A transpiring cooling system for thermal protection of aircraft engine blades

Abstract

The present invention discloses a transpiration cooling system for thermal protection of aircraft engine blades. A cooling layer and a composite layer are provided on the outside of the inner cavity of the turbine blade. The composite layer includes a porous material layer and a metal plate. Liquid coolant is introduced into the inner cavity of the turbine blade and the first cooling channel of the first porous material. The composite layer uses a porous material layer of gradient porous material and is divided by a metal plate. The metal plate is located on the leading edge of the turbine blade, thereby achieving a reasonable distribution of the coolant, causing more coolant to flow to the stagnation point area, improving the uniformity of the temperature distribution on the surface of the turbine blade, reducing the temperature gradient, effectively solving the problem of uneven cooling, and improving the efficiency of coolant use. At the same time, because liquid coolant has a higher specific heat capacity than gas coolant and due to the presence of latent heat of vaporization, the efficiency of transpiration cooling will be greatly improved.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft, in particular to a transpiring cooling system for thermal protection of aircraft engine blades. Background Art

[0002] Hypersonic vehicles encounter extreme aerodynamic heating during flight, requiring their thermal protection systems to withstand and effectively dissipate this heat to ensure structural integrity and functionality. The aerodynamic heating caused by hypersonic flight can cause changes in the structural temperature field, which in turn alters the stiffness, stress, and modal properties of the structure. These challenges significantly increase the complexity of hypersonic vehicle design.

[0003] The thermal environment in typical hypersonic vehicle components has undergone fundamental changes, with total heating reaching as high as 20,000 MJ. In such harsh thermal environments, efficient thermal protection measures must be implemented to prevent the vehicle from burning and to ensure the safety of the instruments and equipment within the cabin. Methods such as ablation, regenerative cooling, film cooling, and transpiration cooling have all been used in the cooling process of aircraft. After comparison, transpiration cooling technology is found to be the most suitable cooling method for harsh thermal environments, making research and development in transpiration cooling necessary. Transspiration cooling is a thermal protection method proposed in the 1940s specifically for high-temperature walls. However, the complex two-phase flow and heat transfer within porous walls, coupled with the extreme conditions of low pressure and shock waves encountered in spacecraft service, make existing models difficult to accurately predict, and reduce transpiration cooling efficiency. These issues pose significant challenges to the development of phase change transpiration cooling.

[0004] During the manufacturing process of aircraft blades, the performance of the blades will be affected differently due to various issues such as the composition of various alloys and the design and control of the solidification process. Turbine blades are extremely complex in terms of production, manufacturing, working environment, and cooling technology. my country has established some research in the fields of hypersonic propulsion and turbine engines, and has begun to take shape. However, due to the complexity of the blades, the research on its basic theory and technological accumulation still deserves further in-depth exploration.

[0005] Prior art discloses a porous material transpiration cooling front edge structure. The transpiration layer of this transpiration cooling system is distributed with a vast number of microscopic, interconnected microchannels. This allows transpiration cooling to transport coolant in all directions and efficiently exchange heat. This creates a uniform and dense low-temperature air film on the hot end surface, effectively insulating the heat and significantly reducing the drag along the path caused by wall friction. Its protection mechanisms primarily include heat absorption by the porous framework and the coolant's heat capacity, heat absorption by the coolant's phase change, and the blocking effect of the liquid and air films. However, transpiration cooling also has its drawbacks, such as poor reliability. Any fault or unexpected change can immediately cause the entire transpiration system to fail, even leading to catastrophic consequences. Currently, transpiration cooling research primarily focuses on three approaches: theoretical research, experimental verification, and numerical simulation. While the development of fluid mechanics relies primarily on theoretical derivation and experimental verification, and experiments are the foundation of all theoretical disciplines and can overturn theoretical derivations, the increasing depth of research in flow space places increasing demands on experimental equipment. Experiments often fail to yield accurate conclusions, leading to the use of computational software for numerical simulation of fluid flow.

[0006] When numerically simulating the transpiration cooling of a two-dimensional flat plate, it was found that local high temperature and high pressure would occur under non-uniform heat flow boundaries and non-uniform pressure boundaries, and the formation of this phenomenon would inevitably lead to the phenomenon of "heat transfer deterioration". It is particularly noteworthy that in real applications, these two types of non-uniform boundaries often exist at the same time. For example, the end of a hypersonic aircraft and the stagnation position of the wing leading edge are very prone to the phenomenon of "heat transfer deterioration". In order to improve the reliability of transpiration cooling technology and provide a basis for the engineering application of transpiration cooling technology, on the one hand, it can be optimized by controlling the physical properties of the porous medium and the distribution of the coolant. On the other hand, structural optimization can be carried out to achieve improved transpiration cooling efficiency. Summary of the Invention

[0007] The purpose of the present invention is to provide a transpiration cooling system for thermal protection of aircraft engine blades, so as to solve the problems existing in the above-mentioned related technologies, improve the cooling uniformity of the blades, and improve the transpiration cooling efficiency.

