CN115163332A - Bypass ratio variable gas turbine engine utilizing rotary knocking - Google Patents
Bypass ratio variable gas turbine engine utilizing rotary knocking Download PDFInfo
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- CN115163332A CN115163332A CN202210875265.0A CN202210875265A CN115163332A CN 115163332 A CN115163332 A CN 115163332A CN 202210875265 A CN202210875265 A CN 202210875265A CN 115163332 A CN115163332 A CN 115163332A
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
- F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
- F02K3/075—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type controlling flow ratio between flows
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
- F23R3/286—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply having fuel-air premixing devices
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
- F23R3/52—Toroidal combustion chambers
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- Combustion & Propulsion (AREA)
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Abstract
The invention discloses a variable bypass ratio gas turbine engine utilizing rotary detonation, which comprises a rotary detonation combustion chamber, a core engine, an inner bypass, an outer bypass and a support plate. The high-speed flow of the rotary detonation combustion along the axial direction is used for ejecting external gas into the outer duct, and the external gas and the high-speed gas behind the turbine are separated or mixed and discharged. The invention realizes the improvement of the air flow of the outer duct by a smaller windward area, thereby increasing the exhaust flow and realizing the thrust increasing effect. The gas after the gas compressor is divided, so that part of kinetic energy of high-speed exhaust after the turbine is used for realizing the injection effect of rotary detonation combustion, the redundant energy loss is avoided, and the propulsion efficiency is effectively improved. In addition, the rotary detonation combustion intensity is changed, the injection airflow flow can be adjusted, the bypass ratio is changed under the fixed geometric configuration, and the flight envelope curve is widened. The invention effectively avoids the complex structure and the processing difficulty of the turbofan engine and improves the working performance of the aero-engine. The invention can be used in the technical field of detonation combustion and aviation propulsion.
Description
Technical Field
The invention belongs to the technical field of detonation combustion and aviation propulsion, and particularly relates to a variable bypass ratio gas turbine engine utilizing rotary detonation.
Background
The rotary detonation combustion technology is one of research hotspots in the field of aerospace propulsion. Rotary detonation combustion is typically organized in an annular combustion chamber, with one or more detonation waves formed in the head of the combustion chamber and propagating circumferentially, with the products of combustion being expelled at high velocity from the open end. The combined power cycle is combined with the traditional jet propulsion technology to form a new combined power cycle, which is beneficial to improving the propulsion performance and promoting the application of the rotary detonation combustion technology.
When the jet engine flies at subsonic speed, the exhaust speed of the jet engine is too high, and the propelling efficiency is low. In addition, when the thermal efficiency is improved (the turbine front gas temperature is increased), the exhaust speed is increased, and the propulsion efficiency is further decreased. Therefore, to improve aircraft engine performance, it is desirable to reduce the average exhaust velocity of the aircraft engine while increasing the turbine front gas temperature.
In view of the above problems, turbofan engines are the existing solutions. In order to improve the defects of a turbojet engine, in the turbofan engine, high-temperature and high-pressure gas in an inner duct converts part of kinetic energy into shaft work to drive a fan by impacting a low-pressure turbine, the flow is respectively led to an outer duct and a gas compressor (the inner duct), the gas flow flowing through the outer duct is compressed and expanded and is discharged downstream at a lower speed, the average exhaust speed of the aero-engine is reduced, the exhaust kinetic energy is reduced, the exhaust flow is increased, the thrust is increased, and the propulsion efficiency is improved. Therefore, by adjusting the design of the low-pressure turbine and the fan, the propulsion efficiency can be improved while the thermal efficiency is improved, and the performance of the aircraft engine is improved. However, the large-diameter fan of the turbofan engine has high design requirement and high processing difficulty, increases the windward area of the engine, has large flow resistance, and is not beneficial to high-speed flight; the design of the low-pressure compressor and the low-pressure turbine changes a single shaft of the aircraft engine into multiple shafts, and the structural complexity and the design and processing difficulty are increased. The turbofan engine technology is high in research and development cost and only mastered by a few countries.
Therefore, it is particularly critical to design a novel propulsion device with simple structure, high propulsion efficiency and wide flight bag line. The invention provides a variable bypass ratio gas turbine engine utilizing rotary detonation, which can just meet the requirements and has important value on the practical application of a rotary detonation combustion mode.
