RU2315261C2 - Stabilizing device of aircraft winged missile - Google Patents

Abstract

FIELD: rocketry, in particular, device for stabilization of aircraft winged missiles at the initial stage of independent flight.

SUBSTANCE: the body of the stabilizing device is made in the form of a convex cover sealing the nozzle of the missile solid-propellant acceleration engine, and the controls-in the form of four pairs of kinematically linked with one another aerodynamic control surfaces and jet vanes of the solid-propellant acceleration engine, each of them is positioned in the plane of installation of the aerodynamic control surface of the missile sustainer stage and linked with the drive of the mentioned control surface by means of control rods. The control mechanism of the stabilizing device control surfaces is made for separation of the links controlling the position of the device aerodynamic control surface, as well as for separation with the control rods from the sustainer stage control surface drive.

EFFECT: simplified structure and reduced overall dimensions of the stabilizing device, enhanced aerodynamic characteristics of the winged missile.

7 dwg

Description

The invention relates to the field of rocket technology, and more particularly to aircraft-based cruise missiles (CD). The invention describes a device that provides stabilization and control of aircraft in the initial phase of flight, immediately after separation from the aircraft, and can find application mainly on missiles originally developed for other types of carriers, for example, sea or ground, and adaptable to air-based conditions.

One of the main requirements for aviation KR is their most compact placement on carrier aircraft. For this purpose, various technical solutions are used, such as: folding the wing of missiles when placed on carriers, for example - KP 3M80 (Moskit) and X-35 (A.V. Karpenko, S.M. Ganin “Domestic tactical missiles” , military-technical collection "Bastion" No. 1, 2000, St. Petersburg, pp. 56-59 and 73-74); placing missiles on an airplane in a position deployed around a longitudinal axis, ensuring a rational configuration of the RS on the suspension, in combination with a post-launch turn to the initial (flight) position described in the application for the invention, "Method for the use of aviation RS with an air-propelled propulsion system", filed simultaneously with this application.

However, folding the wing, as a rule, deprives the missile of controllability until it is fully deployed, and in order to turn the aircraft into flight position, measures may be required to increase the efficiency of the controls, for example, in the case of an unfavorable initial alignment characteristic of missiles with a tandem or integrated solid-fuel booster engine. In such cases, which, as a rule, arise when adapting to the conditions of aviation-based missiles previously developed for other types of carriers, special devices are used to stabilize the RS during the transition process.

The closest in combination of features with the claimed invention is the stabilization device used on the aviation version of the Kyrgyz Republic “Meteorite” (FSUE NPO Mashinostroyeniya. 60 years of dedicated work for peace, publishing house “Arms and Technologies of Russia”, M., 2004 , p. 81, 82, 229). The known stabilization device (see photo from the specified source, given in the appendix) is made in the form of a tail fairing detachable from a rocket, equipped with three aerodynamic rudders. Its functioning is provided by drives and rudder control mechanisms located in the fairing cavity, and the separation from the rocket is provided by special quick-release couplings. As such, as a rule, mechanical pyrotechnic means (pyro-bolts or pyro-pushers) and detachable electrical connectors are used.

However, in relation to small-sized RCs, the joint placement of the necessary equipment within a limited diameter leads to an increase in the length and irrational external forms of the stabilization device. The implementation of such layout solutions, as applied to the KR with combined propulsion systems based on ramjet engines, for example, made according to the layout scheme described in RF patents No. 2117907 and No. 2215981 (IPC F42B 15/00), is associated with additional difficulties.

Firstly, the design features of the combined propulsion system of this KR, which is a direct-flow marching engine (ramjet) with a bearing (that is, an element of the fuselage) body and a solid fuel accelerating rocket engine (RDTT) built into it, make it impossible to conduct electrical communications to the stabilization device within the diameter of the fuselage of the rocket. The layout of quick disconnect joints at the junction of such a switchgear with a stabilization device is also significantly complicated. This leads to the necessity of organizing gargrots and blisters, which are essentially in favor of the contours of the rocket, which means an increase in its aerodynamic drag as a whole.

Secondly, most of the fuel of the built-in booster solid propellant rocket engine is concentrated in the tail part of the CR, which, in turn, allows us to judge the unfavorable initial alignment of the rocket. In combination with a relatively small shoulder of the installation of aerodynamic rudders (see FIG. 1 of the patent specification No. 2215981), this is fraught with a situation of possible static instability of the RC in a configuration without additional stabilizing surfaces. This circumstance that occurs after the stabilization device is separated and is valid until the launch of the accelerated solid propellant rocket engine (that is, the use of its effective gas-dynamic control elements) is the basis for "tightening requirements" for synchronizing the execution of these operations.

The problem solved by the invention is:

- reducing the total number of airborne units necessary to ensure stabilization and control of the rocket at the initial stage of its autonomous flight, including equipment placed in the body of the stabilization device,

- providing continuous control of the CR at the indicated stage of the flight by minimizing the transit time,

- limitation of the negative impact on the aerodynamic characteristics of the Kyrgyz Republic from the introduction of a stabilization device in the composition of the rocket when the latter is adapted to the conditions of aviation based.

