Turbine rotor blade
By designing a double-walled cooling structure in the high-pressure turbine blades and utilizing a combination of impact channels and convection channels for cooling, the problems of insufficient reliability and uneven temperature of the blades under high-intensity mechanical environments were solved, achieving a highly efficient and uniform cooling effect.
Patent Information
- Application Number
- CN202520209465.1
- Authority / Receiving Office
- CN · China
- Patent Type
- Utility models(China)
- Current Assignee / Owner
- Filing Date
- 2025-02-10
- Publication Date
- 2026-01-16
- Estimated Expiration
- 2035-02-10
AI Technical Summary
The existing double-wall structure of high-pressure turbine blades is not reliable enough under high-intensity mechanical conditions, and the uneven distribution of gas temperature leads to local overheating of the blades, affecting structural strength and cooling efficiency.
Design a turbine rotor blade with a double-wall cooling structure, including an outer wall, an inner wall and a baffle, forming an impact channel and a convection channel. The distribution and heat exchange of cooling gas are carried out by using impact holes and film vents to optimize the temperature field.
It improves the structural reliability and cooling efficiency of the blades, uniformly cools the blade surface, reduces stress concentration, and ensures stable operation in high-temperature environments.
Smart Images
Figure CN223806188U_ABST
Abstract
Description
TECHNICAL FIELD
[0001] The utility model relates to the technical field of aero-engine, especially a turbine rotor blade. BACKGROUND
[0002] In the development process of the aviation technology, the technology level of the aero-engine and each component is on the way of continuous progress, and the corresponding technical index is continuously improved. It is worth noting that the increase of the combustion chamber outlet temperature brings an unprecedented thermal load challenge to the high-pressure turbine component in the downstream position. Under this background, the cooling design of the turbine blade naturally becomes the focus of researchers, because it is directly related to whether the high-pressure turbine component can stably and efficiently operate in a high-temperature environment.
[0003] At present, the cooling technology used by the turbine blade presents a diversified situation, and methods such as impingement cooling, convection heat transfer and film cooling show their unique skills. In order to achieve a qualitative leap in cooling efficiency, researchers often ingeniously integrate these different cooling forms for use, and then creatively propose a double-wall cooling structure with a composite cooling mode. Generally speaking, this double-wall cooling structure is composed of an inner wall and an outer wall. In actual operation, the cold gas first enters the channel space between the inner wall and the outer wall through the impingement holes on the inner wall, and first cools the outer wall; then, in this relatively closed double-wall channel, the cooling process continues to proceed with the help of the convection heat transfer mechanism, so that the cold gas fully absorbs heat, and finally, the cooled gas flows out through the film holes distributed on the outer wall.
[0004] However, focusing on the high-pressure turbine moving blade, many difficult problems need to be solved. On the one hand, the moving blade is in a high-speed rotating state, and the huge centrifugal force generated thereby makes the blade root bear an extraordinary load, and the impingement holes and the spoiler column structure in the traditional double-wall structure become "potential risk points" that affect the reliability of the blade in this high-strength mechanical environment. They may cause stress concentration and other problems, thereby weakening the overall structural strength of the blade. On the other hand, the high-temperature gas generated from the combustion chamber is not uniformly distributed, and if viewed from the radial distribution of the blade, it can be found that the middle region of the blade seems to be in a high-temperature zone, and the temperature of the blade root and the blade tip region is relatively moderate.
[0005] Considering the above-mentioned complex and key factors, it is not difficult to see that they have a decisive influence on the cooling design of the high-pressure turbine moving blade. In view of this, it has become an urgent task in the field of aero-engine technology research and development to closely focus on the actual engineering needs of the high-pressure turbine blade, explore new double-wall cooling structures in all directions and at a deep level, and apply them to design and application practice. UTILITY MODEL CONTENT
[0006] In view of the above problems of the prior art, the utility model provides a turbine rotor blade suitable for aero-engine high pressure turbine, improves the reliability of double wall structure, guarantees sufficient heat exchange and optimizes temperature field.
[0007] Specifically, the utility model provides a turbine rotor blade suitable for aero-engine high pressure turbine, the turbine rotor blade has double wall cooling structure inside, including the outer wall and the inner wall of setting up along the turbine rotor blade radial, set up the baffle between the outer wall and the inner wall, open the film hole on the outer wall,
[0008] The outer wall, the inner wall, the baffle and the tip of the turbine rotor blade form the impact channel, the inner wall of forming the impact channel is opened with the impact hole, and the inner chamber of the turbine rotor blade is communicated with the impact channel through the impact hole.
