Cyclic fatigue test device for aero-engine
By designing a cyclic fatigue testing device for aero-engines, and combining laser measurement and timing drive mechanisms, the problem that existing tests cannot accurately simulate the high- and low-cycle combined fatigue of aero-engines has been solved, resulting in more efficient and accurate test results.
Patent Information
- Application Number
- CN202510626210.X
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-05-15
- Publication Date
- 2025-08-01
AI Technical Summary
Existing single high-cycle fatigue or low-cycle fatigue tests cannot effectively simulate the high- and low-cycle combined fatigue damage process of aero-engines under actual installed conditions, resulting in inaccurate test results.
A cyclic fatigue testing device for aero-engines was designed, comprising a power control mechanism, a timing drive mechanism, and a laser measurement mechanism. The device measures the blade disk rotation speed using laser, calculates the number of cycles and dwell time of the resonant speed, and combines the timing drive to move the power push rod to achieve automated testing and efficient heat dissipation.
It improves the accuracy and efficiency of the test, reduces human error, lowers the workload of test personnel, and reduces test data error, enabling better simulation of fatigue damage of aero engines under compound cyclic fatigue.
Smart Images
Figure CN120404167A_ABST
Abstract
Description
Technical Field
[0001] The present invention relates to the technical field of aero-engine tests, and particularly to a cyclic fatigue test device for an aero-engine. Background Technique
[0002] The low-cycle fatigue test is to test the ability of a part to resist fatigue damage under cyclic loads with high stress and low frequency, and the high-cycle fatigue test is to test the ability of a part to resist fatigue damage under cyclic loads with small stress amplitude and high frequency. The test purposes and test methods of the two are different; During actual installation and use, some parts of the engine are simultaneously subjected to the action of low-cycle stress and high-cycle stress. However, a single high-cycle fatigue test or low-cycle fatigue test cannot effectively reflect the damage process of the parts, resulting in the fact that the high-low cycle composite fatigue method at the part level cannot fully simulate the whole-engine working conditions. Therefore, a cyclic fatigue test device for an aero-engine is needed to conduct high-low cycle composite fatigue tests at the whole-engine level. Summary of the Invention
[0003] The purpose of the present invention is to provide a cyclic fatigue test device for an aero-engine to solve the problems raised in the above background technique.
[0004] To achieve the above purpose, the present invention provides the following technical solution: A cyclic fatigue test device for an aero-engine, including a test platform, an aero-engine is arranged at the top end of the test platform, a compressor blade disk and a gas turbine blade disk are arranged inside the aero-engine, and further includes a power control mechanism, a timing drive mechanism and a laser measurement mechanism: The power control mechanism, the power control mechanism includes a support frame, a power control box and a transmission shaft. The support frame is arranged at the top end of the test platform, the power control box is arranged at the top end of the support frame, one end of the transmission shaft is arranged at one end of the power control box close to the aero-engine, and the other end is arranged inside the aero-engine; The timing drive mechanism, the timing drive mechanism includes a mounting frame, and the mounting frame is arranged above the power control box; The laser measurement mechanism, the laser measurement mechanism includes a laser speed measurement device and a laser detection head. The laser speed measurement device is arranged at the top end of the power control box. There are two laser detection heads, which are respectively arranged inside the aero-engine and at the top end of the power control box. The laser detection head is connected to the laser speed measurement device through a wire.
[0005] Preferably, the power control mechanism includes a power control box, a power push rod, a fixed frame, and a translation screw rod. The power control box is fixedly installed at the top of the power control box. The bottom end of the power push rod is movably installed inside the power control box. The fixed frame is fixedly installed at the top of the power control box and is located at one end of the power control box away from the aeroengine. The translation screw rod is movably installed at the top of the fixed frame.
