RU2685002C2 - Missile with air-jet engine - Google Patents

Abstract

FIELD: astronautics.SUBSTANCE: invention relates to the field of rocket equipment, namely to rockets with an air-jet engine - AJE. Device comprises front air intake device. It includes central body with cone of external-internal compression, shell ring with undercut, fixed on central body by means of pylons uniformly located in circumferential direction, combustion chamber with injectors of fuel supply, annular supersonic nozzle at combustion chamber outlet, gas generator, warhead, starting engine of solid fuel, stabilizer. According to invention, AJE is equipped with cone position adjustment system. It is located in central body and is made in form of hydraulic cylinder with piston. Hydraulic cylinder sub-piston cavity is connected with injectors of fuel supply into combustion chamber at gas generator start-up. Hydraulic cylinder above-piston cavity is connected with differential throttle, providing cone movement along axis by specified value relative to central body and shell. Pylons are fixed on the central body at an angle of attack to the longitudinal axis of the missile and are turned towards its rotation. Annular nozzle is located on the shell and is made in the form of an expanding fan extending beyond the rocket launcher during its flight.EFFECT: technical result is increase in speed and range of missile flight, expansion of traction-aerodynamic characteristics of AJE.1 cl, 5 dwg

Description

The invention relates to the field of rocket technology, namely to missiles with an air-jet engine (WFD).

The missile with the WFD contains a frontal air intake device, including the central body and the shell, the combustion chamber, the gas generator, the warhead, the starting solid fuel engine (RDTT) and a stabilizer with drop-down blades.

Air-jet engine mounted in the nose of the rocket is designed to maintain or increase the speed of the rocket in order to increase its flight range after the end of the accelerating solid-fuel engine (RDTT).

One of the ways to increase the effectiveness of rocket armament models is the development of missiles with an increased range. Increasing the flight range of missiles can be achieved by using combined propulsion systems, namely a combination of accelerating solid-propellant (solid propellant rocket propulsion) and a cruising air-jet engine (WFD).

Improvement of these missiles is developing in the direction of searching for rational design and technological solutions when developing WFDs that ensure improved reliability of operation at various flight speeds, thermal stability of the product during long-term operation, stability of intraballistic characteristics in the combustion chamber, as well as the implementation of limit or close-to-economic characteristics .

The subject of the invention is a rocket with a multi-mode air-jet engine with enhanced traction and economic characteristics due to the use of an adjustable air intake, a fan nozzle, improved ballistic characteristics in the combustion chamber of the WFD and increased reliability of the gas-dynamic path of the engine during long-term operation.

Known design of active-projectile with the WFD, located in the bow, having a profiled central body (Sorokin V. A., Yanovsky L.S. and others. Rocket-direct-flow engines on solid and pasty fuels. Nauka, M., 2010 ., p. 31), containing a front air intake device comprising a central body and a shell, a combustion chamber, pylons, an annular nozzle, a gas generator.

Stabilization of the projectile with WFD in flight is due to rotation around the longitudinal axis at a speed of ~ 170 rev / s. The pylons, which rigidly connect the central body with the shell, are parallel to the longitudinal axis of the engine. When the rocket rotates due to the significant skew of the air flow in the gas-dynamic path of the engine and due to disruption of the flow of the mixture of gases and flames by rotating pylons, the resistance of the motor path increases, the loss of total pressure increases, the variation of ballistic characteristics in the combustion chamber of the WFD increases. As a result, the aeroballistic characteristics of the rocket are violated as a whole, and the range and accuracy of the flight are reduced.

The presence of an air intake device (OVC) with a fixed position of a cone of mixed compression on the central body suggests the presence of a large area air intake device for successful launch of the OVC at the start of a projectile or rocket. This leads to a weak compression of the incoming air in the intake device and to a significant limitation of the amount of heat supplied to the combustion chamber, which leads in general to a deterioration of the engine traction characteristics.

