CN112901279B - Turbine blade adopting bolt-fixed ceramic armor - Google Patents
Turbine blade adopting bolt-fixed ceramic armor Download PDFInfo
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- CN112901279B CN112901279B CN202110125484.2A CN202110125484A CN112901279B CN 112901279 B CN112901279 B CN 112901279B CN 202110125484 A CN202110125484 A CN 202110125484A CN 112901279 B CN112901279 B CN 112901279B
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- blade
- ceramic
- ceramic armor
- pin
- armor
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- 239000000919 ceramic Substances 0.000 title claims abstract description 37
- 239000002184 metal Substances 0.000 claims abstract description 26
- 229910052751 metal Inorganic materials 0.000 claims abstract description 26
- 239000011153 ceramic matrix composite Substances 0.000 claims abstract description 19
- 239000011159 matrix material Substances 0.000 claims abstract description 16
- 239000000463 material Substances 0.000 claims abstract description 12
- 239000000112 cooling gas Substances 0.000 claims abstract description 4
- 238000001816 cooling Methods 0.000 claims description 16
- 238000012423 maintenance Methods 0.000 abstract description 7
- 239000000758 substrate Substances 0.000 abstract description 3
- 239000000956 alloy Substances 0.000 description 7
- 229910045601 alloy Inorganic materials 0.000 description 6
- 238000004519 manufacturing process Methods 0.000 description 5
- 239000002737 fuel gas Substances 0.000 description 3
- 239000007789 gas Substances 0.000 description 3
- 239000000567 combustion gas Substances 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 230000006872 improvement Effects 0.000 description 2
- 230000009467 reduction Effects 0.000 description 2
- 238000002679 ablation Methods 0.000 description 1
- 238000005275 alloying Methods 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000005266 casting Methods 0.000 description 1
- 230000008859 change Effects 0.000 description 1
- 238000002485 combustion reaction Methods 0.000 description 1
- 239000002131 composite material Substances 0.000 description 1
- 238000007796 conventional method Methods 0.000 description 1
- 238000005260 corrosion Methods 0.000 description 1
- 230000007797 corrosion Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 230000007613 environmental effect Effects 0.000 description 1
- 230000003628 erosive effect Effects 0.000 description 1
- 238000010438 heat treatment Methods 0.000 description 1
- 230000003116 impacting effect Effects 0.000 description 1
- 238000009434 installation Methods 0.000 description 1
- 239000007769 metal material Substances 0.000 description 1
- 150000002739 metals Chemical class 0.000 description 1
- 230000003647 oxidation Effects 0.000 description 1
- 238000007254 oxidation reaction Methods 0.000 description 1
- 230000002035 prolonged effect Effects 0.000 description 1
- 239000000126 substance Substances 0.000 description 1
- 229910000601 superalloy Inorganic materials 0.000 description 1
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
- F01D5/282—Selecting composite materials, e.g. blades with reinforcing filaments
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
- F01D5/284—Selection of ceramic materials
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Materials Engineering (AREA)
- Ceramic Engineering (AREA)
- Composite Materials (AREA)
- Architecture (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
本发明属于航空发动机涡轮叶片技术领域，涉及一种采用插销固定式陶瓷铠甲的涡轮叶片。本发明的结构方案针对涡轮叶片的前缘区域，可实现陶瓷基复合材料构件与叶片金属基体快捷稳固相连，并且在不破坏叶片原有气动外型的前提下，有效的保护涡轮叶片前缘区域和减少对冷却气体的使用，以提高叶片的耐高温能力，以及发动机的推力性能。同时，该陶瓷铠甲在受损后，可在维修中方便更换，并使高价值的叶片金属基体重复利用延长寿命，从而使发动机的使用维护经济性得到大幅度提高。
The invention belongs to the technical field of aero-engine turbine blades, and relates to a turbine blade using a bolt-fixed ceramic armor. The structural scheme of the present invention is aimed at the leading edge area of the turbine blade, which can realize the quick and stable connection between the ceramic matrix composite material component and the metal substrate of the blade, and effectively protect the leading edge area of the turbine blade without destroying the original aerodynamic shape of the blade. And reduce the use of cooling gas to improve the high temperature resistance of the blades, as well as the thrust performance of the engine. At the same time, after the ceramic armor is damaged, it can be easily replaced during maintenance, and the high-value metal matrix of the blade can be reused to prolong the service life, thereby greatly improving the economical use and maintenance of the engine.
Description
Technical Field
The invention belongs to the technical field of turbine blades of aero-engines, and relates to a turbine blade adopting bolt-fixed ceramic armor.
