Dynamic stress measuring system

Abstract

The invention relates to the technical field of aero-engine detection, and discloses a dynamic stress measurement system, which comprises a strain gauge arranged on a blade of an aero-engine; the slip ring electricity leading device is connected with the strain gauge; the cooling assembly is provided with multiple layers of cooling channels, and the slip ring electricity leading device is arranged in the innermost layer of channel of the cooling assembly. According to the technical scheme, the multiple layers of cooling channels form multi-layer external cooling on the slip ring electricity leading device, the slip ring electricity leading device is effectively protected, and it is ensured that the slip ring electricity leading device can normally work in a high-temperature environment.

Description

Technical Field

[0001] The present invention relates to the technical field of aero-engine detection, and particularly to a dynamic stress measurement system. Background Art

[0002] In modern aero-engine tests, in order to obtain the stress level of blades under high-speed rotation, a blade dynamic stress measurement test is generally required. The specific steps are briefly described as follows: First, strain gauges are pasted on the blades, and then the strain gauges are connected to a slip ring electrical connector, and then the stress level of the blades under the rotating state is obtained. The slip ring electrical connector belongs to a relatively precise measuring device, and its working temperature is generally not allowed to exceed 130 °C. However, the exhaust temperature at the engine where the blades are located can generally reach 700 °C to 800 °C. The existing slip ring electrical connectors are difficult to work properly in the above high-temperature environment. Summary of the Invention

[0003] In view of this, the present invention provides a dynamic stress measurement system to solve the problem that the existing slip ring electrical connectors are difficult to work properly in a high-temperature environment when applied to blade dynamic stress measurement tests.

[0004] The present invention provides a dynamic stress measurement system suitable for installation on an aero-engine. The dynamic stress measurement system includes:

[0005] Strain gauges, arranged on the blades of the aero-engine;

[0006] Slip ring electrical connectors, connected to the strain gauges;

[0007] A cooling component, provided with multiple cooling channels, and the slip ring electrical connector is arranged in the innermost layer channel of the cooling component. Beneficial effects: By adopting the above technical solutions, the present application forms multiple layers of external cooling for the slip ring electrical connector through multiple cooling channels, effectively protecting the slip ring electrical connector and ensuring that the slip ring electrical connector can work properly in a high-temperature environment.

[0008] Optionally, it further includes:

[0009] Lead wires, connecting the strain gauges and the slip ring electrical connectors;

[0010] An extension section assembly is arranged between the slip ring current collector and the strain gauge, and the extension section assembly is arranged between the cooling assembly and the strain gauge, so that both the slip ring current collector and the cooling assembly are located outside the aeroengine; the lead wire passes through the inside of the extension section assembly. Beneficial effects: By adopting the above technical solution, the slip ring current collector can be further away from the high-temperature exhaust gas by arranging the extension section assembly between the aeroengine and the slip ring current collector, and at the same time, the cooling assembly with a large size can be prevented from blocking the exhaust of the aeroengine. Moreover, the technical solution of the present application does not introduce a new exhaust excitation source and does not affect the measurement accuracy and effectiveness of the blade dynamic stress.

[0011] Optionally, the cooling assembly includes:

[0012] An outer tube provided with an air inlet nozzle;

[0013] An inner tube is sleeved inside the outer tube at intervals, and a first channel is formed between the outer tube and the inner tube; a second channel is formed inside the inner tube; a first ventilation hole communicating the first channel and the second channel is provided on the inner tube; the slip ring current collector is located in the second channel;

[0014] The cooling assembly is adapted to introduce cooling air into the first channel from the air inlet nozzle to form a first layer of cooling air flow, and part of the cooling air enters the second channel from the first channel through the first ventilation hole to form a second layer of cooling air flow. Beneficial effects: By adopting the above technical solution, through the double reflux external cooling method, the cooling air forms two layers of external cooling for the slip ring current collector, preventing too much heat from being transferred to the slip ring current collector and effectively protecting the slip ring current collector.