[0006] Preferably, the power control mechanism includes a translation frame, a limit groove, a limit seat, a limit rod, and a screw sleeve. The translation frame is fixedly installed at one end of the translation screw rod close to the power control box. The top end of the power push rod is movably installed inside the translation frame. The limit groove is opened in the middle of the translation screw rod. The limit seat is fixedly installed at the top of the inner side of the fixed frame close to the power control box and is sleeved outside the translation screw rod. Both ends of the limit rod are fixedly connected to the limit seat. The limit rod is movably installed in the limit groove. The screw sleeve is movably installed at the top of the fixed frame away from the power control box. The screw sleeve is movably sleeved on the translation screw rod through threads.
[0007] Preferably, the timing drive mechanism includes a motor, a linkage shaft, a first gear, and a second gear. The motor is fixedly installed inside the mounting frame. The linkage shaft is movably installed on the side wall of the fixed frame and is coaxial with the output shaft of the motor. There is no connection between the linkage shaft and the output shaft of the motor. The first gear is fixedly installed at one end of the linkage shaft away from the aeroengine. The second gear is fixedly sleeved on the screw sleeve. The first gear is located below the screw sleeve and is movably connected to the second gear through meshing.
[0008] Preferably, the timing drive mechanism includes a coupling sleeve, a synchronous groove, and a connection groove. The coupling sleeve is movably sleeved on the output shaft of the motor and the linkage shaft. The synchronous groove is opened at one end of the coupling sleeve close to the motor. The connection groove is opened at one end of the coupling sleeve close to the first gear.
[0009] Preferably, the timing drive mechanism includes a first synchronous rod, a second synchronous rod, an electric push rod, and a translation plate. The first synchronous rod is fixedly installed on the output shaft of the motor. The first synchronous rod is movably installed in the synchronous groove. The second synchronous rod is fixedly installed on the linkage shaft. The second synchronous rod is movably connected to the connection groove. The electric push rod is fixedly installed inside the mounting frame. An expansion rod is arranged inside the electric push rod. The bottom end of the translation plate is fixedly connected to the expansion rod. The top end of the translation plate is movably sleeved on the outer side of the middle of the coupling sleeve.
[0010] Preferably, the laser measurement mechanism includes a heat dissipation box, a heat dissipation fan, and a first belt. The heat dissipation box is fixedly installed between the laser speed measurement device and the fixed frame. There are two heat dissipation fans in total, and they are movably installed inside the heat dissipation box through rotating shafts. Both ends of the first belt are movably sleeved on the rotating shafts of the two heat dissipation fans.
[0011] Preferably, the laser measurement mechanism includes a central rotating shaft, a first bevel gear, a second bevel gear, and a second belt. The central rotating shaft is movably installed at the top of the heat dissipation box and inside the mounting frame. The first bevel gear is fixedly sleeved on the output shaft of the motor. The second bevel gear is fixedly installed at one end of the central rotating shaft close to the output shaft of the motor. The second bevel gear is movably connected to the first bevel gear through meshing. The top end of the second belt is movably sleeved on the central rotating shaft, and the bottom end of the second belt is movably sleeved on the rotating shaft of a heat dissipation fan.
[0012] Compared with the prior art, the beneficial effects of the present invention are as follows: The present invention measures the resonance speed within the rotation speed range of the compressor blade disk and the gas turbine blade disk through laser, calculates the number of cycle tests and residence time of each resonance speed, and shortens the residence time in the maximum power state and increases the maximum continuous state, making up for the shortcomings that a single low-cycle and high-cycle test cannot fully simulate the actual installation and use, and testing the ability of the aero-engine to resist fatigue damage under compound cycle fatigue conditions. By driving the power push rod to move regularly, it not only reduces the manual operation error but also enables the test to be carried out automatically. Moreover, by efficiently dissipating heat from the laser speed measurement device that works for a long time, the device can be used for the cyclic fatigue test of the aero-engine for a long time, reducing the work intensity of the test personnel while reducing the error of test data and improving the test effect. Description of the Drawings
[0013] Figure 1 It is a schematic diagram of the overall structure provided by an embodiment of the present invention; Figure 2 It is a schematic diagram of the connection between the power control mechanism and the laser measurement mechanism provided by an embodiment of the present invention; Figure 3 It is a schematic diagram of the power control mechanism provided by an embodiment of the present invention; Figure 4 It is a schematic diagram of the cross-sectional structure of the translation threaded rod provided by an embodiment of the present invention; Figure 5 It is a schematic diagram of the timing drive mechanism provided by an embodiment of the present invention; Figure 6 It is a schematic diagram of the laser measurement mechanism provided by an embodiment of the present invention.
