CN214063060U - Monocrystalline alloy turbine blade based on aeroengine is used - Google Patents
Monocrystalline alloy turbine blade based on aeroengine is used Download PDFInfo
- Publication number
- CN214063060U CN214063060U CN202022783148.0U CN202022783148U CN214063060U CN 214063060 U CN214063060 U CN 214063060U CN 202022783148 U CN202022783148 U CN 202022783148U CN 214063060 U CN214063060 U CN 214063060U
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- Prior art keywords
- turbine blade
- liquid nitrogen
- spring
- bending
- single crystal
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- 229910045601 alloy Inorganic materials 0.000 title claims abstract description 14
- 239000000956 alloy Substances 0.000 title claims abstract description 14
- IJGRMHOSHXDMSA-UHFFFAOYSA-N Atomic nitrogen Chemical compound N#N IJGRMHOSHXDMSA-UHFFFAOYSA-N 0.000 claims abstract description 58
- 239000007788 liquid Substances 0.000 claims abstract description 29
- 229910052757 nitrogen Inorganic materials 0.000 claims abstract description 29
- 239000013078 crystal Substances 0.000 claims abstract description 8
- 230000017525 heat dissipation Effects 0.000 claims abstract description 4
- 238000005452 bending Methods 0.000 claims description 26
- 230000007246 mechanism Effects 0.000 claims description 14
- 238000009423 ventilation Methods 0.000 claims description 5
- 239000007789 gas Substances 0.000 description 4
- 238000000034 method Methods 0.000 description 4
- 230000008261 resistance mechanism Effects 0.000 description 4
- 238000009434 installation Methods 0.000 description 3
- 238000001816 cooling Methods 0.000 description 2
- 239000002184 metal Substances 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 239000000446 fuel Substances 0.000 description 1
- 230000007774 longterm Effects 0.000 description 1
- 230000000149 penetrating effect Effects 0.000 description 1
- 239000012720 thermal barrier coating Substances 0.000 description 1
- 238000003466 welding Methods 0.000 description 1
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Abstract
The utility model discloses a single crystal alloy turbine blade based on aeroengine is used relates to turbine blade technical field, and the thermal diffusivity of having solved current turbine blade for aviation is mostly not fine, leads to the blade high temperature, the problem that blade life shortens. A monocrystalline alloy turbine blade for an aircraft engine comprises a turbine blade, wherein the turbine blade is crescent in shape, a heat dissipation hole is formed in the middle section of the left side wall of the turbine blade, and a hole groove is vertically formed in the upper end face of the turbine blade in a through mode; and the upper end surface of the turbine blade is provided with a hole groove, the hole grooves are internally provided with liquid nitrogen pipes in an inserted manner, and the liquid nitrogen pipes are provided with six hole grooves in an annular array. This practical liquid nitrogen intraductal liquid nitrogen that is equipped with can absorb the heat, reduces the temperature on turbine blade surface, avoids turbine blade high temperature and influences its life.
Description
Technical Field
The utility model relates to a turbine blade technical field specifically is a single crystal alloy turbine blade based on aeroengine is used.
Background
Turbine blades are important components of the turbine section of a gas turbine engine. The blades rotating at high speed are responsible for drawing high-temperature and high-pressure airflow into the combustor to maintain the operation of the engine. In order to ensure stable and long-term operation in extreme environments of high temperature and high pressure, turbine blades are often forged from high temperature alloys and are cooled in different ways, such as internal air flow cooling, boundary layer cooling, or thermal barrier coatings protecting the blades, to ensure operational reliability. In steam turbine engines and gas turbine engines, metal fatigue of the blades is the leading cause of engine failure. Strong vibration or resonance can lead to metal fatigue. Engineers often use friction dampers to reduce the damage caused by these factors to the blades, and turbofan engines are one of the aircraft engines, have low fuel consumption and excellent thrust characteristics in subsonic flight, and are widely used in modern passenger aircraft. The turbine blade under high load has very large leakage loss because of the very large pressure difference between the two sides of the turbine blade in the prior turbofan engine because of the gap between the casing and the tip of the turbine blade.
