Turbine cooling gas supply system based on lubricating oil precooling and gas turbine engine

Abstract

The invention belongs to the technical field of gas turbine engines, relates to the design technology of turbine cooling systems, and discloses a turbine cooling gas supply system based on lubricating oil precooling and a gas turbine engine, and the turbine cooling gas supply system comprises a heat exchange flow path and a bearing cavity. The heat exchange flow path is located between the combustor inner casing and the bearing cavity casing, the input end of the heat exchange flow path is communicated with the air compressor in an air entraining mode, and the output end is communicated with the turbine disc cavity and the turbine disc. The bearing cavity comprises a gas compressor bearing seat, a turbine bearing seat, a turbine shaft and the bearing cavity casing, and lubricating oil in the bearing cavity cools the gas medium in the heat exchange flow path. According to the turbine cooling air supply system, the turbine disc face can be cooled after the temperature of the gas medium is reduced in advance, on the premise that an additional cooling structure is not added, the gas medium used for cooling the turbine disc is cooled through heat exchange with bearing cavity lubricating oil, the cooling effect of the turbine disc is improved, and the service life of the turbine disc is prolonged. And meanwhile, the air entraining flow is effectively controlled, and the loss of turbine work is reduced.

Description

Technical Field

[0001] The present invention belongs to the technical field of gas turbine engines, and relates to turbine cooling system design technology, and in particular to a turbine cooling air supply system based on lubricating oil precooling and a gas turbine engine. Background Art

[0002] The turbine disk in the engine works in a harsh environment of high temperature and high speed for a long time, and is subjected to extremely high centrifugal loads and thermal loads. It is one of the key components that affect the life of the hot end of the gas turbine engine. The temperature distribution of the turbine disk directly affects its strength life. If it is overheated, the material properties and strength life will be greatly reduced, and even serious failures such as turbine rupture will occur. Therefore, in the development of gas turbine engines, in order to avoid overheating, the turbine disk is usually cooled by bleed air from the compressor outlet. However, the temperature of the bleed air from the compressor outlet is high and it will waste a lot of turbine work, which has an adverse effect on the overall performance of the engine.

[0003] At present, in order to achieve a good cooling effect on the turbine disk and reduce the waste of turbine work, a turbine cooling air supply path with a pre-swirl cooling structure is usually designed in front of the turbine disk. For example, Chinese patent CN110145374A discloses an engine pre-swirl system, which draws air from the inner ring of the combustion chamber, reduces the relative total temperature of the cold air through the pre-swirl cooling structure, and then purges and cools the turbine disk, taking away the heat from the turbine blades in the main channel to the disk body and the wind resistance temperature rise caused by the disk doing work on the cold air, and finally flows into the main channel from the wheel rim stator gap. The pre-swirl cooling structure of its design is complex in design, has high precision requirements, is difficult to process, and will increase the weight and cost of the engine. Summary of the invention

[0004] In order to address the deficiencies of the prior art and achieve the purpose of effectively controlling the bleed air flow and reducing turbine power loss while ensuring the cooling effect of the turbine disk, the present invention discloses a turbine cooling air supply system based on lubricating oil precooling, wherein the turbine cooling air supply system includes a heat exchange flow path and a bearing cavity.

[0005] The heat exchange flow path is located between the combustion chamber casing and the bearing cavity casing, the input end of the heat exchange flow path is connected to the compressor bleed air, and the output end is connected to the turbine disc cavity and the turbine disc. The bearing cavity is located between the rear side of the compressor final disc and the front side of the turbine disc, at a low radius at the same axial position as the combustion chamber casing, and includes a compressor bearing seat, a turbine bearing seat, a turbine shaft and the bearing cavity casing. The lubricating oil in the bearing cavity cools the gas medium in the heat exchange flow path.

[0006] Furthermore, a middle-layer heat exchange casing is provided between the combustion chamber inner casing and the bearing cavity casing, and the middle-layer heat exchange casing divides the heat exchange flow path into a first heat exchange flow path and a second heat exchange flow path;

[0007] The first heat exchange flow path is located between the combustion chamber inner casing and the middle heat exchange casing, the input end of the first heat exchange flow path is connected to the compressor outlet bleed air, and the output end is connected to the turbine disk;

[0008] The second heat exchange flow path is located between the middle layer heat exchange casing and the bearing cavity casing, the input end of the second heat exchange flow path is connected to the compressor intermediate stage bleed air, and the output end is connected to the turbine disc cavity through the third vent hole on the turbine bearing seat.

