RU2352894C1 - Underwater missile - Google Patents

Abstract

FIELD: weaponry.

SUBSTANCE: invention refers to cruise missiles starting underwater. The missile accommodates contains a main rocket with accelerating system comprising high- and low-thrust rocket engines. The low-thrust rocket engine is located in the rocket nose fairing and designed as single-chamber solid engine comprising a number of the identical thrust nozzles or a number of parallel solid engines of the same unit size. The bellmouth axes of low-thrust rocket nozzles are directed at an acute angle to the missile centre line, while nozzle planes are provided at an angle relative to aerodynamic rudder and main rocket wing planes.

EFFECT: accelerating propulsion weight reduction and maintained fixed runaway speed.

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Description

The invention relates to rocket technology, and more specifically to missiles with underwater launch, mainly supersonic anti-ship cruise missiles (CR) long range (anti-ship KR long range missiles are usually classified as missiles with a firing range of more than 100 km, designed to combat highly protected ship formations (Rodionov B.I., Novichkov N.N. “Cruise missiles in naval combat”, Military Publishing House, 1987, p. 14)), deployed and launched from transport and launch containers (TPK).

Known placed in the TPK and starting from under the water, a supersonic cruise missile ("KR in TPK", RF patent No. 2215981, IPC F42B 15/00), which contains a starting-booster and mid-flight stages, as well as a nose fairing.

The marching step of the famous KR is made in the form of an axisymmetric fuselage equipped with a wing and aerodynamic rudders, and is equipped with a ramjet engine with frontal air intake. The nose fairing (BUT) of the rocket ensures the sealing of the internal cavities of the TPK and the longitudinal unfastening of the rocket launcher in the container before the launch of the rocket, as well as the sealing of the air intake of the rocket launcher during its movement under water. BUT is made in the form of a cylinder-conical body, in the cavity of which there are engines designed to separate and remove the fairing after the CR comes out from under the water. The start-and-boost stage (СРС) is intended for the launch and acceleration of the aircraft to the launch speed of the march ramjet. However, the description of the present invention does not contain information about the device of the CDS, allowing to judge about the method of ensuring the movement of the rocket in various environments (under water and in air).

Known missile that implements both underwater and surface launch ("Method of launching a rocket from TPK and device for its implementation", RF patent No. 2240489, IPC F42B 15/00, B64C 15/00).

The launch-booster stage of this rocket is made on the basis of a propulsion system consisting of two parallel mounted rocket engines of large and small thrust. The high-thrust engine is designed to carry out surface launch of the rocket and accelerate it in the air, and the low-thrust engine is mainly designed to ensure the movement of the rocket under water. Parallel installation of these engines provides the possibility of both joint and sequential operation. So, when the small thrust engine operates in the underwater section of the movement, the rational use of the energy potential of the CDS and the limitation of the loading of the rocket when moving in a "dense" environment are ensured, and when the engine of large thrust (or both engines together) is used, the high "thrust ratio" of the CDS is necessary for implementation of surface launch and dispersal of a rocket in the air. Another feature of the device of the known rocket is the use of a nose cone equipped with a system of pulsed solid-fuel rocket orientation control engines designed to perform quick turns in the direction of the target. The device of this rocket in terms of features is closest to the claimed invention and is considered by the authors as the closest analogue.

In an example implementation of the technical solution according to patent No. 2240489, a CDS based on solid propellant rocket engines (RDTT) is considered, while a small thrust solid propellant rocket engine having a torocylindrical combustion chamber is placed around a main thrust solid propellant nozzle. However, the two cylindrical shells and bottoms of the ring-shaped solid propellant solid propellant rocket motor of low thrust, “working” under conditions of high temperatures (up to 3000–3500 K) and pressure (up to 100 kgf / cm ² ), have a rather large mass. The latter, with the integral attachment of these engines, increases by an appropriate amount the “passive” mass of the body of the SRS acceleration propulsion system, which leads to a decrease in the flight performance of the rocket. For example, in relation to supersonic Raman engines with marching ramjet engines, due to the forced need to accelerate the indicated “passive” mass to the ramjet launch speed (at least 1.5 ÷ 2 M), up to 5 ÷ 6% of the potential velocity is lost, which requires a corresponding increase mass of solid propellant solid propellant.

The problem solved by the invention is to reduce the mass of the propulsion system of the CPC at the stage of acceleration of the rocket in the air.