[0015] Optionally, one end of the slip ring current collector close to the aeroengine is installed inside the first casing through a first flange structure; the first casing is sleeved outside the extension section assembly at intervals, and an annular first chamber is formed between the first casing and the extension section assembly; a second ventilation hole communicating the second channel and the first chamber is provided on the first flange structure; one end of the first casing away from the slip ring current collector is installed in the exhaust casing of the aeroengine through an annular support plate; a third ventilation hole communicating with the outside is provided on the support plate; the first chamber is communicated with the space inside the exhaust casing close to the inner side of the support plate;

[0016] The second layer of cooling air flow is adapted to enter the first chamber from the second channel through the second ventilation hole after cooling the slip ring current collector, and then be discharged to the outside through the third ventilation hole. Beneficial effects: By adopting the above technical solution, by setting the first chamber, etc., the flowing cooling air quickly transfers the heat of the first flange structure, significantly improving the cooling efficiency of the cooling air to effectively protect the slip ring current collector.

[0017] Optionally, an outer cylinder is sleeved outside the outer pipe at a position corresponding to the slip ring current collector at intervals, and the outer cylinder extends and is sleeved outside a part of the first casing at intervals; a third channel is formed between the outer cylinder and the first casing and between the outer cylinder and the outer pipe; one side of the third channel away from the aeroengine communicates with the outside; a second chamber is arranged inside the first casing on the side of the first flange structure close to the inside of the aeroengine; a fourth vent hole communicating the first channel with the second chamber is arranged on the first flange structure; a fifth vent hole communicating the second chamber with the third channel is arranged on the side of the second chamber close to the inside of the aeroengine;

[0018] The first layer of cooling air flow is adapted to enter the second chamber from the first channel through the fourth vent hole after cooling the slip ring current collector, then enter the third channel through the fifth vent hole, form a folded-back third layer of cooling air flow, and finally be discharged to the outside. Beneficial effects: By adopting the above technical solution, the present application forms three-layer external cooling of the cooling air on the slip ring current collector through a triple reflux external cooling method, preventing excessive heat from being transferred to the slip ring current collector and effectively protecting the slip ring current collector; and by arranging the second chamber and the like, the flowing cooling air quickly transfers the heat of the first flange structure, significantly improving the cooling efficiency of the cooling air to further effectively protect the slip ring current collector.

[0019] Optionally, the extension section assembly includes:

[0020] An extension shaft with a first through hole inside, and the lead wire passes through the first through hole; the extension shaft is installed in the second casing through a first bearing; one side of the second casing close to the inside of the aeroengine is connected to the bearing casing; the bearing casing is installed on the support plate through a second flange structure; a plurality of blades are installed on the turbine disk, and a second through hole is arranged at the center of the turbine disk, and the lead wire passes through the second through hole; one side of the turbine disk away from the inside of the aeroengine is supported and installed in the bearing casing through a second bearing; one end of the extension shaft inside the aeroengine is fixedly connected to the turbine disk.

[0021] Optionally, the high-temperature exhaust gas of the aeroengine is discharged into the exhaust ventilation pipe of the vehicle platform.

[0022] Optionally, both the inner pipe and the outer pipe are composed of multiple pipe fittings connected together. Beneficial effects: By adopting the above technical solution, the present application facilitates the processing and installation of the inner pipe and the outer pipe.

[0023] Optionally, a first cover plate for blocking the end is provided at one end of the inner tube away from the aeroengine, and a second cover plate for blocking the end is provided at one end of the outer tube away from the aeroengine; the air inlet nozzle is arranged on the side wall of the outer tube. Beneficial effects: By adopting the above technical solution, the present application prevents the leakage of the cooling gas by providing the first cover plate and the second cover plate for blocking the end.

[0024] Optionally, both the cooling oil circuit of the slip ring electrical connector and the signal line connected to the outside are led out from the first cover plate along the second channel. Description of the Drawings

[0025] In order to more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the following will briefly introduce the drawings required for the description of the specific embodiments or the prior art. Obviously, the drawings in the following description are some embodiments of the present invention. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without creative efforts.

