RU2300007C1 - Vortex rocket motor - Google Patents

Abstract

FIELD: rocketry.

SUBSTANCE: invention can be used in liquid propellant and oxidant motors. According to invention, in proposed vortex rocket motor containing combustion chamber in form of cylinder with bottom at one side and dome to change into neck at other side, and nozzles to deliver propellant and oxidant, cone is made on bottom of chamber to turn vortex, some nozzles are uniformly spaced on dome to provide formation of vortex and designed to deliver propellant and part of oxidant, and one nozzle is located in cone to turn vortex and is designed to deliver remaining part of oxidant. Propellant and oxidant delivery nozzles are arranged tangentially relative to circumference of dome at angle of 60° to dome surface. motor is furnished with deflecting blades to compensate for reactive rotation of chamber secured on cone which is mated with chamber neck.

EFFECT: simplified design, reduced mass.

3 cl, 5 dwg

Description

The invention relates to the field of rocket science and can be used to design traction engines for liquid fuel rockets with an oxidizing agent.

Known rocket engine containing a combustion chamber, the working part of which is intended for combustion of fuel and oxidizer and is connected through the neck with an expanding nozzle for the release of gases resulting from combustion. The working part is fed with an oxidizing agent from the end opposite the neck and is covered by a porous surface layer of a thermally structured composite material, into which fuel is supplied from the outside, opposite the working part. Part of the fuel is supplied to the working part through a porous surface layer and serves to power the engine. Another part of the fuel that does not pass through the porous surface layer is directed to the neck and is designed to cool it. Such a small-mass rocket engine containing a small number of parts is easy to manufacture (see RF patent No. 2266423, IPC F02K 9/64).

The disadvantage is the relatively large nozzle, in addition, the presence of a porous membrane creates a high resistance to high-speed fuel supply.

Known liquid propellant rocket engine containing helium additive, including a combustion chamber with a nozzle, which are equipped with regenerative cooling channels, a turbopump system for supplying oxidizer and fuel to the combustion chamber of the engine, a helium circuit for regenerative cooling of the chamber, including a turbine drive supply unit (see application for invention No. 2004106581, IPC F02K 1/00).

However, this engine also has a large mass, the complexity of the storage conditions of the components of the oxidizer and fuel.

Closest to the proposed solution is an engine, which consists of a vortex chamber for igniting and burning fuel, a magnetic plasma accelerator and a gas-dynamic nozzle connected coaxially and conjugated by end surfaces, on which nozzles for supplying water (steam) and corona electrodes are placed. Moreover, the magnetic accelerator of ion-radiation plasma provides the creation of a rotating magnetic field and the separation of ionized plasma gases by spectral mass, and the nozzle ends with probes for the concentration and confluence of ionized gases (see RF patent No. 2099572, IPC F02K 11/00).

The disadvantage is the large mass and complexity of the design. In addition, the working fluid formed in the combustion chamber begins acceleration from the transition of the combustion chamber into the nozzle and to the edge of the nozzle. Its outflow occurs randomly, which leads to the occurrence of vibrations transmitted to the rocket itself, so the rocket must have excessive strength.

The objective of the invention is to simplify the design of the engine and reduce its mass by transferring the function of the nozzle to the combustion chamber with the exception of oxidation of the walls of the combustion chamber.

The problem is solved in that in a vortex rocket engine containing a combustion chamber in the form of a cylinder with a bottom on one side and a vault for passing into the neck on the other hand, nozzles for supplying fuel and an oxidizing agent, according to the invention, a cone for turning the vortex is made at the bottom of the chamber, part of the nozzles is located evenly on the arch with the possibility of ensuring the formation of a vortex and is intended to introduce fuel and part of the oxidizer, and one nozzle is located in the cone to turn the vortex and is intended to enter the remaining part the oxidizing agent.

The fuel and oxidizer nozzles are tangential to the circumference of the arch at an angle of 60 ° to the surface of the arch.

The engine is equipped with deflecting blades to compensate for the jet rotational movement of the chamber, mounted on a cone, which is paired with the neck of the chamber.