This problem is solved due to the fact that in the known rocket with underwater launch, containing the marching stage, equipped with aerodynamic rudders, the starting and accelerating stage, the propulsion system of which consists of two parallel mounted solid propellant rocket engines of large and small thrust, and a nose fairing, made with the possibility of separation in flight, according to the claimed invention, a thruster, made in the form of a single-chamber solid-fuel engine with several identical jet nozzles or several parallel installed solid propellant engines of the same size are placed in the nose cone of the rocket, while the axis of the nozzles of the nozzles of the thruster are directed at an acute angle to the longitudinal axis of the rocket, and the plane of installation of the nozzles are located at an angle relative to the plane of installation of the aerodynamic rudders of the marching stage.

The technical result of the invention lies in the fact that by reducing the mass of the SRS at the stage of acceleration in air, it improves the flight performance of the rocket. The possibility provided by the proposed solution for separating the passive mass of the low thrust solid-propellant hull at the initial air portion of the trajectory (before the intensive acceleration begins) allows to increase the acceleration speed of the marching stage by means of high-thrust solid-propellant rocket engine or, while maintaining a fixed acceleration speed, increase the mass of the target equipment or rocket propellant rocket by the corresponding value.

At the same time, the operation of the small thrust solid propellant rocket engine provided for by this decision in accordance with the “pulling” scheme allows it to fulfill the functions of the nose cowling engines. As the calculations showed, the thrust required for the rocket to move under water is sufficient to separate the BUT (after removing the mechanical connection with the rocket) and to withdraw it along a trajectory safe for the rocket. At the same time, the increase in the total thrust of a small thrust solid-propellant is not necessary due to the increase in mass “accelerated” by the main solid-state solid rocket thrust necessary for the realization of this function, as well as to compensate for the losses “due to the deviation of the engine thrust vector from the longitudinal axis of the rocket”.

The choice of a thrust engine in the form of a single-chamber solid propellant rocket with several (two to four) peripheral nozzles or several solid propellant rocket motors of the same size is mainly due to the layout conditions and should be made taking into account the shape of the nose of the rocket and the limitations on the dimensions of the latter. As the studies have shown, in relation to various variants of the rocket, the totality of requirements can meet:

- RDTT with a combustion chamber formed by two ball bottoms and a relatively small height cylindrical shell;

- RDTT with a combustion chamber of a torocylindrical or toroconic shape, similar to a thruster, described in patent No. 2240489;

- 4 ÷ 6 solid propellant rocket motors of the same size, which, in order to avoid significant “multi-pulling”, can have a gas connection.

As for the pulsed solid-propellant rocket-propelled rocket-propelled rocket-propelled rocket-propelled rocket steering system located in the nose fairing, for example, as provided in patent No. 2240489, the application of this solution is relevant only for anti-aircraft guided and tactical anti-ship missiles, which are characterized by the implementation of quick post-launch turns, for example, in the direction of the target detected in close proximity to the carrier. Heading and pitch turns, characteristic of long-range anti-ship missiles, can be performed less intensively, for example, by means of long-thrust propeller-driven gas-dynamic controls and (or) marching aerodynamic rudders.

The device of the missile with an underwater launch, as exemplified by a supersonic rocket propulsion system with marching ramjet and a thrust accelerator of small thrust with a toroconic shape, which is “most organically” assembled together with a ramjet frontal air intake, is illustrated in figures 1-4. Figure 1 presents a General view of the rocket with a nose fairing, figure 2 shows the placement of the rocket in the TPK. Figure 3 shows in more detail the front of the rocket in the TPK, including the fairing and the nose of the rocket in partial section, and figure 4 shows the cross section of the rocket in the TPK.

The rocket (1) contains the marching stage (2), the starting-accelerating stage and the nose fairing (3).

The marching stage (2) of the rocket (1) is made in the form of an axisymmetric fuselage with a frontal air intake (4) ramjet and a folding plus-shaped wing (5) and plumage (aerodynamic rudders (6)).

The starting and accelerating stage is made in the form of two solid propellant solid propellant rocket motors of small and high thrust - (7) and (8), respectively.

The large thrust solid propellant rocket propeller (7) is located in the air path of the ramjet marching stage ramp (2) with a partial protrusion of the rear of the engine beyond the slice of the fuselage of the rocket (1). The solid propellant rocket motor (7) is equipped with gas-dynamic controls, for example, gas rudders (not shown).

