CN112160835A

CN112160835A - Combustion chamber of turbofan aircraft engine

Info

Publication number: CN112160835A
Application number: CN202011101339.2A
Authority: CN
Inventors: 谭健
Original assignee: Individual
Current assignee: Individual
Priority date: 2019-11-03
Filing date: 2020-10-09
Publication date: 2021-01-01
Also published as: CN110848026A

Abstract

The combustion chamber of the turbofan aircraft engine is a combustion chamber which can rotate at a relatively high speed and is formed by reforming the original combustion chamber. Namely, the combustion chamber is modified into a rotor part and a stator part. The narrow circular compressed air inlet and the circular gas outlet are used as boundaries, a rotor is arranged inside an inner circle of the narrow circular compressed air inlet, and a stator is arranged outside an outer circle of the narrow circular compressed air inlet. Therefore, the air compressor can directly compress air into the combustion chamber through the narrow circular ring-shaped compressed air inlet without obstacles, and the efficiency of the air compressor is obviously improved. As long as a narrow circular compressed air inlet with a small width is made, the oxygen delivery amount of the combustion chamber is greatly increased. And the thrust can be greatly increased by adding a little fuel quantity. The weight of the combustion chamber is not increased by the modification, and the thrust-weight ratio is also greatly increased. The modification does not add any difficulty to the manufacture and the use of materials. The modification can be suitable for military and civil aircraft engines, and can also be suitable for ships and warships using gas turbines as power. The improvement can realize that the fuel injection direction is vertical to the compressed air injection direction, and is favorable for complete combustion of fuel.

Description

Combustion chamber of turbofan aircraft engine

A combustion chamber of a turbofan aircraft engine relates to the field of aircraft engines.

The basic technology of the combustion chamber of a turbofan aircraft engine is the turbofan aircraft engine.

The combustion chamber of a turbofan aircraft engine is characterized in that: the combustion chambers in the basic technology are modified into combustion chambers capable of rotating at a relatively high speed. The original combustion chamber including a flame tube is transformed into a rotor part and a stator part, a narrow annular compressed air inlet and an annular combustion chamber gas outlet are used as boundaries, the rotor is arranged inside an inner circle of the narrow annular compressed air inlet, and the stator is arranged outside an outer circle of the narrow annular compressed air inlet. Therefore, the air compressor can directly compress air into the annular combustion chamber through the narrow annular compressed air inlet without obstacles. Because of no obstacle, the air compressor can compress air in the multi-direction combustion chamber under the same power, and the efficiency is obviously improved. As long as a narrow annular compressed air inlet with a small width is made, the oxygen delivery amount of the annular combustion chamber is greatly increased. And the thrust of the engine can be greatly increased by adding a little fuel. The weight of the combustion chamber is not increased by the modification, and the thrust-weight ratio is greatly increased. Meanwhile, the modification does not add any difficulty to the manufacturing and the materials.

The above modifications can be adapted to military and civil aircraft engines.

The above modifications may also be applied to gas turbines powered by ships and civil ships and certain generator sets.

The transformation can realize that the direction of fuel oil sprayed into the combustion chamber is mutually vertical to the direction of compressed air sprayed into the combustion chamber, and the annular arrangement of a plurality of fuel injection nozzles is favorable for the complete combustion of the fuel oil of the aircraft engine. The technology can adapt to an aircraft engine; can also adapt to the gas turbine; it can also accommodate non-stroke heat engines.

The details of the combustion chamber of a turbofan aircraft engine will now be described with reference to the accompanying drawings.

FIG. 1 is a schematic diagram of the operating principle of a combustion chamber of a turbofan aircraft engine, and is a cross-sectional view taken along the axis.

Fig. 2 is a rotary exploded view of the combustion chamber of fig. 1.

Fig. 3 is an exploded view of the third pedestal circle adjacent the combustion chamber of fig. 1.

All the figures are schematic.

As shown in fig. 1, fig. 1 mainly includes a fan 1, a first shaft base 2, an inner shaft 3, a low-pressure compressor 4, an outer duct 5, a second shaft base 6, an outer shaft 7, a high-pressure compressor 8, an inner duct 9, a third shaft base 10, a narrow annular compressed air inlet 11, a fuel oil inlet 12, a combustor casing rotor 13 (a black casing part close to a rotating shaft), a combustor liner rotor 13 '(a black broken line in the figure indicates that a white broken line between the black casing and the black rotor casing is a space, and the space indicates whether the white broken line is in accordance with the basic technology, the following is the same), a combustor casing stator 14 (a black casing part close to a casing), a combustor liner stator 14' (a black broken line indicates that a white broken line between the black casing and the black casing is a space), a high-pressure turbine 15, a low-pressure turbine 16, a tail nozzle 17, an annular combustor gas outlet, and guide vanes 18 (guide vanes are not shown in the figure), The combustor comprises a combustor flame tube inner space 19, an inner bypass casing 20, an outer bypass casing 21 and the like. (in FIG. 1, the basic solution is the standard contrary to the basic technology).

Some of the accessory parts are not shown in fig. 1 because of the simplified working principle.

