CN107813114A - The processing method that a kind of aero engine turbine blades remove remelted layer air film hole - Google Patents
The processing method that a kind of aero engine turbine blades remove remelted layer air film hole Download PDFInfo
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- CN107813114A CN107813114A CN201711259799.6A CN201711259799A CN107813114A CN 107813114 A CN107813114 A CN 107813114A CN 201711259799 A CN201711259799 A CN 201711259799A CN 107813114 A CN107813114 A CN 107813114A
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- remelted layer
- air film
- film hole
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- blade
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
- B23P—METAL-WORKING NOT OTHERWISE PROVIDED FOR; COMBINED OPERATIONS; UNIVERSAL MACHINE TOOLS
- B23P15/00—Making specific metal objects by operations not covered by a single other subclass or a group in this subclass
- B23P15/02—Making specific metal objects by operations not covered by a single other subclass or a group in this subclass turbine or like blades from one piece
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
- Laser Beam Processing (AREA)
Abstract
The processing method that a kind of aero engine turbine blades remove remelted layer air film hole, belong to aero engine turbine blades processing technique field, including：Turbo blade compound cooling structure designs, and the processing method of ultrafast laser ring cutting and helical scanning carries out air film hole machined, and abrasive Flow Machining and crucible zone remove solution and remove remelted layer.The present invention carries out air film hole machined using ultrafast laser ring cutting and the processing method of helical scanning, can substantially reduce fuel factor, improve surface integrity, and air film hole bur, groove, the micro-crack processed is few, and remelted layer maximum gauge is not more than 25 μm.Abrasive particle is equably ground to passage surface or corner under pressure, deburring, chamfering, light decorations can also be played a part of by being not only able to removal surplus, remelted layer can also effectively be cut down, it is 0.02mm 0.04mm to control aperture incrementss, can effectively control remelted layer and micro-crack.Solution is removed using remelted layer and further removes remelted layer, air film hole import and export is can reach and is remained without remelted layer.
Description
Technical field
The invention belongs to aero engine turbine blades processing technique field, and in particular to a kind of aero-engine
The processing method that turbo blade removes remelted layer air film hole.
Background technology
Turbo blade is the part of thermic load and mechanical load maximum in aero-engine, and its working environment is severe,
High-temperature high-pressure fuel gas impact after bearing burning, and its manufacturing technology is listed in the key technology of modern aeroengine
Wherein blade foundry engieering and air film hole process technology are most important difficult points.At present, the process technology of blade air film hole is main
The methods of including Long Pulse LASER punching, electric spark-erosion perforation, the punching of electro-hydraulic beam.Wherein first two method belongs to hot melt processing,
Re cast layer and micro-crack can be produced in hole wall, the surface integrity of serious shadow blade, ultimately result in blade material performance and peace
The decline of full service life.
The content of the invention
It is good the invention provides a kind of surface integrity in order to solve problems of the prior art, burr, groove,
Micro-crack is few, and the processing method that the aero engine turbine blades without remelted layer remove remelted layer air film hole.
The present invention uses following technical scheme：
The processing method that a kind of aero engine turbine blades remove remelted layer air film hole, comprises the following steps：
Step 1：Turbo blade employs orientation high temperature casting alloy, and blade precision is cast without surplus, and tenon employs three
To fir-shape tenon, blade inner chamber employs U-typed backflow convection current cooling and blade inlet edge, the Compound cooling of trailing edge gaseous film control
Structure；
Step 2：The datum hole that aperture is 50 μm is processed with femtosecond laser, remelted layer thickness control is below 40 μm；
Step 3：With the direct punching 1s of Long Pulse LASER, a diameter of 300-310 μm of positive cone hole is formed, reuses four
Wedge scanning means ring cutting scans 2-3s, forms a diameter of 350-360 μm of cylindrical hole；
Step 4：With 500fs laser helical scanning 15s, its technological parameter is：Sweep speed is 2400r/min, Duplication
For 12%, the amount of feeding is 5 μm, repetition rate 20kHz, and 0.6Pa coaxially blows；
Step 5：Using abrasive Flow lathe carry out abrasive Flow Machining, by turbo blade be fixed on two abrasive particle rainbow containers it
Passage in, by rainbow inner carrier team squeeze 5 abrasive materials flow back and forth, select B500 graininess abrasive materials, operating pressure 6-7mpa,
Aperture incrementss control is in 0.02mm-0.04mm；
Step 6：Oil removing is carried out with aqueous cleaning agent to turbine blade surface, and with moisture film long run test method check table
Face deoiling effect, ensure that turbine blade surface oil removing is complete, then dried up；
Step 7：Prepare remelted layer and remove solution, the formula of solution is the volume integral of hydrogen peroxide+hydrochloric acid, wherein hydrogen peroxide
Number 30%-60%, surplus is hydrochloric acid, by hydrogen peroxide and mixed in hydrochloric acid and is stirred；
Step 8：Turbo blade is put into container, can not stack or cover between each other, remelted layer removal solution is fallen
Enter container and be totally submerged turbo blade, taken out after stirring, immersion 6-10min；
Step 9：Turbo blade is immediately placed in the circulating water of cleaning and rinsed well, and dried.
