Method for improving transient thermal deformation of aero-engine rotor

Abstract

The invention provides a method for improving transient thermal deformation of an aero-engine rotor, and the method for improving transient thermal deformation of the aero-engine rotor comprises the steps that a rotor assembly of an engine is covered with at least two thermal barrier coatings close to the gas side, and the thicknesses of the thermal barrier coatings at all positions are inconsistent. According to the method, the specific rotor assembly structure is covered with the thermal barrier coating with the non-uniform thickness, thermal deformation, caused by the temperature gradient, of the two sides of the rotor assembly is reduced in the stages of starting, transition, shutdown and the like, and the method is simple and easy to implement and can be applied to baffle sealing structures including but not limited to turbine working blades.

Description

Technical Field

[0001] The invention relates to the field of air systems in aircraft engines, and in particular to a method for improving transient thermal deformation of an aircraft engine rotor. Background Art

[0002] In the field of aircraft engine air systems, during a typical gas turbine operating cycle, during startup, cooling or transition phases, there is significant thermal expansion and contraction of rotor components, and various parts of the rotor components tend to change temperature at different rates. This can cause significant differences in the expansion / contraction of adjacent components, posing a challenge to the normal operation of the rotor components.

[0003] During the start-up, transition and shutdown of aircraft engines, each rotor component will experience thermal expansion and contraction under the action of centrifugal force and thermal stress. Due to differences in materials, weight, contact area, and temperature levels of cooling / heating air, different components present uneven temperature gradients, resulting in significant differences in thermal expansion and contraction, especially for rotor components with the following characteristics:

[0004] a) One side of the rotor is exposed to the gas;

[0005] b) The other side of the rotor contacts a rotor component with a larger heat capacity, such as a rotor disk.

[0006] Figure 1 It is a schematic diagram of a rotor component in the prior art.

[0007] like Figure 1 As shown, in the prior art, the left side of the rotor component contacts the rotor disk 10, and the working blades 11 are mounted on the rotor disk 10. The right side of the rotor component is exposed to the hot gas. The specific flow characteristics of the hot gas are the main gas flow area 20 and the gas flow weak area 21. In transient processes such as emergency downshifting, the temperature of the main gas flow area 20 decreases rapidly, and the temperature of the gas flow weak area 21 decreases slowly. There is a certain temperature gradient along the radial direction. Due to the large heat capacity of the rotor disk 10, the temperature of the left contact disk 12 that contacts the rotor disk 10 decreases slowly. There is also a considerable temperature gradient on both sides, which increases rapidly during transient processes, resulting in a significant amount of thermal deformation.

[0008] In view of this, the inventor of the present application has designed a method for improving transient thermal deformation of an aircraft engine rotor in order to overcome the above-mentioned technical problems. Summary of the invention

[0009] The technical problem to be solved by the present invention is to overcome the defect in the prior art that an aircraft engine rotor is prone to significant thermal deformation during a transient process, and to provide a method for improving the transient thermal deformation of an aircraft engine rotor.

[0010] The present invention solves the above technical problems through the following technical solutions:

[0011] A method for improving transient thermal deformation of an aircraft engine rotor is characterized in that the method for improving transient thermal deformation of an aircraft engine rotor comprises: covering at least two thermal barrier coatings on the rotor assembly of the engine close to the gas side, and the thickness of the thermal barrier coatings at each location is inconsistent.

[0012] According to one embodiment of the present invention, the method for obtaining the thickness of the thermal barrier coating comprises the following steps:

[0013] S1. In the temperature field calculation and evaluation of the transient process, the change of the hot gas temperature with time at the hot gas under the transient process is loaded;

[0014] S2. On the basis of step S1, the thickness of the thermal barrier coating applied is considered and converted into the profile thickness of the rotor assembly in the temperature field calculation, and the thermal conductivity of the thermal barrier coating thickness corresponding to the initial state is applied, and the thermal conductivity in the main gas flow area is lower than the thermal conductivity in the gas flow weak area.

[0015] According to an embodiment of the present invention, the change is obtained through test data obtained through bench testing, or through performance calculation or numerical simulation.

[0016] According to one embodiment of the present invention, the method for obtaining the thickness of the thermal barrier coating further comprises the following steps:

[0017] S3, with the maximum temperature difference as the goal, by applying coatings of different thickness combinations at the main gas flow area and the gas flow weak area, the temperature difference T between the left contact disk on the rotor disk side and the main gas flow area is δ1 and the temperature difference T between the main gas flow area and the gas flow weak area δ2 Lower than the target temperature difference during the transient process.

[0018] According to one embodiment of the present invention, the method for obtaining the thickness of the thermal barrier coating further comprises the following steps:

[0019] S4. The target value of step S3 is the final non-uniform coating thickness. Each thermal barrier coating can keep the transient temperature change rate of the rotor assembly consistent with the transient temperature change rate of the area of ​​the left contact plate during the rapid heating / cooling process of the right airflow.

[0020] According to one embodiment of the present invention, the thermal barrier coating comprises a first thermal barrier coating and a second thermal barrier coating, and the thickness of the first thermal barrier coating is greater than or less than the thickness of the second thermal barrier coating.

