RU2114382C1 - Bicaliber guided missile - Google Patents

Abstract

FIELD: small-sized rocket projectiles. SUBSTANCE: releasable booster engine is telescope-joined to the tip section of the sustainer stage. The tip section of the sustainer stage and location for it in the booster engine are made cylindrical, stepped with lowering to the rear end. The booster engine in its front section is furnished with a destabilizer and attenuators in the form of through slots. EFFECT: reduced disturbances of sustainer stage at staging. 2 cl, 2 dwg

Description

 The invention relates to rocket technology and can be used in the construction of small-sized rocket shots.

 In the designs of many missiles (mainly anti-aircraft), to give them a high flight speed in the initial part of the trajectory, a starting engine is used, usually separated from the rocket after burning the starting fuel, having a tandem rear arrangement and, in some cases, a larger caliber (bicaliber design) , for example, in the Wolverine rocket of the company YOU UK [1].

 One of the main requirements for the design of the rocket is the requirement of its minimum length in transport position while maintaining all other characteristics. Fulfillment of this requirement ensures a more rational use of the useful volume of the carrier, eases the requirements for the operation of guidance drives, and reduces the weight of the rocket in the container.

 One of the known solutions to reduce the length of a 2-stage rocket in the transport position is to make it telescopic, i.e. with coaxial placement of the march stage inside the launch stage with extension of the march stage when launching the rocket and the further flight of a 2-stage rocket with a tandem rear position of the launch engine until separation. Such an arrangement defines the aerodynamic design of the marching stage (as a rule, the “duck”).

 A known design of a 2-stage rocket [2], in which the marching stage is placed inside the starting engine prior to launch, while a wing block of the marching stage is fixed in front of the starting engine. When starting, the mid-flight stage is telescopically extended to the entire length of the combustion chamber, being fixed in the extended position relative to the starting stage and fixing the wing block on itself. At the end of the launch phase of the flight, the launch engine is separated.

The disadvantages of the described design are:
- the complexity of the design associated with the need to solve a number of issues - fixing the steps in the folded position, extending the marching step, fixing it in the extended position, etc., each of which is a difficult design task, which ultimately reduces reliability designs in general;
- the inability to provide a sufficiently rigid connection of the extended marching stage with the starting engine due to the need to make gaps on the landing surfaces to ensure relative movement of the marching stage.

 The aforementioned drawbacks are largely devoid of the design of a 2-stage rocket [3], which is the closest analogue (prototype) of the present invention, containing a launch engine, which telescopically includes a cylindrical aft part of the march stage, rigidly connected by a separation unit to the launch engine on the launch site flight and separated from it at the end of engine operation. Partial entry of the marching stage into the starting engine provides a reduction in length.

 The disadvantage of this design is the significant disturbances acting on the march stage of the guided missile during separation. This is due to the presence of a control signal at the launch site, causing deviation of the controls (rudders) and, accordingly, the appearance of a control moment transmitted to the aft part of the marching stage, interacting with the landing diameter of the nest in the starting engine. In this case, the radial interaction force is countered by the starting engine, which is usually stabilized due to its tail unit, during the entire joint launch flight of the steps, and during separation, all the time of the relative movement of the rear end of the aft part of the sustainer stage along the landing diameter of the starting engine jack, those. all the specified time, the steering deviation of the rudders provides control (i.e., rotation around a common center of mass by the angle of attack) of a missile having a large mass and stability margin, which requires large deviations of the rudders. At the moment of disengagement of the rear end of the marching stage with the starting diameter of the starting engine, this deviation of the rudders cannot be instantly reduced, since the separation process lasts milliseconds, therefore, at the moment of disengagement, the marching stage, which has less mass and stability margin than the rocket as a whole before separation, under a large deviation of the rudders makes a sharp angular turn, increasing the angle of attack and produced not relative to its center of mass, but relative to the extreme point of contact of the rear end with the landing diameter, which increases the shoulder of the application of the control force and can cause an increase in the angle of attack of the stage beyond the acceptable range, which may result in the marching stage leaving the control beam and, accordingly, missile loss.

 The aim of the present invention is to reduce perturbations of the march stage of a guided missile during separation, and as a result, increase reliability.

 To achieve this goal, in the known bicalibrated guided missile containing a detachable starting engine, telescopically connected to the aft of the marching stage by means of a cylindrical landing nest, the aft of the marching stage and the landing nest under it are made stepwise with lowering to the rear end, and the starting engine in the front equipped with a destabilizer and attenuation in the form of through longitudinal slots.