Preferably, the abrasive Flow lathe described in step 5 is MLL60D type abrasive Flow lathes.
Preferably, dried up in step 6 using hair-dryer.
Preferably, hydrogen peroxide and hydrochloric acid are that chemistry is pure in step 7.
Preferably, remelted layer removal solution uses after preparation in 60min in step 7.
The beneficial effects of the present invention are：
1) present invention carries out air film hole machined using ultrafast laser ring cutting and the processing method of helical scanning, can drop significantly
Low thermal effect, surface integrity is improved, air film hole bur, groove, the micro-crack processed is few, and remelted layer maximum gauge is not more than
25μm。
2) turbo blade inner chamber employs U-typed backflow convection current cooling and blade inlet edge, the composite cold of trailing edge gaseous film control
But structure, heat transfer of the combustion gas to turbo blade is reduced.
3) abrasive particle is equably ground to passage surface or corner under pressure, and being not only able to removal surplus can also
Play a part of deburring, chamfering, light decorations, abrasive Flow method can also effectively cut down remelted layer.In abrasive Flow Machining process
Middle to control aperture incrementss be 0.02mm-0.04mm, can effectively control remelted layer and micro-crack.
4) solution is removed using remelted layer and further removes remelted layer, can reach air film hole import and export and remained without remelted layer.
Brief description of the drawings
The turbine blade cooling that Fig. 1 is the present invention leads to schematic diagram.
Embodiment
With reference to embodiments, the technical scheme in the present invention is clearly and completely described.Based in the present invention
Embodiment, the every other embodiment that those of ordinary skill in the art are obtained under the premise of creative work is not made, all
Belong to the scope of protection of the invention.
The processing method that a kind of aero engine turbine blades remove remelted layer air film hole, comprises the following steps：
Step 1：Turbo blade employs orientation high temperature casting alloy, and blade precision is cast without surplus, and tenon employs three
To fir-shape tenon, blade inner chamber employs U-typed backflow convection current cooling and blade inlet edge, the Compound cooling of trailing edge gaseous film control
Structure, turbine blade cooling passage such as Fig. 1, cooling air are divided into two-way into behind blade inner chamber from blade tenon bottom, inside
Chamber, respectively from blade inlet edge air film hole and blade trailing edge exhaust seam discharge, while take away heat, reduces blade through U-typed passage
Own temperature, and from blade inlet edge air film hole discharge cooling air after air film hole is discharged, blade basin, back surface implement
Convection current cools down and forms one layer of air film along type face, reduces heat transfer of the combustion gas to blade；
Step 2：The datum hole that aperture is 50 μm is processed with femtosecond laser, remelted layer thickness control is below 40 μm；
Step 3：With the direct punching 1s of Long Pulse LASER, a diameter of 300-310 μm of positive cone hole is formed, reuses four
Wedge scanning means ring cutting scans 2-3s, forms a diameter of 350-360 μm of cylindrical hole；
Step 4：With 500fs laser helical scanning 15s, its technological parameter is：Sweep speed is 2400r/min, Duplication
For 12%, the amount of feeding is 5 μm, repetition rate 20kHz, and 0.6Pa coaxially blows；
Step 5：Using abrasive Flow lathe carry out abrasive Flow Machining, by turbo blade be fixed on two abrasive particle rainbow containers it
Passage in, by rainbow inner carrier team squeeze 5 abrasive materials flow back and forth, select B500 graininess abrasive materials, operating pressure 6-7mpa,
Aperture incrementss control is in 0.02mm-0.04mm；
Step 6：Oil removing is carried out with aqueous cleaning agent to turbine blade surface, and with moisture film long run test method check table
Face deoiling effect, ensure that turbine blade surface oil removing is complete, then dried up；
Step 7：Prepare remelted layer and remove solution, the formula of solution is the volume integral of hydrogen peroxide+hydrochloric acid, wherein hydrogen peroxide
Number 30%-60%, surplus is hydrochloric acid, by hydrogen peroxide and mixed in hydrochloric acid and is stirred；
Step 8：Turbo blade is put into container, can not stack or cover between each other, remelted layer removal solution is fallen
Enter container and be totally submerged turbo blade, taken out after stirring, immersion 6-10min；
Step 9：Turbo blade is immediately placed in the circulating water of cleaning and rinsed well, and dried.
