CN116046404B - Testing system for non-contact dynamic stress of high-temperature-range turbine rotor - Google Patents
Testing system for non-contact dynamic stress of high-temperature-range turbine rotor Download PDFInfo
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- CN116046404B CN116046404B CN202310340750.2A CN202310340750A CN116046404B CN 116046404 B CN116046404 B CN 116046404B CN 202310340750 A CN202310340750 A CN 202310340750A CN 116046404 B CN116046404 B CN 116046404B
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- 238000012360 testing method Methods 0.000 title claims abstract description 54
- 238000001816 cooling Methods 0.000 claims abstract description 127
- 239000013307 optical fiber Substances 0.000 claims abstract description 74
- 238000009434 installation Methods 0.000 claims abstract description 10
- 238000009662 stress testing Methods 0.000 claims abstract 3
- 239000002131 composite material Substances 0.000 claims description 16
- 239000000523 sample Substances 0.000 claims description 7
- 238000003825 pressing Methods 0.000 claims description 5
- 230000035882 stress Effects 0.000 description 38
- 238000005259 measurement Methods 0.000 description 22
- 238000013461 design Methods 0.000 description 11
- 238000005516 engineering process Methods 0.000 description 9
- 230000000694 effects Effects 0.000 description 8
- 238000000034 method Methods 0.000 description 8
- 238000011161 development Methods 0.000 description 6
- 230000008569 process Effects 0.000 description 5
- 238000004458 analytical method Methods 0.000 description 4
- 239000000835 fiber Substances 0.000 description 4
- 238000004364 calculation method Methods 0.000 description 3
- 238000012986 modification Methods 0.000 description 3
- 230000004048 modification Effects 0.000 description 3
- 239000012809 cooling fluid Substances 0.000 description 2
- 230000006872 improvement Effects 0.000 description 2
- 238000005457 optimization Methods 0.000 description 2
- 238000006467 substitution reaction Methods 0.000 description 2
- 230000004083 survival effect Effects 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000001419 dependent effect Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 238000006073 displacement reaction Methods 0.000 description 1
- 238000009413 insulation Methods 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 238000000691 measurement method Methods 0.000 description 1
- 238000012544 monitoring process Methods 0.000 description 1
- 238000011160 research Methods 0.000 description 1
- 230000003068 static effect Effects 0.000 description 1
- 239000013589 supplement Substances 0.000 description 1
- 230000001360 synchronised effect Effects 0.000 description 1
- 238000010998 test method Methods 0.000 description 1
- 230000008646 thermal stress Effects 0.000 description 1
- 238000013024 troubleshooting Methods 0.000 description 1
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M15/00—Testing of engines
- G01M15/02—Details or accessories of testing apparatus
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01L—MEASURING FORCE, STRESS, TORQUE, WORK, MECHANICAL POWER, MECHANICAL EFFICIENCY, OR FLUID PRESSURE
- G01L1/00—Measuring force or stress, in general
- G01L1/24—Measuring force or stress, in general by measuring variations of optical properties of material when it is stressed, e.g. by photoelastic stress analysis using infrared, visible light, ultraviolet
- G01L1/242—Measuring force or stress, in general by measuring variations of optical properties of material when it is stressed, e.g. by photoelastic stress analysis using infrared, visible light, ultraviolet the material being an optical fibre
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- H—ELECTRICITY
- H05—ELECTRIC TECHNIQUES NOT OTHERWISE PROVIDED FOR
- H05K—PRINTED CIRCUITS; CASINGS OR CONSTRUCTIONAL DETAILS OF ELECTRIC APPARATUS; MANUFACTURE OF ASSEMBLAGES OF ELECTRICAL COMPONENTS
- H05K7/00—Constructional details common to different types of electric apparatus
- H05K7/20—Modifications to facilitate cooling, ventilating, or heating
- H05K7/20009—Modifications to facilitate cooling, ventilating, or heating using a gaseous coolant in electronic enclosures
- H05K7/20136—Forced ventilation, e.g. by fans
- H05K7/20145—Means for directing air flow, e.g. ducts, deflectors, plenum or guides
Abstract
