A testing mechanism for static pressure parameters of an aircraft engine, and an aircraft engine

Abstract

The present invention discloses a testing mechanism for static pressure parameters of an aircraft engine. The testing mechanism is arranged in a compressor casing and a turbine casing of the aircraft engine. The testing mechanism includes a data acquisition tube and a sensor connected to the data acquisition tube. The compressor casing and the turbine casing are respectively equipped with the testing mechanism, so that the engine of the present invention can be used to directly measure the static pressure at the compressor outlet and the static pressure behind the first-stage turbine guide vanes under normal working conditions, thereby ensuring that engine data is monitored in real time during the operation of the engine to evaluate the engine status.

Description

Technical Field

[0001] The present invention relates to the field of engines, and in particular to a testing mechanism for static pressure parameters of aviation engines. Background Art

[0002] When conducting aircraft engine tests, the measurement of multiple flow path parameters is involved, including the static pressure at the compressor outlet and the static pressure behind the first-stage turbine guide vanes under normal engine operation. The test requires measuring these static pressure parameters to evaluate the engine's condition. In the prior art, the static pressure at the compressor outlet and the static pressure behind the first-stage turbine guide vanes under normal engine operation cannot be directly measured. Instead, the calculated values can only be obtained through the whole-machine performance calculation model. However, the calculated values obtained through the whole-machine performance calculation model are inaccurate and have certain errors compared to the actual static pressure at the compressor outlet and the static pressure behind the first-stage turbine guide vanes, making it impossible to correctly evaluate the engine's condition. In order to meet the requirements of the static pressure parameter test, the engine's compressor casing and turbine casing and other related parts need to be modified to install test sensors and obtain actual parameter information.

[0003] The invention with publication number CN113864240A discloses an aircraft engine single-duct high-low pressure engine and its intermediate casing component, the generator comprising: a low-pressure compressor (1), wherein the last row of blades at the outlet of the low-pressure compressor (1) are rotor blades; an intermediate casing component (2), wherein the inlet of the intermediate casing component (2) is connected to the outlet of the low-pressure compressor (1), and the leading edge of each support plate (3) therein is bent in the direction opposite to the torsion direction of the last row of blades at the outlet of the low-pressure compressor (1); a high-pressure compressor (4), wherein the inlet of the high-pressure compressor (4) is connected to the outlet of the intermediate casing component (2); the compressor outlet static pressure under normal working state of the engine of this invention cannot be directly measured, and the calculated value obtained by the whole machine performance calculation model is inaccurate. Compared with the actual compressor outlet static pressure, there is a certain error, resulting in an inability to correctly evaluate the engine state. Summary of the Invention

[0004] In response to the above-mentioned deficiencies in the prior art, the present invention provides a testing mechanism for the static pressure parameters of an aircraft engine, which is used to obtain the actual parameter information of the compressor outlet static pressure and the static pressure behind the first-stage turbine guide vanes under normal engine operation to evaluate the engine status.

[0005] The present invention provides the following technical solutions:

[0006] A testing mechanism for static pressure parameters of an aircraft engine, characterized in that the aircraft engine includes a compressor casing and a turbine casing, and the testing mechanism is arranged in the compressor casing and the turbine casing. A rectifier and a working ring are installed inside the compressor casing; the rectifier includes an outer ring and an inner ring, rectifier blades are provided between the outer ring and the inner ring, and a sealing ring is provided inside the inner ring of the rectifier. The testing mechanism in the compressor casing includes a data acquisition tube and a sensor connected to the data acquisition tube. The data acquisition tube is welded to the rectifier. The data acquisition tube includes a tube and a pipe joint. One end of the tube is welded to the rectifier outer ring, and the other end passes through the compressor casing and is covered with a pipe joint welded to the compressor casing. The portion where the tube is connected to the pipe joint constitutes a connector for connecting the sensor. The inner wall of the turbine casing is provided with inner wall ribs, and the inner wall ribs are provided with axial ring grooves. The turbine casing A guide blade is provided at the bottom, and the guide blade consists of an upper edge plate, a lower edge plate and a blade body sandwiched between the upper and lower edge plates. A hook is provided on the upper edge plate, and the hook is hung in the axial ring groove of the inner wall rib through a fixed-distance ring; the test mechanism in the turbine casing includes a data acquisition tube 2 and a sensor connected to the data acquisition tube 2, and an adapter is provided on the turbine casing, and the adapter fixes the data acquisition tube 2 vertically passing through the turbine casing, and the data acquisition tube 2 includes a pipe 2 and a pipe nozzle 2. One end of the pipe 2 of the data acquisition tube 2 is downwardly connected to the upper edge plate of the guide blade, and the end of the pipe 2 higher than the adapter is covered with a pipe nozzle 2, and the part where the pipe 2 and the pipe nozzle 2 are connected constitutes a connector 2 for connecting the sensor.

