Sweat cooling system for thermal protection of aero-engine blade

Abstract

The turbine blade comprises a cooling layer and a composite layer which are sequentially arranged from inside to outside, and the composite layer comprises a porous material layer and a metal plate. Liquid coolants are introduced into the inner cavity of the turbine blade and the first cooling channel of the first porous material, the porous material layer of the gradient porous material is adopted as the composite layer and is divided by the metal plate, and the metal plate is located on one side of the front edge of the turbine blade, so that reasonable distribution of the coolants is achieved, more coolants are promoted to flow to a stagnation point area, and the cooling efficiency is improved. The surface temperature distribution uniformity of the turbine blade is improved, the temperature gradient is reduced, the problem of non-uniform cooling is effectively solved, and the use efficiency of a coolant is improved. Meanwhile, due to the fact that the specific heat capacity of the liquid coolant is higher than that of the gas coolant, and due to the existence of latent heat of vaporization, the sweating cooling efficiency can be greatly improved.

Description

Technical Field

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

[0002] Hypersonic vehicles will encounter extreme aerodynamic heating problems during flight, which requires the aircraft's thermal protection system to be able to withstand and effectively dissipate this heat to ensure the integrity and functionality of the aircraft structure. The aerodynamic heating caused by hypersonic flight will cause changes in the structural temperature field, thereby changing the stiffness, stress and mode of the structure. These difficulties have greatly increased the difficulty of hypersonic aircraft design.

[0003] Nowadays, the thermal environment of typical parts of hypersonic vehicles has undergone fundamental changes, and the total heating can even reach 20,000 MJ. In such a harsh thermal environment, if the aircraft is not burned and the safety of the instruments and equipment in the cabin is guaranteed, efficient thermal protection measures must be taken. Ablation, regenerative cooling, film cooling and transpiration cooling have all been used in the cooling process of aircraft. After comparison, transpiration cooling technology is the most suitable cooling method for harsh thermal environments, so it is necessary to study and develop transpiration cooling. Transpiration cooling is a thermal protection method specifically proposed for high-temperature walls in the 1940s. However, due to the complex two-phase flow and heat transfer laws in porous walls, and the coupling with the extreme conditions such as low pressure and shock waves in the service of aerospace vehicles, the existing models are difficult to accurately predict, and the transpiration cooling efficiency is reduced. These problems have brought great 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 alloy compositions and solidification process design and control issues. 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] The prior art discloses a porous material sweating cooling front edge structure. The sweating layer of the sweating cooling is distributed with a huge number of micro-channels with small size and extending in all directions. Therefore, the sweating cooling can transport the coolant in all directions and exchange heat efficiently, forming a uniform and dense low-temperature air film on the hot end surface to effectively resist heat, and greatly reduce the resistance along the way caused by wall friction. Its protection principles mainly include: heat absorption by the heat capacity of the porous skeleton and the cooling medium, heat absorption by the phase change of the coolant, and the blocking effect of the liquid and air film. However, sweating cooling also has its disadvantages. For example, the reliability of sweating cooling is poor. Any fault or unexpected change may immediately cause the failure of the entire sweating system, and even bring catastrophic consequences. At present, there are three main ways to study sweating cooling: theoretical research, experimental verification and numerical simulation; although the development of fluid mechanics mainly relies on theoretical derivation and experimental verification, and experiments are the basis for the development of all theoretical disciplines and can overturn theoretical derivation, but with the continuous deepening of the study of flow space, the requirements for experimental instruments are getting higher and higher. In many cases, experiments cannot get correct conclusions, so computational software is used to perform numerical simulation of fluids.

[0006] When numerically simulating the transpiring 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 actual applications, these two types of non-uniform boundaries often exist at the same time. For example, the end of a hypersonic vehicle and the stagnation point position of the wing leading edge are prone to the phenomenon of "heat transfer deterioration". In order to improve the reliability of transpiring cooling technology and provide a basis for the engineering application of transpiring 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, and on the other hand, structural optimization can be carried out to improve the efficiency of transpiring cooling. Summary of the invention

[0007] The purpose of the present invention is to provide a transpiring 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 transpiring cooling efficiency.