The invention provides a non-contact dynamic stress testing system for a high-temperature-range turbine rotor, which comprises an optical fiber sensor. The optical fiber sensor is internally provided with a first cooling channel, one end of the first cooling channel is communicated with the air supply system, and the other end of the first cooling channel is communicated with the main flow path of the engine. The optical fiber sensor is wrapped in the cooling sleeve structure, a second cooling channel is formed between the cooling sleeve structure and the optical fiber sensor, and a cold air inlet and an exhaust hole which are used for communicating the second cooling channel with the main flow path of the engine are formed in the cooling sleeve structure. The cooling jacket structure is arranged on the casing, and a third cooling channel is formed between the cooling jacket structure and the casing. The test system designed by the invention realizes the efficient cooling of the sensor and can meet the non-contact dynamic stress test requirement of the turbine rotor of the aeroengine with high temperature and high pneumatic load above 1800K. An integrated installation scheme is established, the influence of the test on the performance and the structure can be reduced, the effective circumferential, axial and radial installation of the sensor is realized, and the test precision is improved.
Description
Technical Field
The invention belongs to the technical field of aeroengine tests, relates to a high-temperature-range turbine rotor dynamic stress measurement technology, and particularly relates to a high-temperature-range turbine rotor non-contact dynamic stress test system.
Background
Because the aeroengine turbine blade has complex modeling and bad working environment, the fault that the turbine blade breaks due to insufficient dynamic stress reserve margin can occur in the process of development, production and use. Meanwhile, in an actual working state of the engine, the working blades and the blade disc are rigidly connected, and the blades are influenced by the composite actions of centrifugal force, thermal stress, aerodynamic force and the like, so that the condition that the vibration stress level of the blades is higher in a rotating state can be caused. Therefore, during engine development and troubleshooting, the dynamic stresses of the engine rotor blades must be measured.
At present, in the aspect of rotor blade dynamic stress measurement, abundant experiences have been accumulated at home and abroad, design capability and test technology are mature, and dynamic stress measurement modes comprise contact type measurement and non-contact type measurement.
The contact dynamic stress measurement technology is already mature and applied to engine development and plays an important role. However, the contact dynamic stress test scheme has the technical characteristics that the structure is more modified, a current collector or a telemetry scheme is required to be adopted for leading out signals, and the whole test system is more complex; the modifying and pasting processes are longer, the survival rate of the strain gauge is lower in the test process, the pasting quality of the strain gauge is seriously dependent on the technical level of pasting operators, and the like, so that the test data cannot be obtained by long-time test, and even the expected purpose of the test cannot be achieved.
The non-contact dynamic stress measurement technology is an important supplement to the contact dynamic stress measurement technology as an important means of dynamic stress test. Compared with the contact type dynamic stress measurement, the non-contact type dynamic stress measurement scheme has the characteristics of small modification amount, simple cooling scheme, high reliability of test signals, short assembly period and the like, can realize real-time monitoring of the dynamic stress of the high-pressure turbine under the complex test procedure and longer test time, and can avoid the problems of long modification period, low survival rate of strain gauges, shortened test period, cost saving, speed up development and the like of the contact type dynamic stress measurement scheme.
However, due to the high requirements of the non-contact dynamic stress measurement technology on the test environment (especially the test environment temperature), the limitation of the refitted space and the like, and the fact that the current aero-engines which are researched at home and abroad bring higher turbine front temperature along with the continuous improvement of the performance, the existing non-contact dynamic stress method/system cannot realize the non-contact dynamic stress measurement on the turbine rotor blade with the turbine front temperature of more than 1800K.
In view of this, it is of great importance to establish a non-contact dynamic stress measurement technique suitable for high temperature range turbine rotors.
