Aero-engine

Abstract

The invention provides an aero-engine. The aero-engine comprises a turbine casing; a combustion chamber casing; the temperature sensor comprises a shell, a sensing part and a fixed connecting part, the shell is of a cylinder or square cylinder structure with a first length; the fixed connecting part is arranged at the first end part of the shell and surrounds one section, corresponding to the fixed connecting part, of the shell; the second end part of the shell is provided with a closed structure; a fixed connecting part of the temperature sensor is arranged on the outer side of the turbine casing and / or the combustion chamber casing, and a shell and a sensing part of the temperature sensor extend into the turbine casing and / or the combustion chamber casing so as to measure the temperature of airflow in the turbine casing and / or the combustion chamber casing; a first overflowing hole pair is formed in the second end part, close to the temperature sensor, of the shell; a second overflowing hole pair is formed in the middle part or the middle upper part of the shell of the temperature sensor; the central axis of the first overflowing hole pair and the central axis of the second overflowing hole pair are parallel to the axial direction of the aero-engine.

Description

Technical Field

[0001] The present invention mainly relates to the field of aviation, and in particular to an aviation engine. Background Art

[0002] As a highly complex and sophisticated thermal machine, aircraft engines are in high temperature, high pressure and high speed when working. For example, the combustion chamber and turbine of an aircraft engine have high temperature airflow passing through them. When the aircraft engine is working, accurately measuring the temperature of the combustion chamber and turbine is an important part of monitoring the operating status of the aircraft engine. Therefore, how to avoid interference from other factors when measuring the temperature of the combustion chamber and turbine of an aircraft engine and obtain accurate temperature values ​​of the combustion chamber and turbine is a problem that needs to be addressed. Summary of the invention

[0003] The technical problem to be solved by the present invention is to provide an aircraft engine, which can achieve accurate measurement of the temperature of the hot air flow in the combustion chamber and turbine part of the aircraft engine.

[0004] In order to solve the above technical problems, the present invention provides an aircraft engine, comprising: a turbine casing; a combustion chamber casing; a temperature sensor, comprising a shell, a sensing component and a fixed connection portion; the shell is a cylindrical or square cylindrical structure with a first length; the fixed connection portion is arranged at the first end of the shell and surrounds a section of the shell corresponding to the fixed connection portion; the second end of the shell has a closed structure; the sensing component is installed in the shell and has a second length; the second length is less than the first length; the fixed connection portion of the temperature sensor is installed on the outside of the turbine casing and / or the combustion chamber casing, and the shell and the sensing component of the temperature sensor extend into the interior of the turbine casing and / or the combustion chamber casing to measure the temperature of the airflow in the turbine casing and / or the combustion chamber casing; a first pair of flow holes is opened at the second end of the shell close to the temperature sensor, and a second pair of flow holes is opened in the middle or upper middle part of the shell of the temperature sensor; the central axes of the first pair of flow holes and the second pair of flow holes are parallel to the axial direction of the aircraft engine.

[0005] In one embodiment of the present invention, the fixed connection part includes a first threaded structure; the turbine casing and / or the combustion chamber casing has a third mounting hole structure corresponding to the fixed connection part, and the inner wall of the third mounting hole structure has a second threaded structure to engage with the first threaded structure of the fixed connection part by screwing.

[0006] In one embodiment of the present invention, the fixed connection portion includes a flange structure and a fourth mounting hole structure located on both sides of the flange structure; the turbine casing and / or the combustion chamber casing has a fifth mounting hole structure corresponding to the fourth mounting hole structure; the flange structure is installed on the outside of the turbine casing and / or the combustion chamber casing by means of bolts or screw components passing through the fourth mounting hole structure and the fifth mounting hole structure.

[0007] In one embodiment of the present invention, the aircraft engine comprises a plurality of the temperature sensors; the plurality of temperature sensors are evenly distributed along the circumference of the turbine casing and / or the combustion chamber casing.

[0008] In an embodiment of the present invention, the temperature sensor further comprises an electrical connection portion for connecting an electrical signal generated by the sensing component based on the sensed temperature.

[0009] In one embodiment of the present invention, the material of the housing of the temperature sensor includes a high-temperature alloy; the high temperature includes a temperature above 600° C.

[0010] In one embodiment of the present invention, the sensing component comprises a thermocouple wire.

[0011] In an embodiment of the present invention, the electrical connection portion is disposed at an upper end of the fixed connection portion.

