WO2021139492A1 - 涡轮叶片蜂巢螺旋腔冷却结构 - Google Patents

涡轮叶片蜂巢螺旋腔冷却结构 Download PDF

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Publication number
WO2021139492A1
WO2021139492A1 PCT/CN2020/136325 CN2020136325W WO2021139492A1 WO 2021139492 A1 WO2021139492 A1 WO 2021139492A1 CN 2020136325 W CN2020136325 W CN 2020136325W WO 2021139492 A1 WO2021139492 A1 WO 2021139492A1
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Prior art keywords
spiral cavity
honeycomb
honeycomb spiral
blade
turbine blade
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PCT/CN2020/136325
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English (en)
French (fr)
Inventor
吕东
王楠
王晓放
孔星傲
孙一楠
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大连理工大学
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Priority to US17/433,985 priority Critical patent/US20220170375A1/en
Publication of WO2021139492A1 publication Critical patent/WO2021139492A1/zh

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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186—Film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/13—Two-dimensional trapezoidal
    • F05D2250/132—Two-dimensional trapezoidal hexagonal
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/25—Three-dimensional helical
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/204—Heat transfer, e.g. cooling by the use of microcircuits
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs

Definitions

  • the invention belongs to the technical field of aeroengine and gas turbine turbine cooling technology, and relates to a honeycomb spiral cavity cooling structure of a turbine blade.
  • the current cooling measures for turbine blades generally include internal enhancement of convection heat transfer and external formation of air film isolation.
  • the principle of cooling design is to use the least amount of cold air to take away as much heat as possible, and to protect the parts in a lower temperature range.
  • a current solution to the cooling problem of turbine blades is to adopt a layered structure, as shown in Fig. 1, whose main feature is the use of a multilayer structure to form the outer wall of the turbine blade.
  • the main structure of the laminate includes an air inlet plate located inside the blade and an air outlet plate located outside the blade.
  • the cooling gas enters the laminate from the cold air passage in the inner cavity of the blade through the air inlet hole, and after heat exchange with the inner cavity spoiler column and other structures, it flows out through the air film hole and covers the outer surface of the blade.
  • the main feature of this solution is that it can organically couple heat conduction, convection cooling, impingement cooling and film cooling together. It has the advantages of larger heat exchange area and full utilization of cold air. However, it also has complex structure, difficult manufacturing, and other advantages. Disadvantages such as high flow resistance and weak strength.
  • the honeycomb spiral cavity cooling structure of turbine blades includes a hollow turbine blade, a honeycomb spiral cavity and a spoiler column;
  • the hollow turbine blades are provided with cold air channels inside the cold air channels for low-temperature cooling gas to flow inside the blades to cool the blades.
  • a plurality of arrays of honeycomb spiral cavities are arranged in the wall surface of the hollow turbine blade for cooling gas to enter and perform convective cooling.
  • the center of the honeycomb spiral cavity is provided with a spoiler column, which has a cylindrical structure, which can not only increase the heat exchange area inside the blade, but also has a guiding effect on the cooling airflow.
  • the cold air rotates and flows around the spoiler column in the honeycomb spiral cavity and then exits the blades, and forms an air film on the surface of the blades.
  • Each honeycomb spiral cavity is a unit body, and the shape of each honeycomb spiral cavity is approximately a regular hexagon, and multiple unit bodies are arranged in a honeycomb shape, so that more cooling structures can be arranged in a unit area to make full use of space And form a rich heat exchange area.
  • this relatively independent design of each unit not only ensures the uniform air flow but also avoids the mutual influence between the various channels of cold air.
  • the air inlet hole and the air outlet hole are respectively located on both sides of the wall surface of the blade, and the center line of the air inlet hole and the center line of the air outlet hole are parallel in the same vertical plane.
  • the angles between the centerline of the inlet hole and the centerline of the outlet hole and the horizontal plane are the incident angle ⁇ A1 and the exit angle ⁇ A2, respectively. Both the incident angle ⁇ A1 and the exit angle ⁇ A2 are acute angles.