Abrasive Flow lathe described in step 5 is MLL60D type abrasive Flow lathes.
Drying is using hair-dryer in described step six.
Hydrogen peroxide and hydrochloric acid are that chemistry is pure in described step seven.
Remelted layer removes solution and used after preparation in 60min in described step seven.
The present invention carries out air film hole machined using ultrafast laser ring cutting and the processing method of helical scanning, can substantially reduce
Fuel factor, surface integrity is improved, air film hole bur, groove, the micro-crack processed is few, and remelted layer maximum gauge is not more than 25
μm.Turbo blade inner chamber employs U-typed backflow convection current cooling and blade inlet edge, the compound cooling structure of trailing edge gaseous film control,
Reduce heat transfer of the combustion gas to turbo blade.Abrasive particle is equably ground to passage surface or corner under pressure, not only
Surplus, which can be removed, can also play a part of deburring, chamfering, light decorations, and abrasive Flow method can also effectively cut down remelted layer.
It is 0.02mm-0.04mm that aperture incrementss are controlled during abrasive Flow Machining, can effectively control remelted layer and micro-crack.
Solution is removed using remelted layer and further removes remelted layer, air film hole import and export is can reach and is remained without remelted layer.
Claims (5)
1. the processing method that a kind of aero engine turbine blades remove remelted layer air film hole, it is characterised in that comprise the following steps：
Step 1：Turbo blade employs orientation high temperature casting alloy, and blade precision is cast without surplus, and tenon employs three pairs of firs
Tree-like tenon tooth, blade inner chamber employ U-typed backflow convection current cooling and blade inlet edge, the Compound cooling knot of trailing edge gaseous film control
Structure；
Step 2：The datum hole that aperture is 50 μm is processed with femtosecond laser, remelted layer thickness control is below 40 μm；
Step 3：With the direct punching 1s of Long Pulse LASER, a diameter of 300-310 μm of positive cone hole is formed, reuses four wedges
Scanning means ring cutting scans 2-3s, forms a diameter of 350-360 μm of cylindrical hole；
Step 4：With 500fs laser helical scanning 15s, its technological parameter is：Sweep speed is 2400r/min, and Duplication is
12%, the amount of feeding is 5 μm, repetition rate 20kHz, 0.6Pa and coaxially blown；
Step 5：Abrasive Flow Machining is carried out using abrasive Flow lathe, by turbo blade be fixed on two abrasive particle rainbow containers it is logical
In road, 5 abrasive materials are squeezed by rainbow inner carrier team and flowed back and forth, select B500 graininess abrasive materials, operating pressure 6-7mpa, aperture
Incrementss are controlled in 0.02mm-0.04mm；
Step 6：Oil removing is carried out with aqueous cleaning agent to turbine blade surface, and examines surface to remove in moisture film long run test method
Oily effect, ensure that turbine blade surface oil removing is complete, then dried up；
Step 7：Prepare remelted layer and remove solution, the formula of solution is the volume fraction of hydrogen peroxide+hydrochloric acid, wherein hydrogen peroxide
30%-60%, surplus are hydrochloric acid, by hydrogen peroxide and mixed in hydrochloric acid and are stirred；
Step 8：Turbo blade is put into container, can not stack or cover between each other, remelted layer removal solution is poured into appearance
Device is simultaneously totally submerged turbo blade, is taken out after stirring, immersion 6-10min；
Step 9：Turbo blade is immediately placed in the circulating water of cleaning and rinsed well, and dried.
2. the processing method that a kind of aero engine turbine blades according to claim 1 remove remelted layer air film hole, it is special
Sign is：Abrasive Flow lathe described in step 5 is MLL60D type abrasive Flow lathes.
3. the processing method that a kind of aero engine turbine blades according to claim 1 remove remelted layer air film hole, it is special
Sign is：Drying is using hair-dryer in step 6.