[0012] In one embodiment of the present invention, the first flow hole pair and the second flow hole pair account for 1 / 8 to 1 / 4 of the circumference of the shell.

[0013] In one embodiment of the present invention, when the second pair of flow holes is located in the middle of the shell of the temperature sensor, a third pair of flow holes is also opened in the middle and upper part of the shell of the temperature sensor; when the second pair of flow holes is located in the middle and upper part of the shell of the temperature sensor, a third pair of flow holes is also opened in the middle of the shell of the temperature sensor.

[0014] In one embodiment of the present invention, the diameters of the flow holes of the first flow hole pair and the second flow hole pair are 3-10 mm.

[0015] In one embodiment of the present invention, each of the turbine casing or the combustion chamber casing is circumferentially equipped with 4 to 8 temperature sensors.

[0016] Compared with the prior art, the present invention has the following advantages: the technical solution of the present application, by setting a plurality of flow hole pairs on the housing of the temperature sensor of the aircraft engine, can realize accurate temperature measurement of the airflow in the turbine casing and / or combustion chamber casing of the aircraft engine; at the same time, it can avoid the influence of heat conduction of the turbine casing and / or combustion chamber casing on the temperature measurement of the hot air flow. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] The accompanying drawings are provided to provide a further understanding of the present application. They are included in and constitute a part of the present application. The accompanying drawings illustrate embodiments of the present application and together with the description serve to explain the principles of the present application.

[0018] In the attached figure:

[0019] Figure 1 It is a schematic structural diagram of a turbine casing or a combustion chamber casing of an aircraft engine according to an embodiment of the present application.

[0020] Figure 2 It is a schematic diagram of the structure of a temperature sensor for an aircraft engine according to an embodiment of the present application.

[0021] Figure 3 It is a schematic diagram of the structure of a temperature sensor for an aircraft engine according to another embodiment of the present application.

[0022] Figure 4 It is a schematic diagram of the first cross-sectional structure of a temperature sensor for an aircraft engine according to an embodiment of the present application.

[0023] Figure 5 It is a schematic diagram of the first cross-sectional structure of a temperature sensor for an aircraft engine according to another embodiment of the present application.

[0024] Figure 6 It is a schematic diagram of the second cross-sectional structure of the temperature sensor of the aircraft engine according to one embodiment of the present application.

[0025] Figure 7 It is a schematic diagram of the structure of a temperature sensor for an aircraft engine according to another embodiment of the present application.

[0026] Figure 8 It is a schematic diagram of the structure of a temperature sensor for an aircraft engine according to another embodiment of the present application. DETAILED DESCRIPTION

[0027] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the following is a brief introduction to the drawings required for the description of the embodiments. Obviously, the drawings described below are only some examples or embodiments of the present application. For ordinary technicians in this field, the present application can also be applied to other similar scenarios based on these drawings without creative work. Unless it is obvious from the language environment or otherwise explained, the same reference numerals in the figures represent the same structure or operation.

[0028] As shown in this application and claims, unless the context clearly indicates an exception, the words "a", "an", "an" and / or "the" do not refer to the singular and may also include the plural. Generally speaking, the terms "include" and "comprise" only indicate the inclusion of the steps and elements that have been clearly identified, and these steps and elements do not constitute an exclusive list. The method or device may also include other steps or elements.

[0029] Unless otherwise specifically stated, the relative arrangement of the components and steps, numerical expressions and numerical values ​​described in these embodiments do not limit the scope of the present application. At the same time, it should be understood that, for the convenience of description, the sizes of the various parts shown in the drawings are not drawn according to the actual proportional relationship. In the description of the present application, it should be understood that the orientation or position relationship indicated by the orientation words such as "front, back, up, down, left, right", "lateral, vertical, vertical, horizontal" and "top, bottom" are usually based on the orientation or position relationship shown in the drawings, only for the convenience of describing the present application and simplifying the description. In the absence of any contrary instructions, these orientation words do not indicate or imply that the device or element referred to must have a specific orientation or be constructed and operated in a specific orientation, and therefore cannot be understood as a limitation on the scope of protection of the present application; the orientation words "inside and outside" refer to the inside and outside relative to the contours of the components themselves.