[0021] According to one embodiment of the present invention, one side of the rotor assembly is exposed to the gas, and the other side contacts the rotor disk.

[0022] According to an embodiment of the present invention, the thickness of the thermal barrier coating is determined by a heat exchange system at different radial positions of the rotor assembly.

[0023] According to one embodiment of the present invention, the first thermal barrier coating is located in the main flow area, the second thermal barrier coating is located in the weak flow area, and the thickness of the first thermal barrier coating is greater than the thickness of the second thermal barrier coating.

[0024] The positive and progressive effects of the present invention are:

[0025] The method for improving transient thermal deformation of an aircraft engine rotor of the present invention reduces thermal deformation on both sides of the rotor assembly due to temperature gradients during startup, transition, shutdown and other stages by covering a thermal barrier coating of non-uniform thickness on a specific rotor assembly structure. The method is simple and easy to implement and can be applied to baffle sealing structures including but not limited to turbine working blades. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] The above and other features, properties and advantages of the present invention will become more apparent through the following description in conjunction with the accompanying drawings and embodiments, in which the same reference numerals always represent the same features, wherein:

[0027] Figure 1 It is a schematic diagram of a rotor component in the prior art.

[0028] Figure 2 The present invention is a schematic structural diagram of a rotor component in a method for improving transient thermal deformation of an aero-engine rotor.

[0029] [Reference Signs]

[0030] Rotor disc 10

[0031] Working blade 11

[0032] Gas main circulation area 20

[0033] Gas flow weak area 21

[0034] First Thermal Barrier Coating 100

[0035] Second thermal barrier coating 200

[0036] Rotor assembly 300

[0037] Left contact plate 12 DETAILED DESCRIPTION

[0038] In order to make the above-mentioned objects, features and advantages of the present invention more obvious and easy to understand, the specific embodiments of the present invention are described in detail below with reference to the accompanying drawings.

[0039] Embodiments of the present invention will now be described in detail with reference to the accompanying drawings. Reference will now be made in detail to preferred embodiments of the present invention, examples of which are shown in the accompanying drawings. Wherever possible, the same reference numerals will be used throughout the drawings to represent the same or similar parts.

[0040] Furthermore, although the terms used in the present invention are selected from well-known and commonly used terms, some terms mentioned in the present invention specification may be selected by the applicant at his or her discretion, and their detailed meanings are explained in the relevant parts of the description of this document.

[0041] Furthermore, it is required that the present invention be understood not only by the actual terms used but also by the meanings connoted by each term.

[0042] Figure 2 The present invention is a schematic structural diagram of a rotor component in a method for improving transient thermal deformation of an aero-engine rotor.

[0043] like Figure 2 As shown, the present invention discloses a method for improving transient thermal deformation of an aircraft engine rotor, which comprises: covering at least two thermal barrier coatings close to the gas side on a rotor assembly 300 of the engine, wherein the thickness of the thermal barrier coatings at each location is inconsistent.

[0044] Preferably, the method for obtaining the thickness of the thermal barrier coating comprises the following steps:

[0045] Step S1, in the temperature field calculation and evaluation of a transient process (such as an emergency shutdown transient process), the change of the hot gas temperature over time at the hot gas (such as the main gas flow area 20 and the gas flow weak area 21) under the transient process is loaded.

[0046] Step S2, on the basis of step S1, the thickness of the thermal barrier coating applied is considered and converted into the profile thickness of the rotor assembly 300 in the temperature field calculation, and the thermal conductivity of the thermal barrier coating thickness corresponding to the initial state is applied, and the thermal conductivity at the main gas flow area 20 is lower than the thermal conductivity at the gas flow weak area 21.

[0047] The changes described here are generally test data obtained through bench tests, or obtained through performance calculations or numerical simulations.

[0048] Furthermore, the method for obtaining the thickness of the thermal barrier coating further comprises the following steps:

[0049] Step S3, with the maximum temperature difference as the target, by applying coatings of different thickness combinations at the main gas flow area 20 and the gas flow weak area 21, the temperature difference T between the left contact disk 12 on the side contacting the rotor disk 10 and the main gas flow area 20 is δ1 and the temperature difference T between the main gas flow area 20 and the gas flow weak area 21 δ2 Lower than the target temperature difference during the transient process.

[0050] The method for obtaining the thickness of the thermal barrier coating further comprises the following steps:

[0051] Step S4: The target value of step S3 is the final non-uniform coating thickness. Each thermal barrier coating can keep the transient temperature change of the rotor assembly consistent with the transient temperature change rate of the area of ​​the left contact disk 12 during the rapid heating / cooling process of the right airflow. This can effectively avoid the generation of a large temperature gradient on both sides, thereby improving thermal deformation in the transient process.

[0052] Preferably, in this embodiment, the thermal barrier coating includes a first thermal barrier coating 100 and a second thermal barrier coating 200 , and the thickness of the first thermal barrier coating 100 is greater than or less than the thickness of the second thermal barrier coating 200 .

[0053] One side of the rotor assembly 300 is exposed to the gas, and the other side contacts the rotor disk 10 .