[0007] Furthermore, the outer ring and the inner ring of the rectifier are semi-ring structures.

[0008] Furthermore, the rectifier in the compressor casing is fixedly connected to the bottom of the compressor casing by bolts provided on the outer ring.

[0009] Furthermore, a precision positioning bolt is provided in the middle of the rectifier outer ring, and the rectifier outer ring passes through the compressor casing through the precision positioning bolt to fix the connection boss, and the number of the boss is multiple.

[0010] Furthermore, a plurality of data acquisition tubes are provided in the compressor casing. Generally, four data acquisition tubes are installed in the compressor casing for installing test sensors to realize the test of static pressure parameters. The position where the data acquisition tube is installed in the compressor casing needs to be selected to avoid the position where the boss is located.

[0011] Furthermore, the distance ring is clamped between the rear end surface of the upper edge plate and the inner wall rib to limit the axial movement of the guide blade.

[0012] Furthermore, the second pipe nozzle is fixed to the adapter seat by a first bolt.

[0013] Furthermore, an adjustment gasket is provided between the second pipe joint and the adapter seat to ensure that the pipe is flush with the upper edge plate of the guide blade.

[0014] Furthermore, a liquid sealing gasket is applied to the connection portion between the second pipe and the upper edge plate of the guide blade to provide a seal.

[0015] Furthermore, an aircraft engine using any of the above-mentioned testing mechanisms.

[0016] Compared with the prior art, the present invention has the following beneficial effects:

[0017] The test mechanism of the present invention is arranged in the compressor casing and turbine casing of the aircraft engine. The test mechanism includes a data acquisition tube and a sensor connected to the data acquisition tube. The compressor casing and turbine casing are respectively equipped with test mechanisms, which facilitates the use of the engine of the present invention to directly measure the compressor outlet static pressure and the static pressure parameters behind the first-stage turbine guide blades under normal working conditions, ensuring real-time monitoring of engine data during engine operation to evaluate the engine status.

[0018] After the test structure of the present invention is assembled on the compressor casing and the turbine casing, it does not affect the subsequent normal use of the engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0020] Figure 1 This is a schematic diagram of the compressor casing;

[0021] Figure 2 Schematic diagram of a rectifier without a data acquisition tube 1;

[0022] Figure 3 This is a schematic diagram of the turbine casing;

[0023] Figure 4 Schematic diagram of data acquisition tube 2 installed on the adapter.

[0024] Among them, the figure marks are: 1. compressor casing; 2. working ring; 3. outer ring; 4. inner ring; 5. straightening blade; 6. sealing ring; 7. pipe one; 8. data acquisition pipe one; 9 pipe nozzle; 10. turbine casing; 11. internal rib; 12. axial ring groove; 13. data acquisition pipe two; 14. upper edge plate; 15. lower edge plate; 16. blade; 17. positioning ring; 18. adapter; 19. pipe two; 20. pipe nozzle two; 21. precision positioning bolt; 22. boss; 23. adjustment gasket. DETAILED DESCRIPTION

[0025] In order to make the technical problems, technical solutions and beneficial effects to be solved by the present invention more clearly understood, the present invention is further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not intended to limit the present invention.

[0026] Example 1

[0027] like Figures 1 to 4 A testing mechanism for static pressure parameters of an aircraft engine is shown. The aircraft engine includes a compressor casing 1 and a turbine casing 10. The aircraft engine includes a compressor casing 1 and a turbine casing 10, and the testing mechanism is arranged inside the compressor casing 1 and the turbine casing 10.

[0028] A rectifier and a working ring 2 are installed inside the compressor casing 1; the rectifier includes an outer ring 3 and an inner ring 4, with rectifier blades 5 provided between the outer ring 3 and the inner ring 4, and a sealing ring 6 provided inside the rectifier inner ring 4. The testing mechanism inside the compressor casing 1 includes a data acquisition tube 8 and a sensor connected to the data acquisition tube 8. The data acquisition tube 8 is welded to the rectifier. The data acquisition tube 8 includes a tube 7 and a pipe joint 9. One end of the tube 7 is welded to the rectifier outer ring 3, and the other end passes through the compressor casing 1 and is covered with a pipe joint 9 welded to the compressor casing 1. The portion where the tube 7 and the pipe joint 9 are connected constitutes a connector 1 for connecting a sensor for data acquisition.

