CN113123833B - Turbine outer ring block air supply structure with separated air supply - Google Patents

Turbine outer ring block air supply structure with separated air supply Download PDF

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CN113123833B
CN113123833B CN202110323611.XA CN202110323611A CN113123833B CN 113123833 B CN113123833 B CN 113123833B CN 202110323611 A CN202110323611 A CN 202110323611A CN 113123833 B CN113123833 B CN 113123833B
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cavity
impact
turbine
throttling
cooling unit
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CN113123833A (en
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邱天
丁水汀
高自强
徐阳
刘传凯
刘晓静
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Beihang University
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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
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—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
    • F01D5/186—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
    • 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

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

Abstract

The invention belongs to the field of thermal protection of aero-engines, and particularly relates to a turbine outer ring block air supply structure capable of accurately supplying air in a cavity. According to the invention, a layer of throttling cavity is additionally arranged between the air supply cavity and the impact cavity of the turbine outer ring block, the throttling cavity is axially divided into a plurality of independent composite cooling throttling sub-cavities along the turbine, and the throttling cavity corresponds to the throttling cavity, and the composite cooling impact cavity is also axially divided into a plurality of independent impact sub-cavities along the turbine, so that the whole outer ring block is axially divided into a plurality of cooling units which are not interfered with each other, the problem that the outer ring block is poor before cooling and rich after cooling due to a single impact cavity is avoided, a pointed cooling structure is formed, the utilization rate of cold air is improved, and the performance of an engine is enhanced.

Description

一种分腔供气的涡轮外环块供气结构A turbine outer ring block gas supply structure for gas supply by cavity

技术领域technical field

本发明属于航空发动机热防护领域，特别涉及一种精确分腔供气的涡轮外环块供气结构。The invention belongs to the field of aero-engine thermal protection, and in particular relates to a turbine outer ring block air supply structure for air supply by precise cavities.

背景技术Background technique

目前航空发动机涡轮前温度达到了2000K左右，根据热力循环来看，未来航空发动机涡轮前温度会随着性能的提升不断增高，因而涡轮外环的冷却极为重要。现有发动机空气系统引气流量达到了25％，未来可能会更高，空气系统引气流量的增大极大程度地抑制了发动机性能的提高，因而如何降低空气系统引气流量成为发动机热端部件冷却设计的关键问题。At present, the temperature before the turbine of the aero-engine has reached about 2000K. According to the thermodynamic cycle, the temperature before the turbine of the aero-engine will continue to increase with the improvement of performance in the future, so the cooling of the outer ring of the turbine is extremely important. The bleed air flow of the existing engine air system has reached 25%, and it may be higher in the future. The increase of the bleed air flow of the air system greatly inhibits the improvement of engine performance. Therefore, how to reduce the bleed air flow of the air system has become the hot end of the engine. Critical issues in component cooling design.

