An aeroengine with a knock-on afterburner

Abstract

The application provides an aero-engine with knock-on afterburning chamber, comprising: a compressor, a combustion chamber, a turbine and a nozzle, the compressor and the turbine are coaxially connected, compressed air from the outlet of the compressor enters the combustion chamber, combustion gas from the outlet of the combustion chamber enters the turbine to drive the turbine to work, and the outlet gas after the turbine works enters the nozzle; the knock-on afterburning chamber is arranged in the nozzle, the knock-on afterburning chamber comprises an afterburning outer ring and an afterburning inner ring, the afterburning inner ring is coaxially sleeved in the inner cavity of the afterburning outer ring, and the annular wall of the afterburning outer ring and the annular wall of the afterburning inner ring form a knock-on ring cavity for continuous rotating knock-on combustion; the outlet gas entering the nozzle enters the knock-on ring cavity from the air inlet end of the knock-on afterburning chamber, mixes with fuel in the knock-on ring cavity and performs continuous rotating knock-on combustion. The application can effectively improve the combustion efficiency of the aero-engine and reduce fuel consumption.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of aero-engines, in particular to an aero-engine with a knock-on afterburner. BACKGROUND

[0002] Jet engine: a jet engine is an engine that produces thrust by expelling a jet of gas at high speed.

[0003] Afterburner: an afterburner is a device in a jet engine that injects fuel into the hot exhaust gases from the turbine to increase the temperature and thus the thrust of the engine. Afterburners are commonly used in the engines of supersonic aircraft, such as fighter jets.

[0004] Due to the low pressure and high flow rate of the gas flow in the afterburner, the ignited mixture needs to complete the combustion process in a long cylinder. In the current afterburner, only 85% to 90% of the heat content of the fuel can be converted into useful heat energy, and the remaining part is either discharged from the engine due to the failure of the fuel droplets to burn, or is lost through heat dissipation in the cylinder. Therefore, improving the afterburner efficiency is of great significance to reducing fuel consumption. In addition, the current afterburner has the problem of increasing the size and weight of the engine due to the large axial length of the combustion zone in the structural design. SUMMARY

[0005] Therefore, the present application provides an aero-engine with a knock-on afterburner to improve the combustion efficiency of the afterburner of the aero-engine and reduce fuel consumption.

[0006] The present application provides the following technical solution: an aero-engine with a knock-on afterburner, comprising: a compressor, a combustion chamber, a turbine and an exhaust nozzle, the compressor and the turbine are coaxially connected, the compressed air at the outlet of the compressor is introduced into the combustion chamber for mixing with fuel and then combusting, the combustion flue gas at the outlet of the combustion chamber is introduced into the turbine for driving the turbine to work, and the flue gas at the outlet after the turbine works is introduced into the exhaust nozzle.

[0007] The knock-on afterburner is arranged in the exhaust nozzle, the knock-on afterburner comprises a supplementary combustion outer ring and a supplementary combustion inner ring, the supplementary combustion inner ring is coaxially sleeved in the inner cavity of the supplementary combustion outer ring, and the annular wall of the supplementary combustion outer ring and the annular wall of the supplementary combustion inner ring form a knock-on annular cavity for continuous knock-on combustion; the flue gas at the outlet introduced into the exhaust nozzle enters the knock-on annular cavity from the gas inlet end of the knock-on afterburner, mixes with the fuel in the knock-on annular cavity and performs continuous knock-on combustion.

[0008] According to an embodiment of the present application, the afterburning outer ring comprises a converging section and a diverging section in the direction of the flue gas flow, and the joint between the converging section and the diverging section is a stepped abrupt expansion transition, forming a stepped expansion channel, and an oil supply ring for supplying fuel to the detonation ring cavity is arranged inside the stepped expansion channel, and the oil supply ring is connected with an external oil supply pipeline.

[0009] According to an embodiment of the present application, an atomizing nozzle is connected to the oil supply ring, and the spray area of the atomizing nozzle corresponds to the detonation ring cavity of the diverging section, so as to form a rotating detonation combustion zone at the corresponding position of the diverging section in the detonation ring cavity.

[0010] According to an embodiment of the present application, the number of atomizing nozzles is multiple, and the atomizing nozzles are circumferentially distributed on the inner wall surface of the stepped expansion channel.

[0011] According to an embodiment of the present application, the rotating detonation ramjet combustor further comprises an isolation ring, the isolation ring is coaxially sleeved between the afterburning outer ring and the afterburning inner ring, and the position of the isolation ring corresponds to the position of the converging section of the afterburning outer ring.