[0014] In the figure: 1. Test platform; 2. Aeroengine; 3. Power control mechanism; 301. Support frame; 302. Power control box; 303. Transmission shaft; 304. Power control box; 305. Power push rod; 306. Fixed frame; 307. Translation screw rod; 308. Translation frame; 309. Limit groove; 310. Limit seat; 311. Limit rod; 312. Nut sleeve; 4. Timing drive mechanism; 401. Mounting frame; 402. Motor; 403. Linkage shaft; 404. First gear; 405. Second gear; 406. Coupling sleeve; 407. Synchronous groove; 408. Connection groove; 409. First synchronous rod; 410. Second synchronous rod; 411. Electric push rod; 412. Translation plate; 5. Laser measurement mechanism; 501. Laser velocimeter; 502. Laser detection head; 503. Heat dissipation box; 504. Heat dissipation fan; 505. First belt; 506. Intermediate rotating shaft; 507. First bevel gear; 508. Second bevel gear; 509. Second belt. Specific embodiments
[0015] The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all the embodiments. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the protection scope of the present invention.
[0016] A cyclic fatigue test device for an aeroengine in this embodiment is as Figures 1 to 6 shown, including a test platform 1. An aeroengine 2 is arranged at the top end of the test platform 1. A compressor blade disk and a gas turbine blade disk are arranged inside the aeroengine 2. It further includes a power control mechanism 3, a timing drive mechanism 4 and a laser measurement mechanism 5. The power control mechanism 3 includes a support frame 301, a power control box 302 and a transmission shaft 303. The support frame 301 is arranged at the top end of the test platform 1. The power control box 302 is arranged at the top end of the support frame 301. One end of the transmission shaft 303 is arranged at one end of the power control box 302 close to the aeroengine 2, and the other end is arranged inside the aeroengine 2.
[0017] In this embodiment, as Figure 2 and Figure 3As shown in the figure, the power control mechanism 3 includes a power control box 304, a power push rod 305, a fixed bracket 306, and a translation screw rod 307. The power control box 304 is fixedly installed at the top of the power control box 302. The bottom end of the power push rod 305 is movably installed inside the power control box 304. The fixed bracket 306 is fixedly installed at the top of the power control box 302 and is located at one end of the power control box 304 away from the aeroengine 2. The translation screw rod 307 is movably installed at the top of the fixed bracket 306; By controlling the power push rod 305 to adjust the aeroengine 2 from the idle state to the maximum power state. During this process, it is necessary to drive the translation screw rod 307 to translate to drive the translation frame 308 to move, and drive the power push rod 305 to rotate through the translation frame 308.
[0018] In this embodiment, as Figure 3 and Figure 4 shown, the power control mechanism 3 includes a translation frame 308, a limit groove 309, a limit seat 310, a limit rod 311, and a screw sleeve 312. The translation frame 308 is fixedly installed at one end of the translation screw rod 307 close to the power control box 304. The top end of the power push rod 305 is movably installed inside the translation frame 308. The limit groove 309 is opened in the middle of the translation screw rod 307. The limit seat 310 is fixedly installed at the top of the inner side of the fixed bracket 306 close to the power control box 304 and is sleeved outside the translation screw rod 307. Both ends of the limit rod 311 are fixedly connected to the limit seat 310. The limit rod 311 is movably installed in the limit groove 309. The screw sleeve 312 is movably installed at the top of the fixed bracket 306 away from the power control box 304. The screw sleeve 312 is movably sleeved on the translation screw rod 307 through threads; By inserting the limit rod 311 into the limit groove 309, the limit of the translation screw rod 307 is realized. When the screw sleeve 312 rotates, it drives the translation screw rod 307 to move in the direction close to the aeroengine 2, so that when the translation screw rod 307 moves, it drives the power push rod 305 to rotate through the translation frame 308.
