CN111207415A - Flame tube of combustion chamber of aircraft engine - Google Patents
Flame tube of combustion chamber of aircraft engine Download PDFInfo
- Publication number
- CN111207415A CN111207415A CN202010050542.5A CN202010050542A CN111207415A CN 111207415 A CN111207415 A CN 111207415A CN 202010050542 A CN202010050542 A CN 202010050542A CN 111207415 A CN111207415 A CN 111207415A
- Authority
- CN
- China
- Prior art keywords
- flame tube
- combustion chamber
- flame
- guide support
- casing
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
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Classifications
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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
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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
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
- F23R3/10—Air inlet arrangements for primary air
- F23R3/12—Air inlet arrangements for primary air inducing a vortex
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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
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/03043—Convection cooled combustion chamber walls with means for guiding the cooling air flow
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Spray-Type Burners (AREA)
Abstract
The invention provides a flame tube of an aircraft engine combustion chamber, which comprises the following structural designs: a guide support plate and a flame tube step convergence section are additionally arranged in an annular cavity between the casing and the flame tube. The guide support plate is additionally arranged in the inner ring cavity and the outer ring cavity, so that the convective heat transfer time can be prolonged, the wall surface temperature of the flame tube can be effectively reduced, and the service life of the flame tube is prolonged. The convergent section step-shaped structure is used for prolonging the residence time of fuel in the flame tube, a small backflow area is formed at each step, mixing of fuel mixed gas and fresh air is facilitated, uniform temperature and stable combustion in the flame tube are facilitated, flameout is prevented, sufficient time for evaporation mixing and combustion of the fuel mixed gas is guaranteed, combustion efficiency is guaranteed, and the requirements of a high-temperature-rise and even ultrahigh-temperature-rise combustion chamber are met.
Description
Technical Field
The invention belongs to the field of flame tubes of combustion chambers of aero-engines, and particularly relates to a flame tube of a combustion chamber of an aero-engine.
Background
Future military aircraft engines are mainly pursuing high thrust-to-weight ratio targets, meaning that engine turbine front (combustor exit) temperatures need to be significantly increased. Therefore, the development trend of the military engine combustion chamber in the future is to develop a high-temperature rise combustion chamber or even an ultrahigh-temperature rise combustion chamber. The temperature rise of the combustion chamber directly corresponds to the oil-gas ratio, according to the current engine technical development, the engine with the thrust-weight ratio of 12 grades is basically adopted, the total oil-gas ratio is 0.040, the temperature rise is 1250K, the engine with the thrust-weight ratio of 15 grades is adopted, the total oil-gas ratio is more than 0.045, and the temperature rise is about 1400. For the temperature rise level of the ultra-high temperature rise combustor, the typical technical characteristics of the combustor are completely different from those of the conventional combustor. The difficulty of the design of the high-temperature-rise combustion chamber is that the thrust-weight ratio is continuously increased, the temperature rise is continuously improved, the total oil-gas ratio is also continuously increased, the air quantity for combustion is also increased, and the air quantity for cooling is obviously reduced. Therefore, the cooling efficiency of the flame tube can only be improved by technical means to meet the requirement of the service life of the flame tube. In order to develop the technology of the high-temperature-rise combustion chamber better, a novel flame tube must be provided.
Disclosure of Invention
The invention aims to solve the technical problem of providing a flame tube of a combustion chamber of an aircraft engine, which can improve the flowing time of air in an annular cavity between a casing and the flame tube and gas mixture in the flame tube. The technology adopts the design that a flow guide support plate and a flame tube step convergence section are additionally arranged in an annular cavity between a casing and a flame tube. The invention has the advantages that: the structure can prolong the flowing time of air in the annular cavity between the casing and the flame tube, improve the reaction time of fuel oil mixed gas in the flame tube, further enhance the cooling effect of the wall surface of the flame tube and improve the heat-containing strength of the flame tube.
Technical scheme
The invention aims to provide a flame tube of a combustion chamber of an aircraft engine.
