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 68
- 238000012546 transfer Methods 0.000 title claims description 20
- 239000007789 gas Substances 0.000 claims abstract description 42
- 238000001816 cooling Methods 0.000 claims abstract description 39
- XLYOFNOQVPJJNP-UHFFFAOYSA-N water Substances O XLYOFNOQVPJJNP-UHFFFAOYSA-N 0.000 claims abstract description 20
- 230000008646 thermal stress Effects 0.000 claims abstract description 17
- 239000002737 fuel gas Substances 0.000 claims abstract description 12
- 238000005259 measurement Methods 0.000 claims abstract description 11
- 239000000498 cooling water Substances 0.000 claims abstract description 9
- 230000030279 gene silencing Effects 0.000 claims abstract description 8
- 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
- 238000007599 discharging Methods 0.000 claims abstract description 3
- 230000005540 biological transmission Effects 0.000 claims description 7
- 230000001105 regulatory effect Effects 0.000 claims description 6
- 230000008602 contraction Effects 0.000 claims description 4
- 239000012774 insulation material Substances 0.000 claims description 4
- 238000001931 thermography Methods 0.000 claims description 4
- 230000006641 stabilisation Effects 0.000 claims description 3
- 238000011105 stabilization Methods 0.000 claims description 3
- 230000003584 silencer Effects 0.000 claims 2
- 238000002485 combustion reaction Methods 0.000 description 12
- 238000011144 upstream manufacturing Methods 0.000 description 8
- 239000003292 glue Substances 0.000 description 3
- 230000003068 static effect Effects 0.000 description 3
- 230000007704 transition Effects 0.000 description 3
- 239000000567 combustion gas Substances 0.000 description 2
- 238000013461 design Methods 0.000 description 2
- 238000011161 development Methods 0.000 description 2
- 238000005457 optimization Methods 0.000 description 2
- 238000011160 research Methods 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000009530 blood pressure measurement Methods 0.000 description 1
- 230000007797 corrosion Effects 0.000 description 1
- 238000005260 corrosion Methods 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
- 238000013401 experimental design Methods 0.000 description 1
- 239000011888 foil Substances 0.000 description 1
- 239000012535 impurity Substances 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 238000012544 monitoring process Methods 0.000 description 1
- 230000000737 periodic effect Effects 0.000 description 1
- 238000012805 post-processing Methods 0.000 description 1
- 238000004321 preservation Methods 0.000 description 1
- 238000012545 processing Methods 0.000 description 1
- 239000000565 sealant Substances 0.000 description 1
- 238000007789 sealing Methods 0.000 description 1
- 230000000087 stabilizing effect Effects 0.000 description 1
- 230000001052 transient effect 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
The utility model relates to a gas turbine blade end wall heat transfer test system.
Background
The improvement of the initial temperature of the turbine inlet of the gas turbine is an important way for improving the circulating power and the heat efficiency of the gas turbine, and the inlet temperature of the prior advanced heavy-duty gas turbine reaches about 1600 ℃.
The wall surface of the end part of the first stage nozzle blade of the gas turbine is exposed to high-temperature high-turbulence gas, and the end part of the blade bears higher heat load and wall surface temperature. Meanwhile, due to the existence of complex secondary flows such as end horseshoe vortices, channel vortices and the like, heat exchange between high-temperature fuel gas and the end wall is seriously uneven, and heat transfer of local areas of the end parts of the blades is enhanced. Therefore, the interface of the combustor outlet and the turbine first stage stator blade end wall should be naturally smooth to reduce the heat transfer deterioration and aerodynamic loss caused by the secondary flow.
Due to the existence of assembly tolerance and different thermal expansion degrees, the joint surface of the outlet of the combustion chamber of the gas turbine and the end wall of the static blade inevitably has dislocation, and the dislocation brings about the change of the structure of the end wall of the blade and brings about direct influence on the development of secondary flow and the condition of end heat transfer. Meanwhile, in order to avoid direct impact of high-temperature combustion gas on the end wall of the blade, discrete film holes or other combined cooling forms are usually arranged upstream of the front edge of the blade, so that effective film covering on the end wall surface is realized, and the cold air outflow arrangement mode is also related to secondary flow development. Therefore, the research on reasonable gas film layout under the real blade end wall structure is the key point of the end wall heat exchange research.