  • the cross section of the air inlet and the air outlet is rectangular, and the air inlet and the air outlet are smoothly connected with the passage in the honeycomb spiral cavity by using a circular arc slide.
  • incident angle ⁇ A1 and the exit angle ⁇ A2 are both 20 ⁇ 45°.
  • the typical value of incident angle ⁇ A1 and exit angle ⁇ A2 is 30°.
  • the cooling unit is arranged in a dense array like a honeycomb in the blade wall, so that the cooling airflow acts more directly on the hot wall surface, which can improve the effectiveness of cooling measures.
  • the densely arranged hexagonal structure with a central spoiler not only takes into account the heat conduction from the hot wall to the cold wall, but also provides a rich convection heat exchange area, which can comprehensively improve the heat exchange efficiency.
  • the spiral structure of the honeycomb spiral cavity makes the cooling air flow path longer in the plate, and the cold air is more fully utilized.
  • the relatively independent unit structure avoids the mutual interference between the airflows, eliminates the mutual collision and mixing of the cooling airflows in adjacent units, and also avoids the problems of cooling gas reflux and crossflow. While ensuring the full and effective heat exchange, the flow loss is reduced.
  • the air film hole with a rectangular cross section used in the present invention is flatter than the round hole used in the laminate structure, and the air film outflow is better fit and the air film covering effect better.
  • the connection between the air film hole and the inner cavity in the present invention is smoother, and has greater advantages in reducing the flow resistance and improving the air film covering effect.
  • a supporting rib structure is formed between the unit bodies of the honeycomb cavity, which can effectively strengthen the strength of the blades relative to the connection of the spoiler posts in the laminate.
  • the cooling structural elements such as holes and spoiler columns and the unit bodies are arranged in a quadrilateral shape, while in the present invention, the structural elements are arranged in a hexagonal honeycomb shape. As shown in FIG. 3, under the same unit spacing, the number of structural elements per unit area of the present invention can be increased by about 15% relative to the quadrilateral arrangement.
  • the present invention has the beneficial effects that firstly, the flow resistance and loss of the cold air are reduced by about 10-15%, and the efficiency of the whole engine is improved.
  • a typical laminate adopts a structure in which the internal units are connected to each other.
  • the cold air in adjacent units will intersect, impact and mix with each other, and there may also be cross-flow and backflow.
  • each unit is relatively independent and isolated from each other without intermixing of cold air, thereby reducing flow loss.
  • the large expansion and contraction of the cross-sectional area of the flow channel will cause energy loss.
  • the inner cavity space is relatively large in size relative to the hole, which makes the air flow in and out of the laminate. It has undergone two approximate throttling flows; while the cross-sectional area of the cold air flow channel in the present invention is approximately the same along the way, no sudden flow expansion and throttling phenomenon will occur, so the relative resistance of the laminate structure is smaller.
  • Turbine blades are mainly subjected to the following loads during operation: centrifugal load caused by high-speed rotation, aerodynamic load imposed by gas flow, and vibration load caused by vibration. These loads exert tensile, torsion and bending on the blade base. Wait for the deformation tendency and produce the corresponding stress, in addition to the thermal stress caused by the uneven thermal expansion. When these stresses are coupled together and exceed the limit that the material can withstand, failure will occur. As shown in Figure 5, for the laminated cooling structure, it is equivalent to opening a cavity in the original solid wall thickness of the blade. The reduction of this material will result in a reduction in the load resistance of the blade. In order to make up for the loss of blade strength, the layer structure adopts spoiler to connect the inner and outer walls to strengthen it.
  • a hexagonal mesh support rib structure is used in the cavity to connect the inner and outer walls, which can improve the overall structure's anti-compression, bending and torsional load capacity in multiple directions, with an amplitude of more than 20% , which brings improvements in the safety and reliability of the engine.
