CN111894761A - Centripetal turbofan jet engine - Google Patents
Centripetal turbofan jet engine Download PDFInfo
- Publication number
- CN111894761A CN111894761A CN202010670311.4A CN202010670311A CN111894761A CN 111894761 A CN111894761 A CN 111894761A CN 202010670311 A CN202010670311 A CN 202010670311A CN 111894761 A CN111894761 A CN 111894761A
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- Prior art keywords
- turbine
- pressure
- compressor
- fan
- low
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- 238000002485 combustion reaction Methods 0.000 claims abstract description 30
- 230000005540 biological transmission Effects 0.000 claims abstract description 28
- 239000000112 cooling gas Substances 0.000 claims abstract description 5
- 238000001816 cooling Methods 0.000 claims description 7
- 239000000295 fuel oil Substances 0.000 claims description 4
- 238000003466 welding Methods 0.000 claims description 2
- 239000007789 gas Substances 0.000 abstract description 8
- 238000000034 method Methods 0.000 abstract description 5
- 230000002035 prolonged effect Effects 0.000 abstract description 3
- 239000000567 combustion gas Substances 0.000 description 3
- 238000010521 absorption reaction Methods 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000006243 chemical reaction Methods 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 239000000446 fuel Substances 0.000 description 1
- 239000000203 mixture Substances 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 238000007789 sealing Methods 0.000 description 1
Images
Classifications
-
- 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
- F02K3/06—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 fan
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/26—Rotors specially for elastic fluids
- F04D29/32—Rotors specially for elastic fluids for axial flow pumps
- F04D29/38—Blades
- F04D29/388—Blades characterised by construction
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/58—Cooling; Heating; Diminishing heat transfer
- F04D29/582—Cooling; Heating; Diminishing heat transfer specially adapted for elastic fluid pumps
- F04D29/584—Cooling; Heating; Diminishing heat transfer specially adapted for elastic fluid pumps cooling or heating the machine
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Physics & Mathematics (AREA)
- Thermal Sciences (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The invention discloses a centripetal turbofan jet engine which comprises a fan, a gas compressor, a combustion chamber, a turbine, an exhaust assembly and an inner shell, wherein the gas compressor consists of a low-pressure area and a high-pressure area, and the low-pressure area of the gas compressor is connected with the fan; the combustion chamber is communicated with a high-pressure area of the compressor; the turbine consists of a high-pressure area and a low-pressure area, and the high-pressure area of the turbine is communicated with one end of the combustion chamber far away from the compressor; the exhaust assembly is communicated with one end of the turbine far away from the combustion chamber; an outer aerodynamic shell is arranged outside the inner shell. The energy in the combustion chamber can be extracted and absorbed by the turbine after the combustion process, the high-temperature and high-pressure gas after combustion can be depressurized in the process, so that the turbine is driven to rotate, the turbine drives the fan and the gas compressor to run through the transmission shaft, and the circulation of the energy can be realized; in addition, the cooling gas is filled in the fan blades, so that the working temperature of the fan is reduced, and the service life of the fan is prolonged.
Description
Technical Field
The invention relates to the field of engines, in particular to a centripetal turbofan jet engine.
Background
Turbofan engine, also known as "turbofan engine", refers to a gas turbine engine in which the combustion gases ejected from the nozzle and the air discharged from the fan generate a reaction thrust, and a portion of the available energy in the combustion gases flowing from the core of the turbofan engine is used to drive the low pressure turbine to drive the fan, and a portion of the available energy is used to accelerate the ejected combustion gases in the nozzle.
Compared with other engines, the turbofan engine has the advantages of large thrust, low noise, long aircraft range and the like. Meanwhile, the turbofan engine has the following defects: the fan diameter is big, and the frontal area is big, therefore the resistance is big, and engine structure is complicated, and the design degree of difficulty is big to traditional turbofan engine's fan does not have the cooling function, leads to fan life relatively short. To solve the above problems, we have developed a radial fan jet engine.