[0012] According to an embodiment of the present application, the afterburning inner ring comprises a curved section and a straight section in the direction of the flue gas flow, the position of the curved section corresponds to the position of the converging section of the afterburning outer ring, and the position of the straight section corresponds to the position of the diverging section of the afterburning outer ring; wherein, in the direction of the flue gas flow, the intake end of the curved section is curved away from the afterburning outer ring.

[0013] According to an embodiment of the present application, the relative distance between the isolation ring and the converging section of the afterburning outer ring is greater than the minimum relative distance between the isolation ring and the curved section of the afterburning inner ring.

[0014] According to an embodiment of the present application, a plurality of air inlet holes are arranged on the ring wall of the diverging section of the afterburning outer ring, and the air inlet holes are circumferentially distributed on the ring wall of the afterburning outer ring.

[0015] According to an embodiment of the present application, the air inlet holes are obliquely arranged on the ring wall of the afterburning outer ring, and the oblique direction of the air inlet holes is consistent with the direction of the flue gas flow.

[0016] According to an embodiment of the present application, the number of rotating detonation ramjet combustors is multiple, and the rotating detonation ramjet combustors are circumferentially distributed in the tail nozzle.

[0017] Compared with the prior art, the at least one technical scheme adopted by the embodiment of the present application can achieve the beneficial effects at least including: the embodiment of the present application sets the detonation afterburning chamber in the turbine nozzle, greatly increases the airflow temperature through detonation combustion, and increases the engine thrust in a short time, so as to achieve the purposes of improving the fuel combustion efficiency and reducing the fuel consumption. BRIEF DESCRIPTION OF DRAWINGS

[0018] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the drawings needed to be used in the embodiments will be briefly introduced as follows. Obviously, the drawings in the following description are only some embodiments of the present application, and all other drawings obtained by those skilled in the art without creative labor on the premise of not deviating from the spirit of the present application.

[0019] Figure 1 is a schematic diagram of the overall structure of the aero-engine of the embodiment of the present application;

[0020] Figure 2 is a schematic diagram of the local structure of the detonation afterburning chamber of the embodiment of the present application;

[0021] Figure 3 is a schematic diagram of the detonation afterburning chamber layout of the first embodiment of the present application;

[0022] Figure 4 is a schematic diagram of the detonation afterburning chamber layout of the second embodiment of the present application;

[0023] In the figure, 1 is a compressor, 2 is a combustion chamber, 3 is a turbine, 4 is a nozzle, 5 is an afterburning outer ring, 6 is an afterburning inner ring, 7 is a separation ring, 8 is an oil supply ring, 9 is an atomizing nozzle, 10 is an air inlet hole, 11 is a continuous rotating detonation combustion zone, 12 is an outer ring smoke isolation layer, and 13 is an inner ring smoke isolation layer. DETAILED DESCRIPTION

[0024] The embodiments of the present application will be described in detail below with reference to the drawings.

[0025] The embodiments of the present application are described below through specific examples, and those skilled in the art can easily understand other advantages and effects of the present application from the disclosure. Obviously, the described embodiments are only some of the embodiments of the present application, not all. The present application can also be implemented or applied by other different specific embodiments, and the details in the specification can be modified or changed based on different views and applications without departing from the spirit of the present application. It should be noted that the following embodiments and features in the embodiments can be combined with each other without conflict. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative labor are within the scope of protection of the present application.

[0026] As Figures 1-2 shown, the embodiment of the application provides an aero-engine with a detonation afterburner 2, comprising: a compressor 1, a combustion chamber 2, a turbine 3 and a tail nozzle 4, the compressor 1 and the turbine 3 are coaxially connected, the compressed air at the outlet of the compressor 1 enters the combustion chamber 2 for combustion after mixing with fuel, the combustion flue gas at the outlet of the combustion chamber 2 enters the turbine 3 for driving the turbine 3 to work, and the outlet flue gas after the turbine 3 works enters the tail nozzle 4; a detonation afterburner is arranged in the tail nozzle 4, the detonation afterburner comprises a supplementary combustion outer ring 5 and a supplementary combustion inner ring 6, the supplementary combustion inner ring 6 is coaxially sleeved in the inner cavity of the supplementary combustion outer ring 5, and both are tubular structures penetrating front and back; the annular wall of the supplementary combustion outer ring 5 and the annular wall of the supplementary combustion inner ring 6 form a detonation annular cavity for continuous rotating detonation combustion; the outlet flue gas entering the tail nozzle 4 enters the detonation annular cavity from the air inlet end of the detonation afterburner, mixes with fuel in the detonation annular cavity and performs continuous rotating detonation combustion.

