Aero-engine compressor rotor blade dynamic stress measurement structure

Abstract

The invention discloses a dynamic stress measurement structure for a compressor rotor blade of an aero-engine, which is used for measuring the dynamic stress of a compressor rotor blade of an aero-engine. The compressor rotor of the aero-engine comprises a central pull rod and a slip ring lead adapter shaft, a transfer clamping nut, a front shaft neck, a first-stage rotor, a second-stage rotor and a third-stage rotor which are sequentially arranged along the axial direction of the central pull rod. The transfer clamping nut is provided with a plurality of ventilation grooves at intervals along the circumferential direction. The dynamic stress measurement structure for the compressor rotor blade of an aero-engine comprises: a lead ring, which is used to be wrapped around the wheel center of a blade disk spoke plate of the third-stage rotor; a wire pressing ring, which is used to be arranged between the front shaft neck and an inner stop of the transfer clamping nut and is axially clamped into an inner hole of the front shaft neck. The wire pressing ring is provided with a lead channel which is arranged offset with the ventilation groove in the circumferential direction and penetrates the wire pressing ring axially.

Description

Technical Field

[0001] The invention relates to the technical field of compressor dynamic stress measurement, in particular to a dynamic stress measurement structure for a compressor rotor blade of an aero-engine. Background Art

[0002] Military standards and civil aviation airworthiness regulations stipulate that during the development of aviation gas engines, rotor blade dynamic stress measurement tests are required. The measured data provides a basis for judging the strength, life and reliability of the blades.

[0003] like Figure 1 As shown, the dynamic stress test of the compressor rotor blade of an aircraft engine is usually carried out on the compressor rotor blade. The strain gauge is arranged on the surface of the rotor blade. The strain gauge test line is led to the rotor wheel center spoke plate through a hole punched at the hub near the blade measuring point, and is fixed along the wall by pressing the plate or applying glue. It is led to the center tie rod through a hole punched at the center axis near the spoke plate, and extends forward along the center tie rod to the slip ring inductor welding line. Because engine parts have strict weight and strength requirements, there is a long radial suspension section between the wheel center of the three-stage rotor blade disk spoke plate and the outer wall of the center tie rod of the existing compressor. According to the requirements of the air system, the rear bearing cavity of the turbine section needs to be ventilated to the front bearing cavity to balance the cavity pressure. Therefore, a ventilation groove is set on the adapter clamping nut, and the strain gauge test line needs to pass through the ventilation groove, which results in a long axial suspension section in the adapter clamping nut. The excessively long suspension section will cause the strain gauge test line to withstand not only the high temperature in the aircraft engine, but also the centrifugal force after the compressor rotor is running. The longer the suspension section, the greater the centrifugal force the strain gauge test line is subjected to, and the greater the possibility that the strain gauge test line is broken by the centrifugal force. The disconnection and loss of signal of the strain gauge test line is the main reason for the failure of the dynamic stress test of the aircraft engine compressor rotor blade.

[0004] At present, in order to achieve the test objectives, the following two measures are mainly taken: 1. Repeated tests to increase the number of successful tests. 2. Redesign the rotor disc structure, center rod structure and ventilation structure to reduce the suspended section of the strain gauge test line to increase the success rate of the test. However, the above two methods have long test cycles, high test costs, and low test efficiency, which cannot meet existing needs. Summary of the invention

[0005] The invention provides an aero-engine compressor rotor blade dynamic stress measurement structure to solve the technical problems of long compressor dynamic stress measurement test cycle, high test cost and low test efficiency.

[0006] According to one aspect of the present invention, a dynamic stress measurement structure for an aircraft engine compressor rotor blade is provided, which is used for measuring the dynamic stress of an aircraft engine compressor rotor blade. The aircraft engine compressor rotor comprises a central pull rod and a slip ring lead adapter shaft, an adapter clamping nut, a front shaft neck, a first-stage rotor, a second-stage rotor and a third-stage rotor which are arranged in sequence along the axial direction of the central pull rod. The adapter clamping nut is provided with a plurality of ventilation grooves spaced apart along the circumferential direction. The dynamic stress measurement structure for the aircraft engine compressor rotor blade comprises: a lead ring, which is used to wrap around the wheel center of the blade disk spoke plate of the third-stage rotor; a wire pressing ring, which is used to be arranged between the front shaft neck and the inner stop of the adapter clamping nut and axially inserted into the inner hole of the front shaft neck. The wire pressing ring is provided with a lead channel which is arranged offset with the ventilation groove in the circumferential direction and passes through the wire pressing ring axially.