Disclosure of Invention
In order to solve the technical problems, the invention provides the following technical scheme:
the present invention provides a radial fan jet engine comprising: the fan is used for generating high-speed airflow and comprises a plurality of fan blades which are arranged in an annular array and have a cooling function;
the air compressor is used for pressurizing air flow in a stepwise manner, the air compressor consists of a low-pressure area and a high-pressure area, and the low-pressure area of the air compressor is connected with the fan;
the combustor is used for mixing and combusting aviation fuel oil and high-pressure airflow and is communicated with a high-pressure area of the compressor;
the turbine is used for driving the compressor and the fan and cooling and decompressing the airflow, the turbine consists of a high-pressure area and a low-pressure area, and the high-pressure area of the turbine is communicated with one end of the combustion chamber far away from the compressor;
an exhaust assembly for exhausting hot air, the exhaust assembly communicating with an end of the turbine remote from the combustion chamber;
the fan, the compressor, the combustion chamber, the turbine and the exhaust assembly are externally provided with an inner shell, and an outer shell which accords with aerodynamics is arranged outside the inner shell.
As a preferred technical scheme of the invention, the fan blades are internally provided with sealed cavities, and cooling gas is filled in the sealed cavities.
As a preferable technical solution of the present invention, the low pressure region and the high pressure region of the compressor are communicated with each other, and the low pressure region of the compressor includes a three-stage low-pressure compressor impeller, and the high pressure region of the compressor includes a nine-stage high-pressure compressor impeller.
In a preferred embodiment of the invention, the high-pressure region and the low-pressure region of the turbine are connected to one another, wherein the high-pressure region of the turbine comprises a first-stage high-pressure turbine wheel and the low-pressure region of the turbine comprises a fourth-stage low-pressure turbine wheel.
As a preferable technical scheme of the invention, the outer part of the inner shell is welded with a connecting frame, and one end of the connecting frame, which is far away from the inner shell, is fixedly connected with the inner wall of the outer shell.
As a preferred technical solution of the present invention, a first transmission shaft is fixedly connected to an axis of the compressor, a second transmission shaft is fixedly connected to an axis of the turbine, one end of the first transmission shaft, which is far away from the turbine, is coaxially and fixedly connected to the fan, and one end of the second transmission shaft, which is far away from the exhaust assembly, is coaxially and fixedly connected to the first transmission shaft.
As a preferable technical scheme of the invention, a low-pressure bearing is arranged between the fan and the air compressor, and the fan is in transmission connection with the first transmission shaft through the low-pressure bearing.
In a preferred embodiment of the present invention, a high-pressure bearing is disposed between the turbine and the combustion chamber, and the turbine is in transmission connection with the second transmission shaft through the high-pressure bearing.
The invention has the beneficial effects that: the centripetal turbofan jet engine accelerates, pressurizes and mixes air through a fan, an air compressor and a combustion chamber respectively, energy in the combustion chamber can be extracted and absorbed by a turbine after the combustion process, and high-temperature and high-pressure gas after combustion can be depressurized in the process so as to drive the turbine to rotate; in addition, the cooling gas is filled in the fan blades, so that the working temperature of the fan is reduced, and the service life of the fan is prolonged.
Drawings
The accompanying drawings, which are included to provide a further understanding of the invention and are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and together with the description serve to explain the principles of the invention and not to limit the invention. In the drawings:
FIG. 1 is a schematic structural view of a radial fan jet engine of the present invention;
FIG. 2 is a cross-sectional view at the axis of a radial fan jet engine of the present invention;
FIG. 3 is an enlarged schematic view of a radial fan jet engine according to the present invention at A;
FIG. 4 is an enlarged schematic view of a radial fan jet engine B of the present invention;
FIG. 5 is a cross-sectional view of a radial fan jet blade according to the present invention;
in the figure: 1. a fan; 2. a compressor; 21. a low pressure compressor impeller; 22. a high pressure compressor impeller; 3. a combustion chamber; 4. a turbine; 41. a high pressure compressor impeller; 42. a low pressure compressor impeller; 5. an exhaust assembly; 6. an inner shell; 7. a housing; 8. a connecting frame; 9. a first drive shaft; 10. a second drive shaft; 11. a fan blade; 12. sealing the cavity; 13. a low pressure bearing; 14. and a high-pressure bearing.