Background
For an aircraft turbine engine, an increase in pre-turbine combustion gas temperature means an increase in engine thrust and performance. On the basis of the existing 1500 ℃ magnitude, the thrust of the engine can be correspondingly improved by more than 10% when the inlet air temperature of the turbine is improved by 100 ℃. According to prediction, the turbine front temperature of the advanced aeroengine in the future is likely to exceed 1800 ℃, which is already greatly higher than the safe working temperature of the known high-temperature alloy and the temperature reduction capability of the conventional cooling measure. Therefore, the key point for developing the aero-engine with the high thrust-weight ratio is to solve the problem of insufficient temperature bearing capacity of the turbine blade. The leading edge portion of the turbine guide vane is directly impacted by the high-temperature, high-pressure and high-speed combustion gas from the combustion chamber, and is a portion relatively easy to be ablated, and the requirement for cooling is higher. At present, the conventional method for cooling the leading edge area of the turbine blade adopts a composite cooling mode of impact, convection and air film. The typical structure of the turbine blade is shown in fig. 1, namely, the turbine blade is designed in a hollow mode, an air film hole is processed at the front edge of the turbine blade, an impact hole is formed by a cold air guide pipe, convection heat transfer is enhanced by impacting a target surface through cooling air jet flow, and an air film is formed when the cooling air is discharged from the front edge to cover and isolate fuel gas for heating, so that the cooling effect is improved. However, with the further increase of the turbine inlet temperature of the next generation engine, such a structure exposes the disadvantages of low cooling effect and large amount of cold air consumption, and the requirements cannot be met, and the improvement of the blade cooling structure and the reduction of the amount of cold air are urgently needed.
Disclosure of Invention
Compared with high-temperature alloy, the ceramic matrix composite can safely work at higher environmental temperature, has the advantages of stable chemical property, corrosion and oxidation resistance, low heat conductivity coefficient and the like, and is a possible material for replacing the traditional high-temperature alloy to manufacture the turbine blade. However, it is difficult to form a blade because of the lack of strength of the CMC material compared to metals, combined with the large centrifugal and aerodynamic loads to which the turbine blade is subjected during operation. In order to solve the problem and realize the engineering application of the ceramic matrix composite material on the turbine blade, the invention provides the turbine blade adopting the bolt fixed ceramic armor. The structure scheme aims at the front edge area of the turbine blade, the ceramic matrix composite member can be quickly and stably connected with the metal matrix of the blade, and on the premise of not damaging the original aerodynamic shape of the blade, the front edge area of the turbine blade is effectively protected, the use of cooling gas is reduced, and the high temperature resistance of the blade and the thrust performance of an engine are improved. Meanwhile, after being damaged, the ceramic armor can be conveniently replaced in maintenance, and the metal matrix of the high-value blade can be repeatedly utilized to prolong the service life, so that the use and maintenance economy of the engine can be greatly improved.
The technical scheme of the invention is as follows:
as shown in figure 2, the invention is a turbine blade using bolt-fixed ceramic armor, which is composed of a ceramic matrix composite leading edge, a metal matrix and a fixed bolt. The main ceramic matrix composite material front edge comprises two connecting structures for fixing, namely a ceramic armor pin hole boss for connecting the front edge part and a blade body fixing buckle for fixing the pressure surface part and the suction surface part. When the blade is installed, on one hand, the blade body fixing buckle is inserted into the blade body fixing groove hole and rebounded to clamp and stop the blade body fixing groove hole for connection, on the other hand, the ceramic armor pin hole boss is inserted into the front edge fixing groove hole and then the fixing bolt is inserted for connection, and the front edge armor is firmly fixed on the blade metal matrix through the double connection.
The thickness of the ceramic armor pinhole boss becomes thinner along with the increase of the extension distance, and the formed pinhole boss wedge angle theta1Typical values of (a) range from 10 to 20. The structure of the blade body fixing buckle is in a shape of being bent along the chord direction of the blade profile, and the head of the buckle is raised and faces the outer side of the blade body. The boss and the buckle are integrated and smoothly connected.
The metal matrix is generally a duplex turbine guide vane and is formed by integrally and precisely casting high-temperature alloy. An internal cooling channel is reserved in the metal base body, and the heat of the blade body is taken away through cooling gas introduced from the outside. Slotted holes are reserved on the front edge and the blade body of the metal matrix and used for installing the fixing bolt and fixing the ceramic armor buckle.
The fixing bolt is also made of high-temperature alloy material and is a main stressed part, and the diameter phi D of the fixing bolt is1Typical value range of (A) is 1-2 mm. When in use, the ceramic armor and the metal matrix are connected and fixed by simultaneously inserting the pin holes on the ceramic armor and the metal matrix.