For example, patent No. CN109719445A discloses a turbine blade suitable for use in an aircraft engine, having a leading edge, a suction side, a pressure side and a trailing edge; the method is characterized in that: the suction surface of the blade is provided with at least two vortex generators which are vertically arranged; the vortex generator is an 1/4 oval plate-shaped structure, wherein the length of the long axis side L2 and the length of the short axis side L3 are respectively 7% and 5% of the chord length of the turbine blade, the thickness of the long axis side L2 and the length of the short axis side L3 are respectively 0.5% of the chord length of the turbine blade, and the long axis side L2 is fixed on the suction surface of the turbine blade.
The thermal diffusivity of current turbine blade for aviation is mostly not very good, leads to the blade high temperature, and blade life shortens, and in addition, when the blade rotated, its lateral wall was exposed in strong air current, and the lateral wall receives the influence of atmospheric pressure, easily takes place bending fracture's problem.
SUMMERY OF THE UTILITY MODEL
An object of the utility model is to provide a single crystal alloy turbine blade based on aeroengine is used to solve the current aeroengine turbine blade's that proposes in the above-mentioned background art thermal diffusivity for the aircraft mostly not very good, lead to the blade high temperature, blade life shortens, in addition, when the blade rotates, its lateral wall exposes in strong air current, and the lateral wall receives the influence of atmospheric pressure, easily takes place bending fracture's problem.
In order to achieve the above object, the utility model provides a following technical scheme: a monocrystalline alloy turbine blade for an aircraft engine comprises a turbine blade, wherein the turbine blade is crescent in shape, a heat dissipation hole is formed in the middle section of the left side wall of the turbine blade, and a hole groove is vertically formed in the upper end face of the turbine blade in a through mode; liquid nitrogen pipes are inserted and mounted in the hole grooves formed in the upper end face of the turbine blade, and six hole grooves are formed in the liquid nitrogen pipes in an annular array; the upside and the downside position of the left side inner wall of turbine blade all are provided with a backing plate, and equal stull is installed a department bending resistance mechanism on the right side wall of two backing plates.
Preferably, the liquid nitrogen pipe comprises a heat-conducting column sleeve and a ventilation cover, the heat-conducting column sleeve is sleeved and installed on the outer portion of the liquid nitrogen pipe, and six ventilation covers are embedded and installed in an upper side hole groove of the liquid nitrogen pipe.
Preferably, the backing plate comprises a fixing pin, and a fixing pin is welded on the left side wall of the backing plate and is inserted and installed with the turbine blade.
Preferably, the bending-resistant mechanism comprises gaskets and bolts, two gaskets are symmetrically arranged on the left side and the right side of the bending-resistant mechanism, and four bolts are symmetrically arranged on the two gaskets.
Preferably, the bending-resistant mechanism further comprises a spring bolt, a spring and a spring column sleeve, the spring bolt is inserted and installed at the center of the gasket at the right side, the spring is installed between the spring bolt and the gasket at the left side in a jacking mode, and the spring column sleeve is rotatably installed outside the spring.
Compared with the prior art, the beneficial effects of the utility model are that:
1. the utility model discloses the gas pocket of seting up on the turbine blade left side wall can give off the heat, reduces turbine blade's temperature, and the heat conduction column cover can give liquid nitrogen pipe with the heat transfer on turbine blade surface, is equipped with the liquid nitrogen in the liquid nitrogen pipe, can absorb the heat, reduces the temperature on turbine blade surface, avoids turbine blade high temperature and influences its life, and the backing plate passes through the fixed pin to be installed on turbine blade, and the installation can conveniently carry out the dismouting to the backing plate like this.
2. The utility model discloses turbine blade receives strong air current when extrudeing at the rotation in-process, and turbine blade's lateral wall bending under pressure, bending resistance mechanism can avoid turbine blade's lateral wall bending under pressure degree too big and take place to break, and when turbine blade bending under pressure, its lateral wall can roof pressure spring bolt, and spring bolt exerts reaction force for turbine blade, makes the crooked range of turbine blade diminish, makes turbine blade be difficult for taking place to buckle.
Drawings
FIG. 1 is a schematic structural view of the present invention;
FIG. 2 is a schematic view of the upper structure of the present invention;
FIG. 3 is a left side schematic view of the present invention;
FIG. 4 is a schematic view of the bending-resistant mechanism of the present invention;
FIG. 5 is a schematic view of the sectional structure of the anti-bending mechanism of the present invention;
FIG. 6 is an enlarged view of part A of FIG. 1;
in the drawings, the corresponding relationship between the component names and the reference numbers is as follows:
1. a turbine blade; 2. a liquid nitrogen pipe; 201. a heat-conducting column sleeve; 202. a ventilation cover; 3. a base plate; 301. a fixing pin; 4. a bending resistance mechanism; 401. a gasket; 402. a bolt; 403. a spring bolt; 404. a spring; 405. and a spring column sleeve.