[0007] As a further improvement of the above technical solution:

[0008] Furthermore, the lead ring includes an annular lead body, and the outer peripheral wall of the annular lead body is provided with two axial ribs extending outward in the radial direction. The two axial ribs are arranged at both ends of the annular lead body in the axial direction and are surrounded by the outer peripheral wall of the annular lead body to form a wrapping groove on the blade disk spoke wheel center of the three-stage rotor.

[0009] Furthermore, the radial outer edge of the axial rib is provided with a lead wire anti-rotation boss extending inwardly in the axial direction and used for clamping with the blade disk spoke plate of the three-stage rotor to prevent rotation.

[0010] Furthermore, the annular lead body is provided with circumferential notches penetrating in radial and axial directions respectively.

[0011] Furthermore, a radial groove is recessed outwardly in the radial direction on the inner circumferential wall of the annular lead body, the radial groove and the circumferential notch are staggered in the circumferential direction of the annular lead body, and a plurality of radial grooves penetrate the annular lead body axially and are arranged at intervals along the circumference of the annular lead body.

[0012] Furthermore, the wire pressing ring includes an annular wire pressing body, which includes a first ring portion for being arranged between the front shaft neck and the inner stopper of the adapter clamping nut, and a second ring portion extending along the axial direction and for being inserted into the inner hole of the front shaft neck and having an interference fit with the inner hole of the front shaft neck is provided on the radial inner end portion of the first ring portion. A lead hole is axially opened in the first ring portion, and a lead groove is radially recessed on the radial outer wall surface of the second ring portion. The lead groove axially penetrates the second ring portion and is coaxially arranged with the lead hole to be interconnected with the lead hole to form a lead channel.

[0013] Furthermore, a plurality of lead wire holes are arranged at intervals along the circumference of the first ring portion, and a plurality of lead wire grooves are arranged at intervals along the circumference of the second ring portion, and the lead wire holes and the lead wire grooves are arranged in one-to-one correspondence.

[0014] Furthermore, a wire pressing and anti-rotation boss extending in the axial direction and used for engaging with the front journal to prevent rotation is provided on the axial end surface of the first ring portion.

[0015] Furthermore, a wire pressing locking plate groove is radially recessed on the radial outer wall surface of the first ring portion, and the wire pressing locking plate groove axially penetrates the first ring portion and is used to communicate with the locking plate groove on the front journal to jointly accommodate the locking plate.

[0016] Furthermore, a plurality of wire pressing locking plate grooves are arranged at intervals along the circumferential direction of the first ring portion.

[0017] The present invention has the following beneficial effects:

[0018] The invention relates to a structure for measuring dynamic stress of rotor blades of an aero-engine compressor. A lead ring is wrapped around the blade disk spoke wheel center of a three-stage rotor, so that a strain gauge test line corresponding to a measuring point on the three-stage rotor is extended along the blade disk spoke of the three-stage rotor to the lead ring, and is fixed to the lead ring through a fixing structure, thereby greatly shortening a radial suspension section between the blade disk spoke wheel center of the three-stage rotor and an outer wall surface of a central tie rod, greatly reducing the risk of radial suspension breaking, and then entering the central tie rod, and strain gauge test lines corresponding to measuring points on the first-stage rotor and the second-stage rotor enter the central tie rod according to the existing wiring method, the wire pressing ring is arranged between the front shaft neck and the inner stop of the adapter clamping nut and is axially clamped into the inner hole of the front shaft neck, so that the strain gauge test line enters the adapter clamping nut under the guidance of the lead channel, and then enters the adapter shaft of the slip ring induction device. Due to the lead channel and the ventilation groove The lead channel is staggered in the circumferential direction and penetrates the wire pressing ring axially, so that the strain gauge test line passes through the wire pressing ring axially through the lead channel and is staggered with the ventilation groove on the adapter clamping nut, so as to avoid the strain gauge test lead falling into the ventilation groove of the adapter clamping nut and causing axial suspension and breaking. When the aircraft engine compressor is running, the strain gauge test line will be close to the solid part of the adapter clamping nut, so as to fundamentally avoid the risk of axial suspension and breaking. The present scheme shortens the radial suspension section by the lead ring and avoids the axial suspension section by the wire pressing ring. Compared with the prior art, the risk of disconnection and loss of signal of the strain gauge test line is avoided or even eliminated to the greatest extent, thereby improving the success rate of the dynamic stress measurement test of the rotor blade of the aircraft engine compressor, shortening the test cycle, reducing the test cost, ensuring the test efficiency, and having strong practicality, and being suitable for wide promotion and application.

