CN113323769A - Variable cycle engine configuration based on multi-duct intake interstage combustion chamber - Google Patents
Variable cycle engine configuration based on multi-duct intake interstage combustion chamber Download PDFInfo
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- CN113323769A CN113323769A CN202110632005.6A CN202110632005A CN113323769A CN 113323769 A CN113323769 A CN 113323769A CN 202110632005 A CN202110632005 A CN 202110632005A CN 113323769 A CN113323769 A CN 113323769A
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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
- F02K3/065—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 with front and aft fans
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/14—Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
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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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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
- Control Of Turbines (AREA)
Abstract
本发明公开一种基于多涵道进气级间燃烧室的变循环发动机构型，与传统变循环发动机相比，在第一涵道出口与高压压气机出口处设置高压混合室；在低压涡轮进口前设置级间燃烧室。该构型可以增大低压涡轮进口温度，提高推力；同时，大幅度提高了涵道比的变化范围，除此之外还可以大幅度提高低压涡轮功，进而增大风扇的做功能力，增加发动机空气流量，改善现有航空发动机高马赫数下空气流量较小的问题，扩大了航空发动机的速域。
The invention discloses a variable-cycle engine type based on a multi-duct intake interstage combustion chamber. Compared with a traditional variable-cycle engine, a high-pressure mixing chamber is arranged at the outlet of the first duct and the outlet of the high-pressure compressor; The interstage combustion chamber is set before the inlet. This configuration can increase the inlet temperature of the low-pressure turbine and increase the thrust; at the same time, the variation range of the bypass ratio can be greatly improved, and in addition, the power of the low-pressure turbine can be greatly improved, thereby increasing the working capacity of the fan and increasing the The air flow of the engine improves the problem of small air flow under the high Mach number of the existing aero-engine, and expands the speed range of the aero-engine.
Description
Technical Field
The invention relates to a turbine engine, in particular to a variable cycle engine configuration based on a multi-bypass inlet interstage combustor.
Background
The high-Mach-number aircraft engine is a pursuit target of research in the field of aircraft engines in various countries, but the existing aircraft engine has the defect of serious insufficient thrust when flying at a high Mach number, and becomes a bottleneck of the aircraft in a cross-speed range. Meanwhile, the change range of the bypass ratio is limited by the windward area of the engine, so that the change of the bypass ratio cannot meet the requirements of a large bypass ratio during low-speed flight and a small bypass ratio during high-speed flight.
The engine thrust is flow x thrust, and at high mach numbers the flow decays rapidly as mach numbers increase, as shown in fig. 1. Therefore, the main way to increase the thrust augmentation speed range is to increase the combustor exit temperature and thus increase the specific thrust. However, the quality of cold air and the temperature resistance degree of the current material are limited, and the improvement of the thrust-weight ratio and the improvement of the outlet temperature of the combustion chamber form an open relation, so that the thrust of the engine is improved only by improving the outlet temperature of the combustion chamber, and certain limitation exists; on the contrary, if the circulation capacity of the air compressor under the high Mach number can be fully released, the air flow under the high Mach number is increased, the thrust can be increased, the oil consumption rate can be reduced, and the method is an effective way for efficiently realizing the cross-speed-range long-range flight.
In the aspect of the duct is than adjusting (duct is than being outer culvert air flow/connotative air flow), traditional technical path is outside extension duct quantity, but this method has only changed outer culvert air flow, therefore the duct is than the range of variation restricted, if can change the interface of connotation and outer culvert, change the range of variation of culvert air flow when changing outer culvert air flow, then will greatly improve the range of variation of duct ratio.
Disclosure of Invention
In order to expand the speed range of the existing aeroengine, the invention provides a variable cycle engine configuration based on a multi-duct air inlet interstage combustion chamber, and the method of adding a mixing chamber behind a first duct outlet and a high-pressure turbine and additionally arranging the interstage combustion chamber in front of a low-pressure turbine greatly improves the air flow especially under high Mach number and increases the thrust of the engine under the high Mach number. Meanwhile, the bypass ratio adjusting range is expanded, so that the engine can adapt to different flight state requirements, and the economy of the engine is improved.
The invention is based on a variable cycle engine configuration of a multi-duct air inlet interstage combustor, a high-pressure mixing chamber is arranged at the outlet of a first duct and the outlet of a high-pressure turbine, and the interstage combustor is arranged behind the high-pressure mixing chamber and in front of the low-pressure turbine.