[0027] The embodiment of the application sets the detonation afterburner in the turbine tail nozzle 4, greatly increases the airflow temperature through rotating detonation combustion, increases the engine thrust in a short time, improves the fuel combustion efficiency, and reduces the purpose of fuel consumption. The problem of low combustion efficiency is solved. At present, in the afterburner, only 85% to 90% of the heat content of the fuel can be converted into useful heat energy, the rest is discharged from the engine due to the failure of the fuel droplets to burn, or is lost through the cylinder heat dissipation.

[0028] Rotating detonation combustion is a combustion technology that uses rotating detonation waves to achieve combustion. Detonation waves are shock waves that can propagate through a detonation reaction, allowing reactants and oxygen to react quickly after contact. Rotating detonation combustion technology can be used to improve combustion efficiency, reduce pollution, and operate at high pressures and temperatures. Rotating detonation combustion is a relatively new technology that is currently being used in the automotive, aerospace, and aerospace industries, primarily to improve combustion efficiency and reduce pollution.

[0029] Rotating detonation engines are engines that use detonation to organize combustion. The combustion is organized in a circumferential direction, so that the length of the rotating detonation engine combustion chamber is lower than that of traditional engines, which can greatly reduce the volume and weight of the rotating detonation engine.

[0030] The working principle of the rotary detonation engine is as follows: firstly, air enters the detonation ring cavity from the annular channel, fuel is injected by the nozzle, and the two are uniformly mixed after a short distance and then enter the detonation ring cavity together. Then, the initial detonation wave is formed in the detonation ring cavity by the igniter. Secondly, the initial detonation wave propagates along the circumferential direction of the annular detonation ring cavity. Since the fuel and air flow in the axial direction and the combustion products also flow out in the axial direction, when the detonation wave propagates one circle back to the initial position, the original combustion products have flowed away, and the fresh fuel and air mixture has flowed in, so the detonation wave can continue to maintain. In this way, the detonation wave continues to rotate and propagate in the annular detonation ring cavity, and the high-temperature and high-pressure combustion products are discharged at high speed through the tail nozzle, thereby generating thrust. Since the frequency of rotary detonation combustion is as high as several kilohertz, the rotary detonation engine can generate stable thrust.

[0031] In an embodiment of the present application, in the direction of the flue gas flow, the afterburning outer ring 5 comprises a contraction section and an expansion section, and the junction between the contraction section and the expansion section is a stepped abrupt transition, forming a stepped expansion channel. The structure of the contraction section and the expansion section can stabilize the flame and enhance combustion. The stepped expansion channel is internally provided with a fuel supply ring 8 for supplying fuel to the detonation ring cavity. The fuel supply ring 8 is connected with an external fuel supply pipeline. In this embodiment, the contraction section serves as the gas inlet end of the afterburning outer ring 5, and the outlet flue gas of the turbine enters the detonation ring cavity from the contraction section. The fuel supply ring 8 is arranged inside the stepped expansion channel and is used to introduce fuel from the external fuel supply pipeline into the detonation ring cavity. The fuel mixes with the outlet flue gas, and after ignition by the igniter, continuous rotary detonation combustion occurs. In order to achieve detonation and produce detonation effect, fuel burns rapidly and releases energy, so the size of the annular surface of the afterburning chamber needs to be designed to be relatively small. Due to the extremely short detonation reaction time, the axial length of the combustion channel is small, and compared with the traditional afterburning chamber, the size and weight of the engine are reduced.

[0032] In a preferred scheme of the present embodiment, the fuel supply ring 8 is connected with an atomizing nozzle 9, which atomizes the fuel. The spray area of the atomizing nozzle 9 corresponds to the detonation ring cavity of the expansion section, so as to form a continuous rotary detonation combustion zone 11 at the corresponding position of the expansion section in the detonation ring cavity. The number of atomizing nozzles 9 is multiple, and they are circumferentially distributed on the inner wall surface of the stepped expansion channel to ensure the uniformity and stability of fuel injection.