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

[0009] The present invention provides a transpiring cooling system for thermal protection of aircraft engine blades. The turbine blade is a hollow structure, and the inner cavity of the turbine blade is connected to an external coolant. The turbine blade includes the following components arranged in sequence from the inside to the outside:

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

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

[0012] Preferably, the metal plate is arranged at a leading edge stagnation area 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 at the leading edge side of the turbine blade is smaller than the porosity of the porous material layer at the trailing edge side of the turbine blade.

[0015] Preferably, the porosity of the porous material layer changes gradually 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 side walls 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 with the external coolant, the turbine blade includes a cooling layer and a composite layer arranged in sequence from the inside to the outside, the cooling layer has a first cooling channel, and the first cooling channel is connected with the external liquid coolant; the composite layer includes a porous material layer and a metal plate, the porous material layer is made of a gradient porous material, 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.

[0020] The present invention is a transpiring cooling system for thermal protection of aircraft engine blades. The inner cavity of the turbine blade and the first cooling channel of the first porous material are both connected to liquid coolant. The composite layer adopts 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, prompting more coolant to flow to the stagnation 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 use efficiency of the coolant. At the same time, since the liquid coolant has a higher specific heat capacity than the gas coolant, and due to the existence of latent heat of vaporization, the efficiency of transpiring cooling will be greatly improved. 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 drawings required for use in the embodiments will be briefly introduced below. 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 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 be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. 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 creative work are within the scope of protection of the present invention.

[0028] The purpose of the present invention is to provide a transpiring 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 transpiring 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.

[0030] Embodiment 1

[0031] The present embodiment provides a transpiring 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 an external coolant. The turbine blade 100 includes a cooling layer 1 and a composite layer 2 arranged in sequence from the inside to the outside. The cooling layer 1 has a first cooling channel 3, and the first cooling channel 3 is connected to an external liquid coolant; the composite layer 2 includes a porous material layer 4 and a metal plate 5. The porous material layer 4 is made of a gradient porous material. The metal plate 5 is arranged 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.

[0032] In the sweating cooling system for thermal protection of aircraft engine blades of the present invention, the inner cavity of the turbine blade 100 and the first cooling channel 3 of the first porous material are both passed with liquid coolant, the composite layer 2 is made of gradient porous material to form a porous material layer 4, and is divided by a metal plate 5, and the metal plate 5 is located on the leading edge side of the turbine blade 100, thereby achieving a reasonable distribution of the coolant, prompting more coolant to flow to the stagnation area, improving the uniformity of the surface temperature distribution of the turbine blade 100, reducing the temperature gradient, effectively solving the problem of uneven cooling, and improving the use efficiency of the coolant. At the same time, because the liquid coolant has a higher specific heat capacity than the gas coolant, and because of the existence of latent heat of vaporization, the efficiency of sweating cooling can be greatly improved.

[0033] Due to the large-area spreading of the porous medium, under the conditions of non-uniform heat flow boundary or non-uniform pressure boundary, local high temperature and high pressure will be generated, and "heat transfer deterioration" will occur successively. The present invention sets a cooling layer 1 with a first cooling channel 3 at the rear of the leading edge of the turbine blade 100 away from the high-temperature mainstream 6. After the liquid coolant is injected into the first cooling channel 3 and sprayed out of the leading edge surface of the turbine blade 100, the coolant is forced to perform convective heat exchange when flowing through the composite layer 2, 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 in the stagnation area of ​​the leading edge of the turbine blade 100, separating the leading edge of the turbine blade 100 from the heat flow. The porous material layer 4 is set, and the porous internal microchannel has a large specific surface area, which can make the coolant get more sufficient heat exchange.

[0034] 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 preferably selected; the coolant passed into the inner cavity of the turbine blade 100 can be a gaseous or liquid coolant.