  • the present invention improves both the internal cooling and external cooling of the turbine blades, and the overall cooling effect is improved by about 8%.
  • the present invention utilizes the cooling air more fully. As shown in Figure 4, after the cooling air in the laminate structure enters the inner cavity, it usually flows around the spoiler column for half a circle and then flows out through the film hole. In the honeycomb spiral cavity structure, the cooling air needs to flow around the spoiler column for more than one week before it can flow out. The flow path is longer, and the total heat exchange with the wall is also greater.
  • the layer structure mainly uses a spoiler column structure to guide the heat flow heated by gas to the cold wall, and its heat conduction ability is related to the total cross-sectional area of the spoiler column.
  • the supporting rib structure formed between the unit bodies is used for heat conduction. The total cross-sectional area of the column and the rib is larger, so Can strengthen the heat conduction between the hot and cold walls.
  • the present invention is also superior to the existing laminate structure.
  • the air film hole of the honeycomb spiral cavity is smoothly transferred to the inner cavity, and its cross section is therefore approximately rectangular, as shown in Figure 6.
  • this relatively flat structure can make the air film outflow more compact, so it can cover a larger area of the blade surface under the same flow rate and cool The effect is better.
  • Figure 1 shows the cooling structure of conventional turbine blades and laminates.
  • Figure 2 (a) is a schematic diagram of the cooling structure of the honeycomb spiral cavity of the turbine blade.
  • Figure 2(b) is a partial enlarged view of the honeycomb spiral cavity cooling structure of the turbine blade.
  • Figure 3 is a comparison diagram of the arrangement of quadrilateral and hexagonal unit bodies.
  • Figure 4 is a comparison diagram of the cooling gas flow state inside the laminate and the honeycomb spiral cavity.
  • Figure 5 is a comparison diagram of the cross-sectional shape of the two laminate structures.
  • Figure 6 is a comparison diagram of the two kinds of air film pore coverage area.
  • Figure 7(a) is a three-dimensional numerical simulation result of the cooling gas flow inside the conventional laminate.
  • Figure 7(b) shows the results of the three-dimensional numerical simulation of the cooling gas flow inside the honeycomb spiral cavity.
  • the present invention conducts a comparative study on the flow state of the internal cooling air between the conventional laminate structure and the honeycomb spiral cavity structure of the present invention through three-dimensional numerical simulation, as shown in Figure 7(a) and Figure 7(b) comparative analysis, it can be seen that the present invention
  • the cross-sectional area of the intercooler air channel is roughly the same along the way, there is no flow protrusion and throttling phenomenon, and the airflow turning angle is small, and there is no mutual collision or mixing, so the relative resistance of the laminate structure is smaller.
  • the honeycomb spiral cavity cooling structure of turbine blades includes a hollow turbine blade 1, a honeycomb spiral cavity 3 and a spoiler 4;
  • the hollow turbine blade 1 is provided with a cold air channel 2 inside, a plurality of arrays of honeycomb spiral cavities 3 are arranged in the blade wall of the hollow turbine blade 1, and the center of the honeycomb spiral cavity 3 is provided with a spoiler 4, the spoiler column 4 is a cylindrical structure, each honeycomb spiral cavity 3 is a unit body, each honeycomb spiral cavity 3 is approximately a regular hexagon, and multiple unit bodies are arranged in a honeycomb shape.
  • the air intake The hole 5 and the air outlet 6 are respectively located on both sides of the blade wall surface, and the center line 9 of the air inlet hole and the center line 10 of the air outlet hole are parallel in the same vertical plane.
  • the cross section of the air inlet hole 5 and the air outlet hole 6 is rectangular, and the air inlet hole 5 and the air outlet hole 6 are smoothly connected with the channel in the honeycomb spiral cavity 3 by using an arc-shaped slideway.