Another disadvantage of the system under consideration is the absence of an effective annular nozzle, which would reduce the lateral loads on the projectile due to the short outer wall of the annular nozzle, possible heterogeneity of the flow in the nozzle and improve the engine thrust characteristics.

The above artillery projectile is characterized by a small range of fire, and therefore, low speed and duration of flight.

The closest in technical terms is the design of the rocket with the WFD, located in the nose of the rocket (patent RU No. 2585211, IPC F42B 12/46, 2015), adopted by the author for the prototype.

Common features with the proposed rocket with the WFD, is the presence in the prototype of a frontal air intake device comprising a central body and a shell with undercuts, a combustion chamber, pylons, nozzles, a gas generator, a warhead, a solid propellant propulsion engine, stabilizers.

WFD contains a frontal air intake device (OVC), including the central body and the shell, the pylons rigidly connecting the central body with the shell, the combustion chamber, injectors, annular nozzle, gas generator.

The cone of external-internal (mixed) compression of the OVC is in a fixed position relative to the central body and to the sidewall, providing a rigid connection between them. The ratio of the minimum flow area (throat) of the air intake device to the area of the mid-section of the rocket is 0.3 ... 0.5. With these ratios, the OVS is launched when the rocket leaves the guide, its throughput is ensured at the acceleration site of the rocket with the help of a solid propellant rocket motor and is activated in the work of the WFD when the optimum speed is reached. This leads to a weak compression of the incoming air in the air intake device and to a significant limitation of the amount of heat supplied in the combustion chamber due to an increase in pressure and the threat of disruption of air inflow into the air intake device, which generally leads to a deterioration in the engine traction characteristics.

It is also known that the aeroballistic characteristics of a rocket are influenced by the processes occurring in the propulsion system of the WFD (the variation of the intraballistic characteristics of the motor tract - the poor mixing of fuel and oxidizer, the variable degree of combustion in the volume of the motor tract, the design features of the combustion chamber and the OVC).

The stability of the aeroballistic characteristics of the rocket is ensured when the rocket is spinned in a tubular guide using U-shaped spiral grooves (as a rule, the angle of rotation is 240 ° with a guide length of ~ 7 meters), and in flight by stabilizers installed in the tail of the rocket at a slight angle attacks to the longitudinal axis and defining the rotational motion of the rocket. If we assume that the stabilizers support the initial angular velocity of rotation of the rocket as it leaves the guide, then the angular velocity at the final stage of acceleration with the help of a solid propellant rocket motor can be ~ (40-90) rev / s.

Although the angular velocity of rotation of the rocket with the WFD placed in the nose is smaller than that of the projectile, the linear speed of rotation of the pylons of the combustion chamber is very significant. Therefore, due to the swirling flow of air entering the combustion chamber, the pylons must be rotated at a certain angle in the direction of rotation of the rocket with the WFD to create less resistance to the air flow in the combustion chamber. It is not possible to unequivocally determine the angle of inclination of the pylons with respect to the longitudinal axis of the rocket, since the speed of the rocket from the launch to the termination of the marching WFD is significantly different. There is also a dependence of the speed of rotation of the rocket on the flight speed. But this feature should be taken into account to improve the intraballistic characteristics of the combustion chamber and the aeroballistic characteristics of the product as a whole. Most likely, the angle of installation of the pylons to the longitudinal axis of the rocket will be determined from the maximum flight speed of the rocket at the end of the acceleration solid propellant rocket operation and may be 3-6 degrees at speed M = 3.

In the prototype under consideration, the pylons are located parallel to the longitudinal axis of the WFD and the rocket, which does not in the best way affect the product characteristics.

High-quality operation of the nozzle assumes a uniform pressure field in the critical section. In the case of using an annular nozzle, obtaining a uniform pressure field becomes unrealistic due to the complexity of the processes in the combustion chamber and the chamber design, despite the subsonic flow regime before the critical nozzle section. Therefore, the interaction of a heterogeneous jet at the exit of the critical section with the rocket cone creates a reaction almost normal to the longitudinal axis of the rocket and when the length of the outer part of the nozzle belonging to the shell is significantly shorter than the inner part belonging to the central body or the rocket.