The invention is illustrated by drawings. In Fig.1 shows a General view of the engine in section, in Fig.2 is a transverse section along aa, in Fig.3 is a transverse section along bb, Fig.4 is a transverse section along bb, in Fig. 5 - the location of one of the nozzles 2. The positions in the drawings indicate:

1 - combustion chamber;

2 - injection nozzles for 100% fuel and 60% oxidizer;

3 - a cone in the form of a recess in the bottom of the combustion chamber;

4 - nozzle input 40% of the oxidizing agent;

5 - deflecting blades;

6 - cone fastening deflecting blades;

7 - arch;

8 - region of the "cold" ascending vortex;

9 - "hot" torch of the final acceleration of the working fluid.

The engine contains a combustion chamber 1, made in the form of a cylinder with a smooth transition (vault) 7 to the neck of the chamber to exit the working fluid. The cylinder also has a smooth transition to the bottom of the chamber, in the center of which there is a cone 3, on top of which a nozzle 4 for introducing 40% of the oxidizer is mounted. It has a dummy valve (not shown in the drawing). The combustion chamber is equipped with nozzles 2 for introducing 100% fuel and 60% of the oxidizer, located evenly on the arch along its perimeter. The nozzles 2 are mounted tangentially to the circumference of the vault 7 and at an angle of 60 ° to the surface of the vault 7. The engine is equipped with deflecting vanes 5 having, for example, the profile of the turbine blades and designed to compensate for the jet rotational movement of the chamber 1 and create additional traction. The blades are mounted on a cone 6, which is paired with the neck of the chamber 1.

The device operates as follows. When the nozzle 4 is muffled, fuel and oxidizer enter through the nozzles 2 and are ignited. In this case, a vortex torch ("cold" vortex) is formed, scattered along the walls. The components of the torch contain a lot of fuel and do not allow the walls of the combustion chamber to oxidize during operation. The “cold” vortex 8, rising to the bottom of the chamber, turns by a cone 3 to the axis and then the missing 40% of the oxidizer is injected into the nozzle 4. There is a second "hot" torch 9 of the final acceleration of the working fluid, accelerating from cone 3 to the neck.

It should be noted that the acceleration of the working fluid is possible to the top of the "hot" torch. Therefore, it is necessary that the entire torch is in the combustion chamber. This is achieved by experimental selection of the length of the combustion chamber for each type of fuel and oxidizer. The calculation boils down to the fact that the entire torch of the "hot" vortex is inside the combustion chamber, while maximum thrust is achieved. The vortex emerging from the neck has a parasitic twist, which is partially removed by introducing deflecting blades 5 into the edges of the vortex. The number and profile of the blades depends on the number of nozzles 2, the flow rate of the working fluid, and is selected experimentally. The engine can work without a cone with deflecting blades, but in this case it is necessary to install two engines on a rocket or their paired number with right- and left-rotating vortices.

The invention allows to reduce the vibration of the engine and, accordingly, the aircraft, and also to increase traction by increasing the speed of the expiration of the working fluid from the combustion chamber, as well as reduce the weight and dimensions of the engine. The structural rigidity requirements of the aircraft can be reduced. The combustion chamber is protected from the aggressive effects of the oxidizing agent due to two-stage combustion of fuel in a vortex space.

Claims (3)

1. A vortex rocket engine containing a combustion chamber in the form of a cylinder with a bottom on one side and a vault for passing into the neck on the other side, nozzles for supplying fuel and an oxidizing agent, characterized in that a cone for turning the vortex is made at the bottom of the chamber, part of the nozzles is located evenly on the arch with the possibility of providing a vortex and is intended to enter fuel and part of the oxidizer, and one nozzle is located in the cone to turn the vortex and is intended to enter the remaining part of the oxidizer. 2. The vortex rocket engine according to claim 1, characterized in that the fuel and oxidizer nozzles are arranged tangentially to the circumference of the arch at an angle of 60 ° to the surface of the arch. 3. The vortex rocket engine according to claim 1, characterized in that it is equipped with deflecting blades to compensate for the jet rotational movement of the chamber, mounted on a cone, which is paired with the neck of the chamber.

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