[0009] The outer wall, the inner wall, the baffle and the root of the turbine rotor blade form the convection channel, and the inner chamber is communicated with the bottom of the convection channel.
[0010] According to an embodiment of the utility model, the impact channel includes a plurality of impact flow channels arranged radially along the turbine rotor blade.
[0011] According to an embodiment of the utility model, the chord-wise width of the impact flow channel along the turbine rotor blade is 3-8mm.
[0012] According to an embodiment of the utility model, the aperture of the impact hole is 0.8-1.5mm, and the ratio of the impact distance of the double wall cooling structure to the aperture of the impact hole is 0.5:1-1:1.
[0013] According to an embodiment of the utility model, a plurality of impact holes corresponding to each impact flow channel are arranged radially along the turbine rotor blade.
[0014] According to an embodiment of the utility model, a plurality of impact holes corresponding to each impact flow channel are arranged radially along the turbine rotor blade.
[0015] According to an embodiment of the utility model, the convection channel includes a plurality of convection flow channels arranged radially along the turbine rotor blade, and the top or bottom of adjacent convection flow channels is communicated.
[0016] According to an embodiment of the utility model, the bottom of the first convection flow channel along the chord direction of the turbine rotor blade is communicated with the inner chamber.
[0017] According to one embodiment of the present application, a flow structure is formed on the outer wall of the convection passage towards the inner wall.
[0018] According to one embodiment of the present application, the flow structure is a rib, a boss or a pit.
[0019] According to one embodiment of the present application, the flow structure is a rib or a boss, and the ratio of the protrusion height of the rib or the boss to the height of the convection passage is 0.2:1-0.5:1.
[0020] The turbine rotor blade provided by the present application is divided into an independent double-wall impingement passage and a convection passage along the radial direction of the blade, impingement cooling is used in the impingement passage to strengthen internal heat exchange in the corresponding height region of the blade, the cooling demand of the middle region of the blade is met, convection heat exchange is implemented in the convection passage, internal heat exchange is strengthened, cold air distribution is optimized, and the cooling demand of the root region of the blade is met.
[0021] It should be understood that the above general description and the following detailed description of the present application are exemplary and illustrative, and are intended to provide further explanation of the present application. BRIEF DESCRIPTION OF DRAWINGS
[0022] The accompanying drawings are included to provide a further understanding of the present application, and they are collected and constitute a part of the present application, the drawings show the embodiments of the present application, and together with the present application, they play a role of explaining the principles of the present application. In the drawings:
[0023] Figure 1 Fig. 1 shows a structure schematic view of a turbine rotor blade according to one embodiment of the present application.
[0024] Figure 2 is Figure 1 the AA sectional view.
[0025] Figure 3 is Figure 1 the three-dimensional structure schematic view.
[0026] Figure 4 is Figure 1 the local schematic view of the impingement flow channel and the convection flow channel in
[0027] Figure 5 is Figure 4 the AA sectional view.
[0028] Figure 6 Fig. 6 shows a structure schematic view of the impingement flow channel and the convection flow channel of the turbine rotor blade according to another embodiment of the present application.
[0029] Figure 7 is Figure 6AA cross-sectional view.
[0030] Figure 8 The structure diagram of the impingement passage and the convection passage of the turbine rotor blade of another embodiment of the utility model is shown.
[0031] Figure 9A The structure diagram of the turbine rotor blade of another embodiment of the utility model is shown.
[0032] Figure 9B The structure diagram of the turbine rotor blade of another embodiment of the utility model is shown.
[0033] Among them, the above-mentioned drawing includes the following figure marks:
[0034] Turbine rotor blade 100
[0035] Tenon 101
[0036] Leading edge 102
[0037] Trailing edge 103
[0038] Pressure surface 104
[0039] Suction surface 105
[0040] Outer wall 106
[0041] Inner wall 107
[0042] Partition 108
[0043] Impingement passage 109
[0044] Impingement hole 110
[0045] Inner chamber 111
[0046] Film hole 112
[0047] Convection passage 113
[0048] Impingement passage 114
[0049] Side wall 115
[0050] Convection passage 116
[0051] Rotary passage 117
[0052] Flow structure 118
[0053] Inlet 119 DETAILED DESCRIPTION
[0054] It should be noted that the embodiments and the features in the embodiments in the present application can be combined with each other in the case of no conflict.
