RU2496090C1 - Anti-aircraft missile and liquid-propellant rocket - Google Patents

Abstract

FIELD: weapons and ammunition.

SUBSTANCE: proposed missile comprises axially symmetric body with oxidiser and propellant tanks and liquid-propellant rocket with combustion chamber and turbo pump unit, and four radial control nozzles. Combustion two-zone chamber comprises cylindrical part with oxidiser atomisers, nozzle with propellant main manifold, extra propellant atomisers, annular and extra manifolds. Turbo pump unit comprises turbine, oxidiser and propellant pumps, extra propellant pump, top and bottom ball hinges. Ball hinge with inner ball and outer spherical shell is arranged between combustion chamber and oxidiser pump. Said four control nozzles are communicated with combustion chamber first zone via pipes.

EFFECT: higher reliability of launching and specific characteristics of liquid-propellant engine.

6 cl, 7 dwg

Description

The invention relates to rocket technology and liquid-propellant rocket engines - LRE, operating on three components of the fuel; oxidizer and two combustibles, and is aimed at improving the specific characteristics and reducing the cost of launching the rocket on which it is installed and at significantly improving its many characteristics: flight range, etc. The most optimal option is to use liquid oxygen as an oxidizing agent, the first fuel - kerosene, the second fuel - liquid hydrogen.

Known multi-stage anti-aircraft missile according to the patent of the Russian Federation for the invention No. 2380648, IPC F42B 15/00, publ. January 27, 2010

The disadvantages are the complexity of the design and the high cost.

Known anti-aircraft missile according to the patent of the Russian Federation for the invention No. 2380650, IPC F42B 15/10, publ. 01/07/2010 This missile contains a warhead, an axisymmetric body, oxidizer and fuel tanks, and a liquid rocket engine.

The disadvantages are the large transverse dimensions of the rocket and difficulties in management.

Known liquid rocket engine according to the patent of the Russian Federation for invention No. 2095607, intended for use in space booster blocks, stages of rocket launchers and as the main engine of spacecraft, includes a combustion chamber with a regenerative cooling path, pumps for supplying components - fuel and an oxidizer with a turbine one shaft into which the capacitor is inserted. The condenser outlet through the refrigerant line is connected to the entrance to the combustion chamber and to the entrance to the regenerative cooling path of the combustion chamber. The exit from the condenser through the coolant line is connected to the inlet to the pump of one of the components. The output from the pump of the same component is communicated with the condenser inlet through the refrigerant line. The second input of the condenser is in communication with the output of the turbine. The pump output of the other component is in communication with the entrance to the combustion chamber.

The disadvantage of the engine is the deterioration of the cavitation properties of the pump during condensate bypass.

A known method of operation of the liquid propellant rocket engine and a liquid rocket engine according to the patent of the Russian Federation for the invention No. 2187684. The method of operation of a liquid-propellant rocket engine is to supply fuel components to the combustion chamber of the engine, gasify one of the components in the cooling path of the combustion chamber, supply it to the turbine of the turbopump unit, and then discharge it into the nozzle head of the combustion chamber. Part of the flow rate of one of the fuel components is directed to the combustion chamber, and the remaining part is gasified and directed to turbines of turbopump units. The gaseous component spent on the turbines is mixed with the liquid component entering the engine at a pressure higher than the saturated vapor pressure of the resulting mixture. A liquid propellant rocket engine contains a combustion chamber with a regenerative cooling path, fuel component feed pumps, and a turbine. The engine comprises a booster turbopump pump and a mixer installed in series in front of the feed pump of one of the fuel components of the main turbopump assembly. The pump outlet of the main turbopump assembly is connected both to the nozzle head of the combustion chamber and to the regenerative cooling path of the combustion chamber. The regenerative cooling circuit, in turn, is connected with the turbines of the main and booster turbopump units, the outputs of which are connected to the mixer.

The disadvantage of this scheme is that the thermal energy removed during cooling of the combustion chamber may not be enough to drive a turbopump engine unit of very high power.

