CN114294123A

CN114294123A - Combined propulsion system of split-exhaust air turbine rocket

Info

Publication number: CN114294123A
Application number: CN202210099718.5A
Authority: CN
Inventors: 杜金峰; 史新兴; 陈玉春
Original assignee: Northwestern Polytechnical University
Current assignee: Northwestern Polytechnical University
Priority date: 2022-01-27
Filing date: 2022-01-27
Publication date: 2022-04-08

Abstract

The invention relates to a split-exhaust air turbine rocket combined propulsion system, which comprises a core machine of an air turbine rocket engine, wherein the core machine consists of a rocket gas generator, a turbine and a gas compressor which are coaxially arranged; the rocket fuel generator is connected with the rocket fuel conveying device, rich fuel gas generated when the rocket fuel generator works drives the turbine to drive the gas compressor, the gas compressor sucks air and boosts the pressure, the boosted air directly enters the first combustion chamber and is combusted with fuel injected into the first combustion chamber in the first combustion chamber to release heat, and the fuel gas expands through the outer spray pipe to do work to generate thrust; meanwhile, the gas after the turbine enters a second combustion chamber for combustion, and then is directly sprayed out from the inner spray pipe to generate thrust. The invention can solve the problem of high pressure intensity of the gas generator, replaces the traditional mixing of gas and air, and widens the working speed range of the engine.

Description

Combined propulsion system of split-exhaust air turbine rocket

Technical Field

The invention relates to an air suction type propulsion system, in particular to an air turbine rocket combined propulsion system.

Background

With the pursuit of speed by human, the working range of the engine is wider and wider, the requirement of the aircraft on the engine is higher and higher, the aircraft has a wide speed range, starts at zero speed, has good performance and the like. The adaptability of the engine to different speeds is required to be improved, and an air turbine rocket engine (ATR) is proposed by the father Godard of the United states rocket in 1932 and has certain application on missiles as an engine with a wider speed range. An engine operating at Ma0-Ma3 is not necessarily capable of meeting the wide range of operating conditions present. This is mainly due to the fact that the chamber pressure of the gas generator is too high when the engine gas generator is operating at high mach, which leads to increased engine design difficulties. This problem of excessive chamber pressure limits the operating speed range of ATR engines.

The turbine rocket engine of US3237400A was applied in 1966 in the united states and requires only liquid hydrogen fuel to be carried, air near the intake port is drawn into the supply system to be cooled with liquid hydrogen and then the cooled air and hydrogen are used as fuel for the rocket engine, and no oxidant is carried as in the conventional ATR engine, increasing the operating time of the engine and thus raising the problem of high pressure in the gas generator.

The invention disclosed in the patent number CN103437914A provides a variable cycle ATR engine which can be maneuvered at a high speed in a short time and can economically cruise at a low speed for a long time, and has a relatively quick starting performance. The invention is based on an aircraft engine, and a gas generator is added in front of a turbine of the aircraft engine. In a high-speed maneuvering state, gas generated by the gas generator drives the blade tip turbine, and the blade tip turbine drives the gas compressor to work. The long-time low-speed economical cruise condition operates in the form of an aircraft engine. This solution does not fundamentally solve the problem of gasifier pressure mismatch.

The patent number CN 106050472A discloses an ATR parallel combination engine, which works with ATR engine at low speed and RBCC engine at high speed, and widens the working range of the engine. But does not essentially address the problem of high ATR engine gas generator pressures.

The problem of overhigh pressure of a gas generator exists in the air-breathing turbine rocket engine or the modification, the working speed range of the engine is seriously influenced, the overhigh pressure increases the manufacturing difficulty, and the academic papers and academic papers about the ATR propulsion system exhausting gas respectively do not exist at present.

Disclosure of Invention

The invention aims to avoid the defects of the prior art and provide a separate exhaust air turbine rocket combined propulsion system which can solve the problem of high pressure of a fuel gas generator, replaces the traditional mixing of fuel gas and air and widens the working speed range of an engine.