[0033] In an embodiment of the present application, the detonation ramjet combustor further comprises a separation ring 7, which is coaxially sleeved between the outer annular combustor 5 and the inner annular combustor 6, and the position of the separation ring 7 corresponds to the position of the converging section of the outer annular combustor 5. In this embodiment, the separation ring 7 separates the outer annular combustor 5 and the inner annular combustor 6 into two gas inlet channels. The flue gas entering the detonation ring cavity through the gas inlet channel between the separation ring 7 and the outer annular combustor 5 flows to the continuous rotating detonation combustion zone 11 to participate in the rotating detonation combustion. The flue gas entering the detonation ring cavity through the gas inlet channel between the separation ring 7 and the inner annular combustor 6 flows along the inner wall of the inner annular combustor 6 to form an inner annular flue gas separation layer 13 on the inner wall of the inner annular combustor 6. The inner annular flue gas separation layer 13 adheres to the inner wall of the inner annular combustor 6 as a cooling gas film, forming a good cooling and separation gas flow, which has a good cooling and temperature reduction effect on the detonation ramjet combustor.

[0034] In a preferred scheme of the present embodiment, in the direction of flue gas flow, the inner annular combustor 6 comprises a curved section and a straight section. The position of the curved section corresponds to the position of the converging section of the outer annular combustor 5, and the position of the straight section corresponds to the position of the diverging section of the outer annular combustor 5. In the direction of flue gas flow, the gas inlet end of the curved section is curved away from the outer annular combustor 5. The structure design of the inner annular combustor 6 forms a converging channel between the inner annular combustor 6 and the separation ring 7, accelerates the flow of flue gas, and has a flow guiding effect, which can suck in more flue gas and better form a separation gas flow.

[0035] Since the flue gas participating in detonation combustion and the flue gas forming a separation gas flow require different flow rates and speeds, the present embodiment uses the separation ring 7 to form two separate channels, and through the geometric size design of the channels, different flow rates and speeds are achieved. In specific implementation, the flue gas between the separation ring 7 and the outer annular combustor 5 needs a large flow rate, and the flue gas between the separation ring 7 and the inner annular combustor 6 needs a high flow speed. Therefore, in the present embodiment, it is further preferred that the relative distance between the separation ring 7 and the converging section of the outer annular combustor 5 is greater than the minimum relative distance between the separation ring 7 and the curved section of the inner annular combustor 6, so that more flue gas enters the gas inlet channel between the separation ring 7 and the outer annular combustor 5 and reaches the continuous rotating detonation combustion zone 11 to participate in combustion, and the flue gas entering the gas inlet channel between the separation ring 7 and the inner annular combustor 6 passes through the converging section at the minimum relative distance to increase the flow speed, so as to form a separation gas flow.

[0036] In one embodiment of the present application, a plurality of air inlet holes 10 are formed on the ring wall of the expansion section of the afterburning outer ring 5, and the air inlet holes 10 are circumferentially distributed on the ring wall of the afterburning outer ring 5. The flue gas outside the detonation afterburning chamber enters the detonation ring cavity through the air inlet holes 10 on the ring wall of the afterburning outer ring 5, flows along the inner wall of the afterburning outer ring 5, and forms an outer ring flue gas isolation layer 12 on the inner wall of the afterburning outer ring 5. The outer ring flue gas isolation layer 12 adheres to the inner wall of the afterburning outer ring 5 as a cooling gas film, forms a good cooling isolation gas flow, and has a good cooling and temperature reduction effect on the detonation afterburning chamber.

[0037] The detonation afterburning chamber of the present embodiment is configured to allow a part of the flue gas to enter the detonation ring cavity to participate in detonation combustion, improve the combustion efficiency, and increase the engine thrust. A part of the flue gas enters the detonation ring cavity to form flue gas isolation layers on the inner walls of the afterburning outer ring 5 and the afterburning inner ring 6, and cool the inner walls of the afterburning outer ring 5 and the afterburning inner ring 6. The remaining flue gas surrounds the outer wall of the detonation afterburning chamber to cool the outer walls of the afterburning outer ring 5 and the afterburning inner ring 6, thereby ensuring the cooling effect of the detonation afterburning chamber.

[0038] In a preferred embodiment of the present embodiment, the air inlet holes 10 are obliquely formed on the ring wall of the afterburning outer ring 5, and the oblique direction of the air inlet holes 10 is consistent with the flue gas flow direction. The obliquely arranged air inlet holes 10 are consistent with the flue gas flow direction, which is conducive to the flue gas entering the detonation ring cavity through the air inlet holes 10 and flowing along the inner wall of the afterburning outer ring 5 under the guidance of the air inlet holes 10, and better forming a flue gas isolation layer.