The technical scheme of the invention is as follows:
a flame tube of a combustion chamber of an aero-engine comprises a combustion chamber casing, a flame tube, a diffuser, a cap cover, a cooling hole and a flow guide support plate. The outer casing and the outer flame tube of the combustion chamber form an outer ring cavity channel; the inner step convergent section casing and the inner step convergent section flame barrel form an inner ring cavity channel; the air channels of the inner and outer annular cavities are provided with guide supporting plates; the inner flame tube and the outer flame tube form a flame tube, and the fuel oil mixed gas is combusted in the flame tube; the arc structures at the front sections of the inner flame tube and the outer flame tube form a cap; the inner casing and the convergent section of the flame tube are designed in a step shape.
Wherein, the inner and outer ring cavity channels are provided with arc-shaped flow guide support plates which form a certain angle with the axial flow direction, and the angle is between 40 and 60 degrees;
wherein, the inner and outer flame tubes are provided with cooling holes, the direction of the cooling holes is consistent with the direction of the arc-shaped flow guide support plate, and the diameter of the cooling holes is between 1.5mm and 3 mm;
wherein, the convergent section of the inner flame tube is designed in a ladder shape, the number of steps is kept between 3 and 5, and the height of each step is between 10 and 15 mm.
The invention has the following beneficial effects:
after air entering from the head diffuser passes through the flow guide support plates arranged in the inner ring cavity and the outer ring cavity, the flow direction is changed from the original axial flow to spiral flow around the outer wall of the flame tube, and the flow time of the air in the ring cavities can be prolonged by the flow mode of changing the direction; meanwhile, a part of cooling gas enters the flame tube through the cooling holes with the same angle as the flow guide support plate, and a gas film is formed on the inner wall surface of the flame tube. The inner ring cavity and the outer ring cavity are provided with the flow guide support plates, so that the convection heat exchange time can be prolonged, the wall surface temperature of the flame tube can be effectively reduced, and the service life of the flame tube is prolonged. The convergent section step-shaped structure is used for prolonging the residence time of fuel in the flame tube, ensures that fuel gas mixture has enough time for evaporation, mixing and combustion, ensures the combustion efficiency and meets the requirements of a high-temperature rise and even an ultrahigh-temperature rise combustion chamber. Due to the step-shaped design of the convergence section, a small backflow area is formed at each step of the step, so that the mixing of fuel oil mixed gas and fresh air is facilitated, the uniform temperature and stable combustion in the flame tube are facilitated, and flameout is prevented.
Drawings
The above features and advantages of the present invention will be better understood upon reading the detailed description of the invention in conjunction with the following drawings. In the drawings, the various features are not necessarily drawn to scale.
FIG. 1 is a schematic diagram of a prior art configuration;
FIG. 2 is a schematic view of the angle of installation of the guide;
FIG. 3 is a cooling hole flow schematic;
description of the symbols:
Detailed Description
The invention will now be further described with reference to the accompanying drawings in which:
with reference to fig. 1, the present invention provides a new aero-engine combustor basket.
The specific process comprises the following steps: after part of high-pressure air entering from a head diffuser 1 flows through a flow guide support plate 2 arranged in an inner ring cavity and an outer ring cavity, the flow direction is changed from the original axial flow (X direction) to spiral flow around the outer walls of flame tubes 6 and 7, and the flow time of the air in the ring cavities can be prolonged by changing the axial flow into the spiral flow; meanwhile, most of high-pressure air enters the flame tube through the cooling holes 8 with the same angle as the flow guide support plate 2, and an air film is formed on the inner wall surface of the flame tube, so that the temperature of the wall surface of the flame tube is effectively reduced, and the service life of the flame tube is prolonged. The inner flame tube convergence section ladder-shaped structure 7 is used for prolonging the residence time of fuel in the flame tube 9, ensuring that fuel mixed gas has enough time for evaporation, mixing and combustion, ensuring the combustion efficiency and meeting the requirements of a high-temperature rise and even an ultrahigh-temperature rise combustion chamber.