In order to achieve the purpose, the invention adopts the technical scheme that: a kind of split-exhaust air turbine rocket makes up the propulsion system, including the core machine of the air turbine rocket engine composed of gas generator of rocket and turbine and air compressor that set up coaxially, the said core machine is set up in the engine case;

the rocket fuel gas generator is arranged behind the gas compressor, and a first combustion chamber is correspondingly arranged at the periphery of the rocket fuel gas generator; a second combustion chamber is arranged at the outlet of a turbine arranged behind the rocket gas generator; the rocket gas generator and the gas compressor are arranged independently;

the tail nozzle of the second combustion chamber is an inner nozzle, and the tail nozzle of the first combustion chamber is an outer nozzle; the outer spray pipe is arranged on the periphery of the inner spray pipe;

the rocket fuel gas generator is connected with the rocket fuel conveying device, rich fuel gas generated when the rocket fuel gas generator works drives the turbine to drive the gas compressor, the gas compressor sucks air and boosts the pressure, the boosted air directly enters the first combustion chamber and is combusted with fuel sprayed into the first combustion chamber in the first combustion chamber to release heat, and the gas expands through the outer spray pipe to do work to generate thrust; meanwhile, the gas after the turbine enters a second combustion chamber for combustion, and then is directly sprayed out from the inner spray pipe to generate thrust.

Further, the first combustion chamber is of a double-wall structure, the outer wall of the first combustion chamber is the engine case, and the inner wall of the first combustion chamber is the outer wall of the second combustion chamber.

Furthermore, a throat corresponding to the inner wall of the outer spray pipe is arranged on the inner wall of the inner spray pipe.

Further, the gas compressor, the turbine, the first combustion chamber, the second combustion chamber, the outer spray pipe and the inner spray pipe are all coaxially arranged.

The invention has the beneficial effects that: compared with the prior art, the invention can increase the working speed range of the engine, simplifies the circulation of the engine and the structure of parts, adopts a split-exhaust mode to replace the traditional mixing of fuel gas and air, reduces the balance condition of the engine, reduces the pressure of the fuel gas generator of the engine and reduces the design difficulty of the fuel gas generator and the overall engine. The turbine outlet is discharged at the ambient pressure, the pressure requirement of a gas generator of the engine which is discharged in a split mode is low, and the working speed range of the engine is widened under the same pressure of the gas generator.

Drawings

Fig. 1 is a schematic structural view of the present invention.

1. A compressor; 2. a gas generator; 3. a turbine; 41. a first combustion chamber; 42. a second combustion chamber; 5. an outer jet nozzle; 6. an inner tail nozzle; 7. an engine case; 8. the inner wall of the combustion chamber; 9. inner tail nozzle throat.

Detailed Description

The principles and features of this invention are described below in conjunction with the following drawings, which are set forth by way of illustration only and are not intended to limit the scope of the invention.

ATR engines present problems, mainly in that the chamber pressure limitation in the gas generator leads to an engine flight mach number not higher than 3Ma, which, although adjustable by means of engine geometry and cycle parameter adjustments, does not fundamentally solve the problem of excessive engine gas generator chamber pressure, in order to solve the above problem, the invention provides the following embodiments:

example 1: a kind of split-exhaust air turbine rocket makes up the propulsion system, including the core machine of the air turbine rocket engine composed of gas generator 2 of rocket, turbine 3 and air compressor 1 that set up coaxially, the core machine is set up in the engine case 7;

the rocket gas generator 2 is arranged behind the gas compressor 1, and a first combustion chamber 41 is correspondingly arranged at the periphery of the rocket gas generator 2; a second combustion chamber 42 is provided at the outlet of the turbine 3 provided after the rocket gas generator 2; the rocket gas generator 2 and the gas compressor 1 are arranged independently;

the tail pipe of the second combustion chamber 42 is an inner pipe 6, and the tail pipe of the first combustion chamber 41 is an outer pipe 5; the outer spray pipe 5 is arranged at the periphery of the inner spray pipe 6;

the rocket fuel generator 2 is connected with the rocket fuel conveying device, rich fuel gas generated when the rocket fuel generator 2 works drives the turbine 3 to drive the gas compressor 1, the gas compressor 1 sucks air and boosts the pressure, the boosted air directly enters the first combustion chamber 41 and is combusted with fuel sprayed into the first combustion chamber 41 in the first combustion chamber 41 to release heat, and the gas expands through the outer spray pipe 5 to do work to generate thrust; meanwhile, the combustion gas after the turbine 3 enters the second combustion chamber 42 for combustion, and then is directly ejected from the inner nozzle 6 to generate thrust.