[0026] Figure 1 It is a schematic cross-sectional structure diagram of the dynamic stress measurement system provided in the embodiment of the present invention;

[0027] Figure 2 It is a schematic cross-sectional structure diagram of the cooling component provided in the embodiment of the present invention;

[0028] Figure 3 It is a schematic cross-sectional structure diagram of the extension section component provided in the embodiment of the present invention;

[0029] Figure 4 It is a schematic partial cross-sectional structure diagram of the aeroengine provided in the embodiment of the present invention.

[0030] Description of the Reference Numerals:

[0031] 1. Cooling air; 2. First-layer cooling air flow; 3. Second-layer cooling air flow; 4. Exhaust ventilation pipe; 5. High-temperature exhaust gas; 6. Third-layer cooling air flow; 7. First chamber; 8. Cooling assembly; 9. Cooling oil circuit; 10. First cover plate; 11. Second cover plate; 12. First ventilation hole; 13. Outer tube; 14. Inner tube; 15. First channel; 16. Second channel; 17. Outer cylinder; 18. Slip ring electrical connector; 19. Third channel; 20. Second ventilation hole; 21. First casing; 22. Fourth ventilation hole; 23. Second chamber; 24. Fifth ventilation hole; 25. First flange structure; 26. Air inlet nozzle; 27. Extension section assembly; 28. Aeroengine; 29. First bearing; 30. Extension shaft; 31. Second casing; 32. First through hole; 33. Exhaust casing; 34. Support plate; 35. Turbine disk; 36. Strain gauge; 37. Lead wire; 38. Second through hole; 39. Second bearing; 40. Bearing casing; 41. Engine axis; 42. Second flange structure. Detailed implementation manners

[0032] To make the objectives, technical solutions and advantages of the embodiments of the present invention clearer, 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. Apparently, the described embodiments are some but not all of the embodiments of the present invention. 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.

[0033] As Figures 1 to 4 shown in a specific implementation manner of the dynamic stress measurement system, it includes: a strain gauge 36, a slip ring electrical connector 18 and a cooling assembly 8. The dynamic stress measurement system described in this application is suitable for being installed on the exhaust end of an aeroengine 28 for measuring the stress of the blades of the aeroengine 28. The high-temperature exhaust gas 5 of the aeroengine 28 is discharged to the exhaust ventilation pipe 4 of the vehicle platform through the inside of the exhaust casing 33.

[0034] As Figure 1 shown, the strain gauge 36 is arranged on the blades of the aeroengine 28; specifically, the strain gauge 36 can be pasted on the blades suitable for high-speed rotation. The slip ring electrical connector 18 is connected to the strain gauge 36 to obtain the stress level on the rotating blades. The cooling assembly 8 is provided with multiple layers of cooling channels, and the slip ring electrical connector 18 is arranged in the innermost layer channel of the cooling assembly 8.

[0035] Further, as Figure 1As shown in the figure, the dynamic stress measurement system of the present application further includes: a lead wire 37 and an extension section assembly 27. The lead wire 37 connects the strain gauge 36 and the slip ring electrical connector 18. The extension section assembly 27 is disposed between the slip ring electrical connector 18 and the strain gauge 36, and the extension section assembly 27 is disposed between the cooling assembly 8 and the strain gauge 36, such that both the slip ring electrical connector 18 and the cooling assembly 8 are located outside the aeroengine 28; the lead wire 37 passes through the inside of the extension section assembly 27. The slip ring electrical connector 18 is located in the high-temperature area outside the aeroengine 28.

[0036] Specifically, as Figure 1 and Figure 2 shown in the figure, the cooling assembly 8 includes: an outer tube 13 and an inner tube 14. The outer tube 13 is provided with an air inlet nozzle 26. The inner tube 14 is sleeved inside the outer tube 13 at intervals, and a first channel 15 is formed between the outer tube 13 and the inner tube 14; a second channel 16 is formed inside the inner tube 14; a first ventilation hole 12 communicating the first channel 15 and the second channel 16 is provided on the inner tube 14; the slip ring electrical connector 18 is located in the second channel 16. The cooling assembly 8 is adapted to introduce cooling air 1 from the air inlet nozzle 26 into the first channel 15 to form a first layer of cooling air flow 2, and part of the cooling air 1 enters the second channel 16 from the first channel 15 through the first ventilation hole 12 to form a second layer of cooling air flow 3. The second layer of cooling air flow 3 directly forms the first layer of cooling protection for the slip ring electrical connector 18; the first layer of cooling air flow 2 directly forms the second layer of cooling protection for the slip ring electrical connector 18.

