A method for manufacturing a test piece for a cooling effect test of an aircraft engine turbine blade

Abstract

The present application belongs to the technical field of aircraft engine turbine blade cooling effect test, and specifically relates to a method for manufacturing an aircraft engine turbine blade cooling effect test piece, comprising: sandblasting the surface of an aircraft engine turbine blade actually produced and delivered to obtain an aircraft engine turbine blade without a thermal barrier coating; grooving the blade surface of the aircraft engine turbine blade without a thermal barrier coating to seal the air film holes on the upper and lower edge plate surfaces; setting a thermocouple in the groove on the blade surface of the aircraft engine turbine blade without a thermal barrier coating; spraying a thermal barrier coating on the blade surface of the aircraft engine turbine blade without a thermal barrier coating to obtain an aircraft engine turbine blade cooling effect test piece.

Description

Technical Field

[0001] The present application belongs to the technical field of cooling effect test of aircraft engine turbine blades, and specifically relates to a method for manufacturing a test piece for cooling effect test of aircraft engine turbine blades. Background Art

[0002] In order to evaluate the cooling effect of aircraft engine turbine blades, a corresponding cooling effect test is designed. When conducting the cooling effect test of aircraft engine turbine blades, the aircraft engine turbine blade cooling effect test piece and multiple aircraft engine turbine blade lining parts form a blade cascade, such as Figure 1 As shown, high-temperature combustion gas is introduced into the blade channel, and cooling gas is introduced into the test piece and the supporting piece, the flow ratio and temperature ratio between the high-temperature combustion gas and the cooling gas are changed, the Reynolds number of the high-temperature combustion gas is changed, and the flow rate of the cooling gas is changed. The temperature distribution on the blade of the test piece is measured with a thermocouple embedded in the surface of the blade of the test piece, and the corresponding blade cooling effect curve is obtained.

[0003] At present, the test pieces for the cooling effect test of aircraft engine turbine blades are mostly made according to the following methods:

[0004] Machining of aircraft engine turbine blades without thermal barrier coatings;

[0005] Grooving the blade surface of an aircraft engine turbine blade without a thermal barrier coating;

[0006] A thermocouple is arranged in a slot on the surface of a turbine blade of an aircraft engine without a thermal barrier coating;

[0007] A thermal barrier coating is sprayed on the blade surface of an aircraft engine turbine blade without a thermal barrier coating to obtain an aircraft engine turbine blade cooling effect test piece.

[0008] The aero-engine turbine blade cooling effect test specimen obtained by the above method has the following defects:

[0009] There is a dimensional deviation between the processed aircraft engine turbine blades without thermal barrier coating and the aircraft engine turbine blades actually produced and delivered. Thermal barrier coating is sprayed on the blade surface of the aircraft engine turbine blade without thermal barrier coating to obtain an aircraft engine turbine blade cooling effect test specimen. The air film holes on the blade surface of the aircraft engine turbine blade cooling effect test specimen will have shrinkage holes, and it is difficult to ensure that the aperture of the air film holes on the blade surface of the aircraft engine turbine blade actually delivered is consistent with that of the air film holes on the blade surface of the aircraft engine turbine blade actually produced and delivered, which will affect the accuracy of the test results and cause the distribution of cooling air flow between the upper and lower edge plates of the test specimen and the air film holes on the blade surface to deviate from the cooling air flow distribution between the upper and lower edge plates of the aircraft engine turbine blade actually delivered and delivered, further affecting the accuracy of the test results.

[0010] This application is proposed in view of the above-mentioned technical defects.

[0011] It should be noted that the disclosure of the above background technology content is only used to assist in understanding the inventive concept and technical solution of the present invention, and it does not necessarily belong to the prior art of this patent application. In the absence of clear evidence that the above content has been disclosed on the filing date of this application, the above background technology should not be used to evaluate the novelty and creativity of the present application. Summary of the invention

[0012] The purpose of the present application is to provide a method for manufacturing a test piece for an aircraft engine turbine blade cooling effect test, so as to overcome or alleviate at least one of the known technical defects.

[0013] The technical solution of this application is:

[0014] A method for manufacturing a test piece for an aero-engine turbine blade cooling effect test, comprising:

[0015] Sandblasting is performed on the surface of the aircraft engine turbine blades actually produced and delivered to obtain aircraft engine turbine blades without thermal barrier coatings;

[0016] Grooving the blade surface of an aircraft engine turbine blade without a thermal barrier coating to seal the air film holes on the upper and lower edge surfaces;

[0017] A thermocouple is arranged in a slot on the surface of a turbine blade of an aircraft engine without a thermal barrier coating;

[0018] A thermal barrier coating is sprayed on the blade surface of an aircraft engine turbine blade without a thermal barrier coating to obtain an aircraft engine turbine blade cooling effect test piece.