Disclosure of Invention
In order to solve the problem that the existing non-contact dynamic stress measurement technology is not suitable for measuring the dynamic stress of the turbine rotor blade at the front temperature of the turbine above 1800K, the invention designs a non-contact dynamic stress test system of a high-temperature-range turbine rotor.
The technical scheme for realizing the aim of the invention is as follows: the system for testing the non-contact dynamic stress of the turbine rotor in the high temperature range comprises a composite optical fiber sensor positioned at the position of a turbine rotor blade in a casing, wherein the composite optical fiber sensor comprises an optical fiber sensor, a first cooling channel is arranged in the optical fiber sensor, one end of the first cooling channel is communicated with an air supply system, and the other end of the first cooling channel is communicated with a main flow path of an engine;
the optical fiber sensor is wrapped in a cooling sleeve structure, a second cooling channel is formed between the cooling sleeve structure and the optical fiber sensor, and a cold air inlet and an exhaust hole which are used for communicating the second cooling channel with a main flow path of the engine are formed in the cooling sleeve structure;
the cooling jacket structure is arranged on the casing, and a third cooling channel is formed by a gap between the cooling jacket structure and the casing.
In an alternative embodiment, the first cooling channel is provided with rack bleed air, and the second cooling channel and the third cooling channel are provided with two cold air streams.
In an alternative embodiment, the cooling jacket structure is fixed on the casing through a fixing piece, and comprises a mounting bracket, a cooling jacket and a mounting pressing plate which are assembled in sequence from top to bottom.
In an alternative embodiment, the backflow margin of the cooling jacket outlet is equal to or greater than 1.4, and the backflow margin of the fiber optic sensor outlet is equal to or greater than 1.2.
In an alternative embodiment, a plurality of the exhaust holes are formed in the side wall and/or the bottom wall of the cooling jacket.
More preferably, when the bottom wall of the cooling sleeve is provided with a plurality of exhaust holes, the exhaust holes are distributed in an eccentric array, so that the air film covering effect at the bottom of the cooling sleeve is improved, and the cooling air-entraining amount is reduced.
In an alternative embodiment, the cold air inlet is located at an upper opening position of the cooling jacket.
In an alternative embodiment, the optical fiber sensor is located at a position of a leading edge or a trailing edge of the turbine rotor blade, and an axial distance between a mounting axis of the optical fiber sensor and a chord length center line of the turbine rotor blade is 1.5-2 mm.
More preferably, the axial distance between the installation axis of the optical fiber sensor and the chord length center line of the turbine rotor blade is 1.5-2 mm.
In an alternative embodiment, the clearance between the probe of the optical fiber sensor and the tip of the turbine rotor blade is 3-5 mm.
In an alternative embodiment, a thermocouple is further arranged on the optical fiber sensor, so that the temperature of the optical fiber sensor can be detected in real time, and the safety of the system is improved.
In an alternative embodiment, the plurality of composite fiber sensors is provided in plurality, and the plurality of composite fiber sensors are unevenly distributed in the circumferential direction within the casing.
More preferably, the number of the composite optical fiber sensors is more than or equal to 4.
Compared with the prior art, the invention has the beneficial effects that: the invention is based on the system engineering thought to carry out system optimization on the non-contact dynamic stress measurement technology of the high-pressure turbine of the aeroengine, designs a set of test system, comprises a composite optical fiber sensor (composed of a cooling structure and an optical fiber sensor), breaks through the technical problem of non-contact dynamic stress measurement of the rotor of the high-pressure turbine of the aeroengine in a high-temperature domain by establishing a high-efficiency cooling method and an installation design method on the composite optical fiber sensor, and can ensure the test requirement of non-contact dynamic stress to be realized above 1800K.
The test system designed by the invention is applied to the development process of a certain turbofan engine, and the result shows that the test system can fill the blank of non-contact dynamic stress measurement of the high-pressure turbine rotor of the aeroengine in a high-temperature area environment, and has important technical value.