[0055] The technical solutions in the embodiments of the present application will be described clearly and completely in the present application with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only a part of the embodiments of the present application, not all the embodiments. The following description of at least one exemplary embodiment is actually only illustrative, but not as any limitation on the present application and its application or use. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative labor are within the scope of protection of the present application.
[0056] It should be noted that the terms used herein are only intended to describe specific embodiments, and are not intended to limit the exemplary embodiments according to the present application. As used herein, the singular form is intended to include the plural form unless the context clearly indicates otherwise, and it should also be understood that when the terms "comprise" and / or "include" are used in the specification, there is a feature, step, operation, device, component and / or combination thereof.
[0057] Unless specifically stated otherwise, the relative arrangement of components and steps, numerical expressions, and numerical values set forth in the various embodiments described herein are not meant to limit the scope of the present application. At the same time, it should be understood that the size of each part shown in the drawings is not drawn in accordance with the actual proportion relationship. The technology, method and equipment known to those skilled in the relevant art can not be discussed in detail, but in appropriate cases, the technology, method and equipment should be considered as part of the authorized description. In all examples shown and discussed herein, any specific value should be interpreted as merely exemplary, not as a limitation. Therefore, other examples of exemplary embodiments can have different values. It should be noted that similar reference numbers and letters represent similar items in the following drawings, so once an item is defined in one drawing, it does not need to be further discussed in subsequent drawings.
[0058] In the description of the present application, it should be understood that the orientation words such as "front, back, up, down, left, right", "horizontal, vertical, perpendicular, horizontal" and "top, bottom" and the like indicate the orientation or position relationship based on the orientation or position relationship shown in the drawings, only for the convenience of describing the present application and simplifying the description, without the opposite indication, these orientation words do not indicate and imply that the indicated device or element must have a specific orientation or be constructed and operated in a specific orientation, therefore, it cannot be understood as a limitation on the scope of protection of the present application; the orientation words "inner, outer" refer to the inner and outer relative to the contour of each part itself.
[0059] For purposes of the description hereinafter, spatially relative terms, such as "above", "below", "up", "down", "right", "left", "vertical", "horizontal", "top", "bottom", "lateral", "longitudinal", "front", "rear", "anterior", "posterior", and the like, can be used herein for ease of description to describe one element or feature's relationship to another element(s) or feature(s) as illustrated in the figures. It will be understood that the spatially relative terms are intended to encompass different orientations of the device in use or operation in addition to the orientations depicted in the figures. For example, if a device in the figures is inverted, elements described as "above" or "up" other elements or features would then be oriented "below" or "down" the other elements or features. Thus, the exemplary term "above" can encompass both an orientation of above and below. The device can be otherwise oriented (rotated 90 degrees or at other orientations) and the spatially relative descriptors used herein interpreted accordingly. The terms "first", "second", "third", etc. can be used herein to describe various elements, components, regions and sections but are not intended to be taken literally, unless otherwise indicated. Thus, a first element could be termed a second element or a third element without departing from the teachings of the present application. The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of this application. It is to be understood that the use of certain specific language herein is merely intended to convey the precise information intended and that the use of such specific language is not intended to limit the scope of the present application in any way.
[0060] In addition, it should be pointed out that the use of the terms "first", "second" and the like, to describe various elements, components, regions and sections is merely intended to distinguish a certain element, component, region or section from another element, component, region or section, and does not have any special meaning unless otherwise stated. In addition, although the terms used in the present application are selected from well-known and commonly-used terms, some terms mentioned in the present application may be selected by the applicant according to his or her judgment, and the detailed meanings of the terms are described in the relevant parts of the description. In addition, the present application is required to be understood not only by the actual terms used, but also by the meanings implied by each term.
[0061] Figure 1 A structural schematic diagram of a turbine rotor blade in an embodiment of the present application is shown. Figure 2 is Figure 1 an AA cross-sectional view. Figure 3 is Figure 1 a three-dimensional structural schematic diagram. Figure 4 is Figure 1 a partial schematic diagram of the impingement flow channel and the counter-flow channel in Figure 5 is Figure 4 an AA cross-sectional view. As shown in the figure, the present application provides a turbine rotor blade 100 suitable for a high-pressure turbine of an aero-engine. A tenon 101 is located at the root of the turbine rotor blade 100. The direction of the blade root-blade tip of the turbine rotor blade 100 is defined as the blade radial direction, and the direction of the leading edge 102-tail edge 103 of the turbine rotor blade 100 is defined as the blade chord direction.
