Aero-engine high-pressure rotor testing device and aero-engine test piece

Abstract

The present disclosure relates to an aero-engine high-pressure rotor testing device and an aero-engine test piece, wherein the testing device comprises a slip ring assembly (11) including a slip ring stator (111) and a slip ring rotor (112) coaxially arranged, the slip ring stator (111) is fixed on a casing, the slip ring rotor (112) is coaxially arranged with the slip ring stator (111) and is drivingly connected with a rear end of a high-pressure rotor (10') of the aero-engine; a detection component is arranged on the high-pressure rotor (10') and is configured to detect a state parameter of the high-pressure rotor (10'), a test lead (12) of the detection component is led out through the slip ring assembly (11); a cooling pipeline (13) is configured to cool the slip ring assembly (11) by introducing a cooling medium; and a guide pipe (14) is arranged at a rear end of the slip ring assembly (11) in an axial direction, and one end of the guide pipe penetrates through a side wall of a tail cone (10) of the aero-engine and is configured to lead out the test lead (12) and the cooling pipeline (13) from the aero-engine.

Description

Technical Field

[0001] This disclosure relates to the field of aero-engine testing technology, and in particular to an aero-engine high-pressure rotor testing device and an aero-engine test piece. Background Technology

[0002] To improve the efficiency, reliability, and reduce pollution emissions of aero engines, it is necessary to accurately measure the surface temperature and strain data of key structural components of aero engines.

[0003] Temperature and strain testing sensors need to be mounted on the surface of rotating components. During engine operation, the rotor rotates at high speed, making it difficult to transmit measurement signals from the rotor to ground-based equipment for processing, analysis, and storage. Similarly, providing power excitation to the sensors during data acquisition presents the same challenge. Therefore, for measurement signals in a rotating state, a specialized measuring instrument system is required to efficiently and reliably transmit measurement data from the rotor to ground-based equipment.

[0004] Currently, the industry uses telemetry equipment or slip ring devices to solve this problem, which can transmit measurement signals clearly and without interference to ground data acquisition equipment. In related technologies, due to concerns about the excessively high temperature at the engine's rear end, the engine slip ring testing device is placed at the engine's front end. However, this approach has the following problems:

[0005] The test data leads are relatively long and easily damaged in the working environment. To meet the measurement lead requirements of the rear rotating parts, it is usually necessary to add a lead tube in the engine to lead the rear lead to the front. This increases the complexity of the rotating parts. At the same time, the thin lead tube is prone to deformation, cracks, and chipping when working in high speed and high temperature areas for a long time. In severe cases, it will affect the safety of engine operation. Summary of the Invention

[0006] The embodiments of this disclosure provide a high-pressure rotor testing device and an aero-engine test piece, which can improve testing reliability while reducing the difficulty of high-pressure rotor testing.

[0007] According to a first aspect of this disclosure, a high-pressure rotor testing apparatus for an aero-engine is provided, comprising:

[0008] The slip ring assembly includes a slip ring stator and a slip ring rotor arranged coaxially. The slip ring stator is fixed to the casing of the aero-engine, and the slip ring rotor is arranged coaxially with the slip ring stator and is driven to the rear end of the high-pressure rotor of the aero-engine.

[0009] The detection component, located on the high-voltage rotor, is configured to detect the state parameters of the high-voltage rotor, and the test leads of the detection component are led out through the slip ring assembly;

[0010] Cooling piping is configured to cool the slip ring assembly by introducing a cooling medium; and

[0011] The guide tube, located axially at the rear end of the slip ring assembly, with one end passing through the tail cone sidewall of the aero-engine, is configured to lead the test leads and cooling lines out of the aero-engine.

[0012] In some embodiments, the guide tube is arranged radially.

[0013] In some embodiments, the guide tube includes a flexible section located at least within the tail cone.

[0014] In some embodiments, the tail cone includes a first tail cone segment and a second tail cone segment, the first tail cone segment and the second tail cone segment being detachably joined axially, and a guide tube being located between the first tail cone segment and the second tail cone segment.

