CN221224140U - Aeroengine blade vibration testing device - Google Patents
Aeroengine blade vibration testing device Download PDFInfo
- Publication number
- CN221224140U CN221224140U CN202323123830.7U CN202323123830U CN221224140U CN 221224140 U CN221224140 U CN 221224140U CN 202323123830 U CN202323123830 U CN 202323123830U CN 221224140 U CN221224140 U CN 221224140U
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- 238000012360 testing method Methods 0.000 title claims abstract description 37
- 230000005284 excitation Effects 0.000 claims abstract description 18
- 230000005571 horizontal transmission Effects 0.000 claims abstract description 10
- 230000002093 peripheral effect Effects 0.000 claims abstract description 5
- 239000013307 optical fiber Substances 0.000 claims description 4
- 230000008878 coupling Effects 0.000 claims description 3
- 238000010168 coupling process Methods 0.000 claims description 3
- 238000005859 coupling reaction Methods 0.000 claims description 3
- 230000009286 beneficial effect Effects 0.000 abstract description 4
- 238000005507 spraying Methods 0.000 abstract description 2
- 230000005540 biological transmission Effects 0.000 description 2
- 238000000034 method Methods 0.000 description 2
- 230000001360 synchronised effect Effects 0.000 description 2
- 238000010586 diagram Methods 0.000 description 1
- 239000000835 fiber Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 239000007921 spray Substances 0.000 description 1
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- Testing Of Devices, Machine Parts, Or Other Structures Thereof (AREA)
Abstract
The utility model relates to an aeroengine blade vibration testing device, which comprises a bottom plate, a left support, a middle support, a right support, a horizontal transmission shaft, a casing, a sensor for testing and a gas excitation application assembly, and is technically characterized in that: the gas excitation applying assembly comprises a hollow shaft arranged on the right support, a plurality of radial hollow rods arranged on the outer peripheral surface of the hollow shaft and communicated with the hollow shafts, an annular distribution cavity communicated with the tail ends of the radial hollow rods, a plurality of gas distribution outlets uniformly arranged on one side end surface of the annular distribution cavity, a concentric shaft arranged at the front end of the hollow shaft, a plurality of radial support arms uniformly arranged on the concentric shaft, and movable nozzles arranged on the radial support arms, wherein the movable nozzles are in one-to-one correspondence with the gas distribution outlets, and flexible connecting pipes are arranged between the movable nozzles and the gas distribution outlets. The device flexibly adjusts the high-pressure airflow spraying position of the air excitation applying component on the blade, is suitable for testing the blade discs with different diameters, and is more beneficial to truly simulating the online exciting force environment of the aero-engine blade.
Description
Technical Field
The utility model relates to the technical field of aero-engine blade vibration testing, in particular to an aero-engine blade vibration testing device.
Background
Aeroengines are considered to be the "heart" of an aircraft. The engine blade is used as a core working component, and the running state, the working efficiency and the safety performance of the whole engine system are directly influenced. In actual operation, the engine blade needs to bear extremely complex exciting force, so that the engine blade is easy to vibrate, fatigue is caused, and even faults such as cracks and breaks occur, so that serious safety accidents are caused. It is necessary to perform vibration testing on the engine blade, and a common testing mode off-line is a non-contact vibration test.
CN 115876415A discloses a non-contact blade vibration test bench, which adopts high-pressure air flow or magnet as excitation load to synchronously or asynchronously excite the blade, and can freely adjust the blade tip distance and realize non-contact test on the vibration condition of the blade through an optical fiber sensor. Specifically, the motor drives the blades to rotate or the magnet sleeved on the bolt synchronously excites the blades, the motor drives the nozzle to rotate, and the nozzle sprays high-speed gas to synchronously excite or asynchronously excite the blades. The device can conveniently carry out synchronous excitation and asynchronous excitation on the blisk device, provides powerful support for the test research of the vibration characteristics of the blisk structure, and is convenient to use for the vibration test of the aero-engine blisk. However, the following problems still exist in the test process: the distance between the nozzle in the gas excitation applying assembly and the center of the blade disc is fixed but not adjustable, the radial ejection position of the high-pressure air flow relative to the blade cannot be adjusted, and the use is inflexible.
Disclosure of utility model
The utility model aims to provide the aero-engine blade vibration testing device which is reasonable in structure and reliable in use and is used for solving the problems, the high-pressure airflow spraying position of the air excitation applying assembly to the blade is flexibly adjusted, the device is suitable for testing the blades with different diameters, and the device is more beneficial to truly simulating the online exciting force environment of the aero-engine blade.