[0009] Furthermore, the middle layer heat exchange casing is coaxial with the combustion chamber inner casing, the bearing cavity and the bearing cavity casing.

[0010] Furthermore, a plurality of first vent holes and a plurality of second vent holes are circumferentially formed on the compressor bearing seat, the first vent holes are connected to the second heat exchange flow path, and the second vent holes are connected to the first heat exchange flow path;

[0011] A plurality of third air holes and a plurality of fourth air holes are circumferentially formed on the turbine bearing seat. The third air holes are in communication with the second heat exchange flow path, and the fourth air holes are in communication with the first heat exchange flow path.

[0012] Furthermore, the turbine cooling air supply system also includes a turbine guide, which is provided with an air collecting cavity, the input end of the air collecting cavity is connected with the output end of the heat exchange flow path, the output end of the air collecting cavity is connected with the turbine disk via an impact hole on the guide cooling ring, the gas medium output by the heat exchange flow path enters the air collecting cavity, and is discharged through the impact hole to impact-cool the turbine disk.

[0013] Furthermore, there are a plurality of the impact holes, and the plurality of the impact holes are distributed in an array on the guide cooling ring.

[0014] Furthermore, a groove is provided between the guide inner ring of the turbine guide and the guide inner sealing ring, one end of the groove is connected to the heat exchange flow path, and the other end is connected to the air collecting cavity.

[0015] Furthermore, the outer side of the inner ring of the guide device is connected to the flame tube of the combustion chamber, and the inner side of the inner ring of the guide device, the inner sealing ring of the guide device and the cooling ring of the guide device are fixed to the turbine bearing seat by bolts.

[0016] An embodiment of the present invention further provides a gas turbine engine, comprising the above-mentioned turbine cooling air supply system based on lubricating oil precooling.

[0017] Compared with the prior art, at least one of the above-mentioned technical solutions adopted in the embodiments of this specification can achieve beneficial effects that include at least: the turbine cooling air supply system of the present invention, by designing a heat exchange flow path, uses the lubricating oil in the bearing cavity to cool the gas medium in the heat exchange flow path, thereby realizing pre-cooling of the turbine disk cooling air and improving the cooling effect of the turbine disk.

[0018] Furthermore, by designing the heat exchange flow path, two supply air paths with independent flows and coupled heat exchanges are adopted, and the air bleed positions of the two supply air paths are designed, so that the gas medium exchanges heat with the low-temperature lubricating oil in the bearing cavity when flowing through the annular cavity channel on the outer side of the bearing cavity casing, and at the same time, the gas medium in the turbine cooling supply air path is cooled, so that the gas medium temperature is pre-lowered and then the turbine disk surface is cooled through the impact holes, so that the gas medium used for cooling the turbine disk is pre-cooled by heat exchange with the bearing cavity lubricating oil without adding additional cooling structure, thereby improving the cooling effect of the turbine disk, and effectively controlling the bleed air flow rate and reducing the loss of turbine power. BRIEF DESCRIPTION OF THE DRAWINGS

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

[0020] Figure 1 It is a schematic diagram of a turbine cooling air supply system based on lubricating oil precooling disclosed in the present invention;

[0021] Among them, 1. bearing cavity; 2. air collecting cavity; 3. second heat exchange flow path; 4. first heat exchange flow path; 5. compressor bearing seat; 6. turbine bearing seat; 7. first fulcrum; 8. second fulcrum; 9. combustion chamber inner casing; 91. air inlet hole; 10. middle layer heat exchange casing; 11. bearing cavity casing; 12. guide inner ring; 13. guide inner sealing ring; 14. guide cooling ring; 15. turbine disk; 16. turbine shaft; 61. third air vent; 62. fourth air vent; 51. first air vent; 52. second air vent; 141. impact hole. DETAILED DESCRIPTION

[0022] The embodiments of the present application are described in detail below with reference to the accompanying drawings.

[0023] The following describes the implementation methods of the present application through specific specific examples, and those skilled in the art can easily understand other advantages and effects of the present application from the content disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. The present application can also be implemented or applied through other different specific implementation methods, and the details in this specification can also be modified or changed in various ways based on different viewpoints and applications without departing from the spirit of the present application. It should be noted that the following embodiments and the features of the embodiments can be combined with each other in the absence of conflict. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in the field without making creative work belong to the scope of protection of the present application.