[0008] To achieve the above object, the present invention provides the following solutions:

[0009] The present invention provides a transpiration cooling system for thermal protection of aero-engine blades. The turbine blade is a hollow structure, and the inner cavity of the turbine blade is connected to an external coolant. The outer portion of the inner cavity of the turbine blade is provided with:

[0010] a cooling layer having a first cooling channel, the first cooling channel being in communication with an external liquid coolant;

[0011] A composite layer, wherein the composite layer includes a porous material layer and a metal plate, wherein the porous material layer is arranged outside the cooling layer, and the porous material layer is made of a gradient porous material, and the metal plate is arranged on the leading edge side of the turbine blade and penetrates the porous material layer to achieve the segmentation of the porous material layer.

[0012] Preferably, the metal plate is arranged at a stagnation area of ​​a leading edge of the turbine blade.

[0013] Preferably, the number of the metal plates is at least two.

[0014] Preferably, the porosity of the porous material layer on the leading edge side of the turbine blade is smaller than the porosity of the porous material layer on the trailing edge side of the turbine blade.

[0015] Preferably, the porosity of the porous material layer changes gradiently along the longitudinal direction of the turbine blade.

[0016] Preferably, the inner cavity of the turbine blade is provided with a blade skeleton, and the inner cavity of the turbine blade is separated into a second cooling channel.

[0017] Preferably, the number of the blade skeletons is multiple groups, and the blade skeletons are arranged along the direction from the leading edge to the trailing edge of the turbine blade.

[0018] Preferably, spoiler ribs are provided on the inner cavity sidewalls of the turbine blades.

[0019] Compared with the related art, the present invention has achieved the following technical effects: the sweating cooling system for thermal protection of aircraft engine blades of the present invention, the turbine blade is a hollow structure, and the inner cavity of the turbine blade is connected to the external coolant, and a cooling layer and a composite layer are arranged outside the inner cavity of the turbine blade, the cooling layer has a first cooling channel, and the first cooling channel is connected to the external liquid coolant; the composite layer includes a porous material layer and a metal plate, the porous material layer is arranged outside the cooling layer, the porous material layer is made of gradient porous material, the metal plate is arranged on the leading edge side of the turbine blade, and penetrates the porous material layer to realize the division of the porous material layer.

[0020] The present invention provides a transpiration cooling system for thermal protection of aircraft engine blades. Liquid coolant is introduced into the inner cavity of the turbine blade and into the first cooling channel of the first porous material. The composite layer utilizes a porous layer of gradient porous material, separated by a metal plate positioned on the leading edge of the turbine blade. This ensures proper distribution of the coolant, encouraging more coolant to flow to the stagnation point, improving the uniformity of the turbine blade surface temperature distribution, and reducing temperature gradients. This effectively resolves uneven cooling and improves coolant efficiency. Furthermore, because liquid coolant has a higher specific heat capacity than gaseous coolant and the presence of latent heat of vaporization, the efficiency of transpiration cooling is greatly enhanced. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or related technologies, the following briefly introduces the drawings required for use in the embodiments. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0022] Figure 1 A schematic front view of a turbine blade disclosed in an embodiment of the present invention;

[0023] Figure 2 A schematic top view of a turbine blade disclosed in an embodiment of the present invention;

[0024] Figure 3 It is a schematic diagram of an enlarged local structure of a turbine blade disclosed in an embodiment of the present invention.

[0025] In the figure: 100, turbine blade;

[0026] 1. Cooling layer; 2. Composite layer; 3. First cooling channel; 4. Porous material layer; 5. Metal plate; 6. High-temperature mainstream; 7. Blade skeleton; 8. Second cooling channel; 9. Spoiler ribs; 10. Top end cover. DETAILED DESCRIPTION

[0027] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of the present invention.

[0028] The purpose of the present invention is to provide a transpiration cooling system for thermal protection of aircraft engine blades, so as to solve the problems existing in the above-mentioned related technologies, improve the cooling uniformity of the blades, and improve the transpiration cooling efficiency.

