CN207248535U - Gas turbine blade end wall heat transfer test system - Google Patents
Gas turbine blade end wall heat transfer test system Download PDFInfo
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- CN207248535U CN207248535U CN201721100744.6U CN201721100744U CN207248535U CN 207248535 U CN207248535 U CN 207248535U CN 201721100744 U CN201721100744 U CN 201721100744U CN 207248535 U CN207248535 U CN 207248535U
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- 238000012360 testing method Methods 0.000 title claims abstract description 39
- 238000012546 transfer Methods 0.000 title claims description 38
- 238000001816 cooling Methods 0.000 claims abstract description 36
- 239000007789 gas Substances 0.000 claims abstract description 36
- XLYOFNOQVPJJNP-UHFFFAOYSA-N water Substances O XLYOFNOQVPJJNP-UHFFFAOYSA-N 0.000 claims abstract description 16
- 239000002737 fuel gas Substances 0.000 claims abstract description 15
- 230000008646 thermal stress Effects 0.000 claims abstract description 15
- 238000005259 measurement Methods 0.000 claims abstract description 13
- 239000000498 cooling water Substances 0.000 claims abstract description 9
- UGFAIRIUMAVXCW-UHFFFAOYSA-N Carbon monoxide Chemical compound [O+]#[C-] UGFAIRIUMAVXCW-UHFFFAOYSA-N 0.000 claims abstract description 4
- 239000003546 flue gas Substances 0.000 claims abstract description 4
- 230000008602 contraction Effects 0.000 claims description 8
- 230000005540 biological transmission Effects 0.000 claims description 7
- 239000011888 foil Substances 0.000 claims description 6
- 230000001105 regulatory effect Effects 0.000 claims description 6
- 241000219289 Silene Species 0.000 claims description 5
- 229910052918 calcium silicate Inorganic materials 0.000 claims description 5
- 239000000463 material Substances 0.000 claims description 5
- 230000003750 conditioning effect Effects 0.000 claims description 4
- 238000013016 damping Methods 0.000 claims description 4
- 230000008676 import Effects 0.000 claims description 4
- 230000030279 gene silencing Effects 0.000 abstract 2
- 238000007599 discharging Methods 0.000 abstract 1
- 238000002485 combustion reaction Methods 0.000 description 12
- 238000011144 upstream manufacturing Methods 0.000 description 8
- 238000010586 diagram Methods 0.000 description 6
- 238000009826 distribution Methods 0.000 description 5
- 239000000567 combustion gas Substances 0.000 description 4
- 238000002474 experimental method Methods 0.000 description 4
- 239000000853 adhesive Substances 0.000 description 3
- 230000001070 adhesive effect Effects 0.000 description 3
- 230000008859 change Effects 0.000 description 2
- 238000013461 design Methods 0.000 description 2
- 238000011161 development Methods 0.000 description 2
- 238000000034 method Methods 0.000 description 2
- 238000011160 research Methods 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 239000011230 binding agent Substances 0.000 description 1
- 230000033228 biological regulation Effects 0.000 description 1
- 230000007797 corrosion Effects 0.000 description 1
- 238000005260 corrosion Methods 0.000 description 1
- 230000006837 decompression Effects 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 230000006866 deterioration Effects 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000005538 encapsulation Methods 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 238000013401 experimental design Methods 0.000 description 1
- 238000009422 external insulation Methods 0.000 description 1
- 239000012530 fluid Substances 0.000 description 1
- 239000003292 glue Substances 0.000 description 1
- 238000010438 heat treatment Methods 0.000 description 1
- 230000006872 improvement Effects 0.000 description 1
- 239000012535 impurity Substances 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 239000000203 mixture Substances 0.000 description 1
- 238000012544 monitoring process Methods 0.000 description 1
- 238000005457 optimization Methods 0.000 description 1
- 239000011505 plaster Substances 0.000 description 1
- 238000012805 post-processing Methods 0.000 description 1
- 230000008569 process Effects 0.000 description 1
- 238000012216 screening Methods 0.000 description 1
- 239000007921 spray Substances 0.000 description 1
- 230000006641 stabilisation Effects 0.000 description 1
- 238000011105 stabilization Methods 0.000 description 1
- 230000035882 stress Effects 0.000 description 1
- 238000001931 thermography Methods 0.000 description 1
- 230000001052 transient effect Effects 0.000 description 1
- 230000007704 transition Effects 0.000 description 1