[0054] In particular, the thickness of the thermal barrier coating described herein is determined by the heat exchange system at different radial positions of the rotor assembly 300 .

[0055] When the first thermal barrier coating 100 is located in the main flow region 20 and the second thermal barrier coating 200 is located in the weak flow region 21 , the thickness of the first thermal barrier coating 100 is greater than the thickness of the second thermal barrier coating 200 .

[0056] The above-mentioned specific rotor component structure that meets the characteristics and the method for obtaining the non-uniform thickness thermal barrier coating that meets the target temperature difference are combined according to a certain logic to obtain a method for improving the transient thermal deformation of the aircraft engine rotor.

[0057] According to the above description, the method for improving transient thermal deformation of an aircraft engine rotor of the present invention is to cover the rotor assembly 300 with a thermal barrier coating (e.g., a first thermal barrier coating 100 and a second thermal barrier coating 200) of non-uniform thickness near the gas side, and the difference between the thermal barrier coatings is mainly in the coating thickness, and the specific thickness is determined by the heat transfer coefficient at the radial position of the flowing gas. The thermal barrier coating can keep the transient temperature change of the rotor assembly 300 consistent with the transient temperature change rate of the area of ​​the left contact disk during the rapid heating / cooling process of the right airflow, avoid the generation of a large temperature gradient on both sides, and thus improve the thermal deformation in the transient process.

[0058] In summary, the method of the present invention for improving the transient thermal deformation of an aircraft engine rotor is to cover a specific rotor component structure with a thermal barrier coating of non-uniform thickness, thereby reducing the thermal deformation on both sides of the rotor component due to the temperature gradient during the startup, transition, and shutdown stages. The method is simple and easy to implement, and can be applied to baffle sealing structures including but not limited to turbine working blades.

[0059] 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 implementation of the present application.

[0060] 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.

[0061] Although the specific embodiments of the present invention are described above, it should be understood by those skilled in the art that these are only examples, and the protection scope of the present invention is defined by the appended claims. Those skilled in the art may make various changes or modifications to these embodiments without departing from the principles and essence of the present invention, but these changes and modifications all fall within the protection scope of the present invention.

Claims

1. A method for improving transient thermal deformation of an aero-engine rotor, characterized in that: The method for improving transient thermal deformation of an aircraft engine rotor comprises: covering at least two thermal barrier coatings on the rotor assembly of the engine close to the gas side, wherein the thickness of the thermal barrier coatings at each location is inconsistent.

2. The method for improving transient thermal deformation of an aircraft engine rotor according to claim 1, characterized in that: The method for obtaining the thickness of the thermal barrier coating comprises the following steps: S1. In the temperature field calculation and evaluation of the transient process, the change of the hot gas temperature with time at the hot gas under the transient process is loaded; S2. On the basis of step S1, the thickness of the thermal barrier coating applied is considered and converted into the profile thickness of the rotor assembly in the temperature field calculation, and the thermal conductivity of the thermal barrier coating thickness corresponding to the initial state is applied, and the thermal conductivity in the main gas flow area is lower than the thermal conductivity in the gas flow weak area.

3. The method for improving transient thermal deformation of an aircraft engine rotor according to claim 2, characterized in that: The changes are obtained through test data obtained through bench tests, or through performance calculations or numerical simulations.

4. The method for improving transient thermal deformation of an aircraft engine rotor according to claim 2, characterized in that: The method for obtaining the thickness of the thermal barrier coating further comprises the following steps: S3, with the maximum temperature difference as the goal, by applying coatings of different thickness combinations at the main gas flow area and the gas flow weak area, the temperature difference T between the left contact disk on the rotor disk side and the main gas flow area is δ1 and the temperature difference T between the main gas flow area and the gas flow weak area δ2 Lower than the target temperature difference during the transient process.

5. The method for improving transient thermal deformation of an aircraft engine rotor according to claim 4, characterized in that: The method for obtaining the thickness of the thermal barrier coating further comprises the following steps: S4. The target value of step S3 is the final non-uniform coating thickness. Each thermal barrier coating can keep the transient temperature change rate of the rotor assembly consistent with the transient temperature change rate of the area of ​​the left contact plate during the rapid heating / cooling process of the right airflow.

6. The method for improving transient thermal deformation of an aircraft engine rotor according to claim 1, characterized in that: The thermal barrier coating includes a first thermal barrier coating and a second thermal barrier coating, wherein a thickness of the first thermal barrier coating is greater than or less than a thickness of the second thermal barrier coating.

7. The method for improving transient thermal deformation of an aircraft engine rotor according to claim 1, characterized in that: One side of the rotor assembly is exposed to the gas and the other side contacts the rotor disk.

8. The method for improving transient thermal deformation of an aircraft engine rotor according to claim 1, characterized in that: The thickness of the thermal barrier coating is determined by the heat exchange system at different radial positions of the rotor assembly.

9. The method for improving transient thermal deformation of an aircraft engine rotor according to claim 6, characterized in that: The first thermal barrier coating is located in a main flow area, the second thermal barrier coating is located in a weak flow area, and a thickness of the first thermal barrier coating is greater than a thickness of the second thermal barrier coating.

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