[0029] In order to make the above-mentioned objects, features and advantages of the present invention more obvious and easy to understand, the present invention is further described in detail below with reference to the accompanying drawings and specific embodiments. Example 1

[0030] This embodiment provides a transpiration cooling system for thermal protection of aircraft engine blades. The turbine blade 100 is a hollow structure, and the inner cavity of the turbine blade 100 is connected to the external coolant. A cooling layer 1 and a composite layer 2 are provided outside the inner cavity of the turbine blade 100. The cooling layer 1 has a first cooling channel 3, which is connected to the external liquid coolant; the composite layer 2 includes a porous material layer 4 and a metal plate 5. The porous material layer 4 is provided outside the cooling layer 1. The porous material layer 4 is made of a gradient porous material. The metal plate 5 is provided on the leading edge side of the turbine blade 100 and penetrates the porous material layer 4 to achieve segmentation of the porous material layer 4.

[0031] The present invention provides a transpiration cooling system for thermal protection of aircraft engine blades. Liquid coolant is introduced into the inner cavity of the turbine blade 100 and the first cooling channel 3 of the first porous material. A composite layer 2 is formed of a gradient porous material, forming a porous material layer 4, which is divided by a metal plate 5. The metal plate 5 is located on the leading edge of the turbine blade 100. This achieves reasonable distribution of the coolant, encourages more coolant to flow to the stagnation area, improves the uniformity of the surface temperature distribution of the turbine blade 100, reduces the temperature gradient, effectively solves the problem of uneven cooling, and improves the efficiency of coolant use. Furthermore, because liquid coolant has a higher specific heat capacity than gaseous coolant and the presence of latent heat of vaporization, the efficiency of transpiration cooling can be greatly improved.

[0032] Due to the large-scale spread of the porous medium, local high temperature and high pressure conditions may occur under non-uniform heat flow boundary or non-uniform pressure boundary conditions, and subsequently "heat transfer deterioration" may occur. The present invention provides a cooling layer 1 having a first cooling channel 3 at the rear of the leading edge of the turbine blade 100, away from the high-temperature mainstream 6. Liquid coolant is injected into the first cooling channel 3 and sprayed out of the leading edge surface of the turbine blade 100. When the coolant flows through the composite layer 2, it is forced to undergo convective heat transfer, thereby reducing the temperature of the porous material layer 4 at the leading edge of the turbine blade 100. At the same time, the coolant is injected into the high-temperature mainstream 6 through the micropores of the porous material layer 4, forming a thicker air film covering layer at the stagnation point area of ​​the leading edge of the turbine blade 100, isolating the leading edge of the turbine blade 100 from the heat flow. The provision of the porous material layer 4 and the large specific surface area of ​​the porous internal microchannels enable the coolant to obtain more sufficient heat exchange.

[0033] It should be noted here that, in this specific embodiment, the first cooling channel 3 of the cooling layer 1 passes through a liquid coolant, and the cooling efficiency of the liquid coolant is higher than that of the gaseous coolant. When selecting the coolant, the liquid coolant is preferred; the coolant passed into the inner cavity of the turbine blade 100 can be a gaseous or liquid coolant.

[0034] Specifically, metal plate 5 is positioned at the stagnation point region at the leading edge of turbine blade 100. The placement of metal plate 5 within porous material layer 4 divides the porous material layer 4, ensuring uniform coolant distribution, improving surface temperature uniformity, and reducing temperature gradients, thereby minimizing the continuous increase in thermal stress in the wall material and achieving optimal thermal protection. Properly arranging the coolant distribution at the stagnation point at the leading edge of turbine blade 100 ensures that more coolant flows to the stagnation point region, effectively isolating the leading edge of turbine blade 100 from the heat flow.

[0035] In this embodiment, the number of metal plates 5 is at least two. When the porous material layer 4 is divided by the metal plates 5, n metal plates 5 may be added. In this embodiment, n is set to ≥ 2. In practical applications, the area of ​​the porous material layer 4 is not very large. Too few metal plates 5 will result in poor coolant distribution, while too many metal plates 5 will increase processing difficulty and affect the efficiency of sweating and cooling.

[0036] The porous material layer 4 is made of a gradient porous material. The selection of the gradient porous material can be made according to the size of the heat flow. The porosity of the porous material layer 4 on the leading edge side of the turbine blade 100 is smaller than the porosity on the trailing edge side of the turbine blade 100. That is, a material with a smaller porosity can be used at the front end, and a material with a larger porosity can be used in other parts with small heat flow, which effectively improves the efficiency of coolant use. Moreover, since upstream sweating has a certain cooling effect on downstream components, such a setting can also appropriately reduce the pressure of downstream heat protection.

[0037] In addition, the porous material layer 4 can be made of a high-temperature resistant material, such as steel, a high-temperature alloy, or a ceramic material. The metal plates 5 used to separate the porous material layer 4 are solid substances, and can be a metal material with a different density than the porous material. It should also be noted that the cooling layer 1 can be formed into a hollow structure through processing or other methods, or made of a porous material. In actual applications, the cooling layer 1 can be selected from a suitable structure or material based on the requirements of the first cooling channel 3.