[0030] For ease of description, spatially relative terms such as "above", "above", "on the upper surface of", "above", etc. may be used here to describe the spatial positional relationship between a device or feature and other devices or features as shown in the figure. It should be understood that spatially relative terms are intended to include different orientations of the device in use or operation in addition to the orientation described in the figure. For example, if the device in the accompanying drawings is inverted, the device described as "above other devices or structures" or "above other devices or structures" will be positioned as "below other devices or structures" or "below other devices or structures". Thus, the exemplary term "above" can include both "above" and "below". The device can also be positioned in other different ways (rotated 90 degrees or in other orientations), and the spatially relative descriptions used here are interpreted accordingly.

[0031] In addition, it should be noted that the use of terms such as "first" and "second" to limit components is only for the convenience of distinguishing the corresponding components. Unless otherwise stated, the above terms have no special meaning and therefore cannot be understood as limiting the scope of protection of this application.

[0032] The embodiment of the present application describes an aircraft engine. The aircraft engine includes a turbine casing, a combustion chamber casing and a temperature sensor. The aircraft engine also includes structures such as a fan, a rotating shaft, blades and a compressor.

[0033] Figure 1 It is a schematic structural diagram of a turbine casing or a combustion chamber casing of an aircraft engine according to an embodiment of the present application. Figure 2 It is a schematic diagram of the structure of a temperature sensor for an aircraft engine according to an embodiment of the present application.

[0034] Figure 3 It is a schematic diagram of the structure of a temperature sensor for an aircraft engine according to another embodiment of the present application. Figure 4 It is a schematic diagram of the first cross-sectional structure of a temperature sensor for an aircraft engine according to an embodiment of the present application. Figure 5 It is a schematic diagram of the first cross-sectional structure of a temperature sensor for an aircraft engine according to another embodiment of the present application.

[0035] Figure 6 It is a schematic diagram of the second cross-sectional structure of the temperature sensor of the aircraft engine according to one embodiment of the present application.

[0036] Figure 6 For along Figure 3 The section line AA' in the figure provides a cross-sectional view of the temperature sensor of the aircraft engine.

[0037] refer to Figures 1 to 6 The temperature sensor 200 includes a housing 211, a sensing component 231 and a fixed connection portion 215. The housing 211 is a cylindrical or square cylindrical structure having a first length s1. The fixed connection portion 215 is disposed at a first end 211a of the housing 211 and surrounds a section of the housing 211 corresponding to the fixed connection portion 215. The second end 211b of the housing 211 has a closed structure. The sensing component 231 is installed in the housing 211 and has a second length s2. The second length s2 is less than the first length s1.

[0038] The housing structure of the temperature sensor in the solution of the present application can provide protection for the sensing component to avoid damage to the sensing component during the use of the temperature sensor, thereby preventing abnormal temperature measurement or failure.

[0039] In some embodiments, the fixed connection portion 215 of the temperature sensor 200 is installed on the outer side 101 of the turbine casing and / or the combustion chamber casing, and the housing 211 and the sensing component 231 of the temperature sensor 200 extend into the interior 103 of the turbine casing and / or the combustion chamber casing to measure the temperature of the airflow (or hot airflow) 331 in the turbine casing and / or the combustion chamber casing.

[0040] A first pair of flow holes 221 is provided at the second end 211b of the housing 211 close to the temperature sensor 200, and a second pair of flow holes 223 is provided at the middle or upper middle portion of the housing 211 of the temperature sensor 200. The central axis X2 of the first pair of flow holes 221 and the second pair of flow holes 223 is parallel to the axial direction X1 of the aircraft engine.

[0041] The first pair of flow holes 221 includes a flow hole 221a and a flow hole 221b. The second pair of flow holes 223 includes a flow hole 223a and a flow hole 223b.

[0042] In some embodiments, the first flow hole pair 221 and the second flow hole pair 223 account for 1 / 8 to 1 / 4 of the circumference of the housing 211. The diameter of the flow holes of the first flow hole pair 221 and the second flow hole pair 223 is 3-10 mm (millimeter).

[0043] The aircraft engine includes a plurality of temperature sensors 200. The plurality of temperature sensors 200 are evenly distributed along the circumference of the turbine casing and / or the combustion chamber casing. Each turbine casing or the combustion chamber casing is provided with 4 to 8 temperature sensors along the circumference. Figure 1 It is shown that the turbine casing or the combustion chamber casing 100 is provided with a first temperature sensor 201 , a second temperature sensor 202 , a third temperature sensor 203 and the like along the circumferential direction.