[0015] In some embodiments, the slip ring rotor and the high-pressure rotor are connected by a coupling; wherein, the ends of the slip ring rotor and the high-pressure rotor that are close to each other are provided with mounting holes, and the coupling includes:

[0016] Main body; and

[0017] The first connecting part and the second connecting part are located at both ends of the main body along the axial direction. The first connecting part extends into the mounting hole of the high-pressure rotor and is connected to the high-pressure rotor. The second connecting part extends into the mounting hole of the slip ring rotor and is connected to the slip ring rotor. Both the first connecting part and the second connecting part are hexagonal head structures, and the hexagonal head structure has an arc-shaped side in the longitudinal section, and the arc-shaped side is in contact with the inner wall of the mounting hole.

[0018] In some embodiments, the casing includes an interstage casing, and the slip ring stator is connected to the mounting edge of the front seal of the tail cone via the inside of the interstage casing.

[0019] In some embodiments, the casing includes an interstage casing with a bearing configured to support the high-pressure rotor, the bearing having a sealing gas passage; the aero-engine also includes a transition section, one end of which is connected to the turbine shaft along the axial direction, and the other end of which is connected to the junction of the high-pressure rotor and the slip ring rotor, the transition section having a first through hole through which sealing gas enters the sealing cavity of the bearing cavity and the interstage casing, and exits from the rear end of the tail cone.

[0020] In some embodiments, the casing includes an interstage casing, and the high-pressure rotor test device for aero-engines also includes a heat shield, coaxially sleeved outside the slip ring assembly, configured to isolate the high-temperature airflow transmitted between the interstage casing and the tail cone, with one end of the heat shield connected to the interstage casing and the other end extending to the guide tube.

[0021] In some embodiments, a test bench ground device is also included, with test leads and cooling lines extending from the tail cone and connected to the test bench ground device.

[0022] According to a second aspect of this disclosure, an aero-engine test component is provided, comprising: the aero-engine high-pressure rotor test apparatus of the above embodiments.

[0023] The high-pressure rotor testing apparatus for an aero-engine disclosed in this embodiment places the slip ring assembly at the rear end of the high-pressure rotor. The test leads can be directly extended rearward to connect to the slip ring assembly, significantly shortening the length of the test leads, simplifying layout, and reducing the risk of damage. Furthermore, by incorporating cooling pipes, the slip ring assembly can be cooled when the internal temperature of the high-pressure turbine is high, preventing excessive temperature from affecting the reliability of its signal transmission. Additionally, by providing guide pipes, the test leads and cooling pipes can be led out of the aero-engine together along the guide pipes, providing protection for the test leads and cooling pipes, reducing the risk of damage, and saving layout space. Attached Figure Description

[0024] The accompanying drawings, which are included to provide a further understanding of this disclosure and form part of this application, illustrate exemplary embodiments of this disclosure and are used to explain this disclosure, but do not constitute an undue limitation of this disclosure. In the drawings:

[0025] Figure 1 These are schematic diagrams illustrating the structure of some embodiments of the aero-engine disclosed herein;

[0026] Figure 2 This is a schematic diagram of the structure of some embodiments of the slip ring assembly in the high-pressure rotor test device for aero-engines disclosed herein;

[0027] Figure 3 This is a schematic diagram of some embodiments of the high-pressure rotor test device for aero-engines disclosed herein, mounted on an aero-engine test piece.

[0028] Figure 4 This is a schematic diagram of the connection structure between the slip ring rotor and the high-voltage rotor in the slip ring assembly.

[0029] Figure 5 A schematic diagram showing the flow direction of the bearing sealing air passage;

[0030] Figure 6 This is a schematic diagram of the connection structure of the slip ring stator in the slip ring assembly;

[0031] Figure 7 This is a schematic diagram showing the lead-out method of the test harness and cooling pipes for the slip ring assembly;

[0032] Figure 8 This is a schematic diagram of the heat insulation cylinder in the high-pressure rotor test device for aero-engines disclosed in this paper;

[0033] Figure 9 This is a schematic diagram of the coupling connecting the high-voltage rotor and the slip ring rotor disclosed in this invention;

[0034] Figure 10 This is a schematic diagram of the connection between the coupling and the slip ring rotor.