[0035] Specifically, the metal plate 5 is arranged at the leading edge stagnation point area of ​​the turbine blade 100. The metal plate 5 is arranged at the porous material layer 4 to divide the porous material layer 4, in order to evenly distribute the coolant, improve the uniformity of the surface temperature distribution, reduce the temperature gradient, and continuously increase the thermal stress of the wall material, thereby obtaining an ideal thermal protection effect. Reasonably arrange the distribution amount of the coolant at the leading edge stagnation point of the turbine blade 100 so that more coolant flows to the stagnation point area, thereby effectively isolating the leading edge of the turbine blade 100 from the heat flow.

[0036] In this specific 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 specific embodiment, n is set to be ≥ 2. In practical applications, the area of ​​the porous material layer 4 will not be very large. Too few metal plates 5 added will result in poor coolant distribution. Too many metal plates 5 added will make processing difficult and affect the efficiency of sweating cooling.

[0037] 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, thereby effectively improving the use efficiency of the coolant. Moreover, since the upstream sweating has a certain cooling effect on the downstream components, such a setting can also appropriately reduce the pressure of downstream heat protection.

[0038] In addition, the porous material layer 4 can be made of a high temperature resistant material, such as steel, high temperature alloy or ceramic material, and the metal plate 5 used to separate the porous material layer 4 is a solid material, which can be a metal material with a different density from the porous material. It should also be noted that the cooling layer 1 can be made into a hollow structure by processing, or made of a porous material. In practical applications, the cooling layer 1 can select a suitable structure or material according to the requirements of the first cooling channel 3.

[0039] 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, flows upward from the root of the turbine blade 100, 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 sweating and cooling on the turbine blade 100. Whether from the root of the turbine blade 100 to the upper end, or from the leading edge of the turbine blade 100 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 have a good cooling effect, and overheating will not occur in some places.

[0040] It should also be noted that the inner cavity of the turbine blade 100 is provided with a blade skeleton 7, and the inner cavity of the turbine blade 100 is separated into a second cooling channel 8, and a gas coolant is introduced into the second cooling channel 8; when the turbine blade 100 has a top end cover 10, the coolant flows from the root to the top end cover 10. In other specific embodiments that can be realized by the present invention, the number of blade skeletons 7 is multiple groups, and the blade skeletons 7 are arranged in the direction from the leading edge to the trailing edge of the turbine blade 100. In addition, spoiler ribs 9 are provided on the side wall of the inner cavity of the turbine blade 100 to further improve the cooling efficiency.

[0041] After conducting relevant numerical simulations on existing sweating 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 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 cause more coolant to flow to the stagnation area, effectively solving the problem of uneven cooling and improving the use efficiency of the coolant. Liquid coolant has a higher specific heat capacity than gas coolant, and due to the existence of latent heat of vaporization, it will greatly improve the efficiency of sweating cooling.

[0042] Embodiment 2

[0043] This embodiment provides an aircraft engine, including the transpiring cooling system for thermal protection of aircraft engine blades of the first embodiment.

[0044] Embodiment 3

[0045] This embodiment provides an aircraft, including the aircraft engine of the second embodiment.

[0046] The present invention uses specific examples to illustrate the principles and implementation methods of the present invention. The above examples are only used to help understand the method and core ideas of the present invention. At the same time, for those skilled in the art, according to the ideas of the present invention, there will be changes in the specific implementation methods and application scope. In summary, the content of this specification should not be understood as limiting the present invention.

Claims

1. A sweating 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 turbine blade includes the following components arranged in sequence from the inside to the outside: a cooling layer, the cooling layer having a first cooling channel, the first cooling channel being in communication with an external liquid coolant; The composite layer comprises a porous material layer and a metal plate, wherein the porous material layer is made of a gradient porous material, and the metal plate is arranged on one side of the leading edge of the turbine blade and penetrates the porous material layer to achieve segmentation of the porous material layer.

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

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

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

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

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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