  • the angles between the centerline 9 of the inlet hole and the centerline 10 of the outlet hole and the horizontal plane are the incident angle ⁇ A1 and the exit angle ⁇ A2 respectively, and the incident angle ⁇ A1 and the exit angle ⁇ A2 are both 20°.
  • the honeycomb spiral cavity cooling structure of a turbine blade includes a hollow turbine blade 1, a honeycomb spiral cavity 3 and a spoiler 4; Multiple arrays of honeycomb spiral cavities 3, the center of the honeycomb spiral cavity 3 is provided with a spoiler column 4, the spoiler column 4 is a cylindrical structure, each honeycomb spiral cavity 3 is a unit body, and each honeycomb spiral cavity 3 has a shape of a regular six A polygonal shape, and a plurality of unit bodies are arranged in a honeycomb shape.
  • the air inlet 5 and the air outlet 6 are respectively located on both sides of the blade wall, and the air inlet center line 9 and the air outlet center line 10 Parallel in the same vertical plane.
  • the angles between the centerline of the inlet hole 9 and the centerline of the outlet hole 10 and the horizontal plane are the incident angle ⁇ A1 and the exit angle ⁇ A2, respectively.
  • the typical values of the incident angle ⁇ A1 and the exit angle ⁇ A2 are both 30°.
  • the honeycomb spiral cavity cooling structure of a turbine blade includes a hollow turbine blade 1, a honeycomb spiral cavity 3 and a spoiler 4; Multiple arrays of honeycomb spiral cavities 3, the center of the honeycomb spiral cavity 3 is provided with a spoiler column 4, the spoiler column 4 is a cylindrical structure, each honeycomb spiral cavity 3 is a unit body, and each honeycomb spiral cavity 3 has a shape of a regular six A polygonal shape, and a plurality of unit bodies are arranged in a honeycomb shape.