The airflow is divided into A, B two paths of airflow after flowing through the fan; wherein, the path A of air flow is divided into two paths after flowing through the core fan, one path of air flow passes through the high-pressure compressor and then sequentially enters the main combustion chamber and the high-pressure turbine to enter the combustion chamber, and then enters the high-pressure turbine; the other path of the water enters a first duct; finally, the two paths of air flows flow into a high-pressure mixing chamber together for mixing; the mixed gas enters an interstage combustion chamber for combustion, and the high-temperature gas after combustion enters a low-pressure turbine for expansion and work; and the B path gas flow enters a second duct, then enters a low-pressure mixing chamber together with the gas flowing out of the low-pressure turbine for mixing again, and the mixed gas enters an afterburner and is finally discharged through a tail nozzle.
The invention makes the total pressure of the two gases entering the mixing chamber approach by reasonably selecting the pressure ratio of the compressor, and realizes mixing with smaller total pressure loss. The bypass ratio adjusting range is expanded by changing the opening and closing states of the interstage combustion chambers; meanwhile, the power output range of the low-pressure turbine is expanded by controlling the interstage combustion temperature, and the rotating speed and flow regulating range of a fan driven by the low-pressure turbine are expanded, so that the speed range of the aeroengine is expanded.
The invention has the advantages that:
1. the variable-cycle engine is based on the variable-cycle engine structure of the multi-bypass intake interstage combustion chamber, the loss caused by mixing is effectively reduced due to the layout of the mixing chamber, and the efficiency and the thrust of the engine are improved.
2. The variable-cycle engine is based on a variable-cycle engine structure with the multi-duct intake interstage combustor, and the interstage combustor is arranged in front of the low-pressure turbine, so that the working capacity of the low-pressure turbine can be greatly improved.
3. The variable-cycle engine is based on a variable-cycle engine structure of the multi-bypass intake interstage combustion chamber, and the adjustment range of the bypass ratio can be greatly increased by changing the opening and closing of the interstage combustion chamber.
Drawings
Fig. 1 is a compressor map.
FIG. 2 is a schematic diagram of a prior art adaptive variable cycle engine configuration;
FIG. 3 is an engine cycle temperature entropy diagram with an interstage combustion chamber.
FIG. 4 is a schematic diagram of a variable cycle engine configuration based on a multi-ducted intake interstage combustor of the present invention.
Detailed Description
The present invention will be described in further detail with reference to the accompanying drawings.
The variable-cycle engine is based on a variable-cycle engine structure of a multi-duct air inlet interstage combustor, a mixing chamber is added behind a first duct outlet and a high-pressure turbine, the total pressure of two gases entering the mixing chamber is approximate by reasonably selecting the pressure increase ratio of an air compressor, and mixing is realized with smaller total pressure loss. Meanwhile, an interstage combustion chamber is arranged behind the high-pressure mixing chamber and in front of the low-pressure turbine, and the power output range of the low-pressure turbine and the rotating speed and flow adjusting range of a fan driven by the low-pressure turbine are expanded by controlling the interstage combustion temperature.
The airflow is divided into A, B paths after passing through the fan. Wherein, the path A of air flow is divided into two paths after flowing through the core fan, one path of air flow passes through the high-pressure compressor and then sequentially enters the main combustion chamber and the high-pressure turbine to enter the combustion chamber, and then enters the high-pressure turbine; the other path of the water enters a first duct; finally, the two paths of air flows flow into a high-pressure mixing chamber together for mixing; and the total pressure of the airflow flowing out of the high-pressure turbine is equivalent to the total pressure of the airflow flowing through the first duct by adjusting the fan pressure ratio and the core fan pressure ratio, so that the two airflows are mixed in the high-pressure mixing chamber with smaller total pressure loss, and the total pressure difference of the two mixed airflows is larger in the mixing process because the airflow flowing through the first duct passes through one compressor more than the airflow flowing through the second duct in the mixing chamber of the existing self-adaptive variable-cycle engine shown as a mixing chamber 1 in fig. 2.