 Such a constructive solution provides aerodynamic damping of the disturbing moment acting on the aft part of the march stage when it is separated due to the active influence of the front part of the starting engine on the aft part of the march stage by means of a destabilizer while damping the dynamic impact by attenuation in the front part of the landing seat.

 Comparison of the proposed technical solution with the prototype made it possible to establish this compliance with the criterion of "novelty." In the study of other well-known technical solutions in this technical field, features that distinguish the proposed solution from the prototype were not identified, and therefore they provide the proposed technical solution with the criterion of "inventive step".

 In FIG. 1 shows the appearance of the rocket in its original position; figure 2 - the relative position of the stages of the rocket during the separation process (at the time of collision).

 The proposed bicaliber missile contains a marching stage 1 (Fig. 1) and a starting engine 2, rigidly connected by a fixation-fixing unit 3. At the same time, the aft part 4 of the marching stage 1 is located in the landing slot 5 (Fig. 2) in front of the starting engine. The aft part 4 and the corresponding landing seat 5 are made stepped, consisting of at least two cylindrical surfaces - a larger diameter 6 and a smaller 7. On the outer surface of the front of the engine 2 there is a destabilizer, which in the general case is any aerodynamic surface (flat blades, cylindrical or conical surfaces) located in front of the center of mass (CM2) of the starting engine. In the proposed drawing, the destabilizer is a skirt 8 with a conical outer surface, rigidly connected with the front of the engine. The front part of the engine 2 is made with weakenings in the form of longitudinal through-cuts 9 to a length not exceeding the length of the section 6 of the socket 5. At the front of the sustainer stage, controls — steering wheels 10 are installed, for the duck weft located in front of the center of mass of the sustainer stage.

 The operation of the device is as follows.

At the start of the flight, a bicaliber missile flies with steps 1 and 2 rigidly interconnected, while the control signal of the control system is converted into angular deviations of the rudders 10, causing the appearance of control moment M1, and the rotation of the axis of the entire rocket as a whole relative to the common center of mass of the CMO by angle of attack to the oncoming stream. Upon completion of the operation of the starting engine, the release unit 3 is activated and the march stage 1 starts the relative movement of its stern 4 relative to the landing slot 5, active or passive (under the influence of the difference in braking forces from the incoming flow). At the same time, while sections 6 and 7 are in interaction with the response sections of the landing nest 5, the control action of the rudders occurs, as well as on the starting section (on the march stage and the engine as a whole). As soon as the interaction of sections 6 and 7 with the response sections of nest 5 ceases, under the influence of the control moment M1, the marching stage begins a sharp turn relative to its center of mass CM1, since at the same value of the control moment the mass, the equatorial moment of inertia J _zz and the margin of stability of the unfolding stage significantly less than the original. At the same time, the engine rotates angularly relative to its center of mass of ЦМ2 in the direction opposite to the direction of turn of the marching stage, under the influence of the incoming air flow on the destabilizer (cone 8), located at an angle of attack to the flow. In this case, the sustainer stage and the engine are moved relative to each other in the axial direction. Mutual counter-turn of the steps occurs within the difference 11 between the diameters of the surfaces 6 and 7, and ultimately the end face 12 of the march step collides with the surface 6 of slot 5 (Fig. 2) in the area of the slots 9. This shock suppresses the turn of the march step due to engine turning energy, reducing its angular velocity and the magnitude of the casting along the angle of attack at the time of disengagement.

 The presence of slots 9 on the conical surface reduces the stiffness of the impact.

 The magnitude of the difference in the diameters of the stern, the lengths of sections 6 and 7 of the landing nest 5 and the sizes of the slots are selected in each case by calculation depending on the inertial-mass characteristics of the steps, the maximum deviation of the rudders, speed, design considerations and are specified during mining.

 Thus, the proposed technical solution provides a reduction in disturbances (casting magnitude) of the march stage at the time of separation due to the energy of the engine casing, which prevents its rotation by an unacceptable angle.

 Sources of information.

 1. Bulletin of ONTI N 84, June 1989.

 2. US patent N 3491692, CL. 102-49.4, 1970.

 3. US patent N 5005781, CL. 244-3.26, 1991.

Claims (2)

 1. Bikalibernaya guided missile containing a detachable starting engine, telescopically connected to the aft of the march stage by means of a cylindrical landing nest, characterized in that the aft of the march stage and the landing nest under it are made stepped with a lowering to the rear end, and the starting engine in the front equipped with a destabilizer and attenuation in the form of through longitudinal slots.  2. The missile under item 1, characterized in that the destabilizer is made in the form of a skirt with a conical outer surface, rigidly connected with the front of the engine.

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