[0035] Figure 11A and Figure 11B These are schematic diagrams showing the coupling deflected upwards and downwards at certain angles relative to the rotor axis of the slip ring. Detailed Implementation

[0036] The present disclosure is described in detail below. In the following paragraphs, different aspects of the embodiments are defined in more detail. The aspects so defined may be combined with any other aspect or aspects unless expressly stated otherwise. In particular, any feature considered preferred or advantageous may be combined with one or more other features considered preferred or advantageous.

[0037] The terms "first" and "second" used in this disclosure are merely for ease of description and to distinguish different components with the same name, and do not indicate a sequential or primary / secondary relationship.

[0038] In the description of this invention, it should be understood that the terms "inner," "outer," "upper," "lower," "left," and "right," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings, and are only for the convenience of describing this invention, and do not indicate or imply that the device referred to must have a specific orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation on the scope of protection of this invention. The terms "circumferential," "axial," and "radial" mentioned in subsequent embodiments are defined based on an aero-engine.

[0039] like Figure 1 As shown, the aero-engine comprises, in the axial direction along the airflow direction, the following components in sequence: fan 1, low-pressure compressor 2, intermediate casing 3, high-pressure compressor 4, combustion chamber 5, high-pressure turbine 6, interstage casing 7, low-pressure turbine 8, turbine rear casing 9, and tail cone 10.

[0040] Figure 1 The rectangle in the middle indicates the test piece 100 of the aero-engine core engine; an enlarged view can be found therein. Figure 3 When testing the high-pressure rotor 10' of the high-pressure turbine 6, the test can be performed in the core engine test piece 100, which includes an intermediate casing 3, a high-pressure compressor 4, a combustion chamber 5, a high-pressure turbine 6, and an interstage casing 7. To ensure the integrity of the function during testing, a tail cone 10 is provided at the rear end of the interstage casing 7.

[0041] like Figures 2 to 8As shown, this disclosure provides a high-pressure rotor testing device for an aero-engine, including: a slip ring assembly 11, a testing component, a cooling pipe 13, and a guide pipe 14.

[0042] like Figure 2 As shown, the slip ring assembly 11 includes a slip ring stator 111 and a slip ring rotor 112 coaxially arranged. The slip ring stator 111 is fixed to the casing of the aero-engine, so that the load of the slip ring stator 111 can be transmitted to the casing. The slip ring rotor 112 is coaxially arranged with the slip ring stator 111, and the slip ring stator 111 is sleeved outside the slip ring rotor 112. The slip ring rotor 112 is also driven to the rear end of the high-pressure rotor 10' of the aero-engine, so as to rotate synchronously under the drive of the high-pressure rotor 10'.

[0043] A detection component, mounted on the high-pressure rotor 10', is configured to detect the state parameters of the high-pressure rotor 10'. The test lead 12 of the detection component is led out through the slip ring assembly 11. The state parameters include dynamic stress parameters, temperature parameters, and / or rotational speed parameters. The test lead 12 is transmitted through the high-pressure rotor 10' and the slip ring rotor 112 to the slip ring stator 111, and is led out of the aero-engine through the data connector of the slip ring stator 111 for connection to the data acquisition system in the ground test bench.

[0044] Cooling conduit 13 is configured to cool slip ring assembly 11 by introducing a cooling medium, which can be liquid or gas. When using gas, cooling gases from the aircraft engine can be readily utilized. The cooling medium can be provided by ground-based test bench equipment, and cooling conduit 13 introduces and exits slip ring assembly 11 through the engine's stator structure.

[0045] A guide tube 14 is axially disposed at the rear end of the slip ring assembly 11, with one end of the guide tube 14 passing through the side wall of the tail cone 10 of the aero-engine. It is configured to lead the test lead 12 and cooling line 13 out of the aero-engine for connection to ground equipment on the test bench. The other end of the guide tube 14 is connected to the slip ring stator 111.