  • the air inlet 5 and the air outlet 6 are respectively located on both sides of the blade wall, and the air inlet center line 9 and the air outlet center line 10 Parallel in the same vertical plane.
  • the cross section of the air inlet hole 5 and the air outlet hole 6 is rectangular, and the air inlet hole 5 and the air outlet hole 6 are smoothly connected with the channel in the honeycomb spiral cavity 3 by using an arc-shaped slideway.
  • the angles between the centerline of the inlet hole 9 and the centerline of the outlet hole 10 and the horizontal plane are the incident angle ⁇ A1 and the exit angle ⁇ A2 respectively.
  • the typical values of the incident angle ⁇ A1 and the exit angle ⁇ A2 are both 45°.

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  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

一种涡轮叶片蜂巢螺旋腔冷却结构，包括空心涡轮叶片（1）、蜂巢螺旋腔（3）和扰流柱（4）；所述的空心涡轮叶片（1）的内部设有冷气通道（2），冷气通道（2）供低温冷却气体在叶片内部流动，对叶片进行冷却。在空心涡轮叶片（1）的叶片壁面内设有多个阵列的蜂巢螺旋腔（3），以供冷却气体进入并进行对流冷却。蜂巢螺旋腔（3）中心设有扰流柱（4），扰流柱（4）为圆柱形结构。在每个单元体中，进气孔（5）和出气孔（6）分别位于叶片壁面的两侧，其进气孔中心线（9）和出气孔中心线（10）在同一竖直平面内平行。采用该冷却结构，在单位面积上的结构要素数量可以增加约15％，冷气流动阻力和损失减小约10～15％，进而带来发动机整机效率的提高。

Description

涡轮叶片蜂巢螺旋腔冷却结构 技术领域
本发明属于航空发动机和燃气轮机涡轮冷却技术领域，涉及涡轮叶片蜂巢螺旋腔冷却结构。
 背景技术
在目前的航空发动机和燃气轮机技术领域中，为提高装置效能而采取的措施通常为提高涡轮前燃气温度，然而目前使用材料的承受极限远低于燃气环境温度，故涡轮叶片的冷却问题备受关注。当下对涡轮叶片的冷却措施普遍为内部强化对流换热、外部形成气膜隔绝，其冷却设计的原则是使用最少的冷气量来带走尽可能多的热量，保护零部件处于较低的温度范围，并且具有较小的温度梯度，具体为对叶片采用中空设计，使冷却气在中空通道内流动同时利用丰富的内腔表面积进行强化换热，并在冷却气体排出叶片时形成气膜覆盖以隔绝燃气的直接加热，在此基础上追求“更大的内部换热面积”、“更高的换热效率”、“更好的覆盖效果”、“更小的流动阻力”、“对结构强度破坏更小”、“更好的可造性和可维护性”等。
