valve 16, a rod 17, on which a spring 18 and a drive a plug 19 connected to the valve 10 by a lever 20. The entire valve device is housed in a housing with a drain hole 21. After the rocket 6 leaves the launch tube 5, the valve 10 closes, mainly due to the pressure of the incoming flow, and the pneumatic damper 11 provides shock-free closing in time ≈1.5 s.

The invention was tested on a simplified launcher (see Fig. 7-10) containing a rocket in the upper barrel and a plastic plug installed in front of it at a distance of ≈120 mm from the nose of the rocket. On the video storyboard, you can see that after the rocket engine starts to work, the plug is pushed out of the barrel by the pressure of gas coming from the "breech" part of the block through the gaps between the rocket and the barrel wall into its bow. The pressure sensor installed in the plug recorded a pressure of ≈12 atm inside the barrel. After the plug comes out of the trunk, the nose of the rocket appears, catching up with it, i.e. the plug is ahead of the rocket at the initial stage of rocket movement.

The technical and economic result of the invention is the failed loss of airplanes that receive the regular ability to maneuver after an attack by closing the launch tubes with valves and preventing the penetration of air flow into them.

The economic effect is defined as the difference in the cost of combat losses between aircraft without a barrel jamming system (GHS) on the BIL unit and aircraft with a GHS - B13S5 system.

Symbols in the calculation:

α is the probability of aircraft damage during an attack;

n is the number of aircraft required to attack;

m — aircraft losses;

c is the cost of one aircraft;

K is the number of blocks on the plane;

C is the cost of losses;

B - the cost of one block;

E - economic effect.

Calculation conditions:

the calculation is carried out for two cases in the value of a (probability of defeat) for the serial block B13L (α ₁ ) and the new block B13C5 (α ₂ ), taking into account the reduction of the latter, according to statistics.

Initial data:

α ₁ = 0.3; α ₂ = 0.1; K = 4; n = 10 pcs.; c = 100 million rubles; B ₁ = 2 million rubles; B ₂ = 2.2 million rubles. (B13C5).

We determine the quantitative losses of aircraft for two cases α: m ₁ = n⋅α ₁ = 0.3⋅10 = 3 pcs.

m ₂ = n⋅α ₂ = 0.3⋅10 = 1 pc.

The cost of aircraft loss for each case is α.

Ts ₁ = m ₁ ⋅c + B ₁ ⋅K⋅m ₁ = m ₁ (s + B⋅K)

Ts ₁ = 3 (100 + 4⋅2) = 3⋅108 = 324 million rubles.

Ts ₁ = 324 million rubles.

C ₂ = m ₂ (s + B ₂ ⋅K) = 1⋅ (100 + 2.2⋅4) = 108.8 million rubles.

C ₂ = 108.8 million rubles.

Economic effect:

E = C ₁ -C ₂ = 324-108.8 = 215.2 million rubles.

E = 215.2 million rubles.

Claims (3)

1. Launcher for aircraft missiles, which includes a cylindrical body with suspension units PU to the aircraft, fairings, a set of launch tubes mounted in the end disks of the body, a means of protecting the rockets from aerodynamic heating, an electrical system for supplying launch pulses to the rockets and the shutter for holding, characterized in that the means of protecting the rockets from aerodynamic heating is made in the form of rotary valves pivotally mounted with the possibility of opening and smooth shockless closing of the launch tubes, while Latter equipped with venting device configured to discharge part reactive rocket jet in the front portion of said pipes. 2. The launcher according to claim 1, characterized in that each rotary valve is kinematically connected with a one-way pneumatic damper, the retractable rod of which is made with the possibility of different speeds when opening and closing the rotary valve. 3. The launcher according to claim 1, characterized in that the gas outlet device is made in the form of a deflector and a gas outlet tube, the latter being placed in the housing parallel to the launch tubes, its nose mounted in the socket of the front end disk and connected to the internal volume of the corresponding launch tube, and the tail part - with the internal volume of the corresponding deflector.

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