The first combustion chamber 41 has a double-wall structure, the outer wall of the first combustion chamber 41 is the engine case 7, and the inner wall of the first combustion chamber 41 is the outer wall of the second combustion chamber 42. The inner wall of the inner spray pipe 6 is provided with a throat 9 corresponding to the inner wall of the outer spray pipe 5.

The compressor 1, the turbine 3, the first combustion chamber 41, the second combustion chamber 42, the outer nozzle 5 and the inner nozzle 6 are all coaxially arranged.

The compressor 1 and the turbine 3 are connected through a shaft, the fuel gas generator 2 is arranged between the compressor 1 and the turbine 3, the first combustion chamber 41 is connected with the compressor 1, the first combustion chamber 41 is connected with the outer spray pipe 5 behind the compressor 1, the inner spray pipe 6 is arranged behind the turbine 3, and the shaft centers of all the components are in a straight line.

The working mode of the embodiment is as follows:

a gas generator 2 of the air turbine rocket engine generates high-temperature and high-pressure gas to drive a turbine 3 to rotate, and the turbine 3 drives a compressor 1 to rotate through a shaft to transmit work to the compressor 1. The compressor 1 sucks air and compresses the air. The compressed air flows to the first combustion chamber 41, and heat is released in the first combustion chamber 41 by combustion with the fuel injected into the first combustion chamber 41. The gas expands through the outer nozzle 5 to do work to generate thrust. The combustion gases after the turbine 3 flow directly out of the inner lance 6 through the second combustion chamber 42.

The above description is only for the purpose of illustrating the preferred embodiments of the present invention and is not to be construed as limiting the invention, and any modifications, equivalents, improvements and the like that fall within the spirit and principle of the present invention are intended to be included therein.

Claims

1. The split-exhaust air turbine rocket combined propulsion system is characterized by comprising a core machine of an air turbine rocket engine, wherein the core machine is composed of a rocket gas generator (2), a turbine (3) and a gas compressor (1) which are coaxially arranged, and the core machine is arranged in an engine casing (7);

the rocket gas generator (2) is arranged behind the compressor (1), and a first combustion chamber (41) is correspondingly arranged at the periphery of the rocket gas generator (2); a second combustion chamber (42) is arranged at the outlet of a turbine (3) arranged behind the rocket gas generator (2); the rocket gas generator (2) and the gas compressor (1) are arranged independently;

the tail pipe of the second combustion chamber (42) is an inner pipe (6), and the tail pipe of the first combustion chamber (41) is an outer pipe (5); the outer spray pipe (5) is arranged at the periphery of the inner spray pipe (6);

the rocket fuel gas generator (2) is connected with the rocket fuel conveying device, rich fuel gas generated when the rocket fuel gas generator (2) works drives the turbine (3) to drive the gas compressor (1), the gas compressor (1) sucks air and pressurizes the air, the pressurized air directly enters the first combustion chamber (41) and burns with fuel sprayed into the first combustion chamber (41) in the first combustion chamber (41) to release heat, and the gas expands through the outer spray pipe (5) to do work to generate thrust; meanwhile, the combustion gas after the turbine (3) enters a second combustion chamber (42) for combustion, and then is directly sprayed out from the inner spray pipe (6) to generate thrust.

2. The combined propulsion system of a divided exhaust air turbine and rocket as recited in claim 1, wherein said first combustion chamber (41) is of a double-walled structure, an outer wall of said first combustion chamber (41) being said engine case (7), an inner wall of said first combustion chamber (41) being an outer wall of said second combustion chamber (42).

3. The combined propulsion system of a divided exhaust air turbine rocket as claimed in claim 1, characterized in that a throat (9) corresponding to the inner wall of the outer nozzle (5) is provided on the inner wall of the inner nozzle (6).

4. The split-exhaust air turbine rocket combined propulsion system according to any one of claims 1 to 3, characterized in that the compressor (1), the turbine (3), the first combustion chamber (41), the second combustion chamber (42), the outer nozzle (5) and the inner nozzle (6) are all coaxially arranged.