4. the processing method that a kind of aero engine turbine blades according to claim 1 remove remelted layer air film hole, it is special
Sign is：Hydrogen peroxide and hydrochloric acid are that chemistry is pure in step 7.
5. the processing method that a kind of aero engine turbine blades according to claim 1 remove remelted layer air film hole, it is special
Sign is：Remelted layer removes solution and used after preparation in 60min in step 7.
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| CN201711259799.6A CN107813114A (en) | 2017-12-04 | 2017-12-04 | The processing method that a kind of aero engine turbine blades remove remelted layer air film hole |
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| CN201711259799.6A CN107813114A (en) | 2017-12-04 | 2017-12-04 | The processing method that a kind of aero engine turbine blades remove remelted layer air film hole |
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Cited By (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN108868895A (en) * | 2018-06-08 | 2018-11-23 | 南京赛达机械制造有限公司 | A kind of high strength titanium alloy blade of aviation engine |
| CN110202277A (en) * | 2019-04-25 | 2019-09-06 | 青岛理工大学 | Aeroengine blade air film hole machining device and working method |
| CN110842751A (en) * | 2019-11-28 | 2020-02-28 | 中国航发沈阳黎明航空发动机有限责任公司 | Method for rounding orifice of guide vane air film hole |
| CN111761149A (en) * | 2020-06-24 | 2020-10-13 | 中国航发北京航空材料研究院 | Method for eliminating high-temperature alloy electric spark hole-making hole wall remelted layer |
| CN112008262A (en) * | 2020-07-30 | 2020-12-01 | 华东师范大学 | Method for intelligently machining special-shaped hole by annular rotating laser |
| CN113909600A (en) * | 2021-10-09 | 2022-01-11 | 中国航发北京航空材料研究院 | Quality evaluation method for turbine blade electrosparking gas film hole |
| CN114083146A (en) * | 2021-10-27 | 2022-02-25 | 中国航发北京航空材料研究院 | Ultrafast laser processing method for double-wall ultra-air cooling turbine blade air film cooling tank |
| CN117139752A (en) * | 2023-10-26 | 2023-12-01 | 中国航发沈阳黎明航空发动机有限责任公司 | Control method for hole making of gas film holes of turbine working blades without remelting layer |
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Cited By (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN108868895A (en) * | 2018-06-08 | 2018-11-23 | 南京赛达机械制造有限公司 | A kind of high strength titanium alloy blade of aviation engine |
| CN110202277A (en) * | 2019-04-25 | 2019-09-06 | 青岛理工大学 | Aeroengine blade air film hole machining device and working method |
| CN110842751A (en) * | 2019-11-28 | 2020-02-28 | 中国航发沈阳黎明航空发动机有限责任公司 | Method for rounding orifice of guide vane air film hole |
| CN111761149A (en) * | 2020-06-24 | 2020-10-13 | 中国航发北京航空材料研究院 | Method for eliminating high-temperature alloy electric spark hole-making hole wall remelted layer |
| CN111761149B (en) * | 2020-06-24 | 2022-05-13 | 中国航发北京航空材料研究院 | Method for eliminating single crystal high temperature alloy electric spark hole-making hole wall remelted layer |
| CN112008262A (en) * | 2020-07-30 | 2020-12-01 | 华东师范大学 | Method for intelligently machining special-shaped hole by annular rotating laser |
| CN113909600A (en) * | 2021-10-09 | 2022-01-11 | 中国航发北京航空材料研究院 | Quality evaluation method for turbine blade electrosparking gas film hole |
| CN114083146A (en) * | 2021-10-27 | 2022-02-25 | 中国航发北京航空材料研究院 | Ultrafast laser processing method for double-wall ultra-air cooling turbine blade air film cooling tank |
| CN114083146B (en) * | 2021-10-27 | 2024-04-09 | 中国航发北京航空材料研究院 | Ultra-fast laser processing method for double-wall ultra-air cooling turbine blade air film cooling groove |
| CN117139752A (en) * | 2023-10-26 | 2023-12-01 | 中国航发沈阳黎明航空发动机有限责任公司 | Control method for hole making of gas film holes of turbine working blades without remelting layer |
| CN117139752B (en) * | 2023-10-26 | 2024-01-16 | 中国航发沈阳黎明航空发动机有限责任公司 | Control method for hole making of gas film holes of turbine working blades without remelting layer |
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