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- Testing Of Devices, Machine Parts, Or Other Structures Thereof (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A heat exchange test system for the end wall of a turbine blade of a gas turbine comprises a blade grid test section and the like, wherein the turbine blade for measurement is arranged in the blade grid test section; the high-temperature gas supply system is communicated with the blade grid test section and used for supplying high-temperature gas to the blade grid test section, and the secondary air supply system is used for providing high-pressure and temperature-adjustable secondary cooling air to the blade grid test section; the cooling water system is divided into a high-pressure water system and a low-pressure water system, the high-pressure water system is used for cooling high-temperature flue gas, and the low-pressure water system is used for cooling an air compressor in the secondary air supply system; the measuring and data collecting system comprises a flow parameter measuring device and a thermal stress measuring device, wherein the flow parameter measuring device is used for measuring the temperature, the pressure and the flow in the cascade test section, and the thermal stress measuring device is used for measuring the thermal stress of the turbine blade; the exhaust system comprises an exhaust adjusting system and a silencing system, wherein the exhaust adjusting system is used for adjusting the exhaust pressure of the cascade test section and discharging high-temperature fuel gas after being cooled through the silencing system.
Description
Technical field
It the utility model is related to a kind of gas turbine turbine blade endwall heat transfer pilot system.
Background technology
It is to improve the important channel of gas turbine cycle power and the thermal efficiency to improve gas turbine turbine import initial temperature, at present
The inlet temperature of advanced heavy duty gas turbine turbine reaches 1600 DEG C or so.
Gas turbine first order nozzle vane end wall be exposed to the high turbulent flow combustion gas of high temperature under, blade tip subject compared with
High thermic load and wall surface temperature.Simultaneously because there are the Secondary Flow of the complexity such as end horse shoe vortex and Passage Vortex, high-temperature fuel gas with
Heat exchange between end wall is seriously uneven, and the heat transfer of blade tip regional area is strengthened.Therefore, combustor exit and turbine
First stage stator blades end wall composition surface answers transition naturally smooth, heat transfer deterioration caused by reduce Secondary Flow and aerodynamic loss.
Due to the presence of build-up tolerance and the difference of degree of thermal expansion, gas-turbine combustion chamber outlet and the engagement of stator blade end wall
Unavoidably there is dislocation in face, this misplace brings the change of blade endwall structure, development and end heat transfer feelings to Secondary Flow
Condition brings direct influence.At the same time in order to avoid directly impact, usually in front of the blade edge upstream of the high-temperature fuel gas to blade end wall
Placement of discrete air film hole or other combination methods for cooling, are realized to the effective film overcast of end wall surface, and cold air goes out flow arrangement side
Formula is also related to Secondary Flow development.Therefore, it is endwall heat transfer research that reasonable air film layout is studied under real blade endwall structure
Emphasis.
There are the following problems for country's turbine endwall heat transfer pilot system at present：Endwall heat transfer test measurement base most of first
In the condition of adiabatic hypothesis, test temperature is relatively low, its endwall heat transfer rule obtained and actual combustion engine turbine endwall heat transfer rule
There is certain difference；Secondly endwall heat transfer experiment is less considers actual combustion engine combustion chamber and first stage stator blades end wall junction geometry
Mismatch the influence to end wall heat transfer characteristic；Aerodynamic heat transfer effect is only considered in heat transfer experiments before again, not in blade
End wall and surface set thermal stress measuring point.
Utility model content
The purpose of this utility model is that needed for existing above-mentioned gas turbine turbine endwall heat transfer system defect or improvement
Ask, there is provided a kind of gas turbine turbine blade endwall heat transfer pilot system, to meet the experiment of steady state heat transfer and Transient Heat Transfer
It is required that.The system is used as cascade test section, experimental condition and the actual fortune of gas turbine by the use of gas turbine turbine first stage stator blades
Row condition is similar.