The mixed gas enters the interstage combustor for combustion, and the oxygen content of the gas is increased due to the fact that the first duct outlet gas and the high-pressure turbine outlet gas are mixed in the mixing chamber and then enter the interstage combustor, and heating capacity of the interstage combustor is facilitated. On one hand, the temperature of the airflow flowing into the mixing chamber from the first duct is lower, so that the temperature of the airflow mixed in the mixing chamber can be reduced, and thus the heating amount and the circulating work can be improved after the lower-temperature gas flows into the interstage combustion chamber. On the other hand, the oxygen content in the gas from the high-pressure compressor is very low, and if the gas directly enters the interstage combustion chamber, the fuel oil in the interstage combustion chamber is not favorably combusted fully, so that the air flowing in from the first duct provides more oxygen for the interstage combustion chamber, the combustion efficiency is favorably improved, more chemical energy of the fuel is converted into gas internal energy, and the gas entering the low-pressure turbine can have more work capacity. As shown in fig. 3, which is an engine cycle temperature entropy diagram with interstage combustion chambers, the shaded portion may represent increased cycle work.
The high-temperature combustion gas after combustion enters the low-pressure turbine to do work through expansion, so that the expansion work of the low-pressure turbine is greatly increased. The working capacity of the low-pressure turbine can be greatly improved as the airflow in front of the low-pressure turbine is combusted in the interstage combustion chamber, and more importantly, the fan is directly driven by the low-pressure turbine, so that from the overall perspective of the engine, the working capacity of the fan is enhanced due to the arrangement of the interstage combustion chamber, particularly, under the condition of high Mach number, the fan can suck more mass gas, the airflow is greatly improved, the problem that the airflow of the existing aircraft engine is greatly attenuated along with the increase of the Mach number under the condition of high Mach number is solved, the speed range of the aircraft engine is further expanded, and the engine constructed according to the principle can theoretically realize the flight of 3.5 Ma.
And the B path gas flow enters a second duct, then enters a low-pressure mixing chamber together with the gas flowing out of the low-pressure turbine for mixing again, and the mixed gas enters an afterburner and is finally discharged through a tail nozzle.
During low-speed flight, the interstage combustion chamber is closed, gas flowing through the first duct does not participate in combustion at the moment, and a Brayton cycle is not formed, so that the outer duct is formed by adding the first duct and the second duct during low-speed flight, the duct ratio is large, and the oil consumption rate is low. During high-speed flight, the interstage combustion chamber is opened, gas flowing through the first bypass enters the interstage combustion chamber to be combusted, so that the gas flowing through the first bypass actually participates in a Brayton cycle, the first bypass is changed from one part of the outer bypass to one part of the inner bypass, only the second bypass remains in the outer bypass, the bypass ratio is greatly reduced, and the requirement of small bypass ratio during high-speed flight is met. Therefore, in this configuration, the adjustment range of the bypass ratio can be greatly increased by changing the opening and closing of the interstage combustor.
The variable-cycle engine is based on the variable-cycle engine structure of the multi-bypass intake interstage combustion chamber and combines with an advanced component design, and under the condition that the interstage combustion chamber is ignited, the 1.5Ma super-cruise oil consumption rate can be reduced by about 19% compared with a fourth-generation engine and can be reduced by about 7% compared with an American self-adaptive variable-cycle engine; the 2.0Ma super-cruise oil consumption rate can be reduced by about 20% compared with a four-generation engine and can be reduced by about 11% compared with an American adaptive variable-cycle engine. The self-adaptive bypass ratio adjusting range can be increased from the highest 0-1.3 to 0-4; the self-adaptive pressurization ratio adjusting range can be expanded from the highest 33% -100% to 5% -100%.
Claims (4)
1. A variable cycle engine configuration based on a multi-duct intake interstage combustor, characterized in that: a high-pressure mixing chamber is arranged at the outlet of the first bypass and the outlet of the high-pressure turbine, an interstage combustion chamber is arranged behind the high-pressure mixing chamber and in front of the low-pressure turbine, and the bypass ratio adjusting range is expanded by changing the opening and closing state of the interstage combustion chamber; meanwhile, the power output range of the low-pressure turbine is expanded by controlling the interstage combustion temperature, and the rotating speed and flow regulating range of a fan driven by the low-pressure turbine are expanded, so that the speed range of the aeroengine is expanded.
2. The variable cycle engine concept configuration based on multi-ducted intake interstage combustion chambers of claim 1 wherein: the airflow is divided into A, B two paths of airflow after flowing through the fan; wherein, the path A of air flow is divided into two paths after flowing through the core fan, one path of air flow passes through the high-pressure compressor and then sequentially enters the main combustion chamber and the high-pressure turbine to enter the combustion chamber, and then enters the high-pressure turbine; the other path of the water enters a first duct; finally, the two paths of air flows flow into a high-pressure mixing chamber together for mixing; the mixed gas enters an interstage combustion chamber for combustion, and the high-temperature gas after combustion enters a low-pressure turbine for expansion and work; and the B path gas flow enters a second duct, then enters a low-pressure mixing chamber together with the gas flowing out of the low-pressure turbine for mixing again, and the mixed gas enters an afterburner and is finally discharged through a tail nozzle.