[0029] The inner wall of the turbine casing 10 is provided with an inner wall rib 11, and the inner wall rib 11 is provided with an axial annular groove 12. A guide blade is provided below the turbine casing 10, and the guide blade is composed of an upper edge plate 14, a lower edge plate 15 and a blade body 16 sandwiched between the upper and lower edge plates. The upper edge plate 14 is provided with a hook, and the hook is hung in the axial annular groove 12 of the inner wall rib 11 through a spacing ring 17; an adapter 18 is provided on the turbine casing 10, and the adapter 18 fixes a data acquisition tube 2 13 that vertically passes through the turbine casing 10. The data acquisition tube 2 13 includes a tube 2 19 and a pipe nozzle 2 20. One end of the tube 2 19 of the data acquisition tube 2 13 is downwardly connected to the upper edge plate 13 of the guide blade, and the end of the tube 2 19 that is higher than the adapter 18 is covered with a pipe nozzle 2 20. The part where the tube 2 19 and the pipe nozzle 2 20 are connected constitutes a connector 2 for connecting a sensor for data acquisition.

[0030] The outer ring 3 and the inner ring 4 of the rectifier are semi-ring structures.

[0031] Furthermore, the rectifier in the compressor casing is fixedly connected to the bottom of the compressor casing 1 by bolts provided on the outer ring 3.

[0032] Furthermore, a precision positioning bolt 21 is provided in the middle of the rectifier outer ring 3 , and the rectifier outer ring 3 passes through the compressor casing 1 through the precision positioning bolt 21 and is fixedly connected to a boss 22 , and the number of the boss 22 is multiple.

[0033] Furthermore, a plurality of data acquisition tubes 8 are provided in the compressor casing. Generally, four data acquisition tubes 8 are installed in the compressor casing for installing test sensors to realize the test of static pressure parameters. The position where the data acquisition tube 8 is installed in the compressor casing needs to be selected to avoid the position where the boss 22 is located.

[0034] The test mechanism of the present invention is disposed within the compressor casing 1 and turbine casing 10 of the aircraft engine. The test mechanism includes a data acquisition tube and a sensor connected to the data acquisition tube. The compressor casing 1 and turbine casing 10 are each equipped with a test mechanism. This facilitates direct measurement of compressor outlet static pressure and static pressure behind the first-stage turbine guide vanes during normal engine operation. This ensures real-time monitoring of engine data during operation to assess engine status. Furthermore, once the test structure of the present invention is installed in the compressor casing 1 and turbine casing 10, it does not affect the subsequent normal operation of the engine.

[0035] Example 2

[0036] like Figures 1 to 4A testing mechanism for static pressure parameters of an aircraft engine is shown. The aircraft engine includes a compressor casing 1 and a turbine casing 10. The aircraft engine includes a compressor casing 1 and a turbine casing 10, and the testing mechanism is arranged inside the compressor casing 1 and the turbine casing 10.

[0037] A rectifier and a working ring 2 are installed inside the compressor casing 1; the rectifier includes an outer ring 3 and an inner ring 4, with rectifier blades 5 provided between the outer ring 3 and the inner ring 4, and a sealing ring 6 provided inside the rectifier inner ring 4. The testing mechanism inside the compressor casing 1 includes a data acquisition tube 8 and a sensor connected to the data acquisition tube 8. The data acquisition tube 8 is welded to the rectifier. The data acquisition tube 8 includes a tube 7 and a pipe joint 9. One end of the tube 7 is welded to the rectifier outer ring 3, and the other end passes through the compressor casing 1 and is covered with a pipe joint 9 welded to the compressor casing 1. The portion where the tube 7 and the pipe joint 9 are connected constitutes a connector 1 for connecting a sensor for data acquisition.