To achieve these goals, the utility model uses following technical scheme：
A kind of gas turbine turbine blade endwall heat transfer pilot system, including high-temperature fuel gas feed system, auxiliary air supply
To system, cascade test section, cooling water system, measurement and data collecting system and exhaust system；Wherein,
Turbine blade for measuring is set in cascade test section, transmission window, bottom are offered at the top of cascade test section
Portion is provided with cooling air chamber；
High-temperature fuel gas feed system is connected with cascade test section, secondary for supplying high-temperature fuel gas to cascade test section
Air supply system is used to provide high pressure and the adjustable secondary cooling air of temperature to cascade test section by cooling air chamber；
Cooling water system is divided into high pressure water system and low pressure water system, and high pressure water system is used for cooling high temperature flue gas, low pressure
Water system is used to cool down the air compressor machine in secondary air supply；
Measurement and data collecting system include flow parameter measurement device and thermal stress measuring device, wherein, flow parameter
Measuring device is used to measure temperature, pressure and the flow in cascade test section, and thermal stress measuring device is used to measure turbine blade
Thermal stress；
Exhaust system includes exhaust conditioning system and silene system, and exhaust conditioning system is used to adjust the exhaust of cascade test section
Pressure, and discharged after high-temperature fuel gas is cooled down through silene system.
The utility model, which further improves, to be, high-temperature fuel gas feed system includes sequentially connected expansion segment, stabilization
Section, contraction section, and it is arranged on the turbulent flow generator of contraction section.
The utility model, which further improves, to be, secondary air supply includes sequentially connected air compressor machine, backheat
Device and electric heater.
The utility model, which further improves, to be, secondary air supply, which further includes, is arranged on air compressor machine inlet
Filter, is arranged on the high pressure tank that air compressor machine exit is used to store compressed gas, for adjusting the pressure regulation of compressed air
Valve, mass flowmenter and regulating valve, and the substation for powering for air compressor machine.
The utility model, which further improves, to be, the combustor exit end wall and the turbine first order in contraction section exit are quiet
End of blade wall import junction smoothly transits or there are dislocation.
The utility model, which further improves, to be, cooling water system includes feed pump, and be connected with feed pump and oneself
The cooling tower and cistern set gradually under above.
The utility model, which further improves, to be, flow parameter measurement device includes being arranged on red above transmission window
Outer thermography system, the thermocouple and pressure-measuring-point being arranged on turbine blade end wall, thermal stress measuring device include being arranged on
High temperature resistant foil gauge on flat blade, and flow parameter measurement device and thermal stress measuring device connect with data collecting system
Connect.
The utility model, which further improves, to be, silene system is sound damping tower.
The utility model, which further improves, to be, the end wall of turbine blade offers film cooling holes.
The utility model, which further improves, to be, cascade test section exit is provided with heat-barrier material.
Relative to existing endwall heat transfer pilot system, the utility model has following beneficial effect：
Gas turbine turbine blade endwall heat transfer pilot system provided by the utility model, the blade used is actual combustion engine
Turbine first stage stator blades, operating condition of test are combustion engine operating condition, and the temperature and heat that can more actually reflect turbine end wall surface should
Power is distributed.
Gas turbine turbine blade endwall heat transfer pilot system provided by the utility model, it is contemplated that actual combustion engine combustion chamber
Mismatched with first stage stator blades end wall junction geometry, and influence of the gaseous film control to end wall heat-transfer character, before this is to research
The film cooling holes structure optimization of edge upstream end wall and arrangement have larger meaning.
Gas turbine turbine blade endwall heat transfer pilot system provided by the utility model, can not only directly measure blade table
The Temperature Distribution in face, and the thermal stress distribution of end wall and blade surface is measured, can be that the design of gas turbine turbine blade is excellent
Change and experimental data is provided.
Brief description of the drawings
Fig. 1 is gas turbine turbine blade endwall heat transfer pilot system arrangement schematic diagram；
Fig. 2 is the cascade test section enlarged drawing shown in Fig. 1；
Fig. 3 is not consider upstream film cooling holes and the unmatched end wall platform schematic diagram of geometry；
Fig. 4 is the end wall platform schematic diagram for only considering geometry mismatch；
Fig. 5 is the end wall platform schematic diagram for only considering upstream film cooling holes；
Fig. 6 is to consider upstream film cooling holes and the unmatched end wall platform schematic diagram of geometry；
Fig. 7 is the layout drawing of blade surface and end wall high temperature resistant foil gauge.