发动机外环块处于涡轮动叶外机匣，在发动机工作过程中，涡轮动叶顶部的压力沿着发动机轴向具有较大的压力梯度，如图1所示，因而外环块沿发动机轴向的出口压力存在极大的压力梯度。现有技术一般在外环块进口采用相同进气压力，如图2所示，外环块具有冲击-扰流-气膜复合冷却三层结构，冲击孔进口压力相同，在冲击腔的冲击靶面上冲击作用基本相同，冲击腔内气膜孔进口处在基本相同的压力边界条件下，导致气膜孔出口吹风比前低后高，从而导致涡轮动叶进口对应的外环块位置处吹风比过小，而在动叶出口对应的外环块吹风比远大于进口处。这样导致的后果是外环块表面气膜覆盖效果前劣后优，同时由于沿发动机轴向主流温度梯度较大，造成了外环块冷却前贫后富。基于此，若想提高冷却效果，则需要提高冷气流量，弥补外环块高温区域冷却的短板。这种措施虽然保证了外环块的足够冷却，但提升了空气系统的引气比例，造成发动机性能的降低。The outer ring block of the engine is located in the outer casing of the turbine bucket. During the operation of the engine, the pressure at the top of the turbine bucket has a large pressure gradient along the axial direction of the engine, as shown in Figure 1. Therefore, the outer ring block is located along the axial direction of the engine. There is a great pressure gradient at the outlet pressure. In the prior art, the same inlet pressure is generally used at the inlet of the outer ring block. As shown in Figure 2, the outer ring block has a three-layer structure of impact-turbulence-air film composite cooling, the inlet pressure of the impact hole is the same, and the impact target in the impact cavity The impact effect on the surface is basically the same, and the inlet of the gas film hole in the impact cavity is under basically the same pressure boundary conditions, which causes the air blowing at the outlet of the gas film hole to be lower than before and higher than that of the back, resulting in blowing at the position of the outer ring block corresponding to the inlet of the turbine bucket. The ratio is too small, and the blowing ratio of the outer ring block corresponding to the outlet of the rotor blade is much larger than that at the inlet. The consequence of this is that the air film covering effect on the surface of the outer ring block is inferior before and then excellent, and at the same time, due to the large temperature gradient of the mainstream along the engine axis, the outer ring block is cooled before being lean and then rich. Based on this, if you want to improve the cooling effect, you need to increase the flow of cold air to make up for the short plate of cooling in the high temperature area of the outer ring block. Although this measure ensures sufficient cooling of the outer ring block, it increases the bleed air ratio of the air system, resulting in a decrease in engine performance.

发明内容SUMMARY OF THE INVENTION

为解决在外环块冷却过程中,由于外环块出口轴向压力梯度大，主流侧温度梯度大导致的冷气分配不合理，进而影响外环块冷却以及冷气过度损失的问题，本发明设计了一种精确分腔供气的涡轮外环块供气结构，通过增加节流腔，将外环块现有的冲击-扰流-气膜复合冷却三层结构，改变为节流-冲击-扰流-气膜复合冷却四层结构。本发明在冲击孔前增加节流功能，使得在沿着涡轮轴向方向上的冲击-扰流-气膜复合冷却的进排气压力能够具有降低的趋势，从而避免由于动叶出口所对应的外环块进气压力过高导致气膜孔吹风比过大导致的冷却不足以及流量过大导致的性能下降等问题。In order to solve the problem of unreasonable distribution of cold air due to the large axial pressure gradient at the outlet of the outer ring block and the large temperature gradient on the mainstream side during the cooling process of the outer ring block, which further affects the cooling of the outer ring block and the excessive loss of cold air, the present invention designs a An air supply structure for the outer ring block of the turbine with accurate air supply by cavities. By adding a throttling cavity, the existing three-layer structure of shock-turbulence-air film composite cooling of the outer ring block is changed to throttling-shock-turbulence. Flow-air film composite cooling four-layer structure. The present invention adds a throttling function before the impingement hole, so that the intake and exhaust pressures of the impingement-spoiler-air film composite cooling in the axial direction of the turbine can have a decreasing trend, thereby avoiding the Excessive intake pressure of the outer ring block leads to problems such as insufficient cooling caused by excessive air blowing ratio of the air film holes and performance degradation caused by excessive flow.