[0019] On other levels, this embodiment also provides a timing drive mechanism 4 for timing driving the power push rod 305 to adjust the aeroengine 2 from the idle state to the maximum power state for cyclic fatigue testing. As Figure 3 、 Figure 5 and Figure 6 shown, the timing drive mechanism 4 includes a mounting frame 401, and the mounting frame 401 is arranged above the power control box 302.
[0020] In this embodiment, as Figure 3 and Figure 5As shown in the figure, the timing drive mechanism 4 includes a motor 402, a linkage shaft 403, a first gear 404, and a second gear 405. The motor 402 is fixedly installed inside the mounting bracket 401. The linkage shaft 403 is movably installed on the side wall of the fixed bracket 306 and is coaxial with the output shaft of the motor 402. There is no connection between the linkage shaft 403 and the output shaft of the motor 402. The first gear 404 is fixedly installed at one end of the linkage shaft 403 away from the aero-engine 2. The second gear 405 is fixedly sleeved on the screw sleeve 312. The first gear 404 is located below the screw sleeve 312 and is movably connected to the second gear 405 through meshing. The first gear 404 is driven to rotate by the linkage shaft 403, and the screw sleeve 312 is driven to rotate by the second gear 405.
[0021] In this embodiment, as Figure 5 shown in the figure, the timing drive mechanism 4 includes a coupling sleeve 406, a synchronous groove 407, and a connection groove 408. The coupling sleeve 406 is movably sleeved on the output shaft of the motor 402 and the linkage shaft 403. The synchronous groove 407 is opened at one end of the coupling sleeve 406 close to the motor 402. The connection groove 408 is opened at one end of the coupling sleeve 406 close to the first gear 404. The first synchronous rod 409 on the output shaft of the motor 402 is inserted into the synchronous groove 407, so that the motor 402 drives the coupling sleeve 406 to rotate, and the coupling sleeve 406 rotates and moves at the same time.
[0022] In this embodiment, as Figure 5 shown in the figure, the timing drive mechanism 4 includes a first synchronous rod 409, a second synchronous rod 410, an electric push rod 411, and a translation plate 412. The first synchronous rod 409 is fixedly installed on the output shaft of the motor 402. The first synchronous rod 409 is movably installed in the synchronous groove 407. The second synchronous rod 410 is fixedly installed on the linkage shaft 403. The second synchronous rod 410 is movably connected to the connection groove 408. The electric push rod 411 is fixedly installed inside the mounting bracket 401. A telescopic rod is arranged inside the electric push rod 411. The bottom end of the translation plate 412 is fixedly connected to the telescopic rod. The top end of the translation plate 412 is movably sleeved on the outer side middle of the coupling sleeve 406. The electric push rod 411 drives the translation plate 412 to move towards the direction close to the linkage shaft 403. The translation plate 412 drives the coupling sleeve 406 to move away from the direction of the motor 402. The first synchronous rod 409 on the output shaft of the motor 402 is inserted into the synchronous groove 407, so that the motor 402 drives the coupling sleeve 406 to rotate, and the coupling sleeve 406 rotates and moves at the same time. Through the connection between the connection groove 408 and the second synchronous rod 410, the linkage shaft 403 follows the coupling sleeve 406 to rotate simultaneously through the second synchronous rod 410.
[0023] On other levels, this embodiment also provides a laser measurement mechanism 5 for measuring the rotational speed of the compressor blade disk and the gas turbine blade disk in the aeroengine 2 for a long time during the cyclic fatigue test of the aeroengine 2. As Figure 1 and Figure 6 shown, the laser measurement mechanism 5 includes a laser speed measurement device 501 and a laser detection head 502. The laser speed measurement device 501 is arranged at the top of the power control box 302. There are two laser detection heads 502, which are respectively arranged inside the aeroengine 2 and at the top of the power control box 302. The laser detection head 502 is connected to the laser speed measurement device 501 through a wire.