In figure：1st, filter；2nd, air compressor machine；3rd, high pressure tank；4th, pressure regulator valve；5th, mass flowmenter；6th, regulating valve；7、
Regenerator；8th, electric heater；9th, cooling air chamber；10th, cascade test section；11st, turbine stator blade；12nd, expansion segment；13rd, stablize
Section；14th, contraction section；15th, turbulent flow generator；16th, thermocouple；17th, infrared thermal imagery system；18th, pressure-measuring-point；19th, high temperature resistant should
Become piece；20th, data collecting system；21st, cooling tower；22nd, cistern；23rd, feed pump；24th, substation；25th, sound damping tower；26th, gas
Film cooling hole；27th, blade ante-chamber；28th, blade back cavity；29th, first stage stator blades end wall；30th, combustor exit end wall；31st, it is heat-insulated
Material；32nd, transmission window.
Embodiment
Below in conjunction with drawings and examples, the utility model is described in further detail.
As shown in Figure 1, a kind of gas turbine turbine blade endwall heat transfer pilot system of the utility model, including high temperature combustion
Gas feed system, secondary air supply, cascade test section 10, cooling water system, measurement and data collecting system, exhaust system
System etc..
Referring to Fig. 1, start air compressor machine 2, cooling air passes through 1 impurity screening of filter, is stored in after the pressurization of air compressor machine 2
In high pressure tank 3, the pressure of air is adjusted by pressure regulator valve 6, is then passed through regenerator 7, electric heater 8 reaches necessarily
Temperature, the cooling air after heating is divided into two-way, air mass flow controlled respectively by regulating valve, and cooling air is from stator blade all the way
Root enters blade ante-chamber 27 and is discharged into blade grid passage；Cooling air-flow is discharged into leaf grating from the film cooling holes 26 of stator blade end wall all the way
Passage, air go out stream and form an angle with blade end wall 29.The cooling water of air compressor machine 2 comes from low pressure recycle water, low pressure water
Backwater is sent to outdoor cooling tower 21, stays back cistern 22 after cooling again.
When cooling air is full of cascade test section 10 with certain pressure, passed through by the high-temperature fuel gas of regulating valve control flow
Expansion segment 12, stable section 13, contraction section 14 and turbulent flow generator 15 enter cascade test section 10.Stable section 12 and contraction section 13 are used
In the incoming for producing stable and uniform, turbulent flow generator 15 is used for producing stable turbulent fluid.Discharged from cascade test section 10
High-temperature fuel gas sprays into flue gas by high pressure water makes its temperature drop to 500 DEG C, pressure is down to 0.8MPa using back pressure regulating valve, subtracts
Combustion gas after temperature decompression enters regenerator 7 and heats air, and the combustion gas after heat exchange is discharged into air through sound damping tower 25.
The pipeline of whole pilot system is by with high temperature resistant, the material manufacture of corrosion resistance characteristic, to ensure for a long time in height
The lower work of temperature.In order to reduce the thermal loss of experimental system, whole experimental design is mounted with an external insulation cover, in heat shield
Heat-barrier material 31 is clogged to ensure the heat insulating ability of experimental system and air-tightness in portion gap.Turbine stator blade 11 using high-temperature seal adhesive and
Bolt is installed on blade end wall 29.
Referring to Fig. 2, in order to truly reflect the gas flowing inside turbine, blade end wall 29 is designed to curved wall, arc
Transmission window 32 is arranged in 2~5 blade grid passages of arrangement above wall, blade grid passage surface, passes through transmission window 32 and profit
It can realize the shooting of opposite end wall surface temperature field with infrared thermal imagery system 17, the temperature field image of acquisition can be with by post processing
Obtain the front and rear temperature difference distribution of end wall cooling.The end wall center arrangement pressure-measuring-point 18 and thermocouple 16 of adjacent blade grid passage, examination
Before testing beginning, the pressure by monitoring runner is distributed to judge whether runner meets periodicity condition, by the temperature for reading wall
Angle value calibrates the temperature of thermal infrared imager.In 10 high-temperature fuel gas inlet of cascade test section, arrangement one discharges into a mouthful temperature point,
Mainstream inlet temperature is measured by installing thermocouple 16, simultaneously can be used for examining the uniformity of inlet air flow.Close to leaf
The upstream wall and downstream wall of grid passage set row's pressure-measuring-point 18 respectively, for measuring the stagnation pressure and quiet of leaf grating inlet and outlet
Pressure distribution.