This problem is solved due to the fact that in the known stabilization device of an aircraft cruise missile containing a body mounted with the possibility of separation on the tail of the rocket, aerodynamic rudders mounted on the body, drives and steering mechanisms, according to the claimed invention, the body of the stabilization device is made in the form of a convex a cover that seals the nozzle of a solid propellant booster rocket engine, and the controls of this device are in the form of four pairs of kinematically connected between each other the aerodynamic rudders and the gas rudders of a solid propellant booster engine, each pair of stabilization device rudders being placed in the planes of the rocket aerodynamic rudders installation and connected to the corresponding rocket rudder drive via control rods and the stabilization device rudder control mechanism, which is capable of separating links controlling position aerodynamic steering device, as well as separation with control rods from the rudder drive.

The technical result of the use of the invention lies in the fact that it allows you to create the design of a stabilization device, which, in comparison with traditional ones, is characterized by a simplified structure and significantly smaller mass and dimensions. This is achieved through the use of solutions that allow aviation KR units to combine several functions, including those assigned to the stabilization device, for example:

- the execution of the nozzle plug of the booster RDTT in the form of a convex cover allows it, along with its main purpose, namely, sealing the internal cavity and providing the specified engine starting mode, to fulfill the functions of the stabilization device body. At the same time, there is no need to use any additional units to separate the device’s body from the rocket and the control gear is provided without gaps to “transfer control” from one (aerodynamic) control device to another stabilization device (gas-dynamic);

- the kinematic connection of the control devices of the stabilization device with each other, as well as with the drive placed on board the rocket, allows this drive to control not only the corresponding aerodynamic wheel of the marching stage of the Kyrgyz Republic, but also a pair of steering wheels of the stabilization device installed in the same plane. In this case, traction from the drive, laid on the outer surface of the tail section of the rocket, can be made in the form of cables or metal bands that almost do not protrude over the contours of its fuselage, and the required gear ratios between the rocket controls and the stabilization device can be achieved by selecting the appropriate the ratio of the shoulders of the rocking mechanism of the steering wheel control.

Thanks to the specified set of technical solutions, from the traditional structure of the stabilization device, special pyro means for fastening and separating the device from the rocket and highly specialized drives can be excluded. Accordingly, there is no need for the electrical equipment of the stabilization device, as well as the wiring of communications and the use of electrical connectors that would connect it with the onboard missile control system.

The essence of the proposed device is illustrated in Fig.1-7.

Figure 1 shows a General view of the CR assembled with a stabilization device when placing the rocket on a carrier aircraft (in a specific example, the KR is placed on the plane in accordance with the technical decision on the application for "Method of using aviation KR ...", that is, in the position rotated around a longitudinal axis at an angle of 45 °). Figure 2 is a rear view of the CR in the indicated position. Figure 3 and 4 shows the tail section of the Raman and a stabilization device in partial section. Figure 5-7 shows a diagram of the functioning of the CR with the stabilization device after separation of the rocket from the aircraft.

The cruise missile (1) is made according to the normal aerodynamic scheme and contains an axisymmetric fuselage, wing and tail, which is made in the form of four all-rotating aerodynamic rudders (2). The rudders (2) are installed in the planes I, II, III and IV of the rocket and are controlled by actuators (3) located inside the fuselage.

KR (1) is equipped with a combined propulsion system, which is a ramjet (4) with a bearing body and an accelerating solid propellant rocket (5) built in it, which has gas-dynamic controls made in the form of four gas rudders (6).

On the tail part of the KR (1), a stabilization device (7) is installed, which consists of a body and controls.

The body of the stabilization device (7) is made in the form of a cover (8) that seals the nozzle (9) of the accelerating solid propellant rocket motor (5). The cover (8) has a cylindrical lateral surface that protrudes beyond the nozzle exit (9) and is fastened to the last calibrated bolts (10), which are designed to be damaged by a tensile force of a certain magnitude.

The controls of the stabilization device (7) are made in the form of four pairs of kinematically coupled interlocking aerodynamic rudders (11) and the aforementioned gas rudders (6) of solid propellant rocket motors (5), which are part of the stabilization device functionally.

Aerodynamic rudders (11) are mounted on the side surface of the cover (8) by means of brackets (12), into which the axis of rotation of these rudders are pressed. Moreover, each aerodynamic steering wheel (11) is installed coaxially with the corresponding gas steering wheel (6) and interacts with the rocking chair (13) of the latter. For this purpose, a guide pin (14) is pressed into the plane of the steering wheel (11), a radial groove is made in the bracket (12) to move the pin (14), and the rocker (13) of the gas steering wheel (6) has a carrier with a slot for the specified pin.