目前解决涡轮叶片冷却问题的一类方案是采用层板结构，如图1所示，其主要特征是采用多层结构组成涡轮叶片的外壁。层板的主要结构包括位于叶片内部的进气板、位于叶片外侧的出气板。在工作时，冷却气体从叶片内腔冷气通道经过进气孔进入层板，在与内腔扰流柱等结构进行换热后，再通过气膜孔流出，并对叶片外表面进行气膜覆盖。该方案的主要特点是能够将导热、对流冷却、冲击冷却和气膜冷却等方式有机的耦合在一起，具有较大的换热面积，冷气利用充分等优点，但同时也有结构复杂、制造难度大、流动阻力大、强度较弱等缺点。
 技术问题
针对现有涡轮叶片层板冷却结构存在的不足，发明了一种蜂巢螺旋腔式的冷却结构。
 技术解决方案
本发明的技术方案如下：
如图2所示，涡轮叶片蜂巢螺旋腔冷却结构，包括空心涡轮叶片、蜂巢螺旋腔和扰流柱；
所述的空心涡轮叶片的内部设有冷气通道，冷气通道供低温冷却气体在叶片内部流动，对叶片进行冷却。
在空心涡轮叶片的叶片壁面内设有多个阵列的蜂巢螺旋腔，以供冷却气体进入并进行对流冷却。蜂巢螺旋腔中心设有扰流柱，扰流柱为圆柱形结构，该结构不仅能增大叶片内部换热面积，还对冷却气流具有导向作用。冷气在蜂巢螺旋腔内绕扰流柱旋转流动一周后排出叶片，并在叶片表面形成气膜覆盖。
每一个蜂巢螺旋腔为一个单元体，每一个蜂巢螺旋腔形状为近似正六边形，并且多个单元体呈蜂巢状排布，这样可以在单位面积内布置下较多的冷却结构，充分利用空间并形成丰富的换热面积。相对于典型层板结构，这种每个单元相对独立的设计既保证了气流均匀又可以避免各路冷气间的相互影响。
在每个单元体中，进气孔和出气孔分别位于叶片壁面的两侧，其进气孔中心线和出气孔中心线在同一竖直平面内平行。进气孔中心线和出气孔中心线与水平面的夹角分别为入射角∠A1 和出射角∠A2 ，入射角∠A1 和出射角∠A2 均为锐角。
进一步的，进气孔和出气孔的横截面为矩形，进气孔和出气孔与蜂巢螺旋腔内的通道采用圆弧状滑道光滑相衔接。
进一步的，入射角∠A1 和出射角∠A2 均为20~45°。入射角∠A1和出射角∠A2典型值为30°。
该发明主要解决的技术问题：
将冷却单元在叶片壁内像蜂巢一样细密的阵列排布，使得冷却气流更直接的作用在热壁面，可以提高冷却措施的有效性。紧密排布的带中心扰流柱的六边形结构在兼顾从热壁向冷壁导热的同时，还提供了丰富的对流换热面积，可以综合提高换热效率。相对于传统的层板方案，蜂巢螺旋腔的螺旋结构使得冷却气流在板内流经的路径更长，对冷气的利用更加充分。从减小流动阻力方面，相对独立的单元体式结构避免了各路气流间的相互干涉，消除了相邻单元体内冷却气流的相互撞击和掺混，也避免了冷却气体回流、串流等问题，在保证换热充分有效的同时减少流动损失。从气膜孔的设计上，本发明所用的截面为矩形的气膜孔相对于层板结构中所用的圆形孔更为扁平，其气膜出流的贴体性更好，气膜覆盖效果更优。另外，相对于层板结构，本发明中气膜孔与内腔的衔接更为光顺，在减小流动阻力和提高气膜覆盖效果方面具有更大的优势。此外，蜂巢腔各单元体之间形成了支撑肋板结构，相对于层板中的扰流柱连接，可以有效加强叶片的强度。
 有益效果
该发明的有益效果：
1、         空间利用更充分
已有的典型层板结构中，孔和扰流柱等冷却结构元素及单元体，均呈现四边形排布，而本发明中结构要素采用呈六边形的蜂巢状排布。如图3所示，在相同的单位间距下，相对于四边形的排布方式，本发明单位面积上的结构要素数量可以增加约15%。
2、         减小冷气流动阻力和损失
本发明相对于原有的层板结构，其有益效果首先是冷气流动阻力和损失减小约10~15%，进而带来发动机整机效率的提高。如图4所示，典型的层板采用的是内部各单元相互连通的结构，相邻单元内的冷气会相互交汇、冲击和掺混，并且还有可能出现串流和回流的现象；而本发明中各单元相对独立，之间相互隔绝无冷气间的相互掺混，从而减少了流动损失。
在气流转折角度方面：冷气进入层板结构后需要在狭窄通道内完成90°的方向转折，而在流出层板时，在气膜孔的入口处，部分气流需要转折135~160°后才能进入到气膜孔，这些过大的气流转折角度将引起流动阻力的大幅度增加；本发明中，冷气在进入和流出冷却腔时，转折角度均为20~45°，在数值上近似等于入射角∠A1和出射角∠A2，大幅度小于层板结构。