Detailed Description
The technical solutions in the embodiments of the present invention will be described clearly and completely with reference to the accompanying drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only some embodiments of the present invention, not all embodiments.
Referring to fig. 1 to 6, the present invention provides an embodiment: a monocrystalline alloy turbine blade based on an aeroengine comprises a turbine blade 1, wherein the turbine blade 1 is crescent in shape, a heat dissipation hole is formed in the middle section of the left side wall of the turbine blade 1, and a hole groove is vertically formed in the upper end face of the turbine blade 1 in a penetrating mode; liquid nitrogen pipes 2 are inserted and installed in hole grooves formed in the upper end face of the turbine blade 1, and six hole grooves are formed in the liquid nitrogen pipes 2 in an annular array; the upside and the downside position of turbine blade 1's left side inner wall all are provided with one backing plate 3, equal stull is installed on the right side wall of two backing plates 3 and is located 4 liquid nitrogen pipes 2 of bending resistance mechanism and include heat conduction post cover 201 and ventilative lid 202, the outside of liquid nitrogen pipe 2 is cup jointed and is installed a heat conduction post cover 201, and inlay in the upside hole groove of liquid nitrogen pipe 2 and install six ventilative lids 202, the gas pocket of seting up on the 1 left side wall of turbine blade can give off the heat, reduce turbine blade 1's temperature, heat conduction post cover 201 can give liquid nitrogen pipe 2 with the heat transfer on turbine blade 1 surface, be equipped with the liquid nitrogen in the liquid nitrogen pipe 2, can absorb the heat, reduce the temperature on turbine blade 1 surface, avoid turbine blade 1 high temperature and influence its life.
Further, backing plate 3 includes fixed pin 301, and the welding has a fixed pin 301 on the left side wall of backing plate 3, and this fixed pin 301 is in the same place with turbine blade 1 grafting installation, and backing plate 3 passes through fixed pin 301 to be installed on turbine blade 1, and the installation can conveniently carry out the dismouting to backing plate 3 like this.
Further, the bending-resistant mechanism 4 comprises gaskets 401 and bolts 402, two gaskets 401 are symmetrically arranged on the left side and the right side of the bending-resistant mechanism 4, four bolts 402 are symmetrically arranged on the two gaskets 401 in the center, when the turbine blade 1 is extruded by strong air flow in the rotating process, the side wall of the turbine blade 1 is pressed and bent, and the bending-resistant mechanism 4 can prevent the side wall of the turbine blade 1 from being broken due to overlarge pressed and bent degree.
Further, the bending-resistant mechanism 4 further comprises a spring bolt 403, a spring 404 and a spring column sleeve 405, the center position of the right-side gasket 401 is provided with the spring bolt 403 in an inserting manner, a spring 404 is arranged between the spring bolt 403 and the left-side gasket 401 in an abutting manner, the outer part of the spring 404 is provided with the spring column sleeve 405 in a rotating manner, when the turbine blade 1 is bent under pressure, the side wall of the turbine blade 1 abuts against the spring bolt 403, the spring bolt 403 exerts a reaction force on the turbine blade 1, so that the bending amplitude of the turbine blade 1 is reduced, and the turbine blade 1 is not easy to bend.
The working principle is as follows: the air hole formed on the left side wall of the turbine blade 1 can radiate heat, the temperature of the turbine blade 1 is reduced, the heat conducting column sleeve 201 can transfer the heat on the surface of the turbine blade 1 to the liquid nitrogen pipe 2, liquid nitrogen is filled in the liquid nitrogen pipe 2 and can absorb the heat, the temperature on the surface of the turbine blade 1 is reduced, the influence on the service life of the turbine blade 1 due to overhigh temperature is avoided, the backing plate 3 is arranged on the turbine blade 1 through the fixing pin 301, the backing plate 3 can be conveniently disassembled and assembled, when the turbine blade 1 is extruded by strong airflow in the rotating process, the side wall of the turbine blade 1 is bent under pressure, the bending-resistant mechanism 4 can avoid the fracture of the side wall of the turbine blade 1 due to overlarge bending degree under pressure, when the turbine blade 1 is bent under pressure, the side wall of the spring bolt 403 can be pressed against the spring bolt 403, the spring bolt 403 can apply reaction force to the turbine blade 1, so that the bending amplitude of the turbine blade 1 is reduced, the turbine blade 1 is less likely to be bent.