Known LRE according to the patent of the Russian Federation for the invention No. 2190114, IPC 7 F02K 9/48, publ. 09/27/2002 This LPRE includes a combustion chamber with a regenerative cooling path, a TNA turbopump unit with oxidizer and fuel pumps, the output lines of which are connected to the head of the combustion chamber, the main turbine and the drive circuit of the main turbine. The main turbine drive circuit includes a fuel pump and a regenerative cooling path of the combustion chamber connected in series with each other and connected to the main turbine inlet. The exit from the turbine TNA is connected to the input of the second stage of the fuel pump.

This engine has a significant drawback. Bypassing the fuel combustion chamber heated in the regenerative cooling path to the entrance to the second stage of the fuel pump will lead to cavitation. Most LREs use fuel components such that the oxidizer consumption is almost always greater than the fuel consumption. Therefore, for powerful rocket engines with great thrust and high pressure in the combustion chamber, this scheme is not acceptable, because fuel consumption will not be enough to cool the combustion chamber and drive the main turbine.

In addition, the LRE launch system, the ignition system of the fuel components and the LRE shutdown system and its cleaning of fuel residues in the regenerative cooling path of the combustion chamber have not been developed.

Known liquid rocket engine according to the patent of the Russian Federation for the invention No. 2232915, publ. 09/10/2003 g, which contains a turbo pump unit, a gas generator, a launch system, means for igniting fuel components and fuel lines. The output of the oxidizer pump is connected to the inlet of the gas generator. The output of the first stage of the fuel pump is connected to the channels of regenerative cooling of the chamber and to the mixing head. The output of the second stage of the fuel pump is connected to an electric flow regulator. Another input of the regulator is connected to the starting tank with standard fuel. The output from the regulator is connected to a gas generator. The outlet of the gas generator is connected to the inlet of the turbine pump unit, the outlet of which is connected to the mixing head. The flow regulator is equipped with a hydraulic actuator of the preliminary stage, which is connected through the cavitating nozzle and hydraulic relay to the starting tank with standard fuel. The hydraulic relay is connected to the second stage of the fuel pump. The throttle installed at the output of the first stage of the fuel pump is made in conjunction with a controlled valve of the preliminary stage.

The disadvantage is the complex pneumohydraulic circuit of the engine, the presence of a large number of valves and regulators and piping, and as a result, a lot of weight and low reliability and problems when starting and turning off the engine.

Known: rocket engine with a controlled thrust vector according to the patent of the Russian Federation No. 2412730, IPC F02k 9/42, publ. 02/20/2011, prototype rocket engine. This rocket engine contains a combustion chamber with a nozzle, a gas generator and a turbopump assembly containing oxidizer pumps, fuel and an additional fuel pump and a thrust vector control cylinder.

The disadvantages of this design are as follows

1. Large transverse dimensions of the rocket engine due to the location of the TNA near the combustion chamber.

2. Difficulties in controlling the thrust vector associated with the need to apply great efforts to the force ring to create a certain moment with a small shoulder.

The objective of the invention is to improve the specific characteristics of the rocket engine, increasing its reliability and reducing the cost of launching missiles.

The solution to this problem was achieved in an anti-aircraft missile containing a warhead and an axisymmetric body with oxidizer and fuel tanks and a liquid propellant rocket engine containing a combustion chamber and VT: A, because according to the invention, the combustion chamber and THA are installed in series along the axis of the rocket, and in four control nozzles are radially mounted in the middle part of an anti-aircraft missile.

The solution to this problem was achieved in a liquid propellant rocket engine containing a turbopump assembly, which in turn contains a turbine, oxidizer and fuel pumps and an additional fuel pump, and a combustion chamber having a cylindrical part with oxidizer and fuel nozzles and a nozzle with a main fuel manifold, characterized in that that the turbopump unit and the combustion chamber are installed coaxially, while the turbine is made inside the cylindrical part of the combustion chamber, the combustion chamber is made of two-zone and contains additional e fuel nozzles on its cylindrical part below the turbine, and four control nozzles are connected to the first zone of the combustion chamber by pipelines, while a hinge is made between the combustion chamber and the oxidizer pump, containing an inner ball and an external spherical shell connected to the power brackets .. Exit from the pump the oxidizing agent may be connected by pipelines containing the oxidizing valve to the combustion chamber. In the middle part of the cylindrical part of the combustion chamber, an annular manifold can be made, into the cavity of which additional fuel nozzles exit. Below the annular manifold, an additional manifold may be made, into the cavity of which additional fuel nozzles exit.