The invention has the beneficial effects that:
1. the temperature bearing capacity of the turbine guide blade and the thrust of the aircraft engine are improved.
Under the condition that the inflow temperature of the fuel gas is not changed, the ceramic matrix composite has lower heat conductivity coefficient and higher thermal resistance, so that the heat transferred from the fuel gas to the metal substrate of the blade can be greatly reduced, the temperature of the covered metal area is effectively reduced, and the temperature can be reduced by more than 100 ℃ compared with the conventional cooling state without ceramic armor, thereby playing an obvious protection role. The temperature variation of the blade in the direction perpendicular to the wall surface with and without the ceramic armor is shown in fig. 3. At this time, although the temperature state of the ceramic armor is higher, the ceramic armor still can be in a safe working range because the temperature resistance of the ceramic armor is far higher than that of metal. By utilizing the principle adopted by the invention, the working temperature of the turbine can be further improved, so that the blades are still in a bearable range, and the thrust performance of the engine is improved.
2. Having higher strength compared to full ceramic matrix composite blades
Although the cmc has superior temperature resistance, it is difficult to manufacture turbine blades with high requirements for aerodynamic and mechanical loads because of insufficient strength and toughness, which are significant disadvantages compared to metallic materials. Compared with the scheme of completely using the ceramic matrix composite material, the blade still uses the high-temperature alloy as a main stressed part on the main body, and the ceramic armor does not bear the pneumatic and mechanical loads of the blade and only needs basic mechanical properties and cold and hot fatigue resistance, so that the problem of poor strength and toughness of the ceramic matrix composite material is avoided. The solution of the invention can improve the integral strength of the blade well, and simultaneously increase the possibility of applying the ceramic matrix composite material component in the high thrust-weight ratio aeroengine.
3. Reducing engine manufacturing, use and maintenance costs
In existing turbine guide vanes based on superalloys and conventional cooling solutions, the leading edge of the vane is relatively susceptible to erosion. Once the blade is damaged by ablation, the whole blade cannot be used and needs to be replaced. In addition, the turbine blades are often made of expensive alloying elements and the manufacturing process is complicated, which results in high replacement (i.e., use and maintenance) costs for the engine. Thus, improvements in blade leading edge cooling performance have a direct impact and significance on turbine blade life growth and engine service maintenance economics. The ceramic matrix composite material front edge designed by the invention is fixed by adopting a bolt type and is designed in a detachable way. When the ceramic armor is damaged after being used for a long time or being injured by foreign objects, the damaged armor part can be detached and replaced on the premise of not damaging the metal matrix of the blade, so that the service life of the whole blade is prolonged, the manufacturing, using and maintaining costs of an engine are reduced, and the economical efficiency is improved.
Drawings
Fig. 1 is a schematic diagram of a conventional combined cooling method of impingement + convection + air film.
FIG. 2 is a schematic view of a turbine blade employing bolt-on ceramic armor.
FIG. 3 is a graph showing the temperature change in the thickness direction of the leading edge of the blade with or without the presence of the seizing.
FIG. 4 is a turbine blade configuration view of an offset pin style fixed ceramic armor.
In the figure: 1. a turbine blade; 2. a leading edge; 3. a gas film hole; 4. a cold air duct; 5. an impingement hole; 6. a ceramic matrix composite leading edge; 7. a metal substrate; 8. fixing the bolt; 9. a ceramic armor pin hole boss; 10. the blade body is fixed with a buckle; 11. blade body fixing slotted holes; 12. a leading edge fixing slot; 13. wedge angle theta of pin hole boss1(ii) a 14. An internal cooling channel; 15. diameter of the pin phi D1(ii) a 16. An offset square fixing bolt; 17. the side length L of the offset square fixing bolt.
Detailed Description
Example 1
In order that the present invention may be more readily and clearly understood, a more particular description of the invention briefly described above will be rendered by reference to specific embodiments that are illustrated in the appended drawings.