[0019] According to at least one embodiment of the present application, the above-mentioned method for manufacturing a test piece for a cooling effect test of a turbine blade of an aero-engine further includes:

[0020] Measuring the aperture of film holes on the surface of aircraft engine turbine blades without thermal barrier coating;

[0021] If the aperture of the film holes on the blade surface of an aircraft engine turbine blade without thermal barrier coating does not match the aperture of the film holes on the blade surface of an aircraft engine turbine blade without thermal barrier coating on the process drawing actually delivered during production, the aircraft engine turbine blade cooling effect test specimen shall be remade.

[0022] According to at least one embodiment of the present application, in the above-mentioned method for preparing a test piece for a cooling effect test of an aircraft engine turbine blade, spraying a thermal barrier coating on the blade surface of an aircraft engine turbine blade without a thermal barrier coating to obtain the test piece for the cooling effect test of an aircraft engine turbine blade comprises:

[0023] On the blade surface of the aircraft engine turbine blade without thermal barrier coating, cover the air film holes and spray the thermal barrier coating bottom layer;

[0024] On the surface of the aero-engine turbine blade with a thermal barrier coating bottom layer, the covering of the air film holes is removed, polishing is performed, and a thermal barrier coating surface layer is sprayed to obtain an aero-engine turbine blade cooling effect test piece.

[0025] According to at least one embodiment of the present application, in the above-mentioned method for manufacturing a test piece for an aircraft engine turbine blade cooling effect test, the air film holes are covered on the blade surface of an aircraft engine turbine blade without a thermal barrier coating, specifically by covering the air film holes with metal foil.

[0026] According to at least one embodiment of the present application, in the above-mentioned method for manufacturing a test piece for a cooling effect test of an aircraft engine turbine blade, the bottom layer of the thermal barrier coating is a NiCrAlY coating;

[0027] The surface layer of the thermal barrier coating is a ZrO3 coating.

[0028] According to at least one embodiment of the present application, the above-mentioned method for manufacturing a test piece for a cooling effect test of a turbine blade of an aero-engine further includes:

[0029] Measuring the aperture of the film holes on the polished surface of the turbine blade of an aircraft engine with a thermal barrier coating bottom layer;

[0030] If the aperture of the air film holes on the blade body surface of the aircraft engine turbine blade with the thermal barrier coating bottom layer is smaller than the aperture of the air film holes on the blade body surface of the aircraft engine turbine blade actually produced and delivered, the aircraft engine turbine blade cooling effect test specimen shall be re-made;

[0031] If the aperture of the air film holes on the blade body surface of an aircraft engine turbine blade with a thermal barrier coating bottom layer is not less than the aperture of the air film holes on the blade body surface of an aircraft engine turbine blade actually produced and delivered, the thermal barrier coating surface layer is sprayed to obtain an aircraft engine turbine blade cooling effect test specimen.

[0032] According to at least one embodiment of the present application, the above-mentioned method for manufacturing a test piece for a cooling effect test of a turbine blade of an aero-engine further includes:

[0033] Measuring the diameter of the air film holes on the blade surface of the aircraft engine turbine blade cooling effect test specimen;

[0034] If the aperture of the air film holes on the blade surface of the aircraft engine turbine blade cooling effect test specimen does not match the aperture of the air film holes on the blade surface of the aircraft engine turbine blade actually produced and delivered, the aircraft engine turbine blade cooling effect test specimen shall be remade. BRIEF DESCRIPTION OF THE DRAWINGS

[0035] Figure 1 The schematic diagram is a cascade composed of an aircraft engine turbine blade cooling effect test piece and a plurality of aircraft engine turbine blade lining parts;

[0036] Figure 2 It is a schematic diagram of a method for manufacturing a test piece for an aircraft engine turbine blade cooling effect test provided in an embodiment of the present application.

[0037] In order to better illustrate this embodiment, some parts of the drawings may be omitted, enlarged or reduced, and do not represent the size of the actual product. In addition, the drawings are only used for illustrative purposes and cannot be understood as limitations on this patent. DETAILED DESCRIPTION

[0038] In order to make the technical solution and advantages of the present application clearer, the technical solution of the present application will be further described in detail in detail and in detail in conjunction with the accompanying drawings. It can be understood that the specific embodiments described here are only partial embodiments of the present application, which are only used to explain the present application, not to limit the present application. It should be noted that, for the convenience of description, only the parts related to the present application are shown in the accompanying drawings, and other related parts can refer to the general design. In the absence of conflict, the embodiments in the present application and the technical features in the embodiments can be combined with each other to obtain new embodiments.