Drawings
In order to more clearly illustrate the technical solution of the embodiments of the present invention, the drawings that are needed in the description of the embodiments will be briefly described.
FIG. 1 is a schematic structural diagram of a system for testing non-contact dynamic stress of a high temperature range turbine rotor in accordance with an embodiment;
FIG. 2 is an enlarged view of the position A of FIG. 1;
FIG. 3 is a schematic view of a mounting bracket in a cooling jacket structure in an embodiment;
FIG. 4 is a schematic view of a mounting platen in a cooling jacket structure in an embodiment;
FIG. 5 is a front view of a cooling jacket in a cooling jacket structure according to an embodiment;
FIG. 6 is a top view of a cooling jacket in a cooling jacket structure according to an embodiment;
1, an optical fiber sensor; 2. a cool air inlet; 3. an exhaust hole; 4. a mounting bracket; 5. mounting a pressing plate; 6. a cooling jacket; 10. a first cooling channel; 20. a second cooling channel; 30. a third cooling channel; 100. turbine rotor blades.
Detailed Description
The invention will be further described with reference to specific embodiments, and advantages and features of the invention will become apparent from the description. These examples are merely exemplary and do not limit the scope of the invention in any way. It will be understood by those skilled in the art that various changes and substitutions of details and forms of the technical solution of the present invention may be made without departing from the spirit and scope of the present invention, but these changes and substitutions fall within the scope of the present invention.
The embodiment provides a testing system for non-contact dynamic stress of a turbine rotor in a high temperature range, which is shown in fig. 1 and 2, and comprises a composite optical fiber sensor positioned at a turbine rotor blade 100 position in a casing, wherein the composite optical fiber sensor comprises an optical fiber sensor 1, a first cooling channel 10 is arranged in the optical fiber sensor 1, one end of the first cooling channel 10 is communicated with an air supply system, and the other end of the first cooling channel is communicated with a main flow path of an engine.
Referring to fig. 1 and 2, the optical fiber sensor 1 is wrapped in a cooling jacket structure, a second cooling channel 20 is formed between the cooling jacket structure and the optical fiber sensor 1, and a cold air inlet 2 and an exhaust hole 3 for communicating the second cooling channel 20 with a main flow path of an engine are formed on the cooling jacket structure.
The cooling jacket structure is disposed on the casing, and a gap between the cooling jacket structure and the casing forms a third cooling channel 30.
Referring to fig. 1 and 2, the test system designed in this embodiment can realize cooling of the composite optical fiber sensor through the first cooling channel 10, the second cooling channel 20 and the third cooling channel 30, so as to ensure that the test system can perform non-contact dynamic stress measurement on the turbine rotor blade under the temperature condition before the turbine with the temperature above 1800K.
The design of the test system is described in detail in this embodiment by the following aspects:
1. dynamic stress test instrument selection
According to the specific embodiment, a blade tip timing scheme is adopted, non-contact dynamic stress measurement is carried out on a high-pressure turbine rotor blade of the high-temperature-range aeroengine, a sensor is circumferentially arranged on a relatively static shell of a rotary machine, a sensor senses a pulse signal generated by a rotary blade passing in front of the sensor to obtain information of blade vibration displacement, and further the amplitude and frequency of blade vibration can be calculated. Meanwhile, by means of the reference synchronous signal, the vibration characteristics of each rotary blade can be analyzed.
Because the space where the turbine rotor blade 100 is located is relatively small and the structural forms of different types of engines are different, in this embodiment, the optical fiber sensor 1 with the first cooling channel 10 is selected and cooled by adopting an air cooling mode.
In order to know the cooling effect of the optical fiber sensor 1 in real time, in the specific embodiment, a thermocouple is preferably additionally arranged in the optical fiber sensor 1 and used for detecting the temperature of the optical fiber sensor in real time, so that the safety of the system is improved. Meanwhile, according to the temperature of the test environment, a reasonable light source (for example, when the temperature is higher, a visible light source with red light is mainly used, if an optical fiber sensor adopts a red light source or a similar light source, a test signal cannot be obtained, and a blue light source is selected at the moment) is selected, so that the influence of the pollution of the test environment light on the test effect is reduced.