[0062] The turbine rotor blade 100 has a double-wall cooling structure inside. The double-wall cooling structure can be located on the pressure surface 104 or the suction surface 105 of the turbine rotor blade 100. In the present embodiment, reference is made to Figure 1 and Figure 2The double-wall cooling structure is located on a pressure surface 104 of the turbine rotor blade 100.
[0063] The double-wall cooling structure includes an outer wall 106 and an inner wall 107 arranged radially along the turbine rotor blade 100, and a partition plate 108 arranged between the outer wall 106 and the inner wall 107. Referring to Figure 4 and Figure 5 The outer wall 106, the inner wall 107, the partition plate 108 and a tip of the turbine rotor blade 100 jointly form an impingement passage 109. An impingement hole 110 is formed in the inner wall 107 of the impingement passage 109, and an inner cavity 111 of the turbine rotor blade 100 is in communication with the impingement passage 109 through the impingement hole 110. An air film hole 112 is formed in the outer wall 106. When the aero-engine high-pressure turbine is working, referring to Figure 5 Cooling gas enters the inner cavity 111 of the turbine rotor blade 100 from the tenon 101 along the left upward arrow, enters the impingement passage 109 along the horizontal direction arrow through the impingement hole 110, and the cooling gas is cooled by high-speed flushing of the outer wall surface, and then flows out of the impingement passage 109 to the right side through the air film hole 112. The cooling gas flowing out of the air film hole 112 forms a continuous air film on the outer wall surface of the impingement passage 109 of the turbine rotor blade 100. The air film is similar to a heat shield, which can isolate the outer wall of the blade from the high-temperature gas, greatly reduce the heat transfer of the high-temperature gas to the blade, reduce the surface temperature of the blade, prevent the blade material from being damaged due to high temperature, and thus improve the working reliability and durability of the blade in high-temperature environment.
[0064] Further, the outer wall 106, the inner wall 107, the partition plate 108 and the root of the turbine rotor blade 100 jointly form a convection passage 113, and the inner cavity 111 is in communication with the bottom of the convection passage 113. Cooling gas enters the inner cavity 111 of the turbine rotor blade 100 from the tenon 101, enters the convection passage 113 along the right upward arrow through the bottom of the convection passage 113 to perform convection heat exchange on the outer wall 106. The cooling gas flows out of the convection passage 113 to the right side through the air film hole 112 on the outer wall 106. The cooling gas flowing out of the air film hole 112 forms an air film on the outer wall surface of the convection passage 113 of the turbine rotor blade 100. That is, an air film is formed on the entire outer wall 106 of the turbine rotor blade 100, which isolates the outer wall of the blade from the high-temperature gas.
[0065] The utility model provides a kind of turbine rotor blade 100, two independent passages are separated by baffle 108 into double-wall cooling structure, namely impact passage 109 located above baffle 108 and convection passage 113 located below baffle 108.Cooling gas is high-speed scouring blade high-temperature zone outer wall surface by impact passage 109, cooling gas carries out convection heat exchange with the outer wall surface of blade non-high-temperature zone by convection passage 113.The turbine rotor blade 100 is adjusted by impact passage 109 and convection passage 113 in double-wall cooling structure in the coverage area in blade radial direction, and then the cold gas amount distribution of blade is adjusted and temperature field is optimized.In addition, since the double-wall cooling structure is only provided with impact hole 110 on inner layer wall 107 side of impact passage 109, it can reduce the problem of stress concentration, and has higher reliability.
[0066] Figure 6 The structure diagram of impact flow channel and convection flow channel of turbine rotor blade of another embodiment of the utility model is shown. Figure 7 Figure 6 is the AA direction sectional view. Figure 6 can be regarded as Figure 2 BB direction sectional view of the utility model. As shown in the figure, in some embodiments, impact passage 109 includes multiple impact flow channels 114 arranged along turbine rotor blade 100 radial direction. Figure 2 As shown in the figure, multiple side walls 115 are arranged in inner chamber 111 of turbine rotor blade 100.Combined with the fact that side wall 115 length direction is substantially perpendicular to inner layer wall 107, and passes through inner layer wall 107 until outer layer wall 106, so as to divide impact passage 109 into multiple impact flow channels 114 which are parallel to each other.Meanwhile, multiple side walls 115 extend inward, and divide inner chamber 111 into multiple chambers, wherein part of chambers correspond to impact flow channel 114 one by one.Multiple impact flow channels 114 operate in parallel, which greatly improves cooling efficiency.When aero-engine high-pressure turbine starts to work, cooling gas flows into inner chamber 111, and then is orderly distributed to corresponding impact flow channel 114 of each chamber, can carry out high-intensity impact cooling to blade high-temperature zone outer wall surface simultaneously, can take away a large amount of heat in a short time, and ensure that blade still maintains good performance in high-temperature environment.