另外，在冷气流动过程中，流道截面积的大幅度扩张和收缩都会引起能量的损失，层板结构就是如此，其内腔空间相对孔来说尺寸较大，使得气流在进出层板时要经历两次近似节流的流动；而本发明中冷气流道的截面积沿程大致相同，不会产生流动突扩和节流现象，因此相对层板结构阻力更小。
3、         提高涡轮叶片的抗载荷能力
涡轮叶片在工作中主要承受以下方面的载荷：由高速旋转引起的离心载荷、由燃气气流施加的气动载荷、由于振动引起的振动载荷，这些载荷施加在叶片基体上呈现了拉伸、扭转和弯曲等变形趋势以及产生了相应的应力，另外还有由于热膨胀不均匀而引起的热应力。这些应力耦合在一起并且超出了材料所能承受的极限后，则会发生破坏。如图5所示，对于层板类冷却结构，相当于在叶片原有的实心壁厚中开设空腔，这种材料的减少会导致叶片抗载荷能力的降低。而为了弥补叶片强度上的损失，层板结构采用扰流柱连接内外两层壁，起到加强作用。但是由于这种近似点支撑结构的抗压能力尚可，但抗弯和抗扭性能较弱，因此对结构强化的作用有限。而本发明中在空腔内部采用了呈六边形的网状支撑肋结构连接内外壁，对可以在多个方向上实现结构整体抗压缩、弯曲和扭转载荷能力的提高，幅度达到20%以上，带来发动机整机的安全性和可靠性改善。
4、         提高叶片冷却效果
相对于典型的层板结构，本发明在对涡轮叶片的内部冷却和外部冷却方面均有所提高，综合冷却效果的提高幅度约为8%左右。
首先，本发明对冷却气利用得更充分。如图4所示，层板结构中冷却气在进入内腔后，通常绕扰流柱半周后再通过气膜孔流出。而蜂巢螺旋腔结构中，冷却气在其内部需围绕扰流柱流动一周以上才能流出，其流动的路径更长，与壁面总换热量也就更大。
另外，本发明中从燃气侧的高温热壁向叶片内部冷壁的导热更好。如图5所示，层板结构主要采用依靠扰流柱结构将燃气加热的热流导向冷壁，其导热的能力与扰流柱的总截面积有关。而蜂巢螺旋腔结构中，除了每个单元体内的扰流柱可以进行导热以外，还增加了单元体之间形成的支撑肋结构用于热量的传导，柱和肋的总截面积更大，因此可以强化冷热壁之间的导热。
在叶片外部气膜冷却方面，本发明也优于已有的层板结构。蜂巢螺旋腔的气膜孔与内腔光滑转接，其截面因此为近似矩形，如图6所示。相对于层板结构中普遍采用的截面为圆形的气膜孔，这种较为扁平的结构可以使气膜出流更为贴体，因此可以在相同流量下覆盖更大面积的叶片表面，冷却效果更优。
 附图说明
图1为常规涡轮叶片及层板冷却结构图。
图2（a）为涡轮叶片蜂巢螺旋腔冷却结构示意图。
图2（b）为涡轮叶片蜂巢螺旋腔冷却结构局部放大图。
图3为四边形和六边形单元体排布方式对比图。
图4为层板与蜂巢螺旋腔内部内冷却气体流动状态对比图。
图5为两种层板结构截面形状对比图。
图6为两种气膜孔覆盖面积对比图。
图7（a）为常规层板内部内冷却气体流动三维数值仿真结果图。
图7（b）为蜂巢螺旋腔内部内冷却气体流动三维数值仿真结果图。
图中：1.空心涡轮叶片；2.冷气通道；3.蜂巢螺旋腔；4.扰流柱；5.进气孔；6.气膜孔；7.入射角∠A1；8.出射角∠A2；9.进气孔中心线；10.出气孔中心线。
 本发明的实施方式
为了使本发明的内容更容易被清楚地理解，下面根据具体实施例并结合附图，对本发明作进一步详细的说明。
实施例1
本发明通过三维数值仿真对常规层板结构和本发明中蜂巢螺旋腔结构进行了内部冷却气流动状态的对比研究，如图7（a）和图7（b）对比分析，可以得知本发明中冷气流道的截面积沿程大致相同，不会产生流动突括和节流现象，并且气流转折角度小以及没有相互间的撞击和掺混等，因此相对层板结构阻力更小。
实施例2
如图2所示，涡轮叶片蜂巢螺旋腔冷却结构，包括空心涡轮叶片1、蜂巢螺旋腔3和扰流柱4；