[0046] In this embodiment, the slip ring assembly 11 is located at the rear end of the high-pressure rotor 10'. Its test lead 12 extends directly rearward to connect to the slip ring assembly 11, significantly shortening the length of the test lead 12, simplifying layout, and reducing the risk of damage. Furthermore, by providing a cooling pipe 12, the slip ring assembly 11 can be cooled when the internal temperature of the high-pressure turbine 6 is high, preventing excessive temperature from affecting the reliability and safety of its signal transmission. Additionally, by providing a guide pipe 14, the test lead 12 and the cooling pipe 13 can be led out of the aircraft engine together along the guide pipe 14, providing protection for the test lead 12 and the cooling pipe 13, reducing the risk of damage, and saving layout space.

[0047] In some embodiments, such as Figure 3As shown, the guide tube 14 is arranged radially. This structure can further shorten the length of the test lead 12 and the cooling pipe 13, and occupies little space, which can meet the layout requirements of the test lead 12 and the cooling pipe 13. Moreover, for the rear slip ring assembly 11, since the number of support plates of the interstage casing 7 is limited and mainly used to realize the functions of bearing cavity oil supply and ventilation, the remaining support plates cannot meet the requirements of slip ring test lead, cooling supply / exhaust, etc. By setting the guide tube 14, the fixing of the test lead 12 and the cooling pipe 13 can be simplified.

[0048] In some embodiments, such as Figure 3 As shown, the guide tube 14 includes a flexible tube section, which is located at least within the tail cone 10. When the tail cone 10 vibrates or deforms, the deformation of the flexible tube section can compensate for the vibration, ensuring the reliability of the slip ring accessories during the operation of the aero-engine.

[0049] This type of guide tube 14 can be compensated to ensure the reliability of slip ring accessories such as test leads 12 and cooling pipes 13 during engine operation.

[0050] For example, flexible pipe sections can be made from corrugated pipe sections, which allow for radial and axial deformation.

[0051] like Figure 3 and Figure 7 As shown, the tail cone 10 includes a first tail cone section 101 and a second tail cone section 102, which are detachably connected axially. A guide tube 14 is located between the first tail cone section 101 and the second tail cone section 102. By designing the first tail cone section 101 and the second tail cone section 102 as separate and detachable structures, it is convenient to pass the test lead 12 and the cooling pipe 13 through the guide tube 14, and it is also convenient to install the guide tube 14.

[0052] In some embodiments, such as Figure 4 As shown, the slip ring rotor 112 and the high-pressure rotor 10' are connected by a coupling 17. Figure 9 and Figure 10 As shown, both the slip ring rotor 112 and the high-pressure rotor 10' have mounting holes 1121 at their ends. The coupling 17 includes a main body 171, a first connecting part 172, and a second connecting part 173. The main body 171 may be cylindrical. The first connecting part 172 and the second connecting part 173 are located at opposite ends of the main body 171 along the axial direction. The first connecting part 172 extends into the mounting hole 1121 of the high-pressure rotor 10' and is connected to the high-pressure rotor 10'. The second connecting part 173 extends into the mounting hole 1121 of the slip ring rotor 112 and is connected to the slip ring rotor 112.

[0053] Both the first connecting portion 172 and the second connecting portion 173 have a hexagonal head structure, and the hexagonal head structure has an arc-shaped side surface 174 in the longitudinal section, and the arc-shaped side surface 174 contacts the inner wall of the mounting hole 1121. Figure 11A As shown, this connection method allows the coupling 17 to deflect upwards at a certain angle relative to the axis of the slip ring rotor 112. Figure 11B As shown, this connection method allows the coupling 17 to deflect downwards at a certain angle relative to the axis of the slip ring rotor 112.

[0054] Because the engine rotor experiences circumferential and radial displacement during operation, rigid couplings cannot eliminate the displacement effects caused by the engine. The embodiments disclosed herein employ an arc-shaped hexagonal head coupling, which can transmit torque, ensuring that the high-pressure rotor 10' rotates at the same speed as the slip ring rotor 112 under test conditions, while compensating for axial and radial displacements from the engine rotor. It allows for a certain amount of movement in all directions to prevent the unbalanced force generated by the high-pressure rotor 10' during operation from being transmitted to the slip ring rotor 112, thereby improving measurement accuracy. Furthermore, it ensures a safe, reliable, and convenient connection between the slip ring rotor 112 and the high-pressure rotor 10'.