The technical scheme of the utility model is as follows:
The utility model provides an aeroengine blade vibration testing arrangement, includes the bottom plate, along left support, middle support and the right support that bottom plate length direction arranged in proper order, locates the horizontal transmission shaft on the left support, locates the cartridge receiver on the middle support, locates a plurality of test sensor on the cartridge receiver periphery wall, locates the gas excitation on the right support and applys the subassembly, the horizontal transmission shaft utilizes the shaft coupling to be connected with the pivot, the pivot end is equipped with the leaf disc that is arranged in the cartridge receiver in passing the cartridge receiver center and on it, and its technical essential is: the gas excitation applying assembly comprises a hollow shaft supported on a right support by a bearing, a plurality of radial hollow rods arranged on the peripheral surface of the hollow shaft and communicated with the hollow shaft, an annular distribution cavity communicated with the tail ends of the radial hollow rods, a plurality of gas distribution outlets uniformly arranged on one side end surface of the annular distribution cavity, a concentric shaft arranged at the front end of the hollow shaft, a plurality of radial support arms uniformly fixed on the concentric shaft, and movable nozzles arranged on the radial support arms and facing the leaf disc, wherein the number of the movable nozzles is equal to that of the gas distribution outlets and corresponds to that of the gas distribution outlets one by one, and flexible connecting pipes are arranged between the movable nozzles and the corresponding gas distribution outlets.
According to the aero-engine blade vibration testing device, the gear ring is fixed on the periphery of the rear portion of the hollow shaft, and the center of the rear end face of the hollow shaft is provided with the air inlet interface.
According to the aeroengine blade vibration testing device, the radial support arm is provided with the radial long holes, the outer wall of the movable nozzle is provided with the front clamping plate and the rear clamping plate, the central line of the movable nozzle is parallel to the horizontal plane, and the movable nozzle penetrates through the radial long holes and is fixedly connected with the radial support arm by the front clamping plate and the rear clamping plate.
According to the aero-engine blade vibration testing device, the central lines of the horizontal transmission shaft and the hollow shaft are overlapped.
According to the aero-engine blade vibration testing device, the sensor for testing is an optical fiber sensor.
The beneficial effects of the utility model are as follows:
The movable nozzle can radially move on the radial support arm, so that the radial position of the movable nozzle on the radial support arm can be changed as required, the radial ejection position of high-pressure airflow relative to the blades is changed, the use is flexible, the device is suitable for testing the blade discs with different diameters, and the device is more beneficial to truly simulating the online exciting force environment of the aero-engine blades.
Drawings
FIG. 1 is a schematic diagram of the structure of the present utility model;
FIG. 2 is a view in the direction A of FIG. 1;
Fig. 3 is a sectional view taken along the direction B-B in fig. 1.
In the figure: 1. left support, 2, horizontal drive shaft, 3, shaft coupling, 4, shaft, 5, casing, 6, test sensor, 7, movable nozzle, 8, radial arm, 9, flexible connecting tube, 10, annular distribution chamber, 11, right support, 12, gear ring, 13, air inlet interface, 14, hollow shaft, 15, bottom plate, 16, radial hollow rod, 17, air distribution outlet, 18, rear clamping plate, 19, front clamping plate, 20, concentric shaft, 21, bladed disk, 22, middle support, 23, radial long hole.
Detailed Description
The utility model will be described in detail with reference to the drawings.
As shown in fig. 1 to 3, the vibration testing device for the aeroengine blade comprises a bottom plate 15, a left support 1, a middle support 22 and a right support 11 which are sequentially arranged along the length direction of the bottom plate 15, a horizontal transmission shaft 2 arranged on the left support 1, a casing 5 arranged on the middle support 22, a plurality of test sensors 6 arranged on the peripheral wall of the casing 5, and a gas excitation applying assembly arranged on the right support 11. The horizontal transmission shaft 2 is connected with a rotating shaft 4 by a coupler 3, and the tail end of the rotating shaft 4 passes through the center of the casing 5 and is provided with a leaf disc 21 positioned in the casing 5.
The gas excitation applying assembly comprises a hollow shaft 14 supported on a right support 11 by a bearing, a plurality of radial hollow rods 16 arranged on the peripheral surface of the hollow shaft 14 and communicated with the hollow shaft, an annular distribution cavity 10 communicated with the tail ends of the radial hollow rods 16, a plurality of gas distribution outlets 17 uniformly arranged on one side end surface of the annular distribution cavity 10, a concentric shaft 20 arranged at the front end of the hollow shaft 14, a plurality of radial support arms 8 uniformly fixed on the concentric shaft 20, and movable nozzles 7 arranged on the radial support arms 8 and facing a leaf disk 21. The movable nozzles 7 are equal in number and in one-to-one correspondence with the gas distribution outlets 17, and flexible connecting pipes 9 are arranged between the movable nozzles 7 and the corresponding gas distribution outlets 17.