3. The variable cycle engine concept configuration based on multi-ducted intake interstage combustion chambers of claim 2 wherein: and the fan pressure ratio and the core fan pressure ratio are adjusted to ensure that the total pressure of the airflow flowing out of the high-pressure turbine is equivalent to that of the airflow flowing through the first duct, so that the mixing loss is reduced.
4. The variable cycle engine concept configuration based on multi-ducted intake interstage combustion chambers of claim 2 wherein: during low-speed flight, the interstage combustion chamber is closed, gas flowing through the first duct does not participate in combustion at the moment, and a Brayton cycle is not formed, so that the outer duct is formed by adding the first duct and the second duct during low-speed flight, the duct ratio is large, and the oil consumption rate is low; during high-speed flight, the interstage combustion chamber is opened, gas flowing through the first duct enters the interstage combustion chamber to be combusted, so that the gas flowing through the first duct actually participates in a Brayton cycle, the first duct at the moment is changed from one part of the outer duct to one part of the inner duct, only the second duct is left in the outer duct, the duct ratio is greatly reduced, and the requirement of high-speed flight with a small duct ratio is met.
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| CN202110632005.6A CN113323769A (en) | 2021-06-07 | 2021-06-07 | Variable cycle engine configuration based on multi-duct intake interstage combustion chamber |
| CN202210637969.4A CN114776473B (en) | 2021-06-07 | 2022-06-07 | Variable cycle engine configuration based on multi-duct air inlet interstage combustion chamber |
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| CN202110632005.6A CN113323769A (en) | 2021-06-07 | 2021-06-07 | Variable cycle engine configuration based on multi-duct intake interstage combustion chamber |
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| CN202210637969.4A Active CN114776473B (en) | 2021-06-07 | 2022-06-07 | Variable cycle engine configuration based on multi-duct air inlet interstage combustion chamber |
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Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
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| CN114810350A (en) * | 2022-05-06 | 2022-07-29 | 中国科学院工程热物理研究所 | Methane precooling turbine-based combined cycle engine system with interstage combustion chamber |
| CN114962004A (en) * | 2022-07-06 | 2022-08-30 | 北京航空航天大学 | Self-adaptive flying and launching integrated thermal management system based on third flow and fuel oil heat sink |
| CN115263598A (en) * | 2022-07-28 | 2022-11-01 | 北京航空航天大学 | A double-variable ducted, wide-speed range, high-pass rheological cycle engine type |
| CN116714783A (en) * | 2023-08-10 | 2023-09-08 | 北京航空航天大学 | Flywheel fan engine type aerospace carrier |
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Cited By (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114810350A (en) * | 2022-05-06 | 2022-07-29 | 中国科学院工程热物理研究所 | Methane precooling turbine-based combined cycle engine system with interstage combustion chamber |
| CN114810350B (en) * | 2022-05-06 | 2023-12-22 | 中国科学院工程热物理研究所 | Methane precooling turbine-based combined cycle engine system with interstage combustion chamber |
| CN114962004A (en) * | 2022-07-06 | 2022-08-30 | 北京航空航天大学 | Self-adaptive flying and launching integrated thermal management system based on third flow and fuel oil heat sink |
| CN114962004B (en) * | 2022-07-06 | 2025-03-04 | 北京航空航天大学 | Adaptive aircraft-engine integrated thermal management system based on third flow and fuel heat sink |
| CN115263598A (en) * | 2022-07-28 | 2022-11-01 | 北京航空航天大学 | A double-variable ducted, wide-speed range, high-pass rheological cycle engine type |
| CN116714783A (en) * | 2023-08-10 | 2023-09-08 | 北京航空航天大学 | Flywheel fan engine type aerospace carrier |
| CN116714783B (en) * | 2023-08-10 | 2023-10-10 | 北京航空航天大学 | A flywheel fan engine type aerospace carrier |
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| Publication number | Publication date |
|---|---|
| CN114776473A (en) | 2022-07-22 |
| CN114776473B (en) | 2024-01-19 |
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Application publication date: 20210831 |