[0067] In some examples, the impingement channel 114 has a chord-wise width of 3-8 mm. This width range can ensure that the cooling gas forms a suitable flow rate and pressure within the impingement channel 114. If the width is too narrow, the cooling gas flow is blocked, the flow rate is too fast, and a large air flow resistance is generated, increasing the energy consumption of the engine operation. If the width is too large, the gas flow rate is too slow, and it is difficult to form sufficient impact force to efficiently cool the outer wall of the blade, and the heat cannot be taken away in time, which also threatens the high-temperature working performance of the blade. When the chord-wise width is between 3-8 mm, the cooling gas can impact the outer wall at a proper flow rate to achieve rapid cooling, and also ensure sufficient residence time to fully absorb heat and improve cooling efficiency. In addition, a suitable chord-wise width helps to maintain the structural strength of the impingement channel 114 and ensure good mechanical properties of the blade.
[0068] In some examples, the impingement hole 110 has a diameter of 0.8-1.5 mm, and the ratio of the impingement distance of the double-wall cooling structure to the diameter of the impingement hole 110 is 0.5:1-1:1. The diameter of the impingement hole 110 is directly related to the flow rate and flow rate of the cooling gas. When the diameter is less than 0.8 mm, the gas throughput is limited, although the flow rate may be increased, but the overall cooling gas supply is insufficient, and it is difficult to fully impact and cool the outer wall 106 of the blade, causing heat accumulation in the high-temperature area of the blade, and the blade cannot be effectively cooled. Conversely, if the diameter is greater than 1.5 mm, the flow rate is greatly reduced, and the impact force of the gas after entering the impingement channel 109 is insufficient, and it is difficult to efficiently break the thermal boundary layer, the heat exchange efficiency is low, and the same cannot achieve good cooling effect. The diameter of the impingement hole 110 is within the range of 0.8-1.5 mm, which can ensure a suitable gas flow rate and flow rate, so that the cooling gas can impact the outer wall 106 with sufficient force and quickly remove heat. It should be noted that the impingement distance refers to the distance from the impingement hole 110 to the point of impact on the outer wall 106 (i.e., the height of the impingement channel 109 from the inner wall 107 to the outer wall 106), and when the ratio is less than 0.5:1, it indicates that the impingement distance is too short, and the cooling gas has not fully diffused before hitting the outer wall 106, resulting in insufficient cooling. If the ratio is greater than 1:1, the energy loss of the gas is too much on the way, and the impact force is weak when it reaches the outer wall 106, and it is also difficult to achieve efficient cooling. Therefore, by controlling the ratio of the impingement distance to the diameter of the impingement hole 110 within the range of 0.5:1-1:1, the cooling gas can be ensured to impact the outer wall 106 with optimal kinetic energy at a proper distance after being sprayed from the impingement hole 110, to achieve efficient heat exchange in the impingement channel 109 and improve cooling efficiency. In addition, reasonable diameter and impingement distance design can effectively avoid stress concentration and reduce local stress abnormalities caused by unreasonable layout, and ensure the structural stability of the blade under complex working conditions.