In this embodiment, a gear ring 12 is fixed on the outer periphery of the rear portion of the hollow shaft 14, and an air inlet port 13 is arranged in the center of the rear end face of the hollow shaft 14. The gear ring 12 is connected with a transmission mechanism, so that the hollow shaft 14 is driven to rotate. The radial support arm 8 is provided with a radial long hole 23, the outer wall of the movable nozzle 7 is provided with front and rear clamping plates 19 and 18, the central line of the movable nozzle 7 is parallel to the horizontal plane, and the movable nozzle 7 passes through the radial long hole 23 and is fixedly connected with the radial support arm 8 by the front and rear clamping plates 19 and 18. The center lines of the horizontal transmission shaft 2 and the hollow shaft 14 are coincident. The test sensor 6 is an optical fiber sensor.
Working principle:
In operation, the drive mechanism drives the horizontal drive shaft 2 to rotate, thereby driving the blade disc 21 to rotate through the coupler 3 and the rotating shaft 4.
When synchronous pneumatic excitation is needed, the hollow shaft 14 does not need to rotate, high-pressure gas enters the hollow shaft 14 through the air inlet interface 13, then enters the annular distribution cavity 10 through each radial hollow rod 16, finally reaches each movable nozzle 7 through each air distribution outlet 17 and the flexible connecting pipe 9, and each movable nozzle 7 applies gas excitation to the bladed disk 21. During this process, each fiber optic sensor monitors and tests blade vibration on the blisk 21.
When asynchronous pneumatic excitation is needed, the transmission mechanism and the gear ring 12 drive the hollow shaft 14 to rotate, so that each movable nozzle 7 is driven to rotate, and the asynchronous excitation of the blades on the blade disc 21 is realized after the rotary gas is sprayed out.
The foregoing describes the embodiments of the present utility model in detail, but the description is only a preferred embodiment of the present utility model and should not be construed as limiting the scope of the utility model. All equivalent changes and modifications within the scope of the present utility model are intended to fall within the scope of the present utility model.
Claims (5)
1. The utility model provides an aeroengine blade vibration testing arrangement, includes the bottom plate, along left support, middle support and the right support that bottom plate length direction arranged in proper order, locates the horizontal transmission shaft on the left support, locates the cartridge receiver on the middle support, locates a plurality of test sensor on the cartridge receiver periphery wall, locates the gas excitation on the right support and applys the subassembly, the horizontal transmission shaft utilizes the shaft coupling to be connected with the pivot, the pivot end is equipped with the leaf disc that is arranged in the cartridge receiver in passing the cartridge receiver center and on it, its characterized in that: the gas excitation applying assembly comprises a hollow shaft supported on a right support by a bearing, a plurality of radial hollow rods arranged on the peripheral surface of the hollow shaft and communicated with the hollow shaft, an annular distribution cavity communicated with the tail ends of the radial hollow rods, a plurality of gas distribution outlets uniformly arranged on one side end surface of the annular distribution cavity, a concentric shaft arranged at the front end of the hollow shaft, a plurality of radial support arms uniformly fixed on the concentric shaft, and movable nozzles arranged on the radial support arms and facing the leaf disc, wherein the number of the movable nozzles is equal to that of the gas distribution outlets and corresponds to that of the gas distribution outlets one by one, and flexible connecting pipes are arranged between the movable nozzles and the corresponding gas distribution outlets.
2. The aero-engine blade vibration testing device of claim 1, wherein: a gear ring is fixed on the periphery of the rear part of the hollow shaft, and an air inlet interface is arranged in the center of the rear end face of the hollow shaft.
3. The aero-engine blade vibration testing device of claim 1, wherein: the radial support arm is provided with a radial slot hole, the outer wall of the movable nozzle is provided with a front clamping plate and a rear clamping plate, the central line of the movable nozzle is parallel to the horizontal plane, and the movable nozzle passes through the radial slot hole and is connected and fixed with the radial support arm by the front clamping plate and the rear clamping plate.
4. The aero-engine blade vibration testing device of claim 1, wherein: the center lines of the horizontal transmission shaft and the hollow shaft are overlapped.
5. The aero-engine blade vibration testing device of claim 1, wherein: the test sensor is an optical fiber sensor.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202323123830.7U CN221224140U (en) | 2023-11-20 | 2023-11-20 | Aeroengine blade vibration testing device |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202323123830.7U CN221224140U (en) | 2023-11-20 | 2023-11-20 | Aeroengine blade vibration testing device |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN221224140U true CN221224140U (en) | 2024-06-25 |
Family
ID=91569903
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202323123830.7U Active CN221224140U (en) | 2023-11-20 | 2023-11-20 | Aeroengine blade vibration testing device |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN221224140U (en) |
-
2023
- 2023-11-20 CN CN202323123830.7U patent/CN221224140U/en active Active
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