[0024] In this embodiment, as Figure 6 shown, the laser measurement mechanism 5 includes a heat dissipation box 503, a heat dissipation fan 504 and a first belt 505. The heat dissipation box 503 is fixedly installed between the laser speed measurement device 501 and the fixing frame 306. There are two heat dissipation fans 504, which are movably installed inside the heat dissipation box 503 through a rotating shaft. The two ends of the first belt 505 are respectively movably sleeved on the rotating shafts of the two heat dissipation fans 504; The laser is emitted by the laser detection head 502 to measure the rotational speed of the compressor blade disk and the gas turbine blade disk in the aeroengine 2. The laser speed measurement device 501 is used for data sending and receiving. During the cyclic fatigue test of the aeroengine 2, the heat dissipation box 503 dissipates heat from the laser speed measurement device 501 efficiently.
[0025] In this embodiment, as Figure 6 shown, the laser measurement mechanism 5 includes a middle rotating shaft 506, a first bevel gear 507, a second bevel gear 508 and a second belt 509. The middle rotating shaft 506 is movably installed at the top of the heat dissipation box 503 and inside the mounting frame 401. The first bevel gear 507 is fixedly sleeved on the output shaft of the motor 402. The second bevel gear 508 is fixedly installed at one end of the middle rotating shaft 506 close to the output shaft of the motor 402. The second bevel gear 508 is movably connected to the first bevel gear 507 through meshing. The top end of the second belt 509 is movably sleeved on the middle rotating shaft 506, and the bottom end of the second belt 509 is movably sleeved on the rotating shaft of a heat dissipation fan 504; The output shaft of the motor 402 drives the first bevel gear 507 to rotate. At this time, the second bevel gear 508 drives the middle rotating shaft 506 to rotate. When the middle rotating shaft 506 rotates, the heat dissipation fan 504 in the heat dissipation box 503 is driven to rotate through the second belt 509 and the first belt 505, so as to discharge the heat in the laser speed measurement device 501.
[0026] Working principle: When the present invention is in use, start the aero-engine 2, and then, according to the test requirements, adjust the aero-engine 2 from the idle state to the maximum power state by controlling the power push rod 305. During this process, it is necessary to drive the translation screw rod 307 to translate and drive the translation frame 308 to move, and drive the power push rod 305 to rotate through the translation frame 308; When it is necessary to drive the translation screw rod 307, first drive the translation plate 412 to move towards the direction close to the linkage shaft 403 by the electric push rod 411, drive the coupling sleeve 406 to move away from the motor 402 through the translation plate 412, insert the first synchronizing rod 409 on the output shaft of the motor 402 into the synchronizing groove 407, so that the motor 402 drives the coupling sleeve 406 to rotate. The coupling sleeve 406 rotates and moves at the same time, and is connected to the second synchronizing rod 410 through the connection groove 408, so that the linkage shaft 403 follows the coupling sleeve 406 to rotate simultaneously through the second synchronizing rod 410. Drive the first gear 404 to rotate through the linkage shaft 403, and drive the screw sleeve 312 to rotate through the second gear 405. Insert the limiting rod 311 into the limiting groove 309 to realize the limitation of the translation screw rod 307. When the screw sleeve 312 rotates, drive the translation screw rod 307 to move towards the direction close to the aero-engine 2, so that when the translation screw rod 307 moves, drive the power push rod 305 to rotate through the translation frame 308. According to the test process, by setting and controlling the extension and retraction time of the telescopic rod in the electric push rod 411, the power adjustment of the aero-engine 2 is controlled at regular intervals; In the cyclic fatigue test, emit laser through the laser detection head 502 to measure the rotational speeds of the compressor blade disk and the gas turbine blade disk in the aero-engine 2. The laser speed measurement device 501 is used for data transceiver. In the cyclic fatigue test of the aero-engine 2, the laser speed measurement device 501 needs to work for a long time, so a large amount of heat will be generated. The laser speed measurement device 501 is efficiently cooled by the heat dissipation box 503. During this process, drive the first bevel gear 507 to rotate through the output shaft of the motor 402. At this time, drive the middle rotating shaft 506 to rotate through the second bevel gear 508. When the middle rotating shaft 506 rotates, drive the cooling fan 504 in the heat dissipation box 503 to rotate through the second belt 509 and the first belt 505 to discharge the heat in the laser speed measurement device 501.