Detailed Description
The preferred embodiments of the present invention will be described in conjunction with the accompanying drawings, and it will be understood that they are described herein for the purpose of illustration and explanation and not limitation.
Example (b): as shown in fig. 1-5, a radial-inflow turbofan jet engine includes a fan 1 for generating a high-speed airflow, a compressor 2 for pressurizing the airflow step by step, a combustion chamber 3 for mixing and combusting aviation fuel oil and high-pressure airflow, a turbine 4 for driving the compressor 2 and the fan 1 and cooling and depressurizing the airflow, and an exhaust assembly 5 for exhausting hot air, wherein the fan 1 includes a plurality of blades 11 arranged in an annular array and having a cooling function, the blades 11 are each provided with a sealed cavity 12 inside, and the sealed cavity 12 is filled with cooling gas, so that the temperature of the fan 1 during operation is reduced, and the service life of the fan 1 is prolonged.
The compressor 2 consists of a low-pressure area and a high-pressure area, and the low-pressure area of the compressor 2 is connected with the fan 1, and the method comprises the following steps: the low-pressure region and the high-pressure region of the compressor 2 are communicated with each other, the low-pressure region of the compressor 2 includes a three-stage low-pressure compressor impeller 21, and the high-pressure region of the compressor 2 includes a nine-stage high-pressure compressor impeller 22.
The combustion chamber 3 is communicated with a high-pressure area of the compressor 2; the turbine 4 consists of a high-pressure area and a low-pressure area, and the high-pressure area of the turbine 4 is communicated with one end of the combustion chamber 3 far away from the compressor 2, specifically: the high-pressure region and the low-pressure region of the turbine 4 communicate with each other, wherein the high-pressure region of the turbine 4 comprises a first-stage high-pressure turbine wheel 41 and the low-pressure region of the turbine 4 comprises a fourth-stage low-pressure turbine wheel 42.
The exhaust assembly 5 is communicated with one end of the turbine 4 far away from the combustion chamber 3; the fan 1, the compressor 2, the combustion chamber 3, the turbine 4 and the exhaust assembly 5 are externally provided with an inner casing 6, and the outside of the inner casing 6 is provided with an outer aerodynamic casing 7. The outer welding of inner shell 6 has link 8, the one end of link 8 keeping away from inner shell 6 and the inner wall fixed connection of shell 7.
The axial line of the compressor 2 is fixedly connected with a first transmission shaft 9, the axial line of the turbine 4 is fixedly connected with a second transmission shaft 10, one end, far away from the turbine 4, of the first transmission shaft 9 is fixedly connected with the fan 1 in a coaxial line mode, and one end, far away from the exhaust assembly 5, of the second transmission shaft 10 is fixedly connected with the first transmission shaft 9 in a coaxial line mode.
A low-pressure bearing 13 is arranged between the fan 1 and the compressor 2, and the fan 1 is in transmission connection with the first transmission shaft 9 through the low-pressure bearing 13. A high-pressure bearing 14 is arranged between the turbine 4 and the combustion chamber 3, and the turbine 5 is in driving connection with the second drive shaft 10 via the high-pressure bearing 14.
The working principle is as follows: the invention accelerates air through the fan 1, pressurizes the air current step by step through the air compressor 2, mixes and burns the mixture of aviation fuel oil and high-pressure air current through the combustion chamber 3, cools and depressurizes the air current through the turbine 4 and drives the air compressor 2 and the fan 1, and finally discharges hot air through the exhaust assembly 5. In the process, the energy in the combustion chamber 3 is extracted and absorbed by the turbine 4 after being combusted, the high-temperature and high-pressure gas after being combusted is depressurized in the absorption process, so that the turbine 4 is driven to rotate, the turbine 4 drives the fan 1 and the compressor to operate 2 through the second transmission shaft 10 and the first transmission shaft 9, and the energy circulation is realized. Compared with the traditional engine, the invention has the advantages of high propulsion efficiency, low fuel consumption, long airplane voyage, long service life and the like.