为实现上述目的，本发明提供了一种分腔供气的涡轮外环块供气结构，沿流体流动方向依次包括节流板、节流腔、冲击板、冲击腔和气膜板；所述节流板与所述冲击板间隔设置且两者之间形成所述节流腔，所述节流板上设置有供流体流入所述节流腔的多个阵列排布的节流孔；所述节流腔中成排间隔设置有多个第一柱肋，以将所述节流腔分割成多个相互独立的节流分腔；所述冲击板和所述气膜板间隔设置且两者之间形成所述冲击腔，所述冲击板上设置有供流体从各节流分腔流入所述冲击腔的多个阵列排布的冲击孔；所述冲击腔中成排间隔设置有多个第二柱肋，以将所述冲击腔分割成多个相互独立的冲击分腔；所述气膜板上设置有供流体从各冲击分腔流出的多个阵列排布的气膜孔；各冲击分腔内包括多个扰流柱；所述节流分腔和所述冲击分腔的数量相等且顺排对应，以将所述供气结构分成多个互相独立的冷却单元；各第一柱肋和各第二柱肋的长度方向与涡轮轴向方向垂直。In order to achieve the above purpose, the present invention provides an air supply structure for the outer ring block of the turbine for air supply by cavities, which sequentially includes a throttle plate, a throttle cavity, an impact plate, an impact cavity and an air film plate along the fluid flow direction; The flow plate and the impingement plate are spaced apart and the throttle chamber is formed therebetween, and the throttle plate is provided with a plurality of orifices arranged in arrays for the fluid to flow into the throttle chamber; the A plurality of first column ribs are arranged in a row and spaced in the throttling cavity to divide the throttling cavity into a plurality of independent throttling sub-cavities; the impingement plate and the air film plate are arranged at intervals and both The impact chamber is formed between the impact chambers, and the impact plate is provided with a plurality of impact holes arranged in a plurality of arrays for the fluid to flow into the impact chamber from each throttling sub-chamber; the impact chamber is provided with a plurality of arrays at intervals. The second column rib is used to divide the impact cavity into a plurality of independent impact sub-cavities; the gas film plate is provided with a plurality of gas film holes arranged in an array for the fluid to flow out from each impact sub-cavity; each The impact sub-chamber includes a plurality of spoiler columns; the throttle sub-chambers and the impact sub-chambers are equal in number and correspond in sequence, so as to divide the air supply structure into a plurality of mutually independent cooling units; each first The longitudinal direction of the column rib and each of the second column ribs is perpendicular to the axial direction of the turbine.

在一些实施方式中，各冷却单元的节流孔总面积配置成，使得进入各冷却单元的流体的流量沿涡轮轴向方向成梯度变化。In some embodiments, the total orifice area of each cooling unit is configured such that the flow rate of fluid entering each cooling unit is gradient in the axial direction of the turbine.

在一些实施方式中，各冷却单元的节流孔总面积设置过程如下：In some embodiments, the setting process of the total area of the orifice of each cooling unit is as follows:

1)计算获取涡轮主流与外环块的交界面沿涡轮轴向的压力及温度分布；1) Calculate and obtain the pressure and temperature distribution of the interface between the main flow of the turbine and the outer ring block along the axial direction of the turbine;

2)根据使各冷却单元宽度，将所述供气结构划分成n个冷却单元；同时将涡轮主流与外环块的交界面分割成n个单元，n个单元与n个冷却单元的气膜孔出口位置一一对应，基于步骤1)中获取的涡轮主流与外环块的交界面沿涡轮轴向的压力及温度分布，确定每个单元下的平均温度和平均压力，即确定每个冷却单元出口处的平均温度和平均压力；2) According to the width of each cooling unit, the air supply structure is divided into n cooling units; at the same time, the interface between the main flow of the turbine and the outer ring block is divided into n units, and the air film between the n units and the n cooling units The positions of the holes are in one-to-one correspondence. Based on the pressure and temperature distribution along the turbine axis at the interface between the turbine main flow and the outer ring block obtained in step 1), the average temperature and average pressure under each unit are determined, that is, each cooling unit is determined. Average temperature and average pressure at the outlet of the unit;

3)分析单个冷却单元的流量特性，获取流经单个冷却单元不同出口温度下的进出口压比及流经流量的关系，基于步骤2)中确定的每个冷却单元出口处的平均温度和平均压力，确定每个冷却单元的进出口压力边界以及温度边界；3) Analyze the flow characteristics of a single cooling unit, and obtain the relationship between the inlet and outlet pressure ratios and flow through a single cooling unit at different outlet temperatures, based on the average temperature and average flow rate at the outlet of each cooling unit determined in step 2). Pressure, determine the inlet and outlet pressure boundaries and temperature boundaries of each cooling unit;

4)根据材料许用温度需求确定流经各冷却单元的冷气流量，步骤3)中获取的流经单个冷却单元不同出口温度下的进出口压比及流经流量的关系，确定各冷却单元进口处的节流孔总面积。4) Determine the flow of cold air flowing through each cooling unit according to the allowable temperature requirements of the material, and determine the relationship between the inlet and outlet pressure ratios and flow through a single cooling unit at different outlet temperatures obtained in step 3), and determine the inlet of each cooling unit total orifice area.