[0019] In addition to the above-described purposes, features and advantages, the present invention has other purposes, features and advantages. The present invention will be further described in detail with reference to the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] The drawings constituting a part of this application are used to provide a further understanding of the present invention. The exemplary embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute an improper limitation of the present invention. In the drawings:

[0021] Figure 1 It is a schematic diagram of the structural layout of a dynamic stress measurement structure of an aircraft engine compressor rotor blade in the prior art;

[0022] Figure 2 A schematic diagram of the structural layout of a dynamic stress measurement structure for an aircraft engine compressor rotor blade according to a preferred embodiment of the present invention;

[0023] Figure 3 It is a structural schematic diagram of a lead ring in a dynamic stress measurement structure of an aircraft engine compressor rotor blade according to a preferred embodiment of the present invention;

[0024] Figure 4 It is a schematic diagram of the local structure after the lead ring is assembled in the dynamic stress measurement structure of the compressor rotor blade of an aircraft engine according to the preferred embodiment of the present invention;

[0025] Figure 5 It is a structural schematic diagram of a pressure wire ring in a dynamic stress measurement structure of an aircraft engine compressor rotor blade according to a preferred embodiment of the present invention;

[0026] Figure 6 It is a partial structural schematic diagram of an aircraft engine compressor rotor blade dynamic stress measurement structure after the pressure wire ring is assembled in the preferred embodiment of the present invention.

[0027] Legend:

[0028] 100, lead wire ring; 110, axial rib; 120, wrapping groove; 130, lead wire anti-rotation boss; 140, circumferential notch; 150, radial groove; 200, wire pressing ring; 210, lead wire channel; 220, first ring portion; 230, second ring portion; 240, wire pressing anti-rotation boss; 250, wire pressing locking plate groove. DETAILED DESCRIPTION

[0029] The embodiments of the present invention are described in detail below with reference to the accompanying drawings. However, the present invention can be implemented in many different ways as defined and covered below.

[0030] Figure 1 It is a schematic diagram of the structural layout of a dynamic stress measurement structure of an aircraft engine compressor rotor blade in the prior art; Figure 2 A schematic diagram of the structural layout of a dynamic stress measurement structure for an aircraft engine compressor rotor blade according to a preferred embodiment of the present invention; Figure 3 It is a structural schematic diagram of a lead ring in a dynamic stress measurement structure of an aircraft engine compressor rotor blade according to a preferred embodiment of the present invention; Figure 4It is a schematic diagram of the local structure after the lead ring is assembled in the dynamic stress measurement structure of the compressor rotor blade of an aircraft engine according to the preferred embodiment of the present invention; Figure 5 It is a structural schematic diagram of a pressure wire ring in a dynamic stress measurement structure of an aircraft engine compressor rotor blade according to a preferred embodiment of the present invention; Figure 6 It is a partial structural schematic diagram of an aircraft engine compressor rotor blade dynamic stress measurement structure after the pressure wire ring is assembled in the preferred embodiment of the present invention.