In the description of the present invention, it should be noted that the terms "vertical", "upper", "lower", "horizontal", and the like indicate orientations or positional relationships based on those shown in the drawings, and are only for convenience of describing the present invention and simplifying the description, but do not indicate or imply that the referred device or element must have a specific orientation, be constructed in a specific orientation, and be operated, and thus, should not be construed as limiting the present invention.
In the description of the present invention, it should also be noted that, unless otherwise explicitly specified or limited, the terms "disposed," "mounted," "connected," and "connected" are to be construed broadly and may, for example, be fixedly connected, detachably connected, or integrally connected; can be mechanically or electrically connected; they may be connected directly or indirectly through intervening media, or they may be interconnected between two elements. The specific meanings of the above terms in the present invention can be understood by those skilled in the art according to specific situations.
Finally, it should be noted that: although the present invention has been described in detail with reference to the foregoing embodiments, it will be apparent to those skilled in the art that changes may be made in the embodiments and/or equivalents thereof without departing from the spirit and scope of the invention. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (8)
1. A radial-inflow turbofan jet engine, comprising:
the fan (1) is used for generating high-speed airflow, and the fan (1) comprises a plurality of fan blades (11) which are arranged in an annular array and have a cooling function;
a compressor (2) for pressurizing the air flow in stages, the compressor (2) consisting of a low-pressure region and a high-pressure region, and the low-pressure region of the compressor (2) being connected to the fan (1);
the combustion chamber (3) is used for mixing and combusting aviation fuel oil and high-pressure airflow, and the combustion chamber (3) is communicated with a high-pressure area of the compressor (2);
the turbine (4) is used for driving the compressor (2) and the fan (1) and cooling and decompressing the airflow, the turbine (4) consists of a high-pressure area and a low-pressure area, and the high-pressure area of the turbine (4) is communicated with one end of the combustion chamber (3) far away from the compressor (2);
an exhaust assembly (5) for exhausting hot air, the exhaust assembly (5) communicating with an end of the turbine (4) remote from the combustion chamber (3);
an inner shell (6) is arranged outside the fan (1), the air compressor (2), the combustion chamber (3), the turbine (4) and the exhaust assembly (5), and an outer shell (7) conforming to aerodynamics is arranged outside the inner shell (6).
2. A radial inflow turbofan jet engine in accordance with claim 1 wherein:
the fan blade is characterized in that a sealed cavity (12) is formed in the fan blade (11), and cooling gas is filled in the sealed cavity (12).
3. A radial inflow turbofan jet engine in accordance with claim 1 wherein:
the low-pressure area and the high-pressure area of the compressor (2) are communicated with each other, the low-pressure area of the compressor (2) comprises three stages of low-pressure compressor impellers (21), and the high-pressure area of the compressor (2) comprises nine stages of high-pressure compressor impellers (22).
4. A radial inflow turbofan jet engine in accordance with claim 1 wherein:
the high-pressure region and the low-pressure region of the turbine (4) communicate with each other, wherein the high-pressure region of the turbine (4) comprises a first-stage high-pressure turbine wheel (41) and the low-pressure region of the turbine (4) comprises a fourth-stage low-pressure turbine wheel (42).
5. A radial inflow turbofan jet engine in accordance with claim 1 wherein:
the outer welding of inner shell (6) has link (8), the one end of keeping away from inner shell (6) of link (8) and the inner wall fixed connection of shell (7).
6. A radial inflow turbofan jet engine in accordance with claim 1 wherein:
a first transmission shaft (9) is fixedly connected to the axis of the compressor (2), and a second transmission shaft (10) is fixedly connected to the axis of the turbine (4);
one end, far away from turbine (4), of first transmission shaft (9) is coaxially and fixedly connected with fan (1), and one end, far away from exhaust assembly (5), of second transmission shaft (10) is coaxially and fixedly connected with first transmission shaft (9).
7. A radial inflow turbofan jet engine in accordance with claim 6 wherein:
a low-pressure bearing (13) is arranged between the fan (1) and the compressor (2), and the fan (1) is in transmission connection with the first transmission shaft (9) through the low-pressure bearing (13).