[0024] In the description of this embodiment, it should be understood that the terms "center", "longitudinal", "lateral", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside" and "outside" etc., indicating orientations or positional relationships, are based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as a limitation on the invention.

[0025] In addition, the terms "first", "second", "third", etc. are used for descriptive purposes only and should not be understood as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Therefore, features defined as "first", "second", etc. may explicitly or implicitly include one or more features. In the description of the present invention, unless otherwise specified, the meaning of "plurality" is two or more.

[0026] It should also be noted that the illustrations provided in the following embodiments are only schematic illustrations of the basic concept of the present disclosure. The illustrations only show components related to the present disclosure rather than being drawn according to the number, shape and size of components in actual implementation. In actual implementation, the type, quantity and proportion of each component may be changed arbitrarily, and the component layout may also be more complicated.

[0027] The embodiment of the present invention discloses a turbine cooling air supply system based on lubricating oil precooling, see Figure 1As shown, the turbine cooling air supply system includes a heat exchange flow path and a bearing cavity 1. The heat exchange flow path is located between the combustion chamber casing 9 and the bearing cavity casing 11. The input end of the heat exchange flow path is connected to the compressor bleed air, and the output end is connected to the turbine disc cavity and the turbine disc 15. The bearing cavity 1 is located between the rear side of the compressor final disc and the front side of the turbine disc, at a low radius at the same axial position as the combustion chamber casing, and includes a compressor bearing seat 5, a turbine bearing seat 6, a turbine shaft 16 and the bearing cavity casing 11. The lubricating oil that continuously circulates with the lubricating oil box and the radiator in the bearing cavity 1 is mainly used to cool and lubricate the bearing assembly. Through the design of the present invention, it can also cool the gas medium in the heat exchange flow path. The cooled gas medium is output to cool the turbine disc cavity and the turbine disc 15. The turbine cooling air supply system of the present invention can solve the technical problems of complex design of the pre-swirl air supply structure in the current gas turbine engine turbine cooling air supply system, high precision requirements, high processing difficulty, and additional increase in engine weight and cost.

[0028] In the specific implementation, see Figure 1 As shown, the compressor bearing seat 5 is supported by a first fulcrum 7 , and the turbine bearing seat 6 is supported by a second fulcrum 8 .

[0029] Further, see Figure 1 As shown, a middle layer heat exchange casing 10 is provided between the combustion chamber inner casing 9 and the bearing cavity casing 11 , and the middle layer heat exchange casing 10 divides a heat exchange flow path into a first heat exchange flow path 4 and a second heat exchange flow path 3 .

[0030] Among them, the first heat exchange flow path 4 is located between the combustion chamber inner casing 9 and the middle layer heat exchange casing 10, the input end of the first heat exchange flow path 4 is connected to the compressor outlet bleed air, and the output end is connected to the turbine disk 15. The second heat exchange flow path 3 is located between the middle layer heat exchange casing 10 and the bearing cavity casing 11, the input end of the second heat exchange flow path 3 is connected to the compressor intermediate stage bleed air, and the output end is connected to the turbine disk cavity through the third vent hole 61 on the turbine bearing seat 6. The temperature of the gas medium bleed from the compressor outlet is greater than the temperature of the gas medium bleed from the compressor intermediate stage, and the temperature of the lubricating oil in the bearing cavity 1 is lower than the temperature of the gas medium bleed from the compressor intermediate stage, so the gas medium in the second heat exchange flow path 3 can be cooled by the lubricating oil in the bearing cavity 1, and the gas medium in the second heat exchange flow path 3 can cool the gas medium in the first heat exchange flow path 4.

[0031] In specific implementation, the first heat exchange flow path 4 and the second heat exchange flow path 3 flow independently and heat exchange is coupled with each other, which can realize the pre-cooling of turbine cooling air while meeting the different functions of engine pivot sealing, wheel rim sealing, turbine cooling and ventilation, so as to achieve the purpose of fine design of engine air system flow path layout and effective control of bleed air flow. The second heat exchange flow path 3 can draw air from the root of the compressor third-stage stator outlet, exchange heat with the higher temperature air in the first heat exchange flow path 4 to seal the second pivot 8, and finally flow into the main flow channel from the root of the turbine rotor outlet. The first heat exchange flow path 4 draws air from the compressor outlet, and after being cooled by the second heat exchange flow path 3, it flows into the turbine guide inner gas collecting cavity 2 through the groove between the guide inner ring 12 and the guide inner sealing ring 13, and then impacts and cools the high-temperature part of the turbine disk 15 through the impact hole 141 on the guide cooling ring 14, and finally flows into the main flow channel through the gap between the forward rotor of the turbine disk 15.