Blade tip is processed using piecemeal, is bonded in by high-temperature seal adhesive on blade, this design is convenient for changing, and can be used
To carry out the heat transfer experiments of different endwall structures.Fig. 3-Fig. 6 is the schematic diagram of four kinds of blade endwall structures, and the utility model includes
But it is not limited only to these four endwall structures.
Referring to Fig. 3, the cross section of blade endwall structure one embodiment.Combustor exit end wall 30 and the turbine first order are quiet
29 import junction of end of blade wall smoothly transits.
Referring to Fig. 4, the cross section of blade endwall structure one embodiment.Combustor exit end wall 30 and the turbine first order are quiet
29 junction of end of blade wall produces dislocation (radial deflection), and offset height is several for combustion engine combustion chamber and first stage stator blades end wall junction
Why not the average value of matching degree.
Referring to Fig. 5, the cross section of blade endwall structure one embodiment.Combustor exit end wall 30 and the turbine first order are quiet
29 junction of end of blade wall smoothly transits.In order to strengthen the cooling to blade-cascade end wall, in the air film that end wall leading edge upstream arrangement is discrete
Cooling hole 26 provides cooling air.
Referring to Fig. 6, the cross section combustor exit end wall 30 of blade endwall structure one embodiment is quiet with the turbine first order
29 junction of end of blade wall produces dislocation (radial deflection), and offset height is several for combustion engine combustion chamber and first stage stator blades end wall junction
Why not the average value of matching degree.In order to strengthen the cooling to blade-cascade end wall, in the gaseous film control that end wall leading edge upstream arrangement is discrete
Hole 26 provides cooling air.
Referring to Fig. 7, the layout drawing of blade and end wall high temperature resistant foil gauge.In blade and end wall surface, multiple high temperature resistants are set
Foil gauge 19 is used for the stress field distribution situation of pilot blade and end wall surface.High temperature resistant foil gauge 19 is adopted with high temperature resistant wire
Blade and end wall surface are placed in special binder adhesive plaster.High temperature resistant wire is drawn by the fairlead below runner, is coupled to
Data collector.Fairlead carries out encapsulation process using refractory seals glue, prevents high-temperature gas from leaking.
Claims (10)
1. a kind of gas turbine turbine blade endwall heat transfer pilot system, it is characterised in that including high-temperature fuel gas feed system, two
Secondary air supply system, cascade test section (10), cooling water system, measurement and data collecting system and exhaust system；Wherein,
Turbine blade (11) for measuring is set in cascade test section (10), and cascade test section offers transmission at the top of (10)
Window (32), bottom are provided with cooling air chamber (9)；
High-temperature fuel gas feed system is connected with cascade test section, secondary for supplying high-temperature fuel gas to cascade test section (10)
Air supply system is used to provide high pressure to cascade test section (10) by cooling air chamber (9) and temperature is adjustable secondary cold
But air；
Cooling water system is divided into high pressure water system and low pressure water system, and high pressure water system is used for cooling high temperature flue gas, low pressure water system
System is used to cool down the air compressor machine (2) in secondary air supply；
Measurement and data collecting system include flow parameter measurement device and thermal stress measuring device, wherein, flow parameter measurement
Device is used to measure temperature, pressure and the flow in cascade test section (10), and thermal stress measuring device is used to measure turbine blade
(11) thermal stress；
Exhaust system includes exhaust conditioning system and silene system, and exhaust conditioning system is used to adjust cascade test section (10) exhaust
Pressure, and discharged after high-temperature fuel gas is cooled down through silene system.
A kind of 2. gas turbine turbine blade endwall heat transfer pilot system according to claim 1, it is characterised in that high temperature
Gas supply system includes sequentially connected expansion segment (12), stable section (13), contraction section (14), and is arranged on contraction section
(14) turbulent flow generator (15).
3. a kind of gas turbine turbine blade endwall heat transfer pilot system according to claim 2, it is characterised in that secondary
Air supply system includes sequentially connected air compressor machine (2), regenerator (7) and electric heater (8).