所述的空心涡轮叶片1的内部设有冷气通道2，在空心涡轮叶片1的叶片壁面内设有多个阵列的蜂巢螺旋腔3，蜂巢螺旋腔3中心设有扰流柱4，扰流柱4为圆柱形结构，每一个蜂巢螺旋腔3为一个单元体，每一个蜂巢螺旋腔3形状为近似正六边形，并且多个单元体呈蜂巢状排布，在每个单元体中，进气孔5和出气孔6分别位于叶片壁面的两侧，其进气孔中心线9和出气孔中心线10在同一竖直平面内平行。进气孔5和出气孔6的横截面为矩形，进气孔5和出气孔6与蜂巢螺旋腔3内的通道采用圆弧状滑道光滑相衔接。进气孔中心线9和出气孔中心线10与水平面的夹角分别为入射角∠A1 和出射角∠A2 ，入射角∠A1 和出射角∠A2 均为20°。
实施例3
涡轮叶片蜂巢螺旋腔冷却结构，包括空心涡轮叶片1、蜂巢螺旋腔3和扰流柱4；所述的空心涡轮叶片1的内部设有冷气通道2，在空心涡轮叶片1的叶片壁面内设有多个阵列的蜂巢螺旋腔3，蜂巢螺旋腔3中心设有扰流柱4，扰流柱4为圆柱形结构，每一个蜂巢螺旋腔3为一个单元体，每一个蜂巢螺旋腔3形状为正六边形，并且多个单元体呈蜂巢状排布，在每个单元体中，进气孔5和出气孔6分别位于叶片壁面的两侧，其进气孔中心线9和出气孔中心线10在同一竖直平面内平行。进气孔中心线9和出气孔中心线10与水平面的夹角分别为入射角∠A1 和出射角∠A2 ，入射角∠A1 和出射角∠A2 的典型值均为30°。
实施例4
涡轮叶片蜂巢螺旋腔冷却结构，包括空心涡轮叶片1、蜂巢螺旋腔3和扰流柱4；所述的空心涡轮叶片1的内部设有冷气通道2，在空心涡轮叶片1的叶片壁面内设有多个阵列的蜂巢螺旋腔3，蜂巢螺旋腔3中心设有扰流柱4，扰流柱4为圆柱形结构，每一个蜂巢螺旋腔3为一个单元体，每一个蜂巢螺旋腔3形状为正六边形，并且多个单元体呈蜂巢状排布，在每个单元体中，进气孔5和出气孔6分别位于叶片壁面的两侧，其进气孔中心线9和出气孔中心线10在同一竖直平面内平行。进气孔5和出气孔6的横截面为矩形，进气孔5和出气孔6与蜂巢螺旋腔3内的通道采用圆弧状滑道光滑相衔接。进气孔中心线9和出气孔中心线10与水平面的夹角分别为入射角∠A1 和出射角∠A2 ，入射角∠A1 和出射角∠A2 的典型值均为45°。

Claims (7)

  1. 涡轮叶片蜂巢螺旋腔冷却结构，其特征在于，包括空心涡轮叶片（1）、蜂巢螺旋腔（3）和扰流柱（4）；
    所述的空心涡轮叶片（1）的内部设有冷气通道（2），冷气通道（2）供低温冷却气体在叶片内部流动，对叶片进行冷却；在空心涡轮叶片（1）的叶片壁面内设有多个阵列的蜂巢螺旋腔（3），以供冷却气体进入并进行对流冷却；蜂巢螺旋腔（3）中心设有扰流柱（4），扰流柱（4）为圆柱形结构；
    每一个蜂巢螺旋腔（3）为一个单元体，每一个蜂巢螺旋腔（3）形状为正六边形，并且多个单元体呈蜂巢状排布，在每个单元体中，进气孔（5）和出气孔（6）分别位于叶片壁面的两侧，其进气孔中心线（9）和出气孔中心线（10）在同一竖直平面内平行；进气孔中心线（9）和出气孔中心线（10）与水平面的夹角分别为入射角∠A1 和出射角∠A2 ，入射角∠A1 和出射角∠A2 均为锐角。
  2. 如权利要求1所述的涡轮叶片蜂巢螺旋腔冷却结构，其特征在于，所述的进气孔（5）和出气孔（6）的横截面为矩形。
  3. 如权利要求1或2所述的涡轮叶片蜂巢螺旋腔冷却结构，其特征在于，所述的入射角∠A1 和出射角∠A2 均为20~45°。
  4. 如权利要求1或2所述的涡轮叶片蜂巢螺旋腔冷却结构，其特征在于，所述的进气孔（5）和出气孔（6）与蜂巢螺旋腔（3）内的通道采用圆弧状滑道光滑相衔接。
  5. 如权利要求3所述的涡轮叶片蜂巢螺旋腔冷却结构，其特征在于，所述的进气孔（5）和出气孔（6）与蜂巢螺旋腔（3）内的通道采用圆弧状滑道光滑相衔接。
  6. 如权利要求3所述的涡轮叶片蜂巢螺旋腔冷却结构，其特征在于，所述的入射角∠A1和出射角∠A2典型值为30°。
  7. 如权利要求5所述的涡轮叶片蜂巢螺旋腔冷却结构，其特征在于，所述的入射角∠A1和出射角∠A2典型值为30°。
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