Small thrust solid propellant rocket propeller (8) is located in the nose fairing (3). The solid propellant rocket motor (8) is characterized by a toroconic combustion chamber formed by the outer cylindrical and inner conical shells and four nozzles (9), the axis of the sockets of which are directed at an acute (within 15 ÷ 20 °) angle to the longitudinal axis of the fairing (3) and, accordingly rockets (1).

The nose fairing (3) is fixed on the front of the march stage (2) of the rocket (1) by means of pyro-bolts (10). At the same time, BUT (3) is installed in such a way that the central body and the shell of the frontal air intake (4) of the ramjet rocket propulsion rocket (1) occupies the internal cavity of the small thrust solid propellant rocket engine (8), in turn, the nozzles (9) of this engine are located at an angle of 45 ° relative to the planes of installation of the wing and the feathering of the march stage (2).

The rocket (1) is placed and operated in the TPK (11), made in the form of a cylinder with a blank rear bottom. In the bottom volume of the TPK (11), limited by the bottom of the container and the tail of the rocket, there is a powder pressure accumulator (12). At the same time, an obturator (13) in contact with the inner surface of the TPK (11) is installed on the rear part of the solid propellant rocket hull (7), which protrudes beyond the slice of the fuselage of the rocket (1). TPK (11) is connected to the HO body (3) by means of shear elements (14), while the fairing body is equipped with belts of sealing rings (15) in contact with the inner surface of the container body.

The rocket (1) is equipped with sensors for exiting the TPK (11) and out of the water, limit switches for detachable and expanding elements, etc.

During the underwater launch, the rocket (1) functions as follows.

As a result of the actuation of the powder pressure accumulator (12), excess pressure is created in the bottom volume of the TPK (11) and the rocket (1) under the influence of the “piston effect” destroys the shear elements (14) and moves in the container. At the moment the shutter TPK (11) passes through the shutter (13), a command is formed to launch the small thrust solid propellant rocket engine (8), as well as a command to open and “start control” the aerodynamic rudders (6).

Under the action of the solid propellant rocket propulsion (8) rocket (1) moves to the surface of the water. In this case, the orientation of the thrust vector of the solid propellant rocket engine (8), due to the corresponding nozzle installation angle (9), excludes the thermoerosive effect of the solid propellant rocket engine combustion products on the fuselage, wing and feathering of the march stage (2), and the possible small “asymmetry” of the engine thrust is compensated by control the movement of the rocket (1) under water through the aerodynamic rudders (6) of the marching stage (2).

As the rocket (1) exits from the water, a team is formed to separate BUT (3) and open the wing (5). As a result of the activation of the pyro-bolts (10), the mechanical connection of the fairing body (3) with the rocket (1) and the HO (3) is removed under the action of the solid propellant propulsion rod (8), and is separated forward and downward with respect to the path of the further flight of the rocket (1) .

Immediately after the fact of separation of the BUT (3), a team is formed to launch the solid propellant rocket propulsion vehicle (7), under the influence of the propulsion of which the rocket (1) is sharply accelerated and, by means of the control moments created by the propellant rocket propulsion system (7) and aerodynamic rudders (6), on a given trajectory.

When the rocket (1) reaches the given supersonic speed, the solid propellant rocket engine (7) will separate, the ramjet will be launched and the marching stage (2) will make a further flight to the target.

As a result, due to the early separation of the “passive” mass of the thruster propulsion system while maintaining a fixed acceleration speed, for example, the mass of the ramjet ramjet fuel can be increased, and, therefore, the flight range of the rocket as a whole.

Claims (1)

An underwater launch rocket containing a marching stage with aerodynamic rudders, a starting and accelerating stage, the propulsion system of which consists of two parallel mounted solid propellant rocket engines of large and small thrust, and a nose fairing made with the possibility of separation in flight, characterized in that the small engine thrust placed in the nose fairing of the rocket and is made in the form of a single-chamber solid-fuel engine with several identical jet nozzles or several parallel mounted solid propellant engines of the same size, with the axis of the nozzles of the nozzles of the thrust engine directed at an acute angle to the longitudinal axis of the rocket, and the installation plane of the nozzles are located at an angle relative to the installation planes of the aerodynamic rudders of the marching stage.

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