[0039] In addition, unless otherwise defined, the technical terms or scientific terms used in the description of this application shall be the common meanings understood by the general technicians in the field to which this application belongs. The words "upper", "lower", "left", "right", "center", "vertical", "horizontal", "inside", "outside" and other words indicating orientation used in the description of this application are only used to indicate the relative direction or positional relationship, and do not imply that the device or element must have a specific orientation, be constructed and operated in a specific orientation. When the absolute position of the described object changes, its relative positional relationship may also change accordingly, so it cannot be understood as a limitation on this application. The "first", "second", "third" and similar terms used in the description of this application are only used for descriptive purposes to distinguish different components, and cannot be understood as indicating or implying relative importance. The similar words "one", "one" or "the" used in the description of this application should not be understood as an absolute limitation on quantity, but should be understood as the existence of at least one. The similar words "including" or "comprising" used in the description of this application mean that the elements or objects appearing in front of the word include the elements or objects listed after the word and their equivalents, without excluding other elements or objects.

[0040] In addition, it should be noted that, unless otherwise clearly specified and limited, the words "installed", "connected", "connected" and similar terms used in the description of this application should be understood in a broad sense. For example, the connection can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection, an indirect connection through an intermediate medium, or the internal connection of two components. Technical personnel in the field can understand their specific meanings in this application according to the specific circumstances.

[0041] The following is combined with Figure 1 to Figure 2 This application is described in further detail.

[0042] A method for manufacturing a test piece for an aero-engine turbine blade cooling effect test, comprising:

[0043] Sandblasting is performed on the surface of the aircraft engine turbine blades actually produced and delivered to obtain aircraft engine turbine blades without thermal barrier coatings;

[0044] Grooving the blade surface of an aircraft engine turbine blade without a thermal barrier coating to seal the air film holes on the upper and lower edge surfaces;

[0045] A thermocouple is arranged in a slot on the surface of a turbine blade of an aircraft engine without a thermal barrier coating;

[0046] A thermal barrier coating is sprayed on the blade surface of an aircraft engine turbine blade without a thermal barrier coating to obtain an aircraft engine turbine blade cooling effect test piece.

[0047] As for the method for preparing the test piece for the cooling effect test of the aircraft engine turbine blade disclosed in the above-mentioned embodiment, it can be understood by technicians in the field that the method obtains the aircraft engine turbine blade without the thermal barrier coating by sandblasting the surface of the aircraft engine turbine blade actually produced and delivered to remove the thermal barrier coating on the surface of the aircraft engine turbine blade actually produced and delivered, so as to ensure that the aircraft engine turbine blade without the thermal barrier coating is highly consistent in size with the aircraft engine turbine blade actually produced and delivered, thereby reducing the deviation of the aperture size of the film holes on the blade surface of the aircraft engine turbine blade actually produced and delivered due to the shrinkage phenomenon of the film holes when spraying the thermal barrier coating on the blade body surface of the aircraft engine turbine blade without the thermal barrier coating, thereby ensuring that the aperture size of the film holes on the blade body surface of the aircraft engine turbine blade actually produced and delivered is consistent with the aperture size of the film holes on the blade body surface of the aircraft engine turbine blade actually produced and delivered, thereby ensuring the accuracy of the test results.

[0048] Since only the temperature distribution on the surface of the blade body of the aircraft engine turbine blade is concerned in the cooling effect test of the aircraft engine turbine blade, in the method for manufacturing the cooling effect test piece for the aircraft engine turbine blade cooling effect test disclosed in the above-mentioned embodiment, it is designed to seal the air film holes on the surfaces of the upper and lower edge plates of the cooling effect test piece for the aircraft engine turbine blade cooling effect test. When conducting the cooling effect test, the flow ratio between the high-temperature combustion gas introduced into the blade grid channel and the cooling air introduced into the cooling effect test piece for the aircraft engine blade can be converted into the flow ratio between the high-temperature combustion gas introduced into the blade grid channel and the cooling air introduced into the air film holes on the blade body surface of the cooling effect test piece for the aircraft engine blade, and the test can be conducted. In this way, the influence of the air holes on the upper and lower edge plates of the cooling effect test piece for the aircraft engine blade cooling effect test on the cooling air flow distribution and the accuracy of the test results can be eliminated.