[0044] In some embodiments, reference Figure 2 The fixed connection part 215 includes a flange structure 213 and a fourth mounting hole structure 216 located on both sides of the flange structure 213. The turbine casing and / or the combustion chamber casing has a fifth mounting hole structure 218 corresponding to the fourth mounting hole structure 216. The flange structure 213 is installed on the outer side 101 of the turbine casing and / or the combustion chamber casing by means of a bolt or screw component 219 that passes through the fourth mounting hole structure 216 and the fifth mounting hole structure 218. The flange structure 213 and the outer side 101 of the turbine casing and / or the combustion chamber casing may further include a gasket structure. The gasket structure is, for example, a material with high specific heat capacity, such as aluminum, copper or other metal materials.

[0045] In some embodiments, reference Figure 3 The fixed connection part 215 includes a first thread structure 214. The turbine casing and / or the combustion chamber casing has a third mounting hole structure 410 corresponding to the fixed connection part 215, and the inner wall of the third mounting hole structure 410 has a second thread structure 412 to engage with the first thread structure 214 of the fixed connection part 215 by screwing.

[0046] In some embodiments, the temperature sensor 200 further includes an electrical connection portion 401 to connect the electrical signal generated by the sensing component 231 based on the sensed temperature. An electrical signal transmission interface 402 is provided on one side of the electrical connection portion 401. The electrical connection portion 401 is provided at the upper end 215a of the fixed connection portion 215.

[0047] In some embodiments, the material of the housing 211 of the temperature sensor 200 includes, for example, a high temperature alloy. High temperature includes a temperature above 600°C. High temperature alloy is a type of metal material based on iron, nickel, and cobalt, which works for a long time under high temperature and certain stress, and has excellent high temperature strength, good oxidation resistance and thermal corrosion resistance, good fatigue performance, fracture toughness and other comprehensive properties. The sensing component 231 includes, for example, a thermocouple wire.

[0048] Figure 7 It is a schematic diagram of the structure of a temperature sensor for an aircraft engine according to another embodiment of the present application.

[0049] Figure 8 It is a schematic diagram of the structure of a temperature sensor for an aircraft engine according to another embodiment of the present application.

[0050] refer to Figure 7 and Figure 8 When the second flow hole pair 223 is located in the middle of the housing 211 of the temperature sensor 200, a third flow hole pair 224 is also provided in the middle and upper part of the housing 211 of the temperature sensor 200. When the second flow hole pair 223 is located in the middle and upper part of the housing 211 of the temperature sensor 200, a third flow hole pair 224 is also provided in the middle of the housing 211 of the temperature sensor 200. The size of the third flow hole pair 224 is similar to or the same as that of the first flow hole pair 221 and the second flow hole pair 223.

[0051] The aircraft engine of the present application can achieve accurate temperature measurement of the airflow in the turbine casing and / or combustion chamber casing of the aircraft engine by setting up multiple flow hole pairs on the housing of the temperature sensor; at the same time, by reducing the contact area between the turbine casing and / or combustion chamber casing and the housing, the influence of heat conduction of the turbine casing and / or combustion chamber casing on the temperature measurement of the hot air flow can be avoided.

[0052] The technical solution of the present application can improve the accuracy of temperature measurement of hot air flow in the turbine casing and / or combustion chamber casing of an aircraft engine by tens of degrees Celsius (°C), effectively improving the accuracy of temperature measurement of hot air flow, thereby providing effective data support for engine operation control.

[0053] The basic concepts have been described above. Obviously, for those skilled in the art, the above invention disclosure is only used as an example and does not constitute a limitation of the present application. Although not explicitly stated herein, those skilled in the art may make various modifications, improvements and amendments to the present application. Such modifications, improvements and amendments are suggested in the present application, so such modifications, improvements and amendments still belong to the spirit and scope of the exemplary embodiments of the present application.

[0054] At the same time, the present application uses specific words to describe the embodiments of the present application. For example, "one embodiment", "an embodiment", and / or "some embodiments" refer to a certain feature, structure or characteristic related to at least one embodiment of the present application. Therefore, it should be emphasized and noted that "one embodiment" or "an embodiment" or "an alternative embodiment" mentioned twice or more in different positions in this specification does not necessarily refer to the same embodiment. In addition, some features, structures or characteristics in one or more embodiments of the present application can be appropriately combined.