[0031] like Figure 1-Figure 6As shown, the dynamic stress measurement structure of the compressor rotor blade of an aircraft engine in this embodiment is used for dynamic stress measurement of the compressor rotor blade of an aircraft engine. The compressor rotor of the aircraft engine includes a center pull rod and a slip ring lead adapter shaft, an adapter clamping nut, a front shaft neck, a first-stage rotor, a second-stage rotor and a third-stage rotor arranged in sequence along the axial direction of the center pull rod. The adapter clamping nut is provided with a plurality of ventilation grooves spaced apart along the circumferential direction. The dynamic stress measurement structure of the compressor rotor blade of an aircraft engine includes: a lead ring 100, which is used to be wrapped on the wheel center of the blade disk spoke plate of the third-stage rotor, a wire pressing ring 200, which is used to be arranged between the front shaft neck and the inner stop of the adapter clamping nut and axially inserted into the inner hole of the front shaft neck, and a lead channel 210 is provided on the wire pressing ring 200, which is arranged offset with the ventilation groove in the circumferential direction and axially penetrates the wire pressing ring 200. Specifically, in the dynamic stress measurement structure of the aircraft engine compressor rotor blade of the present invention, the lead ring 100 is wrapped on the blade disk spoke wheel center of the three-stage rotor, so that the strain gauge test line corresponding to the measuring point on the three-stage rotor extends along the blade disk spoke of the three-stage rotor to the lead ring 100, and is fixed on the lead ring 100 through a fixing structure, thereby greatly shortening the radial suspension section between the blade disk spoke wheel center of the three-stage rotor and the outer wall surface of the center pull rod, greatly reducing the risk of radial suspension breaking, and then entering the center pull rod, and the strain gauge test lines corresponding to the measuring points on the first-stage rotor and the second-stage rotor enter the center pull rod according to the existing wiring method, the wire pressing ring 200 is arranged between the front shaft neck and the inner stop of the adapter clamping nut and is axially inserted into the inner hole of the front shaft neck, so that the strain gauge test line enters the adapter clamping nut under the guidance of the lead channel 210, and then enters the slip ring lead adapter shaft, because the lead channel 210 and the through The air groove is staggered in the circumferential direction and axially penetrates the lead channel 210 of the wire pressing ring 200, so that the strain gauge test line passes through the lead ring 100 axially through the lead channel 210 and is staggered with the ventilation groove on the adapter clamping nut to avoid the strain gauge test lead falling into the ventilation groove of the adapter clamping nut and causing axial suspension and breaking. When the aircraft engine compressor is running, the strain gauge test line will be close to the entity part of the adapter clamping nut to fundamentally avoid the risk of axial suspension and breaking; this solution shortens the radial suspension section by the lead ring 100 and avoids the axial suspension section by the wire pressing ring 200. Compared with the prior art, by avoiding or even eliminating the risk of disconnection and loss of signal of the strain gauge test line to the greatest extent, the success rate of the dynamic stress measurement test of the rotor blade of the aircraft engine compressor is improved, so as to shorten the test cycle, reduce the test cost, ensure the test efficiency, and have strong practicality and are suitable for wide promotion and application. It should be understood that the fixed structure is a press sheet or a glue coating. It should be understood that the radial suspension section between the first-stage rotor blade disk spoke wheel center and the outer wall of the center tie rod and the radial suspension section between the second-stage rotor blade disk spoke wheel center and the outer wall of the center tie rod are relatively small and have little effect on the test results, so there is no need to install the lead ring 100.

[0032] like Figure 3 and Figure 4 As shown, in this embodiment, the lead ring 100 includes an annular lead body, and the outer peripheral wall of the annular lead body is provided with two axial ribs 110 extending outward in the radial direction. The two axial ribs 110 are arranged at both ends of the annular lead body in the axial direction and enclosed with the outer peripheral wall of the annular lead body to form a wrapping groove 120 for wrapping on the blade disk spoke wheel center of the three-stage rotor. Specifically, the wrapping groove 120 is formed by enclosing the two axial ribs 110 and the outer peripheral wall of the annular lead body to wrap on the blade disk spoke wheel center of the three-stage rotor, and the outer diameter of the axial ribs 110 is greater than the diameter of the blade disk spoke wheel center of the three-stage rotor, so as to prevent the lead ring 100 from being disengaged when the compressor is running.

[0033] like Figure 3 and Figure 4 As shown, in this embodiment, the radial outer edge of the axial rib 110 is provided with a lead wire anti-rotation boss 130 extending inwardly in the axial direction and used for engaging with the blade disk spoke of the three-stage rotor to prevent rotation. Specifically, the lead wire anti-rotation boss 130 is engaged with the blade disk spoke of the three-stage rotor to prevent the lead wire ring 100 from rotating circumferentially relative to the blade disk spoke of the three-stage rotor and tearing the strain gauge test line. Correspondingly, the blade disk spoke of the three-stage rotor is provided with a lead wire anti-rotation groove for engaging with the lead wire anti-rotation boss 130.

[0034] like Figure 3 and Figure 4 As shown, in this embodiment, the annular lead body is provided with a circumferential notch 140 penetrating in the radial direction and the axial direction. Specifically, by providing the annular lead body with the circumferential notch 140 penetrating in the radial direction and the axial direction, the lead ring 100 can reduce the outer diameter of the axial flange 110 by circumferentially squeezing the circumferential notch 140, so that the lead ring 100 can smoothly wrap the blade disk spoke wheel center of the three-stage rotor, and when the compressor is running, the lead ring 100 will be closely attached to the blade disk spoke wheel center of the three-stage rotor under the action of centrifugal force without being loosened.