Further development of air-jet engines for missiles with a nose-jetting arrangement leads to the need to find technical solutions aimed at improving the engine performance, intraballistic and operational characteristics under the conditions of high braking temperatures.

The objective of the proposed technical solution is to increase the speed and range of the missile, the expansion of the traction and economic characteristics of the WFD, the improvement of the traction and aerodynamic products in general.

The problem is solved due to the fact that the rocket with an air-jet engine (WFD) contains a frontal air intake device that includes a central body with a cone of external-internal compression, a shell with an undercut fixed to the central body by pylons evenly spaced in the circumferential direction, the combustion chamber with fuel supply injectors, an annular supersonic nozzle at the exit from the combustion chamber, a gas generator, a warhead, a solid fuel starting engine (RDTT), a stabilizer. According to the invention, the WFD is equipped with a cone position regulation system located in the central body and made in the form of a hydraulic cylinder with a piston, the piston cavity of which is connected to the fuel injectors of the combustion chamber when the gas generator starts up, and the piston cavity with the differential throttle providing the cone moving along the axis by a predetermined value relative to the central body and the shell, the pylons are fixed on the central body at an angle not equal to zero (angle of attack) to the longitudinal axis of the rocket, and rotated in the direction of Well, its rotation, and the annular nozzle, located on the side of the shell is made in the form of a drop-down fan that goes beyond the mid-section of the rocket during its flight.

A positive result is achieved through a set of measures to improve the characteristics of a ramjet engine.

Unlike the prototype in the proposed rocket with the WFD air intake is made adjustable, by moving the cone external-internal compression along the axis of the central body of the engine and shell. In the pre-launch position, the cone of the external-internal compression section of the engine is in the extreme right position, increasing the throat area of the air intake to the maximum value, ensuring the regular start of the air intake when the solid rocket motor is operating.

Pylons providing a rigid connection between the central body and the shell are rotated relative to the longitudinal axis of the engine at an angle (in the direction of the rocket rotation), ensuring their continuous flow around as the rocket rotates, thereby reducing the resistance of the gas dynamic path of the engine and the loss of total pressure in the engine's combustion chamber.

A distinctive feature of the proposed WFD from the prototype is the presence of a fan nozzle at the exit from the combustion chamber, which is folded in the prelaunch state of the rocket. When the rocket is launched (exiting the tube guide) under the influence of the velocity pressure of the air passing through the engine, the fan nozzle opens and is ready to work in conjunction with the combustion chamber.

FIG. 1 shows a general view of a missile with a WFD; in fig. 2 and 3 - a system that provides movement of the cone relative to the central body and the shell; in fig. 4 shows the angle of inclination of the pylons with respect to the longitudinal axis of the rocket, the section along the EE of FIG. 3; in fig. 5 scheme of the WFD (in isometric).

The rocket with an air-jet engine contains a frontal air intake device 1, including a cone 2 of external-internal compression and a shell 3 with an undercut, pylons 4, evenly spaced in the circumferential direction, are installed (see Fig. 4) at an angle α ≠ 0 to the longitudinal axis rockets and turned in the direction of its rotation, injectors 5, evenly spaced in the circumferential direction, combustion chamber 6, gas generator 7 with program mode of combustion, central body 8, fan-shaped ring nozzle 9, warhead 10, rocket engine solid fuel 11, art. utilizers 12.

The WFD is equipped with a system for adjusting the position of the cone 2 relative to the shell 3. The system for regulating the position of the cone (Fig. 2) contains a differential throttle 13, a hydraulic cylinder 14, a piston 15. The cavity A and a sub-piston cavity B (Fig. 3) of the hydraulic cylinder 14 ensure the opening of fuel injectors 5 in the combustion chamber 6 at the start of the gas generator 7. The cavity B and the over piston cavity G of the hydraulic cylinder 14 are connected with a differential throttle 13, providing a predetermined mode of movement of the cone 2 along the longitudinal axis of the rocket by a predetermined value relatively central body 8 and shell 3.