[0049] In some optional embodiments, the above-mentioned method for manufacturing a test piece for a cooling effect test of an aircraft engine turbine blade further includes:

[0050] The aperture of the air film hole on the blade surface of the aircraft engine turbine blade without thermal barrier coating can be measured before the blade surface of the aircraft engine turbine blade without thermal barrier coating is grooved, so as to avoid affecting the aperture measurement after the blade surface of the aircraft engine turbine blade is grooved;

[0051] If the aperture of the film holes on the blade surface of an aircraft engine turbine blade without thermal barrier coating does not match the aperture of the film holes on the blade surface of an aircraft engine turbine blade without thermal barrier coating on the process drawing of the aircraft engine turbine blade actually delivered during production, and is too large or too small, the aircraft engine turbine blade cooling effect test specimen shall be remade.

[0052] In some optional embodiments, in the above-mentioned method for preparing a test piece for the cooling effect test of an aircraft engine turbine blade, spraying a thermal barrier coating on the blade surface of an aircraft engine turbine blade without a thermal barrier coating to obtain the test piece for the cooling effect test of an aircraft engine turbine blade comprises:

[0053] On the blade surface of the aircraft engine turbine blade without thermal barrier coating, cover the air film holes and spray the thermal barrier coating bottom layer;

[0054] On the surface of the aero-engine turbine blade with a thermal barrier coating bottom layer, the covering of the air film holes is removed, polishing is performed, and a thermal barrier coating surface layer is sprayed to obtain an aero-engine turbine blade cooling effect test piece.

[0055] As for the method for preparing the test piece for the cooling effect test of aircraft engine turbine blades disclosed in the above-mentioned embodiment, it can be understood by technicians in the field that its design covers the air film holes when spraying the thermal barrier coating base layer on the blade surface of an aircraft engine turbine blade without a thermal barrier coating, so as to avoid the spraying of the thermal barrier coating base layer having a significant impact on the aperture size of the air film holes on the blade surface of the aircraft engine turbine blade, and facilitates the control of the aperture size of the air film holes on the blade surface of the aircraft engine turbine blade when subsequently spraying the thermal barrier coating surface layer, so that the aperture size of the air film holes on the blade surface of the aircraft engine turbine blade cooling effect test piece obtained is consistent with the aperture size of the air film holes on the blade surface of the aircraft engine turbine blade actually produced and delivered, thereby ensuring the accuracy of the test results.

[0056] In some optional embodiments, in the above-mentioned method for preparing test pieces for an aircraft engine turbine blade cooling effect test, the air film holes are covered on the blade surface of the aircraft engine turbine blade without a thermal barrier coating, specifically by covering the air film holes with metal foil.

[0057] In some optional embodiments, in the above-mentioned method for preparing a test piece for a cooling effect test of an aircraft engine turbine blade, the bottom layer of the thermal barrier coating is a NiCrAlY coating;

[0058] The surface layer of the thermal barrier coating is a ZrO3 coating.

[0059] In some optional embodiments, the above-mentioned method for manufacturing a test piece for a cooling effect test of an aircraft engine turbine blade further includes:

[0060] Measuring the aperture of the film holes on the polished surface of the turbine blade of an aircraft engine with a thermal barrier coating bottom layer;

[0061] If the aperture of the air film holes on the blade body surface of the aircraft engine turbine blade with the thermal barrier coating bottom layer is smaller than the aperture of the air film holes on the blade body surface of the aircraft engine turbine blade actually produced and delivered, the aircraft engine turbine blade cooling effect test specimen shall be re-made;

[0062] If the aperture of the air film holes on the blade body surface of an aircraft engine turbine blade with a thermal barrier coating bottom layer is not less than the aperture of the air film holes on the blade body surface of an aircraft engine turbine blade actually produced and delivered, the thermal barrier coating surface layer is sprayed to obtain an aircraft engine turbine blade cooling effect test specimen.

[0063] In some optional embodiments, the above-mentioned method for manufacturing a test piece for a cooling effect test of an aircraft engine turbine blade further includes:

[0064] Measuring the diameter of the air film holes on the blade surface of the aircraft engine turbine blade cooling effect test specimen;

[0065] If the aperture of the air film holes on the blade surface of the aircraft engine turbine blade cooling effect test specimen does not match the aperture of the air film holes on the blade surface of the aircraft engine turbine blade actually produced and delivered, or is too large or too small, the aircraft engine turbine blade cooling effect test specimen shall be remade.