[0035] like Figure 3 and Figure 4 As shown, in this embodiment, a radial groove 150 is concavely formed outwardly along the radial direction on the inner circumferential wall of the annular lead body, the radial groove 150 and the circumferential notch 140 are staggered in the circumferential direction of the annular lead body, and a plurality of radial grooves 150 penetrate the annular lead body in the axial direction and are arranged at intervals along the circumference of the annular lead body. Specifically, the circumferential rigidity of the annular lead body is reduced by the plurality of radial grooves 150, so as to circumferentially squeeze the circumferential notch 140 to reduce the outer diameter of the axial retaining edge 110.

[0036] like Figure 5 and Figure 6As shown, in this embodiment, the wire pressing ring 200 includes an annular wire pressing body, which includes a first ring portion 220 for being arranged between the front shaft neck and the inner stop of the adapter clamping nut, and a second ring portion 230 extending in the axial direction and formed on the radial inner end of the first ring portion 220 for being inserted into the inner hole of the front shaft neck and having an interference fit with the inner hole of the front shaft neck, and the first ring portion 220 is provided with a lead hole in the axial direction, and a lead groove is recessed in the radial direction on the radial outer wall surface of the second ring portion 230, and the lead groove passes through the second ring portion 230 in the axial direction and is arranged coaxially with the lead hole to communicate with the lead hole to form a lead channel 210. Specifically, the axial positioning of the annular wire pressing body is achieved by the first ring portion 220, and the radial positioning is achieved by the second ring portion 230, and the lead hole is opened on the first ring portion 220, and the lead groove is opened on the second ring portion 230, so that the lead hole and the lead groove form a lead channel 210 to guide the strain gauge test line to pass. Optionally, the adapter clamping nut includes a nut body and an adapter connected to the nut body, and the vent groove is opened on the nut body and penetrates the nut body radially.

[0037] like Figure 5 and Figure 6 As shown, in this embodiment, a plurality of lead holes are arranged at intervals along the circumference of the first ring portion 220, and a plurality of lead grooves are arranged at intervals along the circumference of the second ring portion 230, and the lead holes and the lead grooves are arranged one by one. Specifically, a plurality of lead channels 210 arranged at intervals along the circumference of the annular wire pressing body are formed by the plurality of lead holes arranged at intervals along the circumference of the first ring portion 220 and the plurality of lead grooves arranged at intervals along the circumference of the second ring portion 230, so as to guide the strain gauge test lines corresponding to the measuring points at different circumferential positions on the compressor rotor.

[0038] like Figure 5 and Figure 6 As shown, in this embodiment, a wire pressing and anti-rotation boss 240 extending in the axial direction and used for engaging and anti-rotation with the front axle neck is provided on the axial end face of the first ring portion 220. Specifically, the wire pressing and anti-rotation boss 240 is engaged with the front axle neck to prevent the first ring portion 220 from rotating relative to the front axle neck, to prevent the strain gauge test line from being torn off, and to determine the circumferential positioning relationship between the wire pressing ring 200 and the front axle neck. Therefore, the circumferential positioning relationship between the wire pressing ring 200 and the adapter clamping nut can be determined, so that the lead wire channel 210 and the lead wire groove are circumferentially staggered. Correspondingly, a wire pressing and anti-rotation groove that engages with the wire pressing and anti-rotation boss 240 is provided on the front axle neck.

[0039] like Figure 5 and Figure 6As shown, in this embodiment, a wire pressing locking plate groove 250 is concavely provided in the radial direction on the radial outer wall surface of the first ring portion 220, and the wire pressing locking plate groove 250 penetrates the first ring portion 220 in the axial direction and is used to communicate with the locking plate groove on the front journal to jointly accommodate the locking plate. Specifically, by providing the wire pressing locking plate groove 250 on the first ring portion 220, so that the wire pressing locking plate groove 250 penetrates the first ring portion 220 in the axial direction and communicates with the locking plate groove on the front journal, the wire pressing locking plate groove 250 and the locking plate groove on the front journal jointly accommodate the locking plate to determine the circumferential positioning relationship between the wire pressing ring 200 and the front journal, and prevent the first ring portion 220 from rotating relative to the front journal.

[0040] like Figure 5 As shown, in this embodiment, a plurality of wire pressing locking plate grooves 250 are arranged at intervals along the circumference of the first ring portion 220. Specifically, a plurality of wire pressing locking plate grooves 250 and a plurality of locking plate grooves on the front journal are used to accommodate a plurality of locking plates, so as to further ensure a reliable assembly relationship between the wire pressing ring 200 and the primary rotor. Optionally, two wire pressing locking plate grooves 250 are arranged.