When storing the rocket and in the pre-launch position, the piston 15 and the cone 2 are in the extreme right position (see Fig. 2). The liquid completely fills the piston cavity G of the hydraulic cylinder 14, and the piston 15 completely blocks the inlets of the fuel injectors 5 into the combustion chamber 6. The fan part of the annular nozzle 9 is folded (position D), fixed to the rocket body and does not stand for the rocket body dimensions.

The missile with the WFD works as follows (in Fig. 2 and 3).

When launching a rocket, the starting engine of solid fuel 11 is turned on (Fig. 1), while the piston 15 and the cone 2 are in the extreme right position, the injectors 5 of the fuel supply to the combustion chamber 6 of the WFD are closed (see Fig. 2), and the air intake throat between shell 3 and cone 2 as open as possible. After the rocket leaves the tubular guide (not shown in the figure) and increases in speed, the air intake device 1 (OVC) begins to function. Under the influence of the velocity head, the fan part of the annular nozzle 9 is opened. For continuous acceleration of the rocket, immediately before the end of the solid propellant rocket operation, a command is issued to start the gas generator 7. The pressure in cavity A quickly increases and displaces the piston 15 to the left, thus forming the injectors B and opening the injectors 5 fuel injection into the combustion chamber 6. At the same time, the pressure in the piston cavity B increases sharply and the piston 15 displaces the fluid from the cavity G through the differential throttle 13 (see Fig. 2 and 3) into the cavity B. At the same time, the cone 2 of the central body 8 is shifted to the left, to a position optimal for the operation of the OVC at the speed provided by the launching solid propellant solid propellant rocket. At the same time, the combustion products of the gas generator 7 are injected through the open injectors 5 into the combustion chamber 6 of the WFD.

Next, the differential throttle 13 moves the cone 2 forward at a nominal speed (in FIG. 3) along the longitudinal axis as the missile accelerates. When moving the cone 2 to the left, the relative area of the throat of the OVC (the ratio of the area of the throat to the capture area of the jet) decreases continuously, ensuring the operation of the OVC at design or close to it modes depending on the free-stream Mach numbers.

In accordance with the flight program should be organized and the flow rate of the combustion products of the gas generator 7 with the program mode of combustion.

The use of the fan part of the annular nozzle 9 extends its outer part belonging to the shell and contributes to obtaining a uniform pressure field throughout the combustion chamber 6.

A positive result is provided by the proposed design solutions of the missile with the WFD, which improve the tactical and technical characteristics of the declared object.

Information sources

1. Sorokin V.A., Yanovsky L.S. And others. Rocket and direct-flow engines on solid and pasty fuels. Science, M., 2010, p. 31

2. Patent RU No. 2585211, IPC F42B 12/46, 2015 - prototype.

Claims (1)

A jet engine with a frontal air intake device comprising a central body with a cone of external-internal compression, an undercut shell fixed on the central body by longitudinal pylons evenly spaced in the circumferential direction, a combustion chamber with fuel injectors, an annular supersonic nozzle at the outlet of the combustion chamber, the gas generator, the warhead, the starting engine of solid fuel, a stabilizer, characterized in that the air-jet engine The gatel is equipped with a cone position control system located in a central body made in the form of a hydraulic cylinder with a piston, the piston cavity of which is connected to the fuel injectors into the combustion chamber when the gas generator starts up, and the over piston cavity - with a differential throttle ensuring the cone moving along the axis by a predetermined value relative to the central body and the shell, the pylons are fixed on the central body at an installation angle, not equal to zero, to the longitudinal axis of the rocket and turned in the direction of its rotation, and tsevoe supersonic nozzle is designed as a drop of the fan, leaving behind the midsection missile during its flight.

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