[0038] The inner wall of the turbine casing 10 is provided with an inner wall rib 11, and the inner wall rib 11 is provided with an axial annular groove 12. A guide blade is provided below the turbine casing 10, and the guide blade is composed of an upper edge plate 14, a lower edge plate 15 and a blade body 16 sandwiched between the upper and lower edge plates. The upper edge plate 14 is provided with a hook, and the hook is hung in the axial annular groove 12 of the inner wall rib 11 through a spacing ring 17; an adapter 18 is provided on the turbine casing 10, and the adapter 18 fixes a data acquisition tube 2 13 that vertically passes through the turbine casing 10. The data acquisition tube 2 13 includes a tube 2 19 and a pipe nozzle 2 20. One end of the tube 2 19 of the data acquisition tube 2 13 is downwardly connected to the upper edge plate 13 of the guide blade, and the end of the tube 2 19 that is higher than the adapter 18 is covered with a pipe nozzle 2 20. The part where the tube 2 19 and the pipe nozzle 2 20 are connected constitutes a connector 2 for connecting a sensor for data acquisition.

[0039] The distance ring 17 is clamped between the rear end surface of the upper edge plate 14 and the inner wall rib 11 to limit the axial movement of the guide blade.

[0040] Furthermore, the adapter 18 is welded to the turbine casing 10 by brazing, and the brazing deformation is small, which can reduce the deformation of the parts.

[0041] In summary, the test mechanism of the present invention is disposed within the compressor casing 1 and turbine casing 10 of the aircraft engine. The test mechanism includes a data acquisition tube and a sensor connected to the data acquisition tube. The compressor casing 1 and turbine casing 10 are each equipped with a test mechanism. This facilitates direct measurement of the compressor outlet static pressure and the static pressure behind the first-stage turbine guide vanes during normal engine operation. This ensures real-time monitoring of engine data during operation to assess engine status. Furthermore, once the test structure of the present invention is assembled on the compressor casing 1 and turbine casing 10, it does not affect the subsequent normal operation of the engine.

[0042] Example 3

[0043] like Figures 1 to 4 A testing mechanism for static pressure parameters of an aircraft engine is shown. The aircraft engine includes a compressor casing 1 and a turbine casing 10. The aircraft engine includes a compressor casing 1 and a turbine casing 10, and the testing mechanism is arranged inside the compressor casing 1 and the turbine casing 10.

[0044] A rectifier and a working ring 2 are installed inside the compressor casing 1; the rectifier includes an outer ring 3 and an inner ring 4, with rectifier blades 5 provided between the outer ring 3 and the inner ring 4, and a sealing ring 6 provided inside the rectifier inner ring 4. The testing mechanism inside the compressor casing 1 includes a data acquisition tube 8 and a sensor connected to the data acquisition tube 8. The data acquisition tube 8 is welded to the rectifier. The data acquisition tube 8 includes a tube 7 and a pipe joint 9. One end of the tube 7 is welded to the rectifier outer ring 3, and the other end passes through the compressor casing 1 and is covered with a pipe joint 9 welded to the compressor casing 1. The portion where the tube 7 and the pipe joint 9 are connected constitutes a connector 1 for connecting a sensor for data acquisition.

[0045] The inner wall of the turbine casing 10 is provided with an inner wall rib 11, and the inner wall rib 11 is provided with an axial annular groove 12. A guide blade is provided below the turbine casing 10, and the guide blade is composed of an upper edge plate 14, a lower edge plate 15 and a blade body 16 sandwiched between the upper and lower edge plates. The upper edge plate 14 is provided with a hook, and the hook is hung in the axial annular groove 12 of the inner wall rib 11 through a spacing ring 17; an adapter 18 is provided on the turbine casing 10, and the adapter 18 fixes a data acquisition tube 2 13 that vertically passes through the turbine casing 10. The data acquisition tube 2 13 includes a tube 2 19 and a pipe nozzle 2 20. One end of the tube 2 19 of the data acquisition tube 2 13 is downwardly connected to the upper edge plate 13 of the guide blade, and the end of the tube 2 19 that is higher than the adapter 18 is covered with a pipe nozzle 2 20. The part where the tube 2 19 and the pipe nozzle 2 20 are connected constitutes a connector 2 for connecting a sensor for data acquisition.

[0046] Furthermore, the second pipe nozzle 20 is fixed to the adapter 18 by bolts.

[0047] Furthermore, the adapter 18 is welded to the turbine casing 10 by brazing, and the brazing deformation is small, which can reduce the deformation of the parts.

[0048] Furthermore, an adjustment gasket 23 is provided between the second pipe joint 20 and the adapter seat 18 to ensure that the pipe is flush with the upper edge plate 14 of the guide blade.

[0049] Furthermore, a liquid sealing gasket is applied to the connection portion between the second pipe 19 and the upper edge plate 14 of the guide blade to provide a seal.