At present, the domestic turbine end wall heat exchange test system has the following problems: firstly, most end wall heat exchange tests measure conditions based on an adiabatic assumption, the test temperature is low, and the obtained end wall heat exchange rule of the test temperature is different from the actual end wall heat exchange rule of the combustion engine turbine; secondly, the influence of geometric mismatching of the joint of the actual combustion engine combustion chamber and the first stage stationary blade end wall on the end wall heat exchange characteristic is less considered in the end wall heat exchange test; in the conventional heat exchange test, only the pneumatic heat transfer effect is considered, and no thermal stress measuring point is arranged on the end wall and the surface of the blade.
SUMMERY OF THE UTILITY MODEL
An object of the utility model is to above-mentioned gas turbine end wall heat transfer system defect or improvement demand now, provide a gas turbine blade end wall heat transfer test system to satisfy the experimental requirement of steady state heat transfer and transient state heat transfer. The system utilizes the first-stage static blade of the turbine of the gas turbine as a blade cascade test section, and the test condition is similar to the actual operation condition of the gas turbine.
In order to achieve the above purpose, the utility model adopts the following technical scheme:
a gas turbine blade end wall heat exchange test system comprises a high-temperature gas supply system, a secondary air supply system, a blade grid test section, a cooling water system, a measurement and data acquisition system and an exhaust system; wherein,
turbine blades for measurement are arranged in the blade grid test section, a transmission window is formed in the top of the blade grid test section, and a cooling air chamber is formed in the bottom of the blade grid test section;
the high-temperature fuel gas supply system is communicated with the blade grid test section and used for supplying high-temperature fuel gas to the blade grid test section, and the secondary air supply system is used for supplying high-pressure and temperature-adjustable secondary cooling air to the blade grid test section through the cooling air cavity;
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.
The utility model discloses further improvement lies in, high temperature gas feed system is including the expansion section, stable section, the shrink section that connect gradually to and set up the turbulence generator at the shrink section.
The utility model discloses further improvement lies in, secondary air supply system is including the air compressor machine, regenerator and the electric heater that connect gradually.
The utility model discloses further improvement lies in, and secondary air supply system still sets up the high-pressure gas holder that is used for storing compressed gas in the air compressor machine exit including setting up the filter in the air compressor machine entrance for adjust compressed air's air-vent valve, mass flow meter and governing valve, and be used for the transformer substation for the air compressor machine power supply.
The utility model discloses further improvement lies in, the smooth transition of combustor export end wall in shrink section exit and the quiet leaf end wall access connection department of turbine first order or have the dislocation.
The utility model discloses further improvement lies in that cooling water system includes the feed pump to and be connected and cooling tower and the cistern that top-down set gradually with the feed pump.
The utility model discloses further improvement lies in, and flow parameter measurement device sets up thermocouple and pressure measurement station on turbine blade end wall including setting up the infrared thermal imaging system in transmission window top, thermal stress measurement device including setting up the high temperature resistant foil gage on turbine blade, and flow parameter measurement device and thermal stress measurement device all are connected with data acquisition system.
The utility model discloses a further improvement lies in, and sound deadening system is the amortization tower.
The utility model discloses a further improvement lies in, and the end wall of turbine blade has seted up the air film cooling hole.
The utility model discloses further improvement lies in, and cascade test section exit is provided with thermal insulation material.
For current end wall heat transfer test system, the utility model discloses following beneficial effect has:
the utility model provides a gas turbine blade end wall heat transfer test system, the blade of adoption be the first order quiet leaf of actual combustion engine turbine, and experimental operating mode is combustion engine operating condition, can be true temperature and the thermal stress distribution on reflection turbine end wall surface.
The utility model provides a gas turbine blade end wall heat transfer test system has considered actual combustion engine combustion chamber and the first order quiet leaf end wall junction geometry and has mismatched to and the influence of gas film cooling to end wall heat transfer characteristic, this gas film cooling hole structure optimization and the arrangement of studying leading edge upstream end wall have great meaning.