[0055] In some embodiments, such as Figure 6 As shown, the casing includes an interstage casing 7, and a slip ring stator 111 is connected to the mounting edge 10' of the front sealing member of the tail cone 10 via the inner side of the interstage casing 7. A mounting flange may be provided on the slip ring stator 111 for easy installation.

[0056] like Figure 5 As shown, the casing includes an interstage casing 7, where a bearing 16 is located and configured to support the high-pressure rotor 10'. The bearing 16 has a sealing gas passage. The aero-engine also includes a transition section 18, which has a conical structure with a diameter that gradually decreases along the airflow direction. One end of the transition section 18 is connected to the turbine shaft along the axial direction, and the other end is connected to the joint between the high-pressure rotor 10' and the slip ring rotor 112, for example, it can be connected to a coupling 17. The transition section 18 rotates together with the high-pressure rotor 10'. The transition section 18 has a first through hole 181, through which sealing gas enters the bearing cavity and the interstage casing 7 from the radially inner side of the transition section 18, and exits from the rear end of the tail cone 10, with its flow direction as shown by arrow K.

[0057] Specifically, such as Figure 5 As shown, a second through hole 101 is provided on the mounting edge 10' of the tail nozzle center cone. The sealing gas passes through the first through hole 181, the sealing cavity P of the bearing cavity, and the second through hole 101 in sequence and is discharged from the rear end of the tail cone 10.

[0058] In this embodiment, by providing a first through hole 181 on the transition section 18, the sealing air path of the bearing 16 can be ensured to be continuous. The airflow enters the sealing cavity after the bearing pivot point through the first through hole 181, and is discharged through the bearing housing structure, the interstage casing structure, and the tail cone. This structure can ensure the sealing of the bearing cavity.

[0059] In some embodiments, such as Figure 3 and Figure 8 As shown, the casing includes an interstage casing 7, and the high-pressure rotor test device for the aero-engine also includes a heat shield 15, coaxially sleeved outside the slip ring assembly 11, configured to isolate the high-temperature airflow transmitted by the interstage casing 7 and the tail cone 10. One end of the heat shield 15 is connected to the interstage casing 7 along the axial direction; specifically, one end of the heat shield 15 is connected to the front seal of the tail cone of the interstage casing 7; the other end of the heat shield 15 extends to the guide tube 14 and can be cantilevered. This structure can isolate the high-temperature airflow transmitted by the interstage casing 7 and the tail cone 10, thereby reducing the impact of the high-temperature working environment on the slip ring assembly 11.

[0060] In some embodiments, the high-pressure rotor testing apparatus for an aero-engine further includes a test bench ground equipment, with test leads 12 and cooling pipes 13 extending from the tail cone 10 and connected to the test bench ground equipment. The test leads 12 can be connected to a data acquisition system in the test bench ground equipment, and the cooling pipes 13 can be connected to a cooling medium supply system in the test bench ground equipment.

[0061] Furthermore, this testing device only requires adaptive modifications to the high-pressure rotor 10' (the rear shaft of the high-pressure turbine) and the rear sealing plate of the bearing cavity of bearing 16. It is easy to replace, requires no spare parts, and ensures reliable measurement of parameters such as dynamic strain and temperature of the entire high-pressure rotor 10'. When measurements are not required, only the rear shaft of the high-pressure turbine, the rear sealing plate of the bearing cavity of bearing 16, and the tail cone 10 need to be replaced.

[0062] Secondly, this disclosure also provides an aero-engine test component, which in some embodiments includes the aero-engine high-pressure rotor testing device described above. For example, the aero-engine test component can be a complete aero-engine or an aero-engine core test component 100.