8. A radial inflow turbofan jet engine in accordance with claim 6 wherein:
a high-pressure bearing (14) is arranged between the turbine (4) and the combustion chamber (3), and the turbine (5) is in transmission connection with the second transmission shaft (10) through the high-pressure bearing (14).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010670311.4A CN111894761A (en) | 2020-07-13 | 2020-07-13 | Centripetal turbofan jet engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010670311.4A CN111894761A (en) | 2020-07-13 | 2020-07-13 | Centripetal turbofan jet engine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN111894761A true CN111894761A (en) | 2020-11-06 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202010670311.4A Pending CN111894761A (en) | 2020-07-13 | 2020-07-13 | Centripetal turbofan jet engine |
Country Status (1)
| Country | Link |
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| CN (1) | CN111894761A (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20230027726A1 (en) * | 2021-07-19 | 2023-01-26 | Raytheon Technologies Corporation | High and low spool configuration for a gas turbine engine |
Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0469827A1 (en) * | 1990-08-01 | 1992-02-05 | General Electric Company | Heat exchange arrangement in a fan duct |
| CN101331057A (en) * | 2005-12-29 | 2008-12-24 | 美蓓亚株式会社 | Fan blade with non-varying stagger and camber angles |
| CN103089317A (en) * | 2011-10-31 | 2013-05-08 | 中航商用航空发动机有限责任公司 | Hollow fan blade and manufacturing method thereof |
| CN104895617A (en) * | 2015-05-19 | 2015-09-09 | 集美大学 | Bladeless turbine engine |
| CN106438104A (en) * | 2016-09-18 | 2017-02-22 | 中国科学院工程热物理研究所 | Fuel-rich pre-burning turbofan engine |
| CN106593694A (en) * | 2016-12-23 | 2017-04-26 | 李可 | Radial turbofan jet engine |
| CN107908816A (en) * | 2017-10-13 | 2018-04-13 | 北京航空航天大学 | Aero-engine cooling and the integrated design method of cooling air based on hollow fan blade |
| CN111232223A (en) * | 2018-11-29 | 2020-06-05 | 通用电气公司 | Propulsion engine thermal management system |
| CN111350603A (en) * | 2018-12-21 | 2020-06-30 | 劳斯莱斯有限公司 | Gas turbine engine for mounting on an aircraft |
-
2020
- 2020-07-13 CN CN202010670311.4A patent/CN111894761A/en active Pending
Patent Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0469827A1 (en) * | 1990-08-01 | 1992-02-05 | General Electric Company | Heat exchange arrangement in a fan duct |
| CN101331057A (en) * | 2005-12-29 | 2008-12-24 | 美蓓亚株式会社 | Fan blade with non-varying stagger and camber angles |
| CN103089317A (en) * | 2011-10-31 | 2013-05-08 | 中航商用航空发动机有限责任公司 | Hollow fan blade and manufacturing method thereof |
| CN104895617A (en) * | 2015-05-19 | 2015-09-09 | 集美大学 | Bladeless turbine engine |
| CN106438104A (en) * | 2016-09-18 | 2017-02-22 | 中国科学院工程热物理研究所 | Fuel-rich pre-burning turbofan engine |
| CN106593694A (en) * | 2016-12-23 | 2017-04-26 | 李可 | Radial turbofan jet engine |
| CN107908816A (en) * | 2017-10-13 | 2018-04-13 | 北京航空航天大学 | Aero-engine cooling and the integrated design method of cooling air based on hollow fan blade |
| CN111232223A (en) * | 2018-11-29 | 2020-06-05 | 通用电气公司 | Propulsion engine thermal management system |
| CN111350603A (en) * | 2018-12-21 | 2020-06-30 | 劳斯莱斯有限公司 | Gas turbine engine for mounting on an aircraft |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20230027726A1 (en) * | 2021-07-19 | 2023-01-26 | Raytheon Technologies Corporation | High and low spool configuration for a gas turbine engine |
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Application publication date: 20201106 |