[0069] In some examples, the plurality of impingement holes 110 corresponding to each impingement passage 114 are arranged along a radial direction of the turbine rotor blade 100. Figure 6 In the embodiment shown in FIG. 1, the plurality of impingement holes 110 are arranged at equal intervals along the radial direction of the turbine rotor blade 100. This arrangement allows the cooling gas to be injected into the impingement passages 114 uniformly and stably. During the operation of the high-pressure turbine of the aero-engine, when the cooling gas flows from the inner chamber 111 into the impingement passages 114 through the impingement holes 110, the flow of the cooling gas at different positions is relatively balanced due to the equal interval arrangement of the impingement holes 110, and there is no local overabundance or shortage of the cooling gas. This ensures that the cooling effect of the impingement passages 114 on the outer wall 106 of the blade is uniformly distributed in the radial direction of the blade, effectively avoiding the problem of thermal stress concentration caused by local uneven cooling of the blade, and thus improving the reliability and durability of the blade. Figure 9A As shown in FIG. 2, it is easy to understand that, considering that the heating conditions of the blade in the radial direction are not completely uniform, the impingement holes 110 arranged at increasing intervals from the blade root to the blade tip can reasonably distribute the cooling gas according to the change in the thermal load in the radial direction of the blade. In the part close to the blade tip, the interval of the impingement holes 110 gradually increases, which means that the number of the impingement holes 110 per unit length decreases, and the flow of the cooling gas decreases accordingly. The appropriate reduction in the flow of the cooling gas avoids excessive cooling and energy waste. In the area of the partition plate 108 close to the middle part of the blade, the thermal load is relatively large, and the interval of the impingement holes 110 gradually decreases, thereby strengthening the cooling effect on the high-temperature area of the blade tip. This flexible allocation of cooling resources according to the actual thermal load of the blade makes the blade cooling more accurate and efficient, optimizes the energy utilization efficiency, and provides strong support for the efficient operation of the aero-engine. As an example but not limitation, the plurality of impingement holes 110 in the impingement passage 114 can also be arranged in other forms. Figure 9B As shown in FIG. 3, each impingement passage 114 has two rows of impingement holes 110 arranged along the radial direction of the blade, and the two rows of impingement holes 110 are staggered along the chord direction of the blade. When the cooling gas flows into the impingement passages 114 from the inner chamber 111, the two rows of staggered impingement holes 110 can guide the cooling gas to impact the outer wall surface of the blade in a more dispersed and diversified path, avoiding the concentration of the cooling gas in a small area. On the other hand, the staggered arrangement of the impingement holes 110 makes the flow of the cooling gas in the impingement passages 114 more complex and variable, prolongs the contact time of the cooling gas with the outer wall surface of the blade, and promotes more sufficient and efficient heat exchange, ensuring that the blade can still maintain good performance and stable operation in a high-temperature working environment.
[0070] In some examples, with reference to Figure 6 The convection passage 113 includes a plurality of convection flow channels 116 arranged radially along the turbine rotor blade 100, and the top or bottom of adjacent convection flow channels 116 are communicated. Preferably, the bottom of the first convection flow channel 116 along the chord direction of the turbine rotor blade 100 is communicated with the inner chamber 111. The plurality of convection flow channels 116 are arranged in order radially along the turbine rotor blade 100, and the communication between adjacent convection flow channels 116, or at the top or at the bottom, is achieved through the rotation passage 117, which builds an efficient convection heat exchange structure. On the one hand, when the cooling gas enters the convection flow channel 116, the cooling gas can flow back along the radial direction of the blade through the rotation passage 117, so that the gas can flow smoothly between adjacent flow channels, and the heat can be transferred and exchanged in a wider area, avoiding local heat accumulation, and strengthening the overall convection heat exchange effect. The bottom of the first convection flow channel 116 along the chord direction of the turbine rotor blade 100 is communicated with the inner chamber 111. In this way, the distribution of cooling gas to the impingement passage 109 and the convection passage 113 can be achieved. During the operation of the high-pressure turbine of the aero-engine, the cooling gas first enters the inner chamber 111, and then part of the cooling gas is injected from the bottom of the first convection flow channel 116. This not only realizes the effective distribution of cooling gas and reduces energy loss, but also ensures that the convection flow channel 116 can obtain sufficient cooling source in time, and ensures that the convection heat exchange process of the outer wall surface of the non-high-temperature area of the blade is stable and efficient.
[0071] In some examples, with reference to Figure 4 and Figure 5 The flow-around structure 118 is formed on the outer wall 106 of the convection passage 113 and faces the inner wall 107. Preferably, the flow-around structure 118 is a rib, a boss or a pit. When the flow-around structure 118 is a rib or a boss, the rib or the boss can guide the airflow to change the flow direction and produce a disturbance effect. The originally relatively smooth cooling airflow will flow around the profile of the rib or the boss after encountering the rib or the boss, increasing the contact between the cooling gas and the outer wall 106, making the heat exchange more sufficient, and thereby optimizing the heat exchange efficiency of the entire convection passage 113. When the flow-around structure 118 is a pit, the cooling gas will form a vortex flow inside the pit after entering the pit, so that the heat exchange is more thorough.