Disclosure of Invention
Technical problem to be solved
Aiming at the problems of low propelling efficiency, complex structure, high design difficulty and large flight resistance of a turbofan engine, the invention provides a variable bypass ratio gas turbine engine utilizing rotary detonation. The external air flow is injected into the outer duct and then separated from or mixed with the high-speed air flow behind the turbine to be discharged. Through the injection of the rotary detonation combustion, the invention realizes the improvement of the air flow of the outer duct with smaller windward area, increases the exhaust flow and realizes the thrust augmentation effect. The split flow behind the air compressor enables a part of kinetic energy of high-speed exhaust behind the turbine to be used for injection of rotary detonation combustion, avoids redundant energy loss, and improves the heat efficiency and the propulsion efficiency. In addition, the adjustment of the injection airflow flow can be realized by changing the combustion intensity of the rotary detonation, so that the bypass ratio is changed under a fixed geometric configuration, and the flight envelope curve is widened.
In order to achieve the purpose, the invention adopts the technical scheme that:
a variable bypass ratio gas turbine engine utilizing rotational detonation comprises a rotational detonation combustion chamber, a core engine, an inner bypass, an outer bypass and a support plate.
The rotary detonation combustion chamber is arranged outside the main combustion chamber of the turbojet engine, and a part of high-pressure gas behind the gas compressor is used as an oxidant for rotary detonation combustion; meanwhile, high-speed airflow generated along the axial direction by the rotary detonation combustion is used for ejecting, and air outside the casing (1) is ejected into the outer duct.
The core machine consists of a compressor, a main combustion chamber and a turbine. After entering the core machine, airflow is compressed into high-pressure gas through the air compressor in a multi-stage mode, then the high-pressure gas participates in combustion in the main combustion chamber, high-temperature and high-pressure combustion products are formed and then discharged to the downstream, the high-temperature and high-pressure combustion products impact turbine blades and provide driving energy for rotation of the air compressor, and finally the combustion products are expanded in the tail nozzle and discharged at a high speed to generate thrust.
The inner duct is an airflow channel formed by the core machine, the casing (1) and the casing (2).
The rotary detonation combustor consists of an injection device, an annular flow passage and an ignition device. The injector comprises an oxidant annular gap and a fuel injector. A part of high-pressure gas formed by compression of the compressor enters the rotary detonation combustor through the annular seam, meanwhile, the fuel injector sprays fuel to be mixed with an oxidant, a backflow area is easy to form during mixing due to the offset design of the injector, and the mixing effect is improved. The casing (1) and the casing (2) form an annular flow passage of the rotary detonation combustor. After the ignition device is started, a circumferentially rotating detonation wave is formed in the annular flow channel of the rotary detonation combustion chamber, and combustion products are discharged at a high speed through a contraction and expansion channel on the rear side of the rotary detonation combustion chamber.
The outer culvert consists of outer culvert wall surfaces, a casing (1) and a casing (2); the outer culvert wall surface and the casing (1) form an outer culvert inlet, and the outer culvert wall surface is fixedly connected with the casing (1) through a plurality of support plates distributed along the circumferential direction. When the rotary detonation combustion is carried out, the combustion products discharged at high speed have an injection effect on external gas. Near the jet nozzle, the low-speed gas flow of the outer duct is separated from or mixed with the high-speed gas flow after the turbine of the inner duct (the separated discharge is taken as an example in the figure). The injection effect of the rotary detonation combustion effectively improves the flow rate of air flowing into the outer duct, replaces the internal complex fan rotor structure of the existing turbofan engine, reduces the wind resistance, simplifies the design, improves the exhaust flow rate of the aero-engine, and realizes the thrust augmentation effect. Through the intensity of adjustment rotatory detonation burning, influence the high-speed exhaust of rotatory detonation combustion chamber open end and draw the penetrating effect to external gas, and then change outer duct exhalant flow, realize the function of becoming the duct ratio under fixed geometric configuration, widen the flight envelope curve.