[0032] Furthermore, the combustion chamber inner casing 9 and the middle layer heat exchange casing 10 rotate 360° around the central axis of the engine, that is, the middle layer heat exchange casing 10 and the combustion chamber inner casing 9 are sleeved on the bearing cavity casing 11, and both are coaxial with the bearing cavity 1 and the bearing cavity casing 11.

[0033] Further, see Figure 1 As shown, the compressor bearing seat 5 is provided with a plurality of first vent holes 51 and a plurality of second vent holes 52 along the circumferential direction, the first vent holes 51 are connected to the second heat exchange flow path 3, and the second vent holes 52 are connected to the first heat exchange flow path 4. A plurality of air inlet holes 91 are provided along the circumferential direction at the front end of the combustion chamber casing 9, and both the air inlet holes 91 and the second vent holes 52 are used to introduce the gas medium at the compressor outlet into the first heat exchange flow path 4.

[0034] The turbine bearing seat 6 is provided with a plurality of third air holes 61 and a plurality of fourth air holes 62 along the circumferential direction. The third air holes 61 are communicated with the second heat exchange passage 3 , and the fourth air holes 62 are communicated with the first heat exchange passage 4 .

[0035] Furthermore, in order to maintain the same impact distance and enhance the impact cooling effect, the overall structural shape of the guide cooling ring 14 is similar to the surface shape of the turbine disk 15, see Figure 1As shown, the turbine cooling air supply system also includes a turbine guide, and the turbine guide is provided with an air collecting cavity 2. The input end of the air collecting cavity 2 is connected to the output end of the heat exchange flow path. The output end of the air collecting cavity 2 is connected to the turbine disk 15 through the impact hole 141 on the guide cooling ring 14. The gas medium output by the heat exchange flow path enters the air collecting cavity 2 and is discharged through the impact hole 141 to impact cool the turbine disk 15. The impact hole 141 is used to accelerate the expansion of the airflow and impact cool the high-temperature parts of the turbine disk surface. In order to increase the circumferential velocity of the impact cooling airflow and enhance the cooling effect, the aperture of the impact hole 141 is ≤0.8mm, and the angle between the axis of the impact hole 141 and the rotation direction of the turbine disk 15 is ≤90°.

[0036] Further, see Figure 1 As shown, there are multiple impact holes 141, and the multiple impact holes 141 are distributed in an array on the guide cooling ring 14.

[0037] Further, see Figure 1 As shown, a groove is provided between the guide inner ring 12 and the guide inner sealing ring 13 of the turbine guide, one end of the groove is connected to the heat exchange flow path, and the other end is connected to the gas collecting cavity 2.

[0038] Further, see Figure 1 As shown, the outer side of the guide inner ring 12 is connected to the combustion chamber flame tube, and the inner side of the guide inner ring 12, the guide inner sealing ring 13 and the guide cooling ring 14 are fixed to the turbine bearing seat 6 by bolts.

[0039] An embodiment of the present invention further provides a gas turbine engine, comprising the above-mentioned turbine cooling air supply system based on lubricating oil precooling.

[0040] The turbine cooling air supply system of the present invention designs a heat exchange flow path and uses the lubricating oil in the bearing cavity to cool the gas medium in the heat exchange flow path, thereby realizing air precooling for cooling the turbine disk and improving the cooling effect of the turbine disk.

[0041] Furthermore, by designing the heat exchange flow path, two supply air paths with independent flows and coupled heat exchanges are adopted, and the air bleed positions of the two supply air paths are designed, so that the gas medium exchanges heat with the low-temperature lubricating oil in the bearing cavity when flowing through the annular cavity channel on the outer side of the bearing cavity casing, and at the same time, the gas medium in the turbine cooling supply air path is cooled, so that the gas medium temperature is pre-lowered and then the turbine disk surface is cooled through the impact holes, so that the gas medium used for cooling the turbine disk is pre-cooled by heat exchange with the bearing cavity lubricating oil without adding additional cooling structure, thereby improving the cooling effect of the turbine disk, and effectively controlling the bleed air flow rate and reducing the loss of turbine power.