[0063] The above provides a detailed description of an aero-engine high-pressure rotor testing device and an aero-engine test piece provided by this disclosure. Specific embodiments have been used to illustrate the principles and implementation methods of this disclosure. The descriptions of these embodiments are merely for the purpose of helping to understand the method and core ideas of this disclosure. It should be noted that those skilled in the art can make various improvements and modifications to this disclosure without departing from its principles, and these improvements and modifications also fall within the protection scope of the claims of this disclosure.

Claims

1. A high-pressure rotor testing device for an aero-engine, characterized in that, include: The slip ring assembly (11) includes a slip ring stator (111) and a slip ring rotor (112) arranged coaxially. The slip ring stator (111) is fixed on the casing of the aero-engine. The slip ring rotor (112) is arranged coaxially with the slip ring stator (111) and is driven to the rear end of the high-pressure rotor (10') of the aero-engine. A detection component, disposed on the high-voltage rotor (10'), is configured to detect the state parameters of the high-voltage rotor (10'), and the test lead (12) of the detection component is led out through the slip ring assembly (11); Cooling conduit (13) is configured to cool the slip ring assembly (11) by introducing a cooling medium; and A guide tube (14) is axially disposed at the rear end of the slip ring assembly (11), and one end of the guide tube (14) passes through the side wall of the tail cone (10) of the aircraft engine, and is configured to lead out of the aircraft engine from the test lead (12) and the cooling pipe (13); The tail cone (10) includes a first tail cone section (101) and a second tail cone section (102), which are detachably connected axially. The guide tube (14) is located between the first tail cone section (101) and the second tail cone section (102). The casing includes an interstage casing (7), and the slip ring stator (111) is connected to the mounting edge of the front seal of the tail cone (10) through the inner side of the interstage casing (7). The high-pressure rotor test device for the aero-engine also includes a heat shield (15), which is coaxially sleeved outside the slip ring assembly (11) and configured to isolate the high-temperature airflow transmitted by the interstage casing (7) and the tail cone (10). One end of the heat shield (15) is connected to the interstage casing (7), and the other end extends to the guide tube (14).

2. The high-pressure rotor testing device for aero-engines according to claim 1, characterized in that, The guide tube (14) is arranged radially.

3. The high-pressure rotor testing device for aero-engines according to claim 2, characterized in that, The guide tube (14) includes a flexible tube segment, which is located at least within the tail cone (10).

4. The high-pressure rotor testing device for aero-engines according to claim 1, characterized in that, The slip ring rotor (112) and the high-pressure rotor (10') are connected by a coupling (17); wherein, the ends of the slip ring rotor (112) and the high-pressure rotor (10') that are close to each other are provided with mounting holes (1121), and the coupling (17) includes: Main body (171); and The first connecting part (172) and the second connecting part (173) are located at both ends of the main body (171) along the axial direction. The first connecting part (172) extends into the mounting hole (1121) of the high-pressure rotor (10') and is connected to the high-pressure rotor (10'). The second connecting part (173) extends into the mounting hole (1121) of the slip ring rotor (112) and is connected to the slip ring rotor (112). Both the first connecting part (172) and the second connecting part (173) are hexagonal head structures, and the hexagonal head structure has an arc-shaped side surface (174) in the longitudinal section, and the arc-shaped side surface (174) contacts the inner wall of the mounting hole (1121).

5. The high-pressure rotor testing device for aero-engines according to claim 1, characterized in that, The casing includes an interstage casing (7), at which a bearing (16) is provided, configured to support the high-pressure rotor (10'), and the bearing (16) is provided with a sealing gas passage; the aero-engine also includes a transition section (18), one end of which is connected to the turbine shaft along the axial direction, and the other end is connected to the joint of the high-pressure rotor (10') and the slip ring rotor (112), and the transition section (18) is provided with a first through hole (181), through which sealing gas enters the sealing cavity (P) of the bearing cavity and the interstage casing (7), and is discharged from the rear end of the tail cone (10).

6. The high-pressure rotor testing device for aero-engines according to claim 1, characterized in that, It also includes a test bench, wherein the test lead (12) and the cooling pipe (13) are led out from the tail cone (10) and connected to the test bench.

7. An aero-engine test component, characterized in that, include: The high-pressure rotor testing device for aero-engines as described in any one of claims 1 to 6.

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