Has the advantages that:
the variable bypass ratio gas turbine engine utilizing the rotary detonation is characterized in that a rotary detonation combustion chamber is arranged outside a main combustion chamber of a turbine jet engine, high-speed airflow generated along the axial direction by the rotary detonation combustion is used for ejecting, and the ejected airflow is separated from or mixed with the high-speed airflow after the turbine through an outer bypass and is discharged. The injection effect of the rotary detonation combustion realizes the improvement of the air flow of the outer duct by a small windward area, so that the exhaust flow is increased, and the thrust increasing effect is realized. Compared with a turbojet engine, the rotary detonation combustion chamber is introduced, so that the kinetic energy of exhaust after a turbine is reduced for the purpose of injection, the redundant energy loss caused by high-speed exhaust is avoided, the propelling efficiency of the aero-engine is improved while the thermal efficiency is improved, and the high-difficulty design and processing requirements of a large-diameter fan and a multi-shaft structure of the turbofan engine are effectively avoided. In addition, the adjustment of the flow of the injection airflow can be realized by changing the combustion intensity of the rotary detonation, so that the bypass ratio is changed under a fixed geometric configuration, the flight envelope curve is widened, and the adaptability of the propelling device under a complex environment is improved. The invention can be used in the technical field of detonation combustion and aviation propulsion.
Drawings
FIG. 1 is an axial half sectional view of a variable bypass ratio gas turbine engine utilizing rotational detonation in accordance with the present invention.
The gas turbine engine comprises a gas compressor 1, a circular seam 2, an annular flow channel 3, a main combustion chamber 4, a casing (2) 5, a turbine 6, a culvert wall 7, a support plate 8, an ignition device 9, an injector 10 and a casing (1) 11.
Detailed Description
The invention will now be further described with reference to the accompanying drawings in which:
referring to fig. 1, the invention relates to a variable bypass ratio gas turbine engine utilizing rotational detonation, which comprises a rotational detonation combustor, a core engine, an inner bypass, an outer bypass and a support plate 8. The method is characterized in that a rotary detonation combustion chamber is arranged outside a main combustion chamber 4 of the turbojet engine, and the open end of an annular flow channel 3 of the rotary detonation combustion chamber is connected with an outer duct. The outer duct consists of an outer duct wall surface 7, a casing (1) and a casing (2) and 5, and the jet air flow is discharged downstream through the outer duct. The inner duct is the air flow channel formed by the core engine and the casing (1, 11, 5) and is similar to the turbojet engine.
When the turbojet engine works, incoming flow is subjected to multistage pressurization through the air compressor 1 to become high-pressure gas, and the high-pressure gas enters the main combustion chamber 4 to participate in combustion for the turbojet engine; however, in the invention, a part of high-pressure gas is introduced into the rotary detonation combustor, and the shunting of the high-pressure gas influences the combustion organization of the main combustor 4, thereby reducing the gas exhaust speed and reducing the energy loss of high-speed exhaust after the turbine 6; meanwhile, the energy reduced by high-speed exhaust is used for rotary detonation combustion, high-pressure gas is used as an oxidant to enter the annular seam 2, fuel is sprayed in through the injector 10 and is mixed with the oxidant, a backflow area is easily formed during mixing due to the offset design of the injector 10, and the mixing effect is effectively improved. After the ignition device 9 is started, a circumferentially rotating detonation wave is formed in the annular flow passage 3 of the rotary detonation combustor, and combustion products are discharged at a high speed through a contraction and expansion channel at the rear side of the rotary detonation combustor. The high-speed flow can be used for ejecting external gas, effectively improves the flow rate of the airflow entering the outer duct, further improves the exhaust flow rate of the aircraft engine, generates thrust gain and improves the propelling efficiency of the aircraft engine.
The invention effectively improves the propelling efficiency of the turbojet engine by utilizing the injection effect of the rotary detonation combustion; the defects of complex structure, large windward area and the like of the turbofan engine are overcome; in addition, the adjustment of the injection airflow flow can be realized by changing the combustion intensity of the rotary detonation, so that the bypass ratio is changed under a fixed geometric configuration, the flight envelope is widened, and the adaptability of the propelling device under a complex environment is improved.
The invention has important value for the practical application of the rotary detonation combustion mode. Various modifications and optimizations may be made to the above-described method by those skilled in the art without departing from the principles of the invention.