[0027] In the cyclic fatigue test of the aero-engine 2, first conduct the low-cycle fatigue test. The test is carried out 1500 times in total, and the time for one cycle is 15 minutes. The test run procedure is as follows: Then conduct the high-cycle fatigue test. Take 25 states on average between the idle state and the maximum power state of the aero-engine 2, including the idle state and the maximum power state, as the test points for the high-cycle fatigue test. Each state point stays for a certain time to ensure that the number of cycles is not less than 107 , the commissioning procedure is as follows: In the cyclic fatigue test of the aero-engine 2, the test is carried out A times in total, and the time for one cycle is 15 min. Among them, the value of A is not less than 2000. The resonance speeds between the idle state and the maximum power state of the compressor blade disk and the gas turbine blade disk are obtained through simulation calculation, a total of n: Ng-1, Ng-2, ···, Ng-n. According to the following formula, in order to reach 3x10 7 cycles, the residence time T1, T2, ···, Tn required for each speed can be obtained; The test is divided into n regions. The number of cycles in each region is the integer value, and the total number of cycles in each region is A. The residence time for each cycle at the resonance speed is .
[0028] The test spectrum for each region is as follows: It should be noted that in this article, relational terms such as first and second are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any actual relationship or order between these entities or operations. Moreover, the term "comprising", "including" or any other variation thereof is intended to cover non-exclusive inclusion, so that a process, method, article or device comprising a series of elements not only includes those elements, but also includes other elements not expressly listed, or also includes elements inherent to such process, method, article or device.
[0029] Although the embodiments of the present invention have been shown and described, those of ordinary skill in the art can understand that various changes, modifications, substitutions and variations can be made to these embodiments without departing from the principles and spirit of the present invention. The scope of the present invention is defined by the appended claims and their equivalents.
Claims
1. A cyclic fatigue test device for an aeroengine, comprising a test platform (1), wherein an aeroengine (2) is arranged at the top of the test platform (1), and a compressor blade disk and a gas turbine blade disk are arranged in the aeroengine (2), characterized in that, It also includes a power control mechanism (3), a timing drive mechanism (4) and a laser measurement mechanism (5): The power control mechanism (3), the power control mechanism (3) includes a support frame (301), a power control box (302) and a transmission shaft (303). The support frame (301) is arranged at the top of the test platform (1), the power control box (302) is arranged at the top of the support frame (301), one end of the transmission shaft (303) is arranged at one end of the power control box (302) close to the aero-engine (2), and the other end is arranged inside the aero-engine (2); The timing drive mechanism (4), the timing drive mechanism (4) includes a mounting frame (401), and the mounting frame (401) is arranged above the power control box (302); The laser measurement mechanism (5), the laser measurement mechanism (5) includes a laser speed measurement device (501) and a laser detection head (502). The laser speed measurement device (501) is arranged at the top of the power control box (302). There are two laser detection heads (502) which are respectively arranged inside the aero-engine (2) and at the top of the power control box (302). The laser detection head (502) is connected to the laser speed measurement device (501) through a wire.
2. The cyclic fatigue test device for an aero-engine according to claim 1, characterized in that: The power control mechanism (3) includes a power control box (304), a power push rod (305), a fixing frame (306) and a translation threaded rod (307). The power control box (304) is fixedly installed at the top of the power control box (302). The bottom end of the power push rod (305) is movably installed inside the power control box (304). The fixing frame (306) is fixedly installed at the top of the power control box (302) and is located at one end of the power control box (304) away from the aero-engine (2). The translation threaded rod (307) is movably installed at the top of the fixing frame (306).