[0072] In some examples, the flow-around structure 118 is a rib or a boss, and a ratio of a protruding height of the rib or boss to a height of the convection passage 113 is 0.2:1 to 0.5:1. The height ratio range ensures that the flow-around structure 118 can function properly. When the ratio of the protruding height to the height of the convection passage 113 is less than 0.2:1, the rib or boss is too low, and the flow disturbance effect on the cooling gas is weak, which cannot effectively change the direction of the gas flow, increase the contact between the gas and the outer wall 106, and lower the heat exchange efficiency, so that the heat of the outer wall surface of the non-high-temperature area of the blade cannot be sufficiently removed, and the local temperature of the blade is increased, which affects the high-temperature working performance of the blade. Conversely, if the height ratio is greater than 0.5:1, the rib or boss is too high, which excessively hinders the smooth flow of the cooling gas and increases the flow resistance, which is also not conducive to the stable operation of the blade. When the height ratio is in the range of 0.2:1 to 0.5:1, the flow-around structure 118 can not only moderately guide the flow of the gas to generate flow disturbance and strengthen heat exchange, but also ensure that the cooling gas flows relatively smoothly to achieve efficient cooling.
[0073] In some examples, the flow-around structure 118 is a pit, and a ratio of a recessed height of the pit to a thickness of the outer wall 106 is 0.2:1 to 0.5:1. It is easy to understand that when the ratio is less than 0.2:1, the pit is too shallow, the vortex flow formed after the cooling gas enters the pit is insufficient in intensity, the heat exchange cannot reach an ideal degree, and the heat of the outer wall surface of the non-high-temperature area of the blade cannot be efficiently removed, which is easy to cause the local temperature of the blade to be increased, and adversely affect the working performance of the blade in a high-temperature environment. Conversely, if the ratio is greater than 0.5:1, the pit is too deep, which can damage the structural strength of the outer wall 106 and affect the mechanical stability of the blade as a whole. When the ratio of the recessed height of the pit to the thickness of the outer wall 106 is controlled in the range of 0.2:1 to 0.5:1, the pit can function properly, which not only sufficiently prolongs the contact between the heat and the cooling gas to achieve complete heat exchange, but also ensures that the outer wall 106 has sufficient structural strength to ensure the stable and reliable operation of the blade.
[0074] Reference Figure 5 and Figure 6In the embodiment, the impingement passage 109 and the convection passage 113 are divided into three impingement flow channels 114 and three convection flow channels 116 along the radial direction of the blade by the side wall 115, and the top or bottom of adjacent convection flow channels 116 are communicated by the turning passage 117. Each of the impingement flow channel 114 and the convection flow channel 116 has the same chord-wise width. The height H1 of the impingement passage 109 is 1 mm, and the width L1 of each of the impingement flow channels 114 is 4 mm. The impingement passage 109 and the convection passage 113 are separated by the partition plate 108, and the height H2 of the convection passage 113 below the impingement passage 109 is 1 mm, and the width L2 of each of the convection flow channels 116 is 4 mm. The root of the first convection flow channel 116 on the left side is communicated with the inner chamber 111. An array of impingement holes 110 with a diameter of 1 mm and a radial spacing of 4 mm is arranged on the inner layer wall 107 of the impingement passage 109. An array of film holes 112 with a diameter of 0.6 mm is arranged on the outer layer wall 106. A flow-around structure 118 is arranged on the outer layer wall 106 of the convection passage 113. The flow-around structure 118 is a rib with a protruding height of 0.4 mm. In operation, the cooling gas first enters the inner chamber 111 from the root of the blade tenon 101, and part of the cooling gas flows into the double-layer wall impingement passage 109 through the impingement holes 1101 on the inner layer wall 107, and high-speed flushes the corresponding outer layer wall 106 for cooling. Another part of the cooling gas enters from the inlet 119 of the convection flow channel 116 close to the blade leading edge 102, and then flows through each of the convection flow channels 116 through the turning passage 117 in turn.
[0075] Figure 8 The structure diagram of the impingement flow channel and the convection flow channel of the turbine rotor blade in another embodiment of the utility model is shown. As shown in the figure, the impingement passage 109 is not divided into multiple impingement flow channels 114, and the height is 1 mm and the width L1 is 8 mm. The impingement passage 109 is communicated with the inner chamber 111 through the impingement hole 110. Below the impingement passage 109, the convection passage 113 is divided into three convection flow channels 116 along the radial direction of the blade, and the top or bottom of adjacent convection flow channels 116 is communicated by the turning passage 117. The width of the convection flow channel 116 is 2 mm, and the height is 1 mm. The impingement passage 109 and the convection passage 113 are separated by the partition plate 108. The inlet 119 of the first convection flow channel 116 on the left side is communicated with the inner chamber 111. Two arrays of impingement holes 110 with a diameter of 1 mm and a radial spacing of 4 mm are arranged on the inner layer wall 107 of the impingement passage 109. An array of film holes 112 with a diameter of 0.6 mm is arranged on the outer layer wall 106.