The utility model provides a gas turbine blade end wall heat transfer test system can not only the temperature distribution on direct measurement blade surface, measures the thermal stress distribution on end wall and blade surface moreover, can provide experimental data for gas turbine blade design optimization.
Drawings
FIG. 1 is a schematic layout of a gas turbine blade endwall heat exchange test system;
FIG. 2 is an enlarged view of the test section of the cascade shown in FIG. 1;
FIG. 3 is a schematic end wall platform without consideration of upstream film cooling holes and geometric mismatches;
FIG. 4 is a schematic view of an end wall platform considering only geometric mismatches;
FIG. 5 is a schematic endwall platform view of the upstream film cooling hole only considered;
FIG. 6 is a schematic end wall platform view taking into account upstream film cooling holes and geometric mismatches;
FIG. 7 is a layout of high temperature resistant strain gages on the blade surface and endwall.
In the figure: 1. a filter; 2. an air compressor; 3. a high pressure gas storage tank; 4. a pressure regulating valve; 5. a mass flow meter; 6. adjusting a valve; 7. a heat regenerator; 8. an electric heater; 9. a cooling air chamber; 10. a cascade test section; 11. a turbine stator vane; 12. an expansion section; 13. a stabilization section; 14. a contraction section; 15. a turbulence generator; 16. a thermocouple; 17. an infrared thermal imaging system; 18. measuring a pressure point; 19. a high temperature resistant strain gauge; 20. a data acquisition system; 21. a cooling tower; 22. a reservoir; 23. a feed pump; 24. a transformer substation; 25. a silencing tower; 26. a film cooling hole; 27. a blade front cavity; 28. a blade rear cavity; 29. a first stage stationary blade endwall; 30. a combustion chamber outlet end wall; 31. a thermal insulation material; 32. a transmission window.
Detailed Description
The present invention will be described in further detail with reference to the accompanying drawings and examples.
As shown in FIG. 1, the utility model discloses a gas turbine blade end wall heat transfer test system, including high temperature gas feed system, secondary air feed system, cascade test section 10, cooling water system, measurement and data acquisition system, exhaust system etc..
Referring to fig. 1, an air compressor 2 is started, cooling air is filtered to remove impurities through a filter 1, the air is pressurized through the air compressor 2 and then stored in a high-pressure air storage tank 3, the pressure of the air is adjusted through a pressure adjusting valve 6, then the air reaches a certain temperature through a heat regenerator 7 and an electric heater 8, the heated cooling air is divided into two paths, the air flow is respectively controlled through adjusting valves, and one path of cooling air enters a blade front cavity 27 from the root of a stationary blade and is discharged into a blade cascade channel; a cooling air flow is discharged from film cooling holes 26 in the vane endwalls into the cascade channels, and the air outflow is at an angle to the blade endwalls 29. The cooling water of the air compressor 2 comes from low-pressure circulating water, the return water of the low-pressure water is sent to the outdoor cooling tower 21, and the return water is cooled and then is reserved in the reservoir 22.
When the cooling air is filled in the cascade test section 10 under a certain pressure, the high-temperature combustion gas whose flow rate is controlled by the regulating valve enters the cascade test section 10 through the expansion section 12, the stabilization section 13, the contraction section 14 and the turbulence generator 15. The stabilizing section 12 and the converging section 13 are used to generate a steady and uniform incoming flow, and the turbulence generator 15 is used to generate a steady turbulent flow. High-temperature fuel gas discharged from the cascade test section 10 is sprayed into flue gas through high-pressure water to reduce the temperature to 500 ℃, the pressure is reduced to 0.8MPa by utilizing a back pressure regulating valve, the fuel gas after temperature and pressure reduction enters a heat regenerator 7 to heat air, and the fuel gas after heat exchange is discharged into the atmosphere through a silencing tower 25.
The pipeline of the whole test system is made of a material with high temperature resistance and corrosion resistance so as to ensure long-time work at high temperature. In order to reduce the heat loss of the experimental system, an external heat shield is installed in the whole experimental design, and the heat shield is internally filled with heat insulation materials 31 to ensure the heat preservation and air tightness of the experimental system. Turbine vanes 11 are mounted on the blade endwalls 29 using high temperature sealing glue and bolts.