The rudders (11) and (6) of the stabilization device (7) paired in this way are placed in the planes of the rocket I, II, III and IV, that is, in the planes of the aerodynamic rudders (2) installation of the CR (1). Moreover, each rocker (13) of the gas rudder (6), which has a gear sector on the shoulder opposite the carrier, interacts with the rocking gear (16) fixed to the rear end of the fuselage of the KR (1) through an intermediate gear (15). In turn, the aforementioned rocking gear (16) by means of tape rods (17), laid on the outer surface of the tail compartment of the KR (1), is connected with the corresponding drive (3) of the rudder (2) of the rocket. Thus, the kinematic system formed by the rocking gear (16), the intermediate gear (15), the rocking chair (13) of the gas rudder (6) and the guide pin (14) of the aerodynamic rudder (11) connected to it is a control mechanism (18 ) a pair of rudders (6) and (11) from the stabilization device (7).

The specified device operates as follows.

On command from the aircraft’s weapon control system, the captures of the aircraft launching device are opened, as a result of which the KR (1) is separated from the carrier and begins its autonomous movement. KR (1) in the planning flight mode moves away from the aircraft, while its position is stabilized according to the on-board control system (BSU) using the rocket's executive bodies (drives (3) and aerodynamic rudders (2)) and a stabilizing device (7) (traction (17), mechanisms (18) and aerodynamic rudders (11)). It should be noted that the rudders (11) of the stabilization device (7) in this section of the flight of the Kyrgyz Republic (1) make the main contribution to the overall efficiency of the controls, since they, in comparison with the rudders (2) of the rocket, have a significantly greater leverage the center of mass of the rocket.

For a given time delay, the control system of the Kyrgyz Republic (1) issues a command to perform a programmatic roll roll at an angle of 45 °, the implementation of which is carried out by joint deflection of the aerodynamic rudders (2) and (11). In this case, the efforts of the drives (3) create on the axes of the rudders (2) of the rocket (1) the control moments, which are through the rods (17) and links of the mechanisms (18), namely, rocking gears (16), intermediate gears (15), rocking (13) and the interacting guide pins (14) are transmitted to the aerodynamic wheels (11) of the stabilization device (7).

As a result of the U-turn, the KR (1) occupies a flight position corresponding to the vertical and horizontal orientation of its planes.

After removing the CR (1) to the distance of a launch vehicle safe for launching an accelerating solid propellant rocket engine (5) (of the order of several tens of meters), the missile control system issues a command to launch it. As a result, the solid propellant igniter of the solid propellant charge is activated (5) and the pressure in its combustion chamber begins to increase rapidly. Upon reaching a predetermined pressure value in the solid propellant combustion chamber (5), the calibrated bolts (10) that hold the cover (8) on the nozzle (9) are destroyed, and the cover (8), together with the aerodynamic rudders fixed on it (11), is separated from the CR (1) ) In this case, the guide pins (14) of the rudders (11) are disengaged from the carriers of the rocking chairs (13), thereby breaking the chains of mechanisms (18) that previously controlled the position of the rudders (11). Further, the control moments created by the drives (3) are applied only to the aerodynamic wheels (2) of the control gear (1) and the gas wheels (6) of the solid propellant rocket motor (5).

Under the thrust of the accelerated solid propellant rocket engine (5), the CR (1) sharply gains speed, while the gas rudders (6), deflected in the jet stream flowing out of the nozzle (9) of the engine, make the main contribution to ensuring rocket stabilization and possible program maneuvers . As the velocity of the rocket accelerator (1) increases (and, therefore, the pressure head acting on the bearing and control surfaces of the rocket) and the fuel burns out of the accelerated solid propellant rocket engine (5) (and, therefore, the forward center of mass of the rocket is shifted forward), the aerodynamic rudder efficiency (2) by the end of the booster section of the flight, the missiles acquire the ability to “independently” control its movement.

When the thrust of the accelerating solid propellant rocket (5) drops to a certain value, by the pressure of the air entering the air path of the rocket propulsion system, the exhaust engine body is ejected from the ramjet cavity (4). In this case, the intermediate gears (15) disengage from the rocking gears (16) and the previously functioning gas steering control links (6) of the mechanisms (18) are separated in the rear fuselage section of the KR (1).

As a result, the ramjet air path (4) is “released” and, after the launch operations, the KR (1) starts flight in march mode, during which the rocket movement is controlled by aerodynamic rudders (2).

Thus, we can conclude that the use of the proposed solution allows us to simplify the structure, reduce the total number of equipment, reduce the size and weight of the stabilization device, which, in turn, reduces the cost of its development and manufacture. At the same time, the positive effect of introducing such a device into the aviation version of a small-sized missile is not accompanied by a decrease in its aerodynamic characteristics.

Claims (1)

A stabilization device for an aircraft cruise missile, comprising a housing mounted with the possibility of separation on the tail of the rocket, controls in the form of rudders mounted on the housing, actuators and rudder control mechanisms, characterized in that the housing of the stabilization device is made in the form of a convex cover, which seals the solid fuel nozzle the rocket’s accelerating engine, and the controls - in the form of four pairs of kinematically connected aerodynamic rudders and gas rudders of a solid propellant accelerating engine a gatel, each pair of rudders of the stabilization device is placed in the planes of installation of the aerodynamic rudders of the rocket and connected to the corresponding drive of the rudder of the rocket by means of control rods and the control mechanism of the rudders of the stabilization device, which is configured to separate the links controlling the position of the aerodynamic rudder of the device control rods from the rocket rudder drive.

Images