Referring to fig. 2, the present embodiment is a turbine blade using bolt-fixed ceramic armor, which is composed of a ceramic matrix composite leading edge 6, a metal matrix 7, and a fixed bolt 8. The ceramic matrix composite leading edge 6 is made of a ceramic matrix composite, and the metal matrix 7 and the fixing bolt 8 are made of high-temperature alloy. When in use, the ceramic armor and the metal matrix 7 are inserted into the pin holes at the same time, so that the ceramic armor and the metal matrix are connected and fixed. Diameter of the fixing pin phi D for smooth installation and sufficient firmness of the structure1The diameter is 1.6mm, and the orientation of the ceramic armor pin hole boss 9 is parallel to the gas direction. Wedge angle of pin hole bossθ1＝16°。
Example 2
As shown in fig. 4, this embodiment is a turbine blade using offset pin fixed ceramic armor, in this embodiment the fixing pin 8 is in an offset position. Wherein, the front edge 6 of the ceramic matrix composite material is partially provided with a half-square pin hole, and the metal matrix 7 is partially provided with a complete square pin hole. The offset square fixing bolt 16 is the main stressed part. In order to ensure the reliability, 17. the side length L of the offset square fixing bolt is 1.2mm, and the angle of the boss of the ceramic armor pin hole is parallel to the gas direction. Wedge angle side length theta of pin hole boss1＝12°。
Claims (3)
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110125484.2A CN112901279B (en) | 2021-01-29 | 2021-01-29 | Turbine blade adopting bolt-fixed ceramic armor |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110125484.2A CN112901279B (en) | 2021-01-29 | 2021-01-29 | Turbine blade adopting bolt-fixed ceramic armor |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN112901279A CN112901279A (en) | 2021-06-04 |
| CN112901279B true CN112901279B (en) | 2022-03-29 |
Family
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202110125484.2A Active CN112901279B (en) | 2021-01-29 | 2021-01-29 | Turbine blade adopting bolt-fixed ceramic armor |
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| Country | Link |
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| CN (1) | CN112901279B (en) |
Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3315941A (en) * | 1965-04-27 | 1967-04-25 | Rolls Royce | Aerofoil blade for use in a hot fluid stream |
| FR2165264A5 (en) * | 1971-12-23 | 1973-08-03 | Onera (Off Nat Aerospatiale) | Turbine blades - of synthetic resin incorporate reinforcing plate which is exposed at the leading edge |
| US3856434A (en) * | 1973-10-18 | 1974-12-24 | Westinghouse Electric Corp | Centrifugal fan wheel |
| JPS59215905A (en) * | 1983-05-25 | 1984-12-05 | Hitachi Ltd | Wear resistant bucket |
| CN1107934A (en) * | 1993-10-27 | 1995-09-06 | 西屋电气公司 | Gas turbine vane |
| KR101955858B1 (en) * | 2018-09-07 | 2019-03-08 | 국방과학연구소 | Flow path forming guide for hollow gas turbine blades and apparatus for manufacturing the flow path forming guide |
| CN110439629A (en) * | 2018-05-02 | 2019-11-12 | 通用电气公司 | With the CMC nozzle of interlocked mechanical connector and manufacture |
| CN110700898A (en) * | 2019-11-21 | 2020-01-17 | 中国科学院工程热物理研究所 | Ceramic-Metal Combined Turbine Guide Vane and Its Gas Turbine |
Family Cites Families (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US8167573B2 (en) * | 2008-09-19 | 2012-05-01 | Siemens Energy, Inc. | Gas turbine airfoil |
| CN211058869U (en) * | 2019-11-21 | 2020-07-21 | 中国科学院工程热物理研究所 | Ceramic-Metal Combined Turbine Guide Vane and Its Gas Turbine |
-
2021
- 2021-01-29 CN CN202110125484.2A patent/CN112901279B/en active Active
Patent Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3315941A (en) * | 1965-04-27 | 1967-04-25 | Rolls Royce | Aerofoil blade for use in a hot fluid stream |
| FR2165264A5 (en) * | 1971-12-23 | 1973-08-03 | Onera (Off Nat Aerospatiale) | Turbine blades - of synthetic resin incorporate reinforcing plate which is exposed at the leading edge |
| US3856434A (en) * | 1973-10-18 | 1974-12-24 | Westinghouse Electric Corp | Centrifugal fan wheel |
| JPS59215905A (en) * | 1983-05-25 | 1984-12-05 | Hitachi Ltd | Wear resistant bucket |
| CN1107934A (en) * | 1993-10-27 | 1995-09-06 | 西屋电气公司 | Gas turbine vane |
| CN110439629A (en) * | 2018-05-02 | 2019-11-12 | 通用电气公司 | With the CMC nozzle of interlocked mechanical connector and manufacture |
| KR101955858B1 (en) * | 2018-09-07 | 2019-03-08 | 국방과학연구소 | Flow path forming guide for hollow gas turbine blades and apparatus for manufacturing the flow path forming guide |
| CN110700898A (en) * | 2019-11-21 | 2020-01-17 | 中国科学院工程热物理研究所 | Ceramic-Metal Combined Turbine Guide Vane and Its Gas Turbine |
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| Publication number | Publication date |
|---|---|
| CN112901279A (en) | 2021-06-04 |
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