Referring to fig. 2, in order to truly reflect the gas flow inside the turbine, the blade end wall 29 is designed into an arc-shaped wall surface, 2-5 blade grid channels are arranged above the arc-shaped wall surface, a transmission window 32 is arranged right above the blade grid channels, the temperature field on the surface of the end wall can be shot through the transmission window 32 and by using the infrared thermal imaging system 17, and the obtained temperature field image can be subjected to post-processing to obtain the temperature difference distribution before and after the end wall is cooled. A pressure measuring point 18 and a thermocouple 16 are arranged in the center of the end wall of the adjacent cascade channel, before the test is started, whether the flow channel meets the periodic condition or not is judged by monitoring the pressure distribution of the flow channel, and the temperature of the thermal infrared imager is calibrated by reading the temperature value of the wall surface. A row of inlet temperature measuring points are arranged at a high-temperature fuel gas inlet of the cascade test section 10, the temperature of a main flow inlet is measured by installing a thermocouple 16, and meanwhile, the inlet temperature measuring points can also be used for detecting the uniformity of inlet airflow. A row of pressure measuring points 18 are respectively arranged on the upstream wall surface and the downstream wall surface close to the cascade channel and are used for measuring the total pressure and static pressure distribution of the inlet and the outlet of the cascade.
The blade tip adopts the block processing, bonds on the blade through high temperature sealed glue, and this kind of design is convenient to be changed, can be used to carry out the heat transfer test of different end wall structures. Fig. 3-6 are schematic views of four blade endwall configurations, including but not limited to the four endwall configurations.
Referring to FIG. 3, a cross-section of one embodiment of a blade endwall structure. The combustor outlet endwall 30 is in smooth transition with the inlet connection of the turbine first stage vane endwall 29.
Referring to FIG. 4, a cross-section of one embodiment of a blade endwall structure. The combustion chamber outlet endwall 30 is offset (radially offset) from the turbine first stage vane endwall 29 by an average of the geometric mismatch at the combustion engine combustion chamber to first stage vane endwall junction.
Referring to FIG. 5, a cross-section of one embodiment of a blade endwall structure. The combustor outlet endwall 30 is in smooth transition with the turbine first stage vane endwall 29 junction. To enhance cooling of the cascade endwall, discrete film cooling holes 26 are disposed upstream of the endwall leading edge to provide cooling air.
Referring to FIG. 6, the cross-sectional combustor exit endwall 30 of an embodiment of the blade endwall configuration is offset (radially offset) from the turbine first stage vane endwall 29 by an average of the geometric mismatch at the combustor first stage vane endwall junction. To enhance cooling of the cascade endwall, discrete film cooling holes 26 are disposed upstream of the endwall leading edge to provide cooling air.
Referring to fig. 7, a layout of the vanes and endwall refractory strain gages. A plurality of high temperature resistant strain gauges 19 are arranged on the surfaces of the blade and the end wall and used for testing the distribution of the thermal stress field of the surfaces of the blade and the end wall. The high-temperature resistant strain gauge 19 and the high-temperature resistant lead are arranged on the surfaces of the blade and the end wall by adopting special bonding glue. The high-temperature-resistant lead is led out through a lead hole below the runner and is connected to the data collector in parallel. The lead hole is sealed by adopting high-temperature-resistant sealant, so that high-temperature gas leakage is prevented.
Claims (10)
1. A gas turbine blade end wall heat exchange test system is characterized by comprising a high-temperature gas supply system, a secondary air supply system, a blade grid test section (10), a cooling water system, a measurement and data acquisition system and an exhaust system; wherein,
a turbine blade (11) for measurement is arranged in the blade grid test section (10), a transmission window (32) is arranged at the top of the blade grid test section (10), and a cooling air chamber (9) is arranged at the bottom of the blade grid test section;
the high-temperature fuel gas supply system is communicated with the blade grid test section and used for supplying high-temperature fuel gas to the blade grid test section (10), and the secondary air supply system is used for supplying high-pressure and temperature-adjustable secondary cooling air to the blade grid test section (10) through the cooling air chamber (9);
the cooling water system comprises a high-pressure water system and a low-pressure water system, wherein 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 (2) in a secondary air supply system;
the measuring and data acquisition 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 (10), and the thermal stress measuring device is used for measuring the thermal stress of the turbine blade (11);
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 (10) and discharging high-temperature fuel gas after being cooled through the silencing system.