[0037] One end of the slip ring electrical connector 18 close to the aeroengine 28 is installed inside the first casing 21 through a first flange structure 25; the first casing 21 is sleeved outside the extension section assembly 27 at intervals, and an annular first chamber 7 is formed between the first casing 21 and the extension section assembly 27; a second ventilation hole 20 communicating the second channel 16 and the first chamber 7 is provided on the first flange structure 25; one end of the first casing 21 far from the slip ring electrical connector 18 is installed in the exhaust casing 33 of the aeroengine 28 through an annular support plate 34; a third ventilation hole communicating with the outside is provided on the support plate 34; the first chamber 7 is communicated with the space inside the exhaust casing 33 close to the inner side of the support plate 34. The second layer of cooling air flow 3 is adapted to enter the first chamber 7 from the second channel 16 through the second ventilation hole 20 after cooling the slip ring electrical connector 18, and then be discharged to the outside through the third ventilation hole. Further, both the inner tube 14 and the outer tube 13 are composed of multiple pipe fittings connected together.

[0038] Further, as Figure 1 and Figure 2As shown, an outer cylinder 17 is sleeved outside the outer tube 13 at intervals corresponding to the position of the slip ring current collector 18, and the outer cylinder 17 extends and is sleeved outside a part of the first casing 21 at intervals; a third channel 19 is formed between the outer cylinder 17 and the first casing 21 and between the outer cylinder 17 and the outer tube 13; one side of the third channel 19 away from the aeroengine 28 is communicated with the outside; a second chamber 23 is provided inside the first casing 21 on the side of the first flange structure 25 close to the inside of the aeroengine 28; a fourth vent hole 22 communicating the first channel 15 with the second chamber 23 is provided on the first flange structure 25; a fifth vent hole 24 communicating the second chamber 23 with the third channel 19 is provided on the side of the second chamber 23 close to the inside of the aeroengine 28. The first layer of cooling air flow 2 is adapted to enter the second chamber 23 from the first channel 15 through the fourth vent hole 22 after cooling the slip ring current collector 18, then enter the third channel 19 through the fifth vent hole 24, form a folded-back third layer of cooling air flow 6, and finally be discharged into the exhaust ventilation pipe 4 of the vehicle platform. After being mixed with the high-temperature exhaust gas 5 of the aeroengine 28, it is discharged to the outside. A part of the first casing 21 is directly exposed to the high-temperature exhaust gas 5, and heat is transferred along the first casing 21 made of metal material to the first flange structure 25; since the existing first flange structure 25 is made of metal material, it is easy to transfer the heat of the high-temperature exhaust gas 5 to the slip ring current collector 18 in the form of metal heat transfer through the first flange structure 25, which is very likely to cause the temperature of the slip ring current collector 18 to rise and result in damage to the slip ring current collector 18. Therefore, the present application provides the above-mentioned first chamber 7 and second chamber 23, etc. During the flow of the cooling air 1, the heat transfer on the first casing 21 can be effectively carried away, and the heat of the first flange structure 25 can be quickly carried away to protect the slip ring current collector 18. The third layer of cooling air flow 6 directly forms the third layer of cooling protection for the slip ring current collector 18.

[0039] Specifically, as Figure 1 and Figure 2 shown, a first cover plate 10 for blocking the end is provided at one end of the inner tube 14 away from the aeroengine 28, and a second cover plate 11 for blocking the end is provided at one end of the outer tube 13 away from the aeroengine 28; the air inlet nozzle 26 is arranged on the side wall of the outer tube 13. The cooling oil circuit 9 of the slip ring current collector 18 and the signal line connected to the outside both lead out along the second channel 16 from the first cover plate 10 and are connected to the corresponding external structures; a sealing arrangement is adopted at the leading-out position. One end of the inner tube 14 and the outer tube 13 close to the aeroengine 28 are both fixedly connected to the first flange structure 25. The inner tube 14 and the outer tube 13 are L-shaped, and one end of the inner tube 14 and the outer tube 13 away from the aeroengine 28 penetrates from inside the exhaust ventilation pipe 4 to the outside.