[0042] Obviously, those skilled in the art should understand that the above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, the embodiments of the present invention may have various modifications and changes. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

[0043] In addition, it should be understood that although the present specification is described according to implementation modes, not every implementation mode contains only one independent technical solution. This description of the specification is only for the sake of clarity. Those skilled in the art should regard the specification as a whole. The technical solutions in each embodiment may also be appropriately combined to form other implementation modes that can be understood by those skilled in the art.

Claims

1. A turbine cooling air supply system based on lubricating oil precooling, characterized in that: The turbine cooling air supply system comprises: A heat exchange flow path, the heat exchange flow path being located between a combustion chamber casing (9) and a bearing chamber casing (11), the input end of the heat exchange flow path being in communication with the compressor bleed air, and the output end of the heat exchange flow path being in communication with the turbine disc cavity and the turbine disc (15); A bearing cavity (1), the bearing cavity (1) is located between the rear side of the compressor final stage disk and the front side of the turbine disk, and is arranged at a low radius at the same axial position as a casing (9) in a combustion chamber, and comprises a compressor bearing seat (5), a turbine bearing seat (6), a turbine shaft (16) and the bearing cavity casing (11), wherein the lubricating oil in the bearing cavity (1) cools the gas medium in the heat exchange flow path.

2. The turbine cooling air supply system based on lubricating oil precooling according to claim 1, characterized in that: A middle-layer heat exchange casing (10) is provided between the combustion chamber inner casing (9) and the bearing chamber casing (11), and the middle-layer heat exchange casing (10) divides a heat exchange flow path into a first heat exchange flow path (4) and a second heat exchange flow path (3); The first heat exchange flow path (4) is located between the combustion chamber inner casing (9) and the middle layer heat exchange casing (10), the input end of the first heat exchange flow path (4) is connected to the compressor outlet bleed air, and the output end is connected to the turbine disk (15); The second heat exchange flow path (3) is located between the middle layer heat exchange casing (10) and the bearing cavity casing (11); the input end of the second heat exchange flow path (3) is connected to the compressor intermediate stage bleed air, and the output end is connected to the turbine disc cavity through the third vent hole (61) on the turbine bearing seat (6).

3. The turbine cooling air supply system based on lubricating oil precooling according to claim 2, characterized in that: The middle layer heat exchange casing (10) is coaxial with the combustion chamber inner casing (9), the bearing cavity (1) and the bearing cavity casing (11).

4. The turbine cooling air supply system based on lubricating oil precooling according to claim 2, characterized in that: A plurality of first vent holes (51) and a plurality of second vent holes (52) are provided on the compressor bearing seat (5) along a circumferential direction, the first vent holes (51) being in communication with the second heat exchange flow path (3), and the second vent holes (52) being in communication with the first heat exchange flow path (4); A plurality of third ventilation holes (61) and a plurality of fourth ventilation holes (62) are provided on the turbine bearing seat (6) along the circumferential direction; the third ventilation holes (61) are connected to the second heat exchange flow path (3); and the fourth ventilation holes (62) are connected to the first heat exchange flow path (4).

5. The turbine cooling air supply system based on lubricating oil precooling according to claim 1 or 2, characterized in that: The turbine cooling air supply system also includes a turbine guide, the turbine guide is provided with an air collecting cavity (2), the input end of the air collecting cavity (2) is connected to the output end of the heat exchange flow path, the output end of the air collecting cavity (2) is connected to the turbine disk (15) via an impact hole (141) on a guide cooling ring (14), the gas medium output from the heat exchange flow path enters the air collecting cavity (2), and is discharged through the impact hole (141) to impact-cool the turbine disk (15).

6. The turbine cooling air supply system based on lubricating oil precooling according to claim 5, characterized in that: There are a plurality of impact holes (141), and the plurality of impact holes (141) are distributed in an array on the guide cooling ring (14).

7. The turbine cooling air supply system based on lubricating oil precooling according to claim 5, characterized in that: A groove is provided between the guide inner ring (12) of the turbine guide and the guide inner sealing ring (13), one end of the groove is connected to the heat exchange flow path, and the other end is connected to the air collecting chamber (2).

8. The turbine cooling air supply system based on lubricating oil precooling according to claim 7, characterized in that: The outer side of the guide inner ring (12) is connected to the combustion chamber flame tube, and the inner side of the guide inner ring (12), the guide inner sealing ring (13) and the guide cooling ring (14) are fixed to the turbine bearing seat (6) via bolts.

9. A gas turbine engine, characterized in that: It comprises a turbine cooling air supply system based on lubricating oil precooling as claimed in any one of claims 1 to 8.

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