The invention is illustrated in the drawings of figures 1 ... 7, where

- figure 1 shows a diagram of an anti-aircraft missile,

- figure 2 shows a second variant of an anti-aircraft missile with control nozzles,

- figure 3 shows a diagram of the rocket engine,

- figure 4 shows the design of the combustion chamber,

- figure 5 shows the TNA,

- figure 6 shows the connection of the oxidizer pump and the combustion chamber,

- Fig.7 shows the power hinge.

The rocket (FIGS. 1 and 2) contains a head part 1, an axisymmetric body 2, an oxidizer tank 3, a fuel tank 4, a liquid rocket engine 5, which in turn contains a combustion chamber 6 with a nozzle 7 and a TNA 8. The combustion chamber 6 and the TNA 7 are arranged sequentially along the axis of the rocket. TNA 8 comprises a turbine 9, an oxidizer pump 10, a fuel pump 11, and an additional fuel pump 12 having a common shaft 13.

The combustion chamber 6 (FIGS. 1 and 2) contains a head 14 and a cylindrical part 15. The nozzle 7 contains a tapering part 16 and an expanding part 17 and a main fuel manifold 18 in the lower part. On the combustion chamber 9, the middle and intermediate manifolds 19 and 20 are made, respectively.

The turbine 9 is installed in the combustion chamber 6, more precisely, at the top of its cylindrical part 15. In this case, the combustion chamber 6 is dual-zone and contains a first zone 21 and a second 22 separated by a turbine 9. In the upper part of the TNA 8 there is a hinge 23 to which rods 24 of cylinders control 25, which are fixed on the upper power ring 26. The lower power ring 27 is also fixed on the housing 2. Between the head 13 of the combustion chamber 6 and the oxidizer pump 10 a power joint 28 is made, which is fixed on the lower power ring 27 using rods 29 (Fig. one).

The rocket can be equipped with four control nozzles 30 mounted radially and connected by pipelines 31 containing valves 32 with the first zone 18 of combustion chamber 6. (FIG. 2).

The turbine 9 includes a nozzle apparatus 33, an impeller 34, a disk 35 and a straightening apparatus 36. The TNA 8 has three bearings 37 ... 39 and a seal 40 (Fig. 4).

As tapering 16 and expanding 17 parts of the nozzle 7 are made with the possibility of regenerative cooling (Fig.1 and 2) and contain two walls; the inner wall 41 and the outer wall 42 with a gap 43 between them for the passage of cooling fuel. The cavity of the gap 43 communicates with the cavity 44 of the main manifold of the fuel 18. Inside the combustion chamber 6 (Fig. 3), an upper plate 45 is made, and a lower plate 46 with a gap (cavity) between them 47. A cavity 48 is made above the upper plate 45. Inside the head 14 combustion chamber 6 has an oxidizer nozzle 49 and a fuel nozzle 50. The oxidizer nozzle 49 communicates a cavity 49 with an internal cavity 51 of the combustion chamber 6. Fuel nozzle 50 communicates a cavity 48 with an internal cavity 51. Ignition devices 52 are installed on the head 13 of the combustion chamber 5. To the head 14 cameras combustion 6 is connected by several pipelines 53 containing oxidizer valves 54 The head 14 of the combustion chamber 6 and the oxidizer pump 16 are connected by a sleeve 55.

A pipe 56 with a valve 57 is connected to the main fuel manifold 18. The other end of the pipeline 55 is connected to the outlet of the fuel pump 11. To the middle manifold 19 is connected to a pipe 58 having a flow regulator 59 and valve 60, the other end of pipe 58 is connected to the outlet of the additional pump fuel 12. The entrance to the additional fuel pump 12 by a pipe 61 is connected to the outlet of the fuel pump 11. The cavity 62 of the additional manifold 20 is connected to the gap 40 and additional fuel nozzles 63.