4. a kind of gas turbine turbine blade endwall heat transfer pilot system according to claim 3, it is characterised in that secondary
Air supply system further includes the filter (1) for being arranged on air compressor machine (2) inlet, is arranged on air compressor machine (2) exit and is used for
The high pressure tank (3) of compressed gas is stored, for adjusting the pressure regulator valve (4), mass flowmenter (5) and regulating valve of compressed air
And the substation (24) for powering for air compressor machine (2) (6),.
5. a kind of gas turbine turbine blade endwall heat transfer pilot system according to claim 2, it is characterised in that shrink
The combustor exit end wall 30 in section (14) exit smoothly transits or exists with 29 import junction of turbine first stage stator blades end wall
Dislocation.
A kind of 6. gas turbine turbine blade endwall heat transfer pilot system according to claim 1, it is characterised in that cooling
Water system includes feed pump (23), and the cooling tower (21) for being connected with feed pump (23) and setting gradually from top to bottom and water storage
Pond (22).
A kind of 7. gas turbine turbine blade endwall heat transfer pilot system according to claim 1, it is characterised in that flowing
Parameter measuring apparatus includes being arranged on the infrared thermal imagery system (17) above transmission window (32), is arranged on turbine blade (11) end
Thermocouple (16) and pressure-measuring-point (18) on wall, thermal stress measuring device include the high temperature resistant being arranged on turbine blade (11)
Foil gauge (19), and flow parameter measurement device and thermal stress measuring device are connected with data collecting system (20).
A kind of 8. gas turbine turbine blade endwall heat transfer pilot system according to claim 1, it is characterised in that noise reduction
System is sound damping tower (25).
A kind of 9. gas turbine turbine blade endwall heat transfer pilot system according to claim 1, it is characterised in that turbine
The end wall of blade (11) offers film cooling holes (26).
A kind of 10. gas turbine turbine blade endwall heat transfer pilot system according to claim 1, it is characterised in that leaf
Grid test section (10) exit is provided with heat-barrier material (31).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201721100744.6U CN207248535U (en) | 2017-08-30 | 2017-08-30 | Gas turbine blade end wall heat transfer test system |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201721100744.6U CN207248535U (en) | 2017-08-30 | 2017-08-30 | Gas turbine blade end wall heat transfer test system |
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| Publication Number | Publication Date |
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| CN207248535U true CN207248535U (en) | 2018-04-17 |
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ID=61883770
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| CN201721100744.6U Expired - Fee Related CN207248535U (en) | 2017-08-30 | 2017-08-30 | Gas turbine blade end wall heat transfer test system |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN108982113A (en) * | 2018-04-26 | 2018-12-11 | 西安交通大学 | A kind of two-phase experimental system for turbine blade leading edge impinging cooling |
| CN109738193A (en) * | 2019-01-08 | 2019-05-10 | 哈尔滨电气股份有限公司 | Gas-turbine combustion chamber test measures segment structure with air-cooled type |
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| CN111830083A (en) * | 2020-08-11 | 2020-10-27 | 沈阳航空航天大学 | System and method for measuring heat exchange coefficient of blade in high-temperature environment |
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| CN109738193B (en) * | 2019-01-08 | 2021-04-06 | 哈尔滨电气股份有限公司 | Air-cooled measuring section structure for gas turbine combustion chamber test |
| CN112098058A (en) * | 2020-08-07 | 2020-12-18 | 上海发电设备成套设计研究院有限责任公司 | Thermal fatigue life analysis method and test system for heavy gas turbine blade |
| CN111830083A (en) * | 2020-08-11 | 2020-10-27 | 沈阳航空航天大学 | System and method for measuring heat exchange coefficient of blade in high-temperature environment |
| CN111830083B (en) * | 2020-08-11 | 2023-04-07 | 沈阳航空航天大学 | System and method for measuring heat exchange coefficient of blade in high-temperature environment |
| CN112254941A (en) * | 2020-09-08 | 2021-01-22 | 中国航发湖南动力机械研究所 | Cold efficiency test piece of turbine blade |
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| CN112683943A (en) * | 2020-12-01 | 2021-04-20 | 西安交通大学 | Turbine experimental apparatus with adjustable pitch |
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