It is obvious to a person skilled in the art that the invention is not restricted to details of the above-described exemplary embodiments, but that it can be implemented in other specific forms without departing from the spirit or essential characteristics of the invention. The present embodiments are therefore to be considered in all respects as illustrative and not restrictive, the scope of the invention being indicated by the appended claims rather than by the foregoing description, and all changes which come within the meaning and range of equivalency of the claims are therefore intended to be embraced therein. Any reference sign in a claim should not be construed as limiting the claim concerned.
Claims (5)
1. A single crystal alloy turbine blade for an aircraft engine, comprising: the turbine blade is characterized by comprising a turbine blade (1), wherein the turbine blade (1) is crescent in shape, a heat dissipation hole is formed in the middle section of the left side wall of the turbine blade (1), and a hole groove is vertically formed in the upper end face of the turbine blade (1) in a through mode; liquid nitrogen pipes (2) are inserted and installed in hole grooves formed in the upper end face of the turbine blade (1), and six hole grooves are formed in the liquid nitrogen pipes (2) in an annular array; the turbine blade is characterized in that a base plate (3) is arranged on the upper side and the lower side of the inner wall of the left side of the turbine blade (1), and a bending-resistant mechanism (4) is installed on the right side wall of the base plates (3) in two positions through a cross brace.
2. A single crystal alloy turbine blade for an aircraft engine based on according to claim 1, wherein: the liquid nitrogen pipe (2) comprises a heat-conducting column sleeve (201) and a ventilation cover (202), the heat-conducting column sleeve (201) is sleeved outside the liquid nitrogen pipe (2), and six ventilation covers (202) are embedded in an upper side hole groove of the liquid nitrogen pipe (2).
3. A single crystal alloy turbine blade for an aircraft engine based on according to claim 1, wherein: the backing plate (3) comprises a fixing pin (301), the left side wall of the backing plate (3) is welded with the fixing pin (301), and the fixing pin (301) is spliced with the turbine blade (1) and installed together.
4. A single crystal alloy turbine blade for an aircraft engine based on according to claim 1, wherein: the bending-resistant mechanism (4) comprises gaskets (401) and bolts (402), the left side and the right side of the bending-resistant mechanism (4) are symmetrically provided with two gaskets (401), and four bolts (402) are symmetrically arranged on the two gaskets (401).
5. A single crystal alloy turbine blade for an aircraft engine according to claim 4, wherein: the bending-resistant mechanism (4) further comprises a spring bolt (403), a spring (404) and a spring column sleeve (405), the spring bolt (403) is installed at the center of the gasket (401) on the right side in an inserting mode, the spring (404) is installed between the spring bolt (403) and the gasket (401) on the left side in a pressing mode, and the spring column sleeve (405) is installed on the outer portion of the spring (404) in a rotating mode.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202022783148.0U CN214063060U (en) | 2020-11-27 | 2020-11-27 | Monocrystalline alloy turbine blade based on aeroengine is used |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202022783148.0U CN214063060U (en) | 2020-11-27 | 2020-11-27 | Monocrystalline alloy turbine blade based on aeroengine is used |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN214063060U true CN214063060U (en) | 2021-08-27 |
Family
ID=77403709
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202022783148.0U Expired - Fee Related CN214063060U (en) | 2020-11-27 | 2020-11-27 | Monocrystalline alloy turbine blade based on aeroengine is used |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN214063060U (en) |
-
2020
- 2020-11-27 CN CN202022783148.0U patent/CN214063060U/en not_active Expired - Fee Related
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| Date | Code | Title | Description |
|---|---|---|---|
| GR01 | Patent grant | ||
| GR01 | Patent grant | ||
| CF01 | Termination of patent right due to non-payment of annual fee |
Granted publication date: 20210827 |
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| CF01 | Termination of patent right due to non-payment of annual fee |