[0055] Similarly, it should be noted that in order to simplify the description of the disclosure of this application and thus help understand one or more embodiments of the invention, in the above description of the embodiments of this application, multiple features are sometimes combined into one embodiment, figure or description thereof. However, this disclosure method does not mean that the features required by the object of this application are more than the features mentioned in the claims. In fact, the features of the embodiments are less than all the features of the single embodiment disclosed above.

[0056] In some embodiments, numbers describing the quantity of components and attributes are used. It should be understood that such numbers used in the description of the embodiments are modified by the modifiers "about", "approximately" or "substantially" in some examples. Unless otherwise specified, "about", "approximately" or "substantially" indicates that the numbers allow a variation of ±20%.

[0057] Although the present application has been described with reference to the current specific embodiments, ordinary technicians in this technical field should recognize that the above embodiments are only used to illustrate the present application, and various equivalent changes or substitutions may be made without departing from the spirit of the present application. Therefore, as long as the changes and modifications to the above embodiments are within the essential spirit of the present application, they will fall within the scope of the claims of the present application.

Claims

1. An aircraft engine, comprising: turbine casing; Combustion chamber casing; A temperature sensor, comprising a housing, a sensing component and a fixed connection portion; the housing is a cylindrical or square cylindrical structure having a first length; the fixed connection portion is arranged at a first end of the housing and surrounds a section of the housing corresponding to the fixed connection portion; the second end of the housing has a closed structure; the sensing component is installed in the housing and has a second length; the second length is less than the first length; The fixed connection portion of the temperature sensor is installed on the outside of the turbine casing and / or the combustion chamber casing, and the housing and the sensing component of the temperature sensor extend into the interior of the turbine casing and / or the combustion chamber casing to measure the temperature of the airflow in the turbine casing and / or the combustion chamber casing; A first pair of flow holes is provided at a second end of the housing close to the temperature sensor, and a second pair of flow holes is provided at a middle or upper middle portion of the housing of the temperature sensor; The central axes of the first flow hole pair and the second flow hole pair are parallel to the axial direction of the aircraft engine.

2. The aircraft engine according to claim 1, characterized in that: The fixed connection part includes a first threaded structure; the turbine casing and / or the combustion chamber casing has a third mounting hole structure corresponding to the fixed connection part, and the inner wall of the third mounting hole structure has a second threaded structure to engage with the first threaded structure of the fixed connection part by screwing.

3. The aircraft engine according to claim 1, characterized in that: The fixed connection part includes a flange structure and a fourth mounting hole structure located on both sides of the flange structure; the turbine casing and / or the combustion chamber casing has a fifth mounting hole structure corresponding to the fourth mounting hole structure; the flange structure is installed on the outside of the turbine casing and / or the combustion chamber casing by means of bolts or screw components passing through the fourth mounting hole structure and the fifth mounting hole structure.

4. The aircraft engine according to claim 1, characterized in that: The aircraft engine includes a plurality of the temperature sensors; the plurality of temperature sensors are evenly distributed along the circumference of the turbine casing and / or the combustion chamber casing.

5. The aircraft engine according to claim 1, characterized in that: The temperature sensor further includes an electrical connection portion for connecting an electrical signal generated by the sensing component based on the sensed temperature.

6. The aircraft engine according to claim 1, characterized in that: The material of the housing of the temperature sensor includes a high-temperature alloy; the high temperature includes a temperature above 600°C.

7. The aircraft engine according to claim 1, characterized in that: The sensing component includes a thermocouple wire.

8. The aircraft engine according to claim 5, characterized in that: The electrical connection portion is arranged at the upper end of the fixed connection portion.

9. The aircraft engine according to claim 1, characterized in that: The first flow hole pair and the second flow hole pair account for 1 / 8 to 1 / 4 of the circumference of the shell.

10. The aircraft engine according to claim 1, characterized in that: When the second pair of flow holes is located in the middle of the housing of the temperature sensor, a third pair of flow holes is further provided in the middle and upper part of the housing of the temperature sensor; When the second pair of flow holes is located in the upper middle portion of the housing of the temperature sensor, a third pair of flow holes is further provided in the middle portion of the housing of the temperature sensor.

11. The aircraft engine according to claim 1, characterized in that: The diameters of the flow holes of the first flow hole pair and the second flow hole pair are 3-10 mm.

12. The aircraft engine according to claim 4, characterized in that: Each turbine casing or combustion chamber casing is provided with 4 to 8 temperature sensors along the circumferential direction.

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