2. Rotational speed reference design
The rotation speed of the aeroengine is generally measured by a magneto-electric rotation speed sensor and an optical fiber rotation speed, and the magneto-electric rotation speed sensor (not shown in the drawing) is selected to measure the rotation speed in the specific embodiment, so that the aeroengine has good signal stability. In this concrete embodiment, magnetoelectric revolution speed transducer installs fixedly through the support, designs the sound teeth of a cogwheel in the pivot simultaneously, and after the sound teeth of a cogwheel was walked around magnetoelectric revolution speed transducer every time, can produce magnetoelectric pulse signal, and then obtains engine rotor rotational speed signal according to pulse signal quantity in the unit time.
3. Fiber optic sensor 1 cooling design
In this embodiment, the optical fiber sensor 1 performs a cooling effect test, and uses the cooling fluid in the first cooling channel 10 as a research object, where the cooling fluid flow can be regarded as adiabatic isentropic flow, and the flow calculation formula is as follows:
，/>
，
wherein->
Maximum air supply pressure for the bench,/->
The maximum air supply temperature for the rack is set,
for the real flow area of the optical fiber sensor 1, < >>
Is a speed factor->
Mach number>
Is a factor (0.0404 when the rack bleed air circulating in the first cooling channel 10 is air)>
Is a thermal insulation index (value 1.4 when the rack bleed air circulating in the first cooling channel 10 is air), is>
As a flow function +.>
For the mass flow of the optical fiber sensor 1,
is the maximum mass flow of the optical fiber sensor.
When the rack bleed air flow is larger, the cooling effect of the optical fiber sensor 1 is better, but the influence on the aerodynamic performance of the engine runner is larger. Furthermore, because the optical fiber sensor 1 is limited in the size of the first cooling channel 10, the air-entraining (i.e. cooling air) at the high pressure and low temperature part of the rack or the engine can be throttled. The present embodiment selects a cooling mode of combining two cold air streams inside the aeroengine with the rack bleed air (or bleed air at a high-pressure and low-temperature part of the engine or other forms of bleed air on the engine) to cool the optical fiber sensor 1.
The present embodiment designs a cooling jacket structure for fixing and cooling the optical fiber sensor 1. Referring to fig. 1-2, the cooling jacket structure is disposed on the casing, and a gap between the cooling jacket structure and the casing forms a third cooling channel 30; the optical fiber sensor 1 is wrapped in the cooling sleeve structure, a second cooling channel 20 is formed between the cooling sleeve structure and the optical fiber sensor 1, and two streams of cold air flow in the third cooling channel 30 and the second cooling channel 20.
Referring to fig. 3, 4, 5 and 6, the cooling jacket structure is fixed on a casing (i.e. a middle casing of a high-pressure turbine) through a fixing piece, and comprises a mounting bracket 4, a cooling jacket 6 and a mounting pressing plate 5 which are assembled sequentially from top to bottom.
Referring to fig. 5 and 6, the upper opening of the cooling jacket 6 is provided with a cold air inlet 2, and the side wall and/or the bottom of the cooling jacket 6 (except the center of the bottom, phi 3.2 through holes are formed to enable the optical fiber sensor 1 to work) are provided with exhaust holes 3 (also called air film holes, the aperture can be 0.4 mm), more preferably, in order to reduce the cold air entering the main flow path through the exhaust holes 3, the exhaust holes 3 of the bottom wall of the cooling jacket 6 are designed into eccentric array hole structures, so that the air film covering effect at the bottom of the cooling jacket can be improved, and the cooling air-entraining amount can be reduced. By arranging the exhaust holes 3 in the direction of the windward flow, two cold air flows can pneumatically form an air film at the bottom rear edge position of the cooling sleeve 6 along with the air flow of the main flow path after passing through the exhaust holes 3.