Claims (6)
1. A variable bypass ratio gas turbine engine utilizing rotational detonation comprises a rotational detonation combustor, a core engine, an inner bypass, an outer bypass and a support plate. The method is characterized in that a rotary detonation combustion chamber is arranged on the outer side of a main combustion chamber of the turbojet engine, and a part of high-pressure gas at the outlet of a compressor is used as an oxidant for rotary detonation combustion. The shunting of the high-pressure gas influences the combustion organization of the main combustion chamber, so that the gas discharge speed is reduced, and the energy loss of high-speed exhaust after a turbine is reduced; meanwhile, the energy reduced by high-speed exhaust is used for rotary detonation combustion, high-speed airflow generated along the axial direction is used for ejecting, external airflow is introduced into the outer duct, and the ejected airflow and the high-speed airflow behind the turbine are separated or mixed and discharged. The injection effect is beneficial to improving the airflow of the outer duct, the exhaust flow of the aircraft engine is increased, and the thrust increasing effect is realized; the propulsion efficiency is improved while ensuring the improvement of the thermal efficiency. In addition, the adjustment of the injection airflow flow can be realized by changing the combustion intensity of the rotary detonation, so that the bypass ratio is changed under a fixed geometric configuration, the flight envelope is widened, and the adaptability of the propelling device under a complex environment is improved.
2. The variable bypass ratio gas turbine engine with rotational detonation of claim 1, wherein: the core machine consists of a gas compressor, a main combustion chamber and a turbine. After entering the core machine, the incoming flow is compressed into high-pressure gas through the gas compressor in a multi-stage mode, then the high-pressure gas participates in combustion in the main combustion chamber, high-temperature and high-pressure combustion products are discharged downstream to impact turbine blades to provide driving energy for the rotation of the gas compressor, and finally the combustion products are expanded in the tail nozzle and discharged at a high speed to generate thrust.
3. The variable bypass ratio gas turbine engine with rotational detonation of claim 1, wherein: the inner duct is a flow channel formed by the core machine, the casing (1) and the casing (2).
4. The variable bypass ratio gas turbine engine with rotational detonation of claim 1, wherein: the rotary detonation combustor consists of an injection device, an annular flow passage and an ignition device. The injection device comprises an oxidant circumferential weld and a fuel injector, a part of high-pressure gas formed by compression of the compressor enters the rotary detonation combustor through the circumferential weld, meanwhile, the fuel injector sprays fuel to be mixed with the oxidant, a backflow area is easily formed during mixing due to the offset design of the injector, and the mixing effect is improved. The casing (1) and the casing (2) form an annular flow passage of the rotary detonation combustor. After the ignition device is started, detonation waves rotating along the circumferential direction are formed in the annular flow channel of the rotary detonation combustion chamber, and combustion products are discharged at a high speed through a contraction and expansion channel on the rear side of the rotary detonation combustion chamber.
5. The variable bypass ratio gas turbine engine with rotational detonation of claim 1, wherein: the outer culvert consists of outer culvert wall surfaces, a casing (1) and a casing (2); the outer culvert wall surface and the casing (1) form an outer culvert inlet, and the outer culvert wall surface is fixedly connected with the casing (1) through a plurality of support plates distributed along the circumferential direction. When the rotary detonation combustion is carried out, the combustion products discharged at high speed have an injection effect on external gas. Near the jet nozzle, the low-speed gas flow of the outer duct is separated from or mixed with the high-speed gas flow after the turbine of the inner duct (the separated discharge is taken as an example in the figure). The injection effect of the rotary detonation combustion can improve the airflow flow entering the outer duct, replace the internal complex fan rotor structure of the existing turbofan engine, reduce the wind resistance and simplify the design, improve the exhaust flow of the aero-engine, and realize the thrust augmentation effect.
6. The variable bypass ratio gas turbine engine with rotational detonation of claim 5, wherein: through adjusting rotatory detonation combustion intensity, influence the high-speed exhaust of rotatory detonation combustion chamber open end and penetrate the effect to the injection of external gas, and then change the air flow of flowing through outer duct, realize the function of becoming the duct ratio under the fixed geometric configuration, widen the flight envelope curve, improve the adaptability under the advancing device complex environment.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210875265.0A CN115163332A (en) | 2022-07-25 | 2022-07-25 | Bypass ratio variable gas turbine engine utilizing rotary knocking |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210875265.0A CN115163332A (en) | 2022-07-25 | 2022-07-25 | Bypass ratio variable gas turbine engine utilizing rotary knocking |
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| Publication Number | Publication Date |
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| CN115163332A true CN115163332A (en) | 2022-10-11 |
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| CN202210875265.0A Pending CN115163332A (en) | 2022-07-25 | 2022-07-25 | Bypass ratio variable gas turbine engine utilizing rotary knocking |
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- 2022-07-25 CN CN202210875265.0A patent/CN115163332A/en active Pending
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