[0041] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, the present invention may have various modifications and variations. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A dynamic stress measurement structure for an aircraft engine compressor rotor blade, used for dynamic stress measurement of an aircraft engine compressor rotor blade, the aircraft engine compressor rotor comprises a central tie rod and a slip ring lead adapter shaft, an adapter clamping nut, a front journal, a first-stage rotor, a second-stage rotor and a third-stage rotor arranged in sequence along the axial direction of the central tie rod, the adapter clamping nut is provided with a plurality of ventilation grooves spaced apart along the circumferential direction, characterized in that: The dynamic stress measurement structure of the aircraft engine compressor rotor blade includes: A lead ring (100) is used for wrapping around the blade disk spoke wheel center of the three-stage rotor; The lead ring (100) comprises an annular lead body, and the outer peripheral wall of the annular lead body is provided with two axial ribs (110) extending outward in the radial direction, the two axial ribs (110) are arranged at both ends of the annular lead body in the axial direction and are enclosed with the outer peripheral wall of the annular lead body to form a wrapping groove (120) for wrapping on the blade disk spoke wheel center of the three-stage rotor; A wire pressing ring (200) is arranged between the front journal and the inner stop of the adapter clamping nut and is axially inserted into the inner hole of the front journal. The wire pressing ring (200) is provided with a wire guide channel (210) which is arranged offset with the ventilation groove in the circumferential direction and axially penetrates the wire pressing ring (200); The wire pressing ring (200) comprises an annular wire pressing body, which comprises a first ring portion (220) for being arranged between the front shaft neck and the inner stop of the adapter clamping nut, and a second ring portion (230) extending in the axial direction and for being inserted into the inner hole of the front shaft neck and having an interference fit with the inner hole of the front shaft neck is provided on the radial inner end portion of the first ring portion (220). The first ring portion (220) is provided with a lead hole in the axial direction, and a lead groove is recessed in the radial direction on the radial outer wall surface of the second ring portion (230). The lead groove passes through the second ring portion (230) in the axial direction and is arranged coaxially with the lead hole to be connected with the lead hole to form a lead channel (210).

2. The structure for measuring dynamic stress of aero-engine compressor rotor blades according to claim 1, characterized in that: A lead wire anti-rotation boss (130) extending inwardly in the axial direction and used for clamping with the blade disk spoke plate of the three-stage rotor to prevent rotation is provided on the radial outer edge of the axial rib (110).

3. The structure for measuring dynamic stress of aero-engine compressor rotor blades according to claim 1, characterized in that: The annular lead body is provided with a circumferential notch (140) penetrating in the radial direction and the axial direction respectively.

4. The structure for measuring dynamic stress of a compressor rotor blade of an aero-engine according to claim 3, characterized in that: A radial groove (150) is recessed outwardly in the radial direction on the inner circumferential wall of the annular lead body. The radial groove (150) and the circumferential notch (140) are staggered in the circumferential direction of the annular lead body. A plurality of radial grooves (150) penetrate the annular lead body in the axial direction and are arranged at intervals along the circumference of the annular lead body.

5. The structure for measuring dynamic stress of a compressor rotor blade of an aero-engine according to any one of claims 1 to 4, characterized in that: A plurality of lead wire holes are arranged at intervals along the circumference of the first ring portion (220), and a plurality of lead wire grooves are arranged at intervals along the circumference of the second ring portion (230), and the lead wire holes and the lead wire grooves are arranged in a one-to-one correspondence.

6. The structure for measuring dynamic stress of a compressor rotor blade of an aero-engine according to any one of claims 1 to 4, characterized in that: A wire pressing and anti-rotation boss (240) extending in the axial direction and used for engaging with the front journal to prevent rotation is provided on the axial end surface of the first ring portion (220).

7. The structure for measuring dynamic stress of a compressor rotor blade of an aircraft engine according to any one of claims 1 to 4, characterized in that: A wire pressing locking plate groove (250) is radially recessed on the radial outer wall surface of the first ring portion (220), the wire pressing locking plate groove (250) axially penetrates the first ring portion (220) and is used to communicate with the locking plate groove on the front journal to jointly accommodate the locking plate.

8. The structure for measuring dynamic stress of aero-engine compressor rotor blades according to claim 7, characterized in that: A plurality of wire pressing locking plate grooves (250) are arranged at intervals along the circumference of the first ring portion (220).

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