[0038] More specifically, along the longitudinal direction of the turbine blade 100, the porosity of the porous material layer 4 changes in a gradient. Figure 1 When the coolant flows from bottom to top, that is, from the root of the turbine blade 100 upward, in order for the coolant to flow smoothly to the upper end, the porous material layer 4 is also designed to be gradient porous in the longitudinal direction to ensure that sweating and cooling are carried out on the turbine blade 100. Whether from the root to the upper end of the turbine blade 100 or from the leading edge to the trailing edge of the turbine blade 100, the porous material layer 4 is designed according to the size of the heat flow and the size of the flow resistance. Any part can achieve a good cooling effect and overheating will not occur in certain places.

[0039] It should also be noted that the inner cavity of the turbine blade 100 is provided with a blade skeleton 7, which separates the inner cavity of the turbine blade 100 into a second cooling channel 8, into which a gas coolant is introduced; when the turbine blade 100 has a top end cap 10, the coolant flows from the root toward the top end cap 10. In other specific embodiments of the present invention, the number of blade skeletons 7 is multiple, and the blade skeletons 7 are arranged from the leading edge to the trailing edge of the turbine blade 100. In addition, the inner cavity sidewalls of the turbine blade 100 are provided with spoiler ribs 9 to further improve cooling efficiency.

[0040] After conducting relevant numerical simulations on existing transpiration cooling, it was found that local high temperature and high pressure will occur under non-uniform heat flow boundaries or non-uniform pressure boundaries, and under uniform heat flow conditions, the coolant will also flow to the porous medium boundary, which will lead to uneven cooling. The composite layer 2 of the present invention uses a metal plate 5 to divide the porous material layer 4 made of gradient porous material, and reasonably distributes the coolant, which can encourage more coolant to flow to the stagnation area, effectively solving the problem of uneven cooling and improving the efficiency of coolant use. Liquid coolant has a higher specific heat capacity than gas coolant, and due to the presence of latent heat of vaporization, it will greatly improve the efficiency of transpiration cooling. Example 2

[0041] This embodiment provides an aircraft engine, including the transpiration cooling system for thermal protection of aircraft engine blades according to the first embodiment. Example 3

[0042] This embodiment provides an aircraft, including the aviation engine of embodiment 2.

[0043] The present invention uses specific examples to illustrate the principles and implementation methods of the present invention. The above examples are only intended to help understand the method and core concept of the present invention. At the same time, those skilled in the art will find that the specific implementation methods and application scopes may vary based on the concept of the present invention. In summary, the contents of this specification should not be construed as limiting the present invention.

Claims

1. A transpiring cooling system for thermal protection of aircraft engine blades, characterized in that: The turbine blade is a hollow structure, and the inner cavity of the turbine blade is connected to the external coolant. The outer portion of the inner cavity of the turbine blade is provided with: a cooling layer having a first cooling channel, the first cooling channel being in communication with an external liquid coolant; A composite layer, wherein the composite layer includes a porous material layer and a metal plate, wherein the porous material layer is arranged outside the cooling layer, and the porous material layer is made of a gradient porous material, and the metal plate is arranged on the leading edge side of the turbine blade and penetrates the porous material layer to achieve the segmentation of the porous material layer.

2. The transpiring cooling system for thermal protection of aircraft engine blades according to claim 1, characterized in that: The metal plate is arranged at a stagnation area of ​​a leading edge of the turbine blade.

3. The transpiring cooling system for thermal protection of aircraft engine blades according to claim 1, characterized in that: The number of the metal plates is at least two.

4. The transpiring cooling system for thermal protection of aircraft engine blades according to claim 1, characterized in that: The porosity of the porous material layer on the leading edge side of the turbine blade is smaller than the porosity of the porous material layer on the trailing edge side of the turbine blade.

5. The transpiring cooling system for thermal protection of aircraft engine blades according to claim 1, characterized in that: The porosity of the porous material layer changes gradiently along the longitudinal direction of the turbine blade.

6. The transpiring cooling system for thermal protection of aircraft engine blades according to any one of claims 1 to 5, characterized in that: The inner cavity of the turbine blade is provided with a blade skeleton, and the inner cavity of the turbine blade is separated into a second cooling channel.

7. The transpiring cooling system for thermal protection of aircraft engine blades according to claim 6, characterized in that: The number of the blade skeletons is multiple, and the blade skeletons are arranged along the direction from the leading edge to the trailing edge of the turbine blade.

8. The transpiring cooling system for thermal protection of aircraft engine blades according to claim 6, characterized in that: The inner cavity side wall of the turbine blade is provided with a spoiler rib.

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