[0076] Turning back to Figure 2In the present embodiment, the inner chamber 111 of the turbine rotor blade 100 is divided into a plurality of chambers I-VI by the side wall 115, wherein the chamber I is adjacent to the blade leading edge 102, the chambers II, III, IV are respectively communicated with the corresponding impingement flow passages 114 through the impingement holes 110, and the chamber VI is adjacent to the blade trailing edge 103. In combination with Figure 1 , Figure 4 and Figure 5 As shown, according to the cooling requirement of the blade, the film holes 112 are arranged on the surface of the blade leading edge 102, the pressure surface 104 and the suction surface 105. When the high pressure turbine of the engine is working, the cooling gas enters the inner chamber 111 through the inlet 119 of the blade tenon 101 in three ways. The first way of the cooling gas firstly enters the chamber II, wherein a part of the cooling gas enters the impingement flow passages 114 and the chamber I of the leading edge 102 through the impingement holes 110 respectively, and then flows out of the blade through the film holes 112; another part of the cooling gas enters the convection passage 113, and then flows out of the blade through the film holes 112. The second way of the cooling gas enters the chambers III and IV respectively, and then enters the corresponding impingement flow passages 114 through the impingement holes 110, and then flows out of the blade from the film holes 112. The third way of the cooling gas firstly enters the chamber V, and then flows to the chamber VI and flows out of the trailing edge 103.
[0077] It is obvious for those skilled in the art that various modifications and variations can be made to the above-mentioned exemplary embodiments of the present application without departing from the spirit and scope of the present application. Therefore, it is intended to cover the modifications and variations of the present application falling within the scope of the appended claims and their equivalents.
Claims
1. A turbine rotor blade suitable for use in a high pressure turbine of an aeroengine, characterised in that, The turbine rotor blade has a double-wall cooling structure, which comprises an outer wall and an inner wall arranged radially along the turbine rotor blade, a partition plate arranged between the outer wall and the inner wall, and air film holes arranged on the outer wall; An impingement channel is formed between the outer wall, the inner wall, the partition plate and the blade tip of the turbine rotor blade, and an impingement hole is arranged on the inner wall forming the impingement channel, and the inner cavity of the turbine rotor blade communicates with the impingement channel through the impingement hole; A convection channel is formed between the outer wall, the inner wall, the partition plate and the blade root of the turbine rotor blade, and the inner cavity communicates with the bottom of the convection channel.
2. The turbine rotor blade of claim 1, wherein, The impingement channel comprises a plurality of impingement flow channels arranged radially along the turbine rotor blade.
3. The turbine rotor blade of claim 2, wherein, The chord-wise width of the impingement flow channel is 3-8 mm.
4. The turbine rotor blade of claim 1, wherein, The diameter of the impingement hole is 0.8-1.5 mm, and the ratio of the impingement distance of the double-wall cooling structure to the diameter of the impingement hole is 0.5:1-1:
1.
5. The turbine rotor blade of claim 2, wherein, The plurality of impingement holes corresponding to each impingement flow channel are arranged radially along the turbine rotor blade.
6. The turbine rotor blade of claim 5, wherein, The plurality of impingement holes corresponding to each impingement flow channel are arranged at equal intervals radially along the turbine rotor blade, or the intervals increase from the blade root to the blade tip of the turbine rotor blade.
7. The turbine rotor blade of claim 1, wherein, The convection channel comprises a plurality of convection flow channels arranged radially along the turbine rotor blade, and the top or bottom of adjacent convection flow channels communicates.
8. The turbine rotor blade of claim 7, wherein, The bottom of the first convection flow channel along the chord-wise direction of the turbine rotor blade communicates with the inner cavity.
9. The turbine rotor blade of claim 1, wherein, A flow-around structure is formed on the outer wall of the convection channel, which faces the inner wall.
10. The turbine rotor blade of claim 9, wherein, The flow-around structure is a rib, a boss or a pit.
11. The turbine rotor blade of claim 10, wherein, The flow-around structure is a rib or a boss, and the ratio of the protrusion height of the rib or boss to the height of the convection channel is 0.2:1-0.5:1.