[0040] Specifically, asFigure 1 and Figure 3 As shown, the extension section assembly 27 includes: an extension shaft 30. A first through hole 32 is provided inside the extension shaft 30, and the first through hole 32 is located at the center of the extension shaft 30. The lead wire 37 passes through the first through hole 32; the extension shaft 30 is installed in the second casing 31 through the first bearing 29; specifically, two first bearings 29 are respectively provided near the two ends of the extension shaft 30. Figure 1 and Figure 3 As shown, the side of the second casing 31 close to the inside of the aircraft engine 28 is connected to the bearing casing 40; specifically, the second casing 31 can be fixed to the flange structure of the bearing casing 40 by bolts. The bearing casing 40 is mounted on the support plate 34 through the second flange structure 42; a plurality of blades are mounted on the turbine disk 35, and a second through hole 38 is provided at the center of the turbine disk 35, and the lead wire 37 passes through the second through hole 38; the side of the turbine disk 35 away from the inside of the aircraft engine 28 is supported and installed in the bearing casing 40 by a second bearing 39; the turbine disk 35 can rotate at high speed along the engine axis 41. The end of the extension shaft 30 located inside the aircraft engine 28 is fixedly connected to the turbine disk 35 to ensure that the extension shaft 30 and the turbine disk 35 rotate synchronously.

[0041] The extension section assembly 27 described in the present application, on the one hand, moves the slip ring inductor 18 further away from the high temperature area, and at the same time prevents the cooling assembly 8 with a larger outer diameter from blocking the exhaust of the aircraft engine 28, thereby affecting the performance of the aircraft engine 28. In the present application, the outer diameter of the cooling assembly 8 is increased due to the design of multiple cooling channels for the cooling assembly 8. The dynamic stress measurement system described in the present application does not have a connection structure or cooling pipeline in the exhaust casing 33 of the aircraft engine 28, and there is no redundant exhaust excitation source, which does not affect the dynamic stress measurement of the blade.

[0042] Although the embodiments of the present invention have been described in conjunction with the accompanying drawings, those skilled in the art may make various modifications and variations without departing from the spirit and scope of the present invention, and such modifications and variations are all within the scope defined by the appended claims.

Claims

1. A dynamic stress measurement system, suitable for installation on an aircraft engine (28), characterized in that: include: A strain gauge (36) is arranged on a blade of an aircraft engine (28); A slip ring inductor (18) connected to the strain gauge (36); The cooling component (8) is provided with multiple layers of cooling channels, and the slip ring eliminator (18) is arranged in the innermost layer of the channel of the cooling component (8).

2. The dynamic stress measurement system according to claim 1, characterized in that: Also includes: A lead wire (37) connecting the strain gauge (36) and the slip ring lead (18); The extended section component (27) is arranged between the slip ring lead (18) and the strain gauge (36), and the extended section component (27) is arranged between the cooling component (8) and the strain gauge (36), so that the slip ring lead (18) and the cooling component (8) are both located outside the aircraft engine (28); the lead wire (37) passes through the interior of the extended section component (27).

3. The dynamic stress measurement system according to claim 2, characterized in that: The cooling assembly (8) comprises: The outer tube (13) is provided with an air inlet nozzle (26); The inner tube (14) is sleeved inside the outer tube (13) at intervals, and a first channel (15) is formed between the outer tube (13) and the inner tube (14); a second channel (16) is formed inside the inner tube (14); a first vent hole (12) is provided on the inner tube (14) and communicates with the first channel (15) and the second channel (16); the slip ring energizer (18) is located in the second channel (16); The cooling component (8) is suitable for introducing cooling air (1) into the first channel (15) from the air inlet nozzle (26) to form a first layer of cooling airflow (2), and part of the cooling air (1) enters the second channel (16) from the first channel (15) through the first air hole (12) to form a second layer of cooling airflow (3).