2. The turbine blade end wall heat exchange test system of claim 1, wherein the high temperature gas supply system comprises an expansion section (12), a stabilization section (13), a contraction section (14) and a turbulence generator (15) arranged at the contraction section (14) which are connected in sequence.
3. The turbine blade end wall heat exchange test system of the gas turbine according to claim 2, wherein the secondary air supply system comprises an air compressor (2), a heat regenerator (7) and an electric heater (8) which are connected in sequence.
4. The turbine blade end wall heat exchange test system of claim 3, wherein the secondary air supply system further comprises a filter (1) disposed at an inlet of the air compressor (2), a high-pressure air storage tank (3) disposed at an outlet of the air compressor (2) for storing compressed air, a pressure regulating valve (4), a mass flow meter (5) and a regulating valve (6) for regulating the compressed air, and a substation (24) for supplying power to the air compressor (2).
5. The turbine blade endwall heat transfer test system of claim 2, wherein the combustor exit endwall 30 at the exit of the convergent section (14) is smoothly blended or misaligned with the inlet connection of the turbine first stage vane endwall 29.
6. The gas turbine blade end wall heat exchange test system according to claim 1, wherein the cooling water system comprises a feed water pump (23), and a cooling tower (21) and a water reservoir (22) which are connected with the feed water pump (23) and are arranged in sequence from top to bottom.
7. The turbine blade end wall heat exchange test system of the gas turbine as claimed in claim 1, wherein the flow parameter measuring device comprises an infrared thermal imaging system (17) arranged above the transmission window (32), a thermocouple (16) and a pressure measuring point (18) arranged on the end wall of the turbine blade (11), the thermal stress measuring device comprises a high temperature resistant strain gauge (19) arranged on the turbine blade (11), and the flow parameter measuring device and the thermal stress measuring device are both connected with the data acquisition system (20).
8. The gas turbine blade end wall heat exchange test system of claim 1, wherein the silencer system is a silencer tower (25).
9. The heat exchange test system for the end wall of the turbine blade of the gas turbine as claimed in claim 1, wherein the end wall of the turbine blade (11) is provided with film cooling holes (26).
10. The gas turbine blade endwall heat transfer test system of claim 1, wherein an insulation material (31) is provided at the exit of the cascade test section (10).
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| Application Number | Priority Date | Filing Date | Title |
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| CN201721100744.6U CN207248535U (en) | 2017-08-30 | 2017-08-30 | Gas turbine blade end wall heat transfer test system |
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| Application Number | Priority Date | Filing Date | Title |
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| CN201721100744.6U CN207248535U (en) | 2017-08-30 | 2017-08-30 | Gas turbine blade end wall heat transfer test system |
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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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| 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 |
| CN109752187A (en) * | 2018-12-12 | 2019-05-14 | 西安航天动力试验技术研究所 | Rail control engine vacuum environment high-speed and high-temperature combustion gas rapid pressure cooling system |
| 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 |
| 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 |
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| 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 |
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| CN112683943A (en) * | 2020-12-01 | 2021-04-20 | 西安交通大学 | Turbine experimental apparatus with adjustable pitch |
| CN112683943B (en) * | 2020-12-01 | 2021-11-16 | 西安交通大学 | Turbine experimental apparatus with adjustable pitch |
| CN113624357A (en) * | 2021-07-26 | 2021-11-09 | 中国船舶重工集团公司第七0三研究所 | Method for measuring temperature of turbine chamber of marine gas turbine |
| CN113758968A (en) * | 2021-09-30 | 2021-12-07 | 西安交通大学 | Experimental system and steady-state experimental method for measuring heat exchange coefficient of blade top of turbine movable blade |
| CN114993638A (en) * | 2022-05-07 | 2022-09-02 | 中国联合重型燃气轮机技术有限公司 | Rotating turbine blade cooling test system and method |
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