Claims (1)
1. A flame tube of an aircraft engine combustion chamber comprises a combustion chamber casing, a flame tube, a diffuser, a cap cover, a cooling hole and a flow guide support plate;
wherein, the outer casing of the combustion chamber and the outer flame tube form an outer ring cavity channel; the inner step convergent section casing and the inner step convergent section flame barrel form an inner ring cavity channel; the air channels of the inner and outer annular cavities are provided with guide supporting plates; the inner flame tube and the outer flame tube form a flame tube, and the fuel oil mixed gas is combusted in the flame tube; the arc structures at the front sections of the inner flame tube and the outer flame tube form a cap; the inner casing and the flame tube convergence section are designed in a step shape;
wherein, the inner and outer ring cavity channels are provided with arc-shaped flow guide support plates which form a certain angle with the axial flow direction, and the angle is between 40 and 60 degrees;
wherein, the inner and outer flame tubes are provided with cooling holes, the direction of the cooling holes is consistent with the direction of the arc-shaped flow guide support plate, and the diameter of the cooling holes is between 1.5mm and 3 mm;
wherein, the convergent section of the inner flame tube is designed in a ladder shape, the number of steps is kept between 3 and 5, and the height of each step is between 10 and 15 mm.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010050542.5A CN111207415A (en) | 2020-01-17 | 2020-01-17 | Flame tube of combustion chamber of aircraft engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010050542.5A CN111207415A (en) | 2020-01-17 | 2020-01-17 | Flame tube of combustion chamber of aircraft engine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN111207415A true CN111207415A (en) | 2020-05-29 |
Family
ID=70789115
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202010050542.5A Pending CN111207415A (en) | 2020-01-17 | 2020-01-17 | Flame tube of combustion chamber of aircraft engine |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN111207415A (en) |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH09196377A (en) * | 1996-01-12 | 1997-07-29 | Hitachi Ltd | Gas turbine combustor |
| JP2000320837A (en) * | 1999-05-06 | 2000-11-24 | Hitachi Ltd | Gas turbine combustor |
| CN1395063A (en) * | 2001-06-20 | 2003-02-05 | 西门子公司 | Combustion chamber of gas turbine and air flow-guiding method |
| CN103994468A (en) * | 2013-02-20 | 2014-08-20 | 株式会社日立制作所 | Gas turbine combustor equipped with heat-transfer device |
| CN205505079U (en) * | 2016-04-11 | 2016-08-24 | 杨庆春 | Supersonic combustion chamber of accurate isothermal |
| CN206600840U (en) * | 2016-12-29 | 2017-10-31 | 中国航发商用航空发动机有限责任公司 | A kind of burner inner liner of combustion chamber |
-
2020
- 2020-01-17 CN CN202010050542.5A patent/CN111207415A/en active Pending
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH09196377A (en) * | 1996-01-12 | 1997-07-29 | Hitachi Ltd | Gas turbine combustor |
| JP2000320837A (en) * | 1999-05-06 | 2000-11-24 | Hitachi Ltd | Gas turbine combustor |
| CN1395063A (en) * | 2001-06-20 | 2003-02-05 | 西门子公司 | Combustion chamber of gas turbine and air flow-guiding method |
| CN103994468A (en) * | 2013-02-20 | 2014-08-20 | 株式会社日立制作所 | Gas turbine combustor equipped with heat-transfer device |
| CN205505079U (en) * | 2016-04-11 | 2016-08-24 | 杨庆春 | Supersonic combustion chamber of accurate isothermal |
| CN206600840U (en) * | 2016-12-29 | 2017-10-31 | 中国航发商用航空发动机有限责任公司 | A kind of burner inner liner of combustion chamber |
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| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| WD01 | Invention patent application deemed withdrawn after publication |
Application publication date: 20200529 |
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| WD01 | Invention patent application deemed withdrawn after publication |