3. The cyclic fatigue test device for an aeroengine according to claim 2, wherein: The power control mechanism (3) includes a translation frame (308), a limit groove (309), a limit seat (310), a limit rod (311) and a screw sleeve (312). The translation frame (308) is fixedly installed at one end of the translation threaded rod (307) close to the power control box (304). The top end of the power push rod (305) is movably installed inside the translation frame (308). The limit groove (309) is opened in the middle of the translation threaded rod (307). The limit seat (310) is fixedly installed at the top of the inner side of the fixing frame (306) close to the power control box (304) and is sleeved outside the translation threaded rod (307). Both ends of the limit rod (311) are fixedly connected to the limit seat (310). The limit rod (311) is movably installed in the limit groove (309). The screw sleeve (312) is movably installed at the top of the fixing frame (306) away from the power control box (304). The screw sleeve (312) is movably sleeved on the translation threaded rod (307) through a thread.
4. A cyclic fatigue test device for an aero-engine according to claim 3, characterized in that: The timing drive mechanism (4) includes a motor (402), a linkage shaft (403), a first gear (404) and a second gear (405). The motor (402) is fixedly installed inside the mounting frame (401). The linkage shaft (403) is movably installed on the side wall of the fixed frame (306) and is coaxial with the output shaft of the motor (402). There is no connection between the linkage shaft (403) and the output shaft of the motor (402). The first gear (404) is fixedly installed at one end of the linkage shaft (403) away from the aero-engine (2). The second gear (405) is fixedly sleeved on the screw sleeve (312). The first gear (404) is located below the screw sleeve (312) and is movably connected to the second gear (405) through meshing.
5. The cyclic fatigue test device for an aeroengine according to claim 4, wherein: The timing drive mechanism (4) includes a coupling sleeve (406), a synchronous groove (407) and a connection groove (408). The coupling sleeve (406) is movably sleeved on the output shaft of the motor (402) and the linkage shaft (403). The synchronous groove (407) is opened at one end of the coupling sleeve (406) close to the motor (402). The connection groove (408) is opened at one end of the coupling sleeve (406) close to the first gear (404).
6. The cyclic fatigue test device for an aero-engine according to claim 5, characterized in that: The timing drive mechanism (4) includes a first synchronous rod (409), a second synchronous rod (410), an electric push rod (411) and a translation plate (412). The first synchronous rod (409) is fixedly installed on the output shaft of the motor (402). The first synchronous rod (409) is movably installed in the synchronous groove (407). The second synchronous rod (410) is fixedly installed on the linkage shaft (403). The second synchronous rod (410) is movably connected to the connection groove (408). The electric push rod (411) is fixedly installed inside the mounting frame (401). An expansion rod is arranged inside the electric push rod (411). The bottom end of the translation plate (412) is fixedly connected to the expansion rod. The top end of the translation plate (412) is movably sleeved on the outer side middle of the coupling sleeve (406).
7. The cyclic fatigue test device for an aeroengine according to claim 6, characterized in that: The laser measurement mechanism (5) includes a heat dissipation box (503), a heat dissipation fan (504) and a first belt (505). The heat dissipation box (503) is fixedly installed between the laser speed measurement device (501) and the fixed frame (306). There are two heat dissipation fans (504) in total, which are movably installed inside the heat dissipation box (503) through a rotating shaft. The two ends of the first belt (505) are respectively movably sleeved on the rotating shafts of the two heat dissipation fans (504).
8. A cyclic fatigue test device for an aeroengine according to claim 7, characterized in that: The laser measurement mechanism (5) includes a middle rotating shaft (506), a first bevel gear (507), a second bevel gear (508) and a second belt (509). The middle rotating shaft (506) is movably installed at the top of the heat dissipation box (503) and inside the mounting frame (401). The first bevel gear (507) is fixedly sleeved on the output shaft of the motor (402). The second bevel gear (508) is fixedly installed at one end of the middle rotating shaft (506) close to the output shaft of the motor (402). The second bevel gear (508) is movably connected to the first bevel gear (507) through meshing. The top end of the second belt (509) is movably sleeved on the middle rotating shaft (506), and the bottom end of the second belt (509) is movably sleeved on the rotating shaft of a heat dissipation fan (504).