More preferably, by performing pneumatic analysis on the fluidity of the whole structure, in order to prevent the two streams of cold air from flowing backward to the head position of the optical fiber sensor 1, the probe of the optical fiber sensor 1 is ablated, and in this embodiment, the backflow margin of the outlet of the cooling jacket 6 is made not lower than 1.4, and the backflow margin of the outlet of the optical fiber sensor 1 is made not lower than 1.2.
In this embodiment, the fixing member may be a bolt or nut, and after the optical fiber sensor 1 is installed in the cooling jacket 6, the mounting bracket 4, the mounting pressing plate 5 and the cooling jacket 6 are assembled to form a cooling jacket structure, and then the assembled cooling jacket structure is fixed to the receiver through the bolt or nut.
4. Positional relationship of cooling jacket structure to turbine rotor blade 100
When the cooling jacket structure is installed, the installation positions of the optical fiber sensor 1 in the circumferential direction, the axial direction and the radial direction in the casing are considered.
Circumferential assembly number: according to the embodiment, through testing and analyzing each component, a plurality of cooling jacket structures are required to be installed on the optical fiber sensor 1 in the circumferential direction, for example, the cooling jacket structures are determined according to the number of rotor blades, the testing radius and other conditions, and 4 or more cooling jacket structures which are unevenly distributed in the circumferential direction (the circumferential testing positions are finally calculated according to the number of engine blades and the rotating speed of interest) are verified, so that the optical fiber sensor has a good testing effect and a good cooling effect.
Axial mounting position: since the vibration amplitude at the leading or trailing edge of the turbine rotor blade 100 is large, the cooling jacket structure is chosen to be mounted at the leading or trailing edge of the turbine rotor blade 100 avoiding being arranged at a mid-chord position of the blade. And the specific embodiment determines that the axial distance between the installation axis of the optical fiber sensor 1 and the blade chord length central line of the turbine rotor blade 100 is 1.5-2 mm through carrying out axial dimension chain calculation and analysis on the cooling jacket structure. The analysis method comprises the following steps: defining the axial distance between the installation axis of the optical fiber sensor 1 and the chord length center line of the blade as
By means of an axial dimension chain calculation analysis, according to the formula +.>
The axial distance +_ of the installation of the optical fiber sensor 1 can be determined>
Wherein->
Mounting the optical fiber sensor 1 with an axial distance of the axis from the leading edge blade tip of the turbine rotor blade 100,/>
For the tip axial dimension of the turbine rotor blade 100, < >>
Is the axial movement of the rotor of the engine.
Radial mountingPosition: in the present embodiment, the probe position of the optical fiber sensor 1 is required to be as close to the blade tip as possible, and the gap between the probe and the blade tip of the optical fiber sensor 1 is determined by performing signal calibration operation on each component
. Specifically, the clearance between the probe and the tip of the optical fiber sensor 1 +.>
By the formula->
Calculated, wherein->
For the thermal state variation of the rotor-stator clearance of the tip of the engine,/-for the engine>
The calculated radial distance for the cold dimension chain can be used to calculate the clearance between the probe of the fibre-optic sensor 1 and the tip of the turbine rotor blade 100>
3 to 5mm.
The invention is based on the system engineering thought to carry out system optimization on the non-contact dynamic stress measurement technology of the high-pressure turbine of the aeroengine, designs a set of test system, comprises a composite optical fiber sensor (composed of a cooling structure and an optical fiber sensor), breaks through the technical problem of non-contact dynamic stress measurement of the rotor of the high-pressure turbine of the aeroengine in a high-temperature domain by establishing a high-efficiency cooling method and an installation design method on the composite optical fiber sensor, and can ensure the test requirement of non-contact dynamic stress to be realized above 1800K.
The test system designed by the invention is applied to the development process of a certain turbofan engine, and the result shows that the test system can fill the blank of non-contact dynamic stress measurement of the high-pressure turbine rotor of the aeroengine in a high-temperature area environment, and has important technical value.