[0066] As for the method for manufacturing the test piece for the cooling effect test of the aircraft engine turbine blade disclosed in the above-mentioned embodiment, it can be understood by technicians in the field that it is designed with the steps of measuring the aperture of the film holes on the blade surface of the aircraft engine turbine blade without thermal barrier coating, measuring the aperture of the film holes on the blade surface of the aircraft engine turbine blade with a thermal barrier coating bottom layer after polishing, and measuring the aperture of the film holes on the blade surface of the test piece for the cooling effect test of the aircraft engine turbine blade, so as to monitor the aperture of the film holes on the blade surface during the whole process of manufacturing the test piece for the cooling effect test of the aircraft engine turbine blade, so as to control the aperture of the film holes on the blade surface of the test piece for the cooling effect test of the aircraft engine turbine blade to be consistent with the aperture of the film holes on the blade surface of the aircraft engine turbine blade actually produced and delivered, so as to ensure the accuracy of the test results.

[0067] The various embodiments in the specification are described in a progressive manner, and each embodiment focuses on the differences from other embodiments. The same or similar parts between the various embodiments can be referenced to each other.

[0068] So far, the technical solution of the present application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the scope of protection of the present application is obviously not limited to these specific embodiments. Without departing from the principles of the present application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will fall within the scope of protection of the present application.

Claims

1. A method for preparing a test piece for an aircraft engine turbine blade cooling effect test, characterized in that: include: Sandblasting is performed on the surface of the aircraft engine turbine blades actually produced and delivered to obtain aircraft engine turbine blades without thermal barrier coatings; Grooving the blade surface of an aircraft engine turbine blade without a thermal barrier coating to seal the air film holes on the upper and lower edge surfaces; A thermocouple is arranged in a slot on the surface of a turbine blade of an aircraft engine without a thermal barrier coating; Spraying a thermal barrier coating on the blade surface of an aircraft engine turbine blade without a thermal barrier coating to obtain an aircraft engine turbine blade cooling effect test piece; Measuring the diameter of the air film holes on the blade surface of aircraft engine turbine blades without thermal barrier coating; If the aperture of the air film holes on the blade body surface of the aircraft engine turbine blade without thermal barrier coating does not match the aperture of the air film holes on the blade body surface of the aircraft engine turbine blade without thermal barrier coating on the process drawing of the aircraft engine turbine blade actually delivered by production, the aircraft engine turbine blade cooling effect test specimen shall be re-made; Measuring the aperture of the film holes on the polished surface of the turbine blade of an aircraft engine with a thermal barrier coating bottom layer; If the aperture of the air film holes on the blade body surface of the aircraft engine turbine blade with the thermal barrier coating bottom layer is smaller than the aperture of the air film holes on the blade body surface of the aircraft engine turbine blade actually produced and delivered, the aircraft engine turbine blade cooling effect test specimen shall be re-made; If the aperture of the air film holes on the blade body surface of the aircraft engine turbine blade with the thermal barrier coating bottom layer is not less than the aperture of the air film holes on the blade body surface of the aircraft engine turbine blade actually produced and delivered, then the thermal barrier coating surface layer is sprayed to obtain an aircraft engine turbine blade cooling effect test specimen; Measuring the diameter of the air film holes on the blade surface of the aircraft engine turbine blade cooling effect test specimen; If the aperture of the air film holes on the blade surface of the aircraft engine turbine blade cooling effect test specimen does not match the aperture of the air film holes on the blade surface of the aircraft engine turbine blade actually produced and delivered, the aircraft engine turbine blade cooling effect test specimen shall be remade.

2. The method for manufacturing a test piece for an aircraft engine turbine blade cooling effect test according to claim 1, characterized in that: The method of spraying a thermal barrier coating on the blade surface of an aero-engine turbine blade without a thermal barrier coating to obtain an aero-engine turbine blade cooling effect test piece comprises: On the blade surface of the aircraft engine turbine blade without thermal barrier coating, cover the air film holes and spray the thermal barrier coating bottom layer; On the surface of the aero-engine turbine blade with a thermal barrier coating bottom layer, the covering of the air film holes is removed, polishing is performed, and a thermal barrier coating surface layer is sprayed to obtain an aero-engine turbine blade cooling effect test piece.

3. The method for manufacturing a test piece for an aircraft engine turbine blade cooling effect test according to claim 2, characterized in that: The air film holes are covered on the blade surface of the aircraft engine turbine blade without thermal barrier coating, specifically by covering the air film holes with metal foil.

4. The method for manufacturing a test piece for an aircraft engine turbine blade cooling effect test according to claim 3, characterized in that: The bottom layer of the thermal barrier coating is a NiCrAlY coating; The surface layer of the thermal barrier coating is a ZrO3 coating.

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