Turning to the combustion chamber of fig. 1 in more detail. The right side of the third shaft seat 10 is filled with black and within black to indicate that the space 19 in the liner is not filled with color and is a combustion chamber. The annular combustion chamber has a gas outlet 18 with guide vanes not shown. Adjacent the shaft is the rotor of the combustion chamber and the other combustion chamber component is the stator of the combustion chamber. The rotor to stator boundary is: a narrow annular compressed air inlet 11 to an annular combustion chamber gas outlet 18. This forms a combustion chamber that can rotate at relatively high speeds. The air compressor can directly compress air into the space 19 in the flame tube of the annular combustion chamber through the narrow annular compressed air inlet 11 without obstacles, and the efficiency of the air compressor is obviously improved because of no obstacles. As long as the width of the narrow annular compressed air inlet 11 is increased a little, the oxygen delivery amount in the annular combustion chamber is greatly increased, and as long as the fuel amount is increased a little, the thrust of the engine is greatly increased. The weight of the combustion chamber is not increased, and the thrust-weight ratio is increased greatly.

As shown in fig. 2, fig. 2 is a rotary round exploded view of the combustion chamber. Fig. 2 mainly comprises an inner shaft 3, an outer shaft 7, a narrow circular ring-shaped compressed air inlet 11, a fuel oil inlet 12, a combustor casing rotor 13, a combustor liner rotor 13 ', a combustor casing stator 14, a combustor liner stator 14', a combustor liner inner space 19 and the like. In this way, the combustion chamber is mainly composed of the rotor of the combustion chamber housing, the stator and the rotor of the liner, the stator and the inner space of the liner of the combustion chamber.

See figure 2 for further details. A plurality of fuel inlets 12 are regularly arranged into a circle shape; it can also be seen from both fig. 1 and 2 that: the direction of the fuel injection inlet 12 is perpendicular to the direction of the narrow circular compressed air inlet 11, which is favorable for complete combustion of the fuel. The inner space 19 from the liner rotor 13 'to the liner stator 14' is the inner space of the liner.

As shown in fig. 3, fig. 3 mainly comprises an inner shaft 3, an outer shaft 7, a third shaft seat 10, an inner ducted casing 20, and the like.

Fig. 3 primarily shows that the third axle seat is air permeable, while the first and second axle seats are air permeable. They are all indicated by dashed lines in fig. 1.

The working principle of the combustion chamber of a turbofan aircraft engine will be described with reference to the accompanying drawings.

The fan 1 absorbs a large amount of air, a part of which enters the outer duct 5 and a part of which enters the inner duct 9. Air entering the inner duct 9 is compressed by the low-pressure compressor 4 and the high-pressure compressor 8, enters the inner space 19 of the flame tube of the annular combustion chamber through the narrow annular inlet 11, is mixed with atomized fuel oil entering the fuel oil inlet 12, is ignited and rapidly combusted, and generated high-temperature and high-pressure fuel gas rushes to the high-pressure turbine 15 under the action of the guide vanes at the outlet 18, then rushes to the low-pressure turbine 16 and finally rushes to the tail nozzle 17, so that thrust is generated at this time. But the speed will be slower after each impact until the next impact. The high-pressure turbine drives the high-pressure compressor, and the low-pressure turbine drives the fan and the low-pressure compressor. The process is repeated in cycles. The aircraft can take off when the aircraft reaches the take-off speed.

Claims

1. The basic technology of the combustion chamber of a turbofan aircraft engine is the turbofan aircraft engine; the combustion chamber of a turbofan aircraft engine is characterized in that: various combustion chambers in the basic technology are transformed into combustion chambers capable of rotating at a relatively high speed; the original combustion chamber including a flame tube is reformed into a rotor and a stator; the narrow circular compressed air inlet and the circular combustion chamber gas outlet are used as boundaries, a rotor is arranged inside an inner circle of the narrow circular compressed air inlet, and a stator is arranged outside an outer circle of the narrow circular compressed air inlet; therefore, the air compressor can directly compress air into the annular combustion chamber through the narrow annular compressed air inlet without obstacles; because of no obstacle, the gas compressor can compress gas in a multi-way combustion chamber under the same power, and the efficiency is obviously improved; the oxygen delivery amount of the annular combustion chamber can be greatly increased only by making a narrow annular compressed air inlet with a small width; the thrust of the engine can be greatly increased as long as a little fuel quantity is added; the weight of the combustion chamber is not increased by the transformation, and the thrust-weight ratio is greatly increased; meanwhile, the modification does not add any difficulty to the manufacturing and the materials.

2. The above modifications can be adapted to military and civil aircraft engines.

3. The above modifications may also be applied to gas turbines powered by ships and civil ships and certain generator sets. (if not appropriate for a gas turbine, when the applicant does not claim such a claim)

4. The transformation can realize that the direction of fuel oil injected into the combustion chamber is mutually vertical to the direction of compressed air injected into the combustion chamber, and the circle arrangement of a plurality of fuel injection nozzles is favorable for the complete combustion of the fuel oil of the aircraft engine, and the technology can adapt to the aircraft engine; it can also be adapted to gas turbines. (if not applicable to gas turbines, the claims of this sentence are not submitted by the applicant).