4. The dynamic stress measurement system according to claim 3, characterized in that: The end of the slip ring energizer (18) close to the aircraft engine (28) is installed inside the first casing (21) through a first flange structure (25); the first casing (21) is sleeved outside the extension section assembly (27) at intervals, and an annular first chamber (7) is formed between the first casing (21) and the extension section assembly (27); a second vent (20) connecting the second channel (16) and the first chamber (7) is provided on the first flange structure (25); the end of the first casing (21) away from the slip ring energizer (18) is installed in the exhaust casing (33) of the aircraft engine (28) through an annular support plate (34); a third vent connected to the outside is provided on the support plate (34); the first chamber (7) is connected to the space in the exhaust casing (33) close to the inner side of the support plate (34); The second layer of cooling airflow (3) is suitable for entering the first chamber (7) from the second passage (16) through the second vent hole (20) after cooling the slip ring inductor (18), and then being discharged to the outside through the third vent hole.

5. The dynamic stress measurement system according to claim 4, characterized in that: An outer tube (17) is sleeved at a position corresponding to the slip ring feeder (18) outside the outer tube (13), and the outer tube (17) is sleeved at a distance from the outer tube (17) outside a portion of the first casing (21); a third channel (19) is formed between the outer tube (17) and the first casing (21) and between the outer tube (17) and the outer tube (13); the third channel (19) is connected to the outside at a side away from the aircraft engine (28); a second chamber (23) is provided inside the first casing (21) at a side of the first flange structure (25) close to the inside of the aircraft engine (28); a fourth vent (22) connecting the first channel (15) and the second chamber (23) is provided on the first flange structure (25); a fifth vent (24) connecting the second chamber (23) and the third channel (19) is provided at a side of the second chamber (23) close to the inside of the aircraft engine (28); The first layer of cooling airflow (2) is suitable for entering the second chamber (23) from the first channel (15) through the fourth vent hole (22) after cooling the slip ring inductor (18), and then entering the third channel (19) through the fifth vent hole (24), thereby forming a return third layer of cooling airflow (6), which is finally discharged to the outside.

6. The dynamic stress measurement system according to claim 4 or 5, characterized in that: The extension section assembly (27) comprises: The extended shaft (30) has a first through hole (32) inside, and the lead wire (37) passes through the first through hole (32); the extended shaft (30) is installed in the second casing (31) through a first bearing (29); the second casing (31) is connected to the bearing casing (40) on the side close to the inside of the aircraft engine (28); the bearing casing (40) is installed on the support plate (34) through a second flange structure (42); a plurality of blades are installed on a turbine disk (35), a second through hole (38) is provided at the center of the turbine disk (35), and the lead wire (37) passes through the second through hole (38); the side of the turbine disk (35) away from the inside of the aircraft engine (28) is supported and installed in the bearing casing (40) through a second bearing (39); the end of the extended shaft (30) located on the inside of the aircraft engine (28) is fixedly connected to the turbine disk (35).

7. The dynamic stress measurement system according to any one of claims 1 to 5, characterized in that: The high-temperature exhaust gas (5) of the aircraft engine (28) is discharged into the exhaust ventilation pipe (4) of the vehicle platform.

8. The dynamic stress measurement system according to any one of claims 3 to 5, characterized in that: The inner tube (14) and the outer tube (13) are both composed of multiple sections of pipe fittings that are connected and arranged.

9. The dynamic stress measurement system according to any one of claims 3 to 5, characterized in that: A first cover plate (10) for sealing the end portion is provided on the inner tube (14) at one end away from the aircraft engine (28), and a second cover plate (11) for sealing the end portion is provided on the outer tube (13) at one end away from the aircraft engine (28); the air inlet nozzle (26) is arranged on the side wall of the outer tube (13).

10. The dynamic stress measurement system according to claim 9, characterized in that: The cooling oil circuit (9) of the slip ring lead (18) and the signal line connected to the outside are both led out from the first cover plate (10) along the second channel (16).

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