The foregoing description of the preferred embodiments of the invention is not intended to be limiting, but rather is intended to cover all modifications, equivalents, alternatives, and improvements that fall within the spirit and scope of the invention.
Furthermore, it should be understood that although the present disclosure describes embodiments, not every embodiment is provided with a separate embodiment, and that this description is provided for clarity only, and that the disclosure is not limited to the embodiments described in detail below, and that the embodiments described in the examples may be combined as appropriate to form other embodiments that will be apparent to those skilled in the art.
Claims (10)
1. The testing system for the non-contact dynamic stress of the turbine rotor in the high temperature range is characterized by comprising a composite optical fiber sensor positioned at the position of a turbine rotor blade in a casing, wherein the composite optical fiber sensor comprises an optical fiber sensor (1), a first cooling channel (10) is arranged in the optical fiber sensor (1), one end of the first cooling channel (10) is communicated with an air supply system, and the other end of the first cooling channel is communicated with a main flow path of an engine;
the optical fiber sensor (1) is wrapped in a cooling sleeve structure, a second cooling channel (20) is formed between the cooling sleeve structure and the optical fiber sensor (1), and a cold air inlet (2) and an exhaust hole (3) which are used for communicating the second cooling channel (20) with a main flow path of an engine are formed in the cooling sleeve structure;
the cooling jacket structure is arranged on the casing, and a gap between the cooling jacket structure and the casing forms a third cooling channel (30).
2. The system for testing the non-contact dynamic stress of the high-temperature-domain turbine rotor according to claim 1, wherein rack bleed air or engine high-pressure and low-temperature part bleed air flows in the first cooling channel (10), and two cold air flows flow in the second cooling channel (20) and the third cooling channel (30).
3. The system for testing the non-contact dynamic stress of the high-temperature-domain turbine rotor according to any one of claims 1 or 2, wherein the cooling jacket structure is fixed on the casing through a fixing piece and comprises a mounting bracket (4), a cooling jacket (6) and a mounting pressing plate (5) which are assembled in sequence from top to bottom.
4. A system for testing non-contact dynamic stress of a high temperature domain turbine rotor according to claim 3, wherein the backflow margin of the outlet of the cooling jacket (6) is not less than 1.4, and the backflow margin of the outlet of the optical fiber sensor (1) is not less than 1.2.
5. A system for testing the non-contact dynamic stress of a high-temperature-domain turbine rotor according to claim 3, wherein a plurality of exhaust holes (3) are formed, and the exhaust holes (3) are formed on the side wall and/or the bottom wall of the cooling sleeve (6);
and when the bottom wall of the cooling sleeve (6) is provided with a plurality of exhaust holes (3), the exhaust holes (3) are distributed in an eccentric array.
6. A system for testing the non-contact dynamic stress of a high temperature domain turbine rotor according to claim 3, wherein the cold air inlet (2) is positioned at the upper opening position of the cooling jacket (6).
7. The high temperature domain turbine rotor non-contact dynamic stress testing system according to claim 1, wherein the optical fiber sensor (1) is located at the front edge or the tail edge of the turbine rotor blade (100), and the axial distance between the installation axis of the optical fiber sensor (1) and the chord length center line of the turbine rotor blade (100) is 1.5-2 mm.
8. The high temperature domain turbine rotor non-contact dynamic stress testing system according to claim 1, wherein a gap between a probe of the optical fiber sensor (1) and a tip of the turbine rotor blade (100) is 3-5 mm.
9. The system for testing the non-contact dynamic stress of the high-temperature-domain turbine rotor according to claim 1, wherein a thermocouple is further arranged on the optical fiber sensor (1).
10. The system for testing the non-contact dynamic stress of the high-temperature-domain turbine rotor according to any one of claims 7 to 9, wherein a plurality of the composite optical fiber sensors are distributed unevenly in the circumferential direction in the casing.
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