[0050] In summary, the test mechanism of the present invention is disposed within the compressor casing 1 and turbine casing 10 of the aircraft engine. The test mechanism includes a data acquisition tube and a sensor connected to the data acquisition tube. The compressor casing 1 and turbine casing 10 are each equipped with a test mechanism. This facilitates direct measurement of the compressor outlet static pressure and the static pressure behind the first-stage turbine guide vanes during normal engine operation. This ensures real-time monitoring of engine data during operation to assess engine status. Furthermore, once the test structure of the present invention is assembled on the compressor casing 1 and turbine casing 10, it does not affect the subsequent normal operation of the engine.

[0051] The foregoing description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. Those skilled in the art will readily appreciate that various modifications and variations of the present invention are possible. Any modifications, equivalent substitutions, or improvements made within the spirit and principles of the present invention are intended to be within the scope of the present invention.

Claims

1. A testing mechanism for static pressure parameters of an aircraft engine, characterized in that: The aircraft engine includes a compressor casing and a turbine casing, and the testing mechanism is arranged in the compressor casing and the turbine casing; A rectifier and a working ring are installed inside the compressor casing; the rectifier includes an outer ring and an inner ring, with rectifier blades provided between the outer and inner rings, and a sealing ring provided inside the inner ring of the rectifier. The testing mechanism inside the compressor casing includes a data acquisition tube 1 and a sensor connected to the data acquisition tube 1, the data acquisition tube 1 being welded to the rectifier, the data acquisition tube 1 including a tube 1 and a pipe connector 1, one end of the tube 1 being welded to the rectifier outer ring, and the other end passing through the compressor casing and being mounted on a pipe connector 1 welded to the compressor casing, the portion where the tube 1 and the pipe connector 1 meet constitutes a connector 1 for connecting the sensor; The inner wall of the turbine casing is provided with inner wall ribs, and the inner wall ribs are provided with axial annular grooves. A guide blade is provided under the turbine casing, and the guide blade is composed of an upper edge plate, a lower edge plate and a blade body sandwiched between the upper and lower edge plates. A hook is provided on the upper edge plate, and the hook is hung in the axial annular groove of the inner wall rib through a fixed-distance ring; the test mechanism in the turbine casing includes a data acquisition tube 2 and a sensor connected to the data acquisition tube 2, and an adapter is provided on the turbine casing, and the adapter fixes the data acquisition tube 2 vertically passing through the turbine casing, and the data acquisition tube 2 includes a pipe 2 and a pipe nozzle 2. One end of the pipe 2 of the data acquisition tube 2 is downwardly connected to the upper edge plate of the guide blade, and the end of the pipe 2 higher than the adapter is provided with a pipe nozzle 2, and the part where the pipe 2 and the pipe nozzle 2 are connected constitutes a connector 2 for connecting the sensor.

2. The testing mechanism for static pressure parameters of an aircraft engine according to claim 1, characterized in that: The outer ring and the inner ring of the rectifier are semi-ring structures.

3. The testing mechanism for static pressure parameters of an aircraft engine according to claim 1, characterized in that: The rectifier in the compressor casing is fixedly connected to the lower side of the compressor casing by bolts provided at the outer ring.

4. The testing mechanism for static pressure parameters of an aircraft engine according to claim 1, characterized in that: A precision positioning bolt is provided in the middle of the rectifier outer ring, and the rectifier outer ring passes through the compressor casing and is fixed to the connection boss via the precision positioning bolt.

5. The testing mechanism for static pressure parameters of an aircraft engine according to claim 1, characterized in that: A plurality of data acquisition tubes are provided in the compressor casing.

6. The testing mechanism for static pressure parameters of an aircraft engine according to claim 1, characterized in that: The distance ring is clamped between the rear end surface of the upper edge plate and the inner wall rib to limit the axial movement of the guide blade.

7. The testing mechanism for static pressure parameters of an aircraft engine according to claim 1, characterized in that: The second pipe joint is fixed on the adapter seat by a second bolt.

8. The testing mechanism for static pressure parameters of an aircraft engine according to claim 7, characterized in that: An adjusting gasket is provided between the second pipe joint and the adapter seat.

9. The testing mechanism for static pressure parameters of an aircraft engine according to claim 1, characterized in that: The connection portion between the second pipe and the upper edge plate of the guide blade is coated with a liquid sealing gasket.

10. An aircraft engine, characterized in that: The test device comprises the test mechanism described in any one of claims 1 to 9.

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