The engine contains a purge system with an inert gas cylinder 64, a pipe 65 and a valve 66. The pipe 65 is connected to the main manifold of the fuel 18. At the bottom of the rocket there is a bottom protection 67. The rocket is equipped with a system of pressurization of the oxidizer tanks 3 and fuel 4. This system contains a cylinder 68 and pressurization ducts 69 and 70 .. In the head part 1, a guidance unit 71, an explosive device 72 and an on-board computer 73 are connected to which an ignition device 52 and valves 54, 60, 57 and 66 and a flow regulator 59 are connected by electrical connections. (FIG. .1 and 6).

Fig. 7 shows a ball joint 28. This joint comprises an inner joint 74 and a hollow outer joint 75. The hollow outer joint 75 consists of two parts: an upper 76 and a lower 77, connected by bolts 78 through a gasket 79. In the upper and lower parts 76 and 77, conical holes 80 are made for the passage of the shaft 13 and ensure its swing at 5 ... 7 deg. The inner hinge 74 is fixed to the shaft 13 with a key 81.

: The launch of the rocket engine is as follows.

The oxidizer valves 48 and valves 52 and 55 are opened.

The oxidizing agent and fuel enter the combustion chamber 2 more precisely in its first zone 16, where they are ignited by ignition devices 47 and burn out at relatively low temperatures of 500 ... 700 ° C. Through additional fuel nozzles 50, a large part of the fuel flow enters the second zone, where it burns at a temperature of 3500 to 4000 ° C, which provides the maximum specific thrust of the rocket engine.

The regulation of the operation of the rocket engine is carried out by the flow regulator 46. In this case, the flow rate of the combustion products through the turbine 11 and the temperature of the combustion products at the entrance to it are changed.

The LRE shutdown is performed in the reverse order. After closing all the valves of the oxidizer 31 and the fuel 34 and 37, the purge valve 43 is opened and the inert gas from the cylinder 41 through the pipe 42 enters the main manifold of the fuel 20 by blowing the combustion chamber 1 of the engine to clear the remaining fuel.

Regulation of the direction of the thrust vector is performed using the control cylinders 25 turning the rocket engine around the power joint 28.

The application of the invention will allow:

1. To reduce the transverse dimensions of the rocket engine due to the placement of the combustion chamber and TNA in series in a single line along one common axis

2. Reduce engine weight due to the lack of a gas generator. The function of the gas generator is performed by the first zone of the combustion chamber.

3. Simplify the engine layout for the same reasons.

4. Reduce the force to control the thrust vector of the rocket engine due to the maximum increase in the shoulder of the lever for applying the force of the hydraulic cylinder.

5. To reduce the harmful mutual influence of hot and cold (working on cryogenic fuel components) components and assemblies.

Claims (6)

1. An anti-aircraft missile containing a warhead and an axisymmetric body with oxidizer and fuel tanks and a liquid propellant rocket engine containing a combustion chamber and TNA, characterized in that the combustion chamber and TNA are installed sequentially along the axis of the missile, and four control radially mounted in its middle part nozzles. 2. A liquid rocket engine containing a turbopump assembly, comprising, in turn, a turbine, oxidizer and fuel pumps and an additional fuel pump, and a combustion chamber having a cylindrical part with oxidizer and fuel nozzles and a nozzle with a main fuel manifold, characterized in that the turbopump unit and the combustion chamber are installed coaxially, while the turbine is made inside the cylindrical part of the combustion chamber, the combustion chamber is dual-zone and contains additional fuel nozzles on its cylinder of the lower part of the turbine, and four control nozzles are connected to the first zone of the combustion chamber by pipelines, while a spherical hinge is made between the combustion chamber and the oxidizer pump, containing an inner ball and an outer spherical shell connected to power rods, and an upper ball hinge is made at the upper end of the TNA with which the rods of the control mechanisms are connected. 3. The liquid rocket engine according to claim 2, characterized in that the outlet of the oxidizer pump is connected by a pipe containing oxidizer valves to the combustion chamber. 4. The liquid rocket engine according to claim 2 or 3, characterized in that an annular manifold is made in the middle part of the cylindrical part of the combustion chamber, into the cavity of which additional fuel nozzles exit. 5. The liquid propellant rocket engine according to claim 2 or 3, characterized in that an additional manifold is made below the annular collector, into the cavity of which additional fuel nozzles exit. 6. The liquid propellant rocket engine according to claim 4, characterized in that an additional manifold is made below the annular collector, into the cavity of which additional fuel nozzles exit.

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