[0039] In one embodiment of the present application, as shown in Figure 3 for a smaller size nozzle 4, the afterburning outer ring 5 and the afterburning inner ring 6 can be a whole ring structure, and the number of detonation afterburning chambers is one. In another embodiment of the present application, as shown in Figure 4 for a larger size nozzle 4, the afterburning outer ring 5 and the afterburning inner ring 6 can be a plurality of ring structures, and a plurality of detonation afterburning chambers are circumferentially distributed in the nozzle 4.

[0040] The aeroengine of the present embodiment is provided with a detonation afterburning chamber in the turbine nozzle, which can effectively improve the fuel combustion efficiency, greatly increase the airflow temperature in a short time, and increase the engine thrust. Through the geometric structure design of the detonation afterburning chamber, the cooling effect of the detonation ring cavity is effectively ensured, and the axial size of the afterburning chamber is also shortened, thereby reducing the size and weight of the engine.

[0041] The above merely provides the specific implementation of the present application, but the protection scope of the present application is not limited to this. Any changes or replacements within the technical scope disclosed by the present application can be easily conceived by the person skilled in the art, and should be covered within the protection scope of the present application. Therefore, the protection scope of the present application should be subject to the protection scope of the claims.

Claims

1. An aircraft engine with a knock augmented combustor, characterized by, The application relates to a turbofan engine. The turbofan engine comprises a compressor, a combustion chamber, a turbine and a nozzle, the compressor and the turbine are coaxially connected, compressed air at the outlet of the compressor is introduced into the combustion chamber for combustion after mixing with fuel, combustion flue gas at the outlet of the combustion chamber is introduced into the turbine for driving the turbine to work, and the flue gas after the turbine works is introduced into the nozzle. The nozzle is provided with a detonation afterburning chamber, the detonation afterburning chamber comprises an afterburning outer ring and an afterburning inner ring, the afterburning inner ring is coaxially sleeved in the inner cavity of the afterburning outer ring, and a detonation ring cavity for continuous rotating detonation combustion is formed between the ring wall of the afterburning outer ring and the ring wall of the afterburning inner ring; the flue gas at the outlet is introduced into the detonation ring cavity from the air inlet end of the detonation afterburning chamber, mixed with fuel in the detonation ring cavity and subjected to continuous rotating detonation combustion. In the flue gas flow direction, the afterburning outer ring comprises a contraction section and an expansion section, the joint between the contraction section and the expansion section is a stepped sudden expansion transition, a stepped expansion channel is formed, an oil supply ring for supplying fuel to the detonation ring cavity is arranged in the stepped expansion channel, and the oil supply ring is connected with an external oil supply pipeline. The detonation afterburning chamber further comprises an isolation ring, the isolation ring is coaxially sleeved between the afterburning outer ring and the afterburning inner ring, and the setting position of the isolation ring corresponds to the position of the contraction section of the afterburning outer ring. In the flue gas flow direction, the afterburning inner ring comprises a curved section and a straight section, the position of the curved section corresponds to the position of the contraction section of the afterburning outer ring, and the position of the straight section corresponds to the position of the expansion section of the afterburning outer ring; wherein in the flue gas flow direction, the air inlet end of the curved section is bent away from the afterburning outer ring. The relative distance between the isolation ring and the contraction section of the afterburning outer ring is greater than the minimum relative distance between the isolation ring and the curved section of the afterburning inner ring. The air inlet hole is obliquely arranged on the ring wall of the afterburning outer ring, and the oblique direction of the air inlet hole is consistent with the flue gas flow direction.

2. The aircraft engine with a knock-on ram combustor of claim 1, wherein, The oil supply ring is connected with an atomizing nozzle, the spraying area of the atomizing nozzle is the detonation ring cavity corresponding to the expansion section, so that a rotating detonation combustion zone is formed in the position corresponding to the expansion section in the detonation ring cavity.

3. The aircraft engine with a knock-on ram combustor of claim 2, wherein, The number of the atomizing nozzles is multiple, and the atomizing nozzles are circumferentially distributed on the inner wall surface of the stepped expansion channel.

4. The aircraft engine with a knock-on ram combustor of claim 1, wherein, A plurality of air inlet holes are arranged on the ring wall of the expansion section of the afterburning outer ring, and the air inlet holes are circumferentially distributed on the ring wall of the afterburning outer ring.

5. The aircraft engine with a knock-on ram combustor of claim 1, wherein, The number of the detonation afterburning chambers is multiple, and the detonation afterburning chambers are circumferentially distributed in the nozzle.

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