在一些实施方式中，各冷却单元的节流孔总面积设置成沿涡轮轴向逐渐递减。In some embodiments, the total area of the orifices of each cooling unit is arranged to gradually decrease along the axial direction of the turbine.

在一些实施方式中，所述节流孔、冲击孔和所述气膜孔为圆形孔，孔径依据冷却单元的冷气需求设置。In some embodiments, the throttle hole, the impact hole and the air film hole are circular holes, and the hole diameters are set according to the cold air demand of the cooling unit.

在一些实施方式中，所述扰流柱为圆柱形扰流柱或人字形扰流柱。In some embodiments, the spoiler column is a cylindrical spoiler column or a herringbone spoiler column.

本发明的有益效果：Beneficial effects of the present invention:

1)本发明通过精确分腔供气改变了外环块冷气分配前贫后富的弊端，保证了冷气的合理分配，降低了由于高温部分冷气量少导致的空气系统引气量增多；1) The present invention has changed the disadvantage that the cold air distribution of the outer ring block is poor before and then rich by accurate air supply by dividing the cavity, ensuring the reasonable distribution of the cold air, and reducing the increase in the air system bleed volume caused by the low amount of cold air in the high temperature part;

2)本发明以更加合理的冷却方式保证了外环块出口汇入主流的吹风比，避免了由于吹风比过大对主流造的影响；2) The present invention ensures the blowing ratio of the outlet of the outer ring block into the mainstream by a more reasonable cooling method, and avoids the influence on the mainstream caused by the excessive blowing ratio;

3)本发明将外环块沿轴向分为多个互不干扰的冷却结构，即在进口处首先通过沿轴向不同节流能力的节流孔对空气系统引气进行分配，调节整个分腔的流阻，从而实现冷气流经各个腔室后具有不同的冷气供给，进而保证了沿着发动机轴线方向上具有一定的流量梯度，保证了外环块各个位置的充足冷却，进而增强外环块整体的冷却效果。3) In the present invention, the outer ring block is divided into a plurality of cooling structures that do not interfere with each other along the axial direction, that is, at the inlet, the bleed air of the air system is firstly distributed through the throttling holes with different throttling capabilities along the axial direction, and the entire cooling structure is adjusted. The flow resistance of the cavity, so that the cold air has different cold air supply after passing through each cavity, thereby ensuring a certain flow gradient along the axis of the engine, ensuring sufficient cooling of each position of the outer ring block, and then strengthening the outer ring. The overall cooling effect of the block.

附图说明Description of drawings

图1是外环块出口涡轮动叶顶端压力分布图；Fig. 1 is the pressure distribution diagram of the top of the turbine bucket at the outlet of the outer ring block;

图2是现有技术的冲击-扰流-气膜复合冷却系统的外环块结构示意图；Fig. 2 is the schematic diagram of the outer ring block structure of the prior art impact-turbulence-air film composite cooling system;

图3是本发明实施例的分腔供气的涡轮外环块供气结构示意图；3 is a schematic diagram of the gas supply structure of the outer ring block of the turbine for gas supply in a divided cavity according to an embodiment of the present invention;

图4为本发明实施例的节流孔分布示意图；FIG. 4 is a schematic diagram of orifice distribution according to an embodiment of the present invention;

图5是本发明实施例的双层壁人字形扰流柱结构示意图；5 is a schematic structural diagram of a double-walled herringbone spoiler column according to an embodiment of the present invention;

图6是本发明实施例的双层壁人字形扰流柱结构基元剖面图。6 is a cross-sectional view of a structural element of a double-walled herringbone spoiler column according to an embodiment of the present invention.

具体实施方式Detailed ways

本发明在涡轮外环块的供气腔与冲击腔间增加了一层节流腔，并将节流腔沿涡轮轴向分为多个独立的复合冷却的节流分腔，同时与节流腔对应，复合冷却的冲击腔也沿涡轮轴向分割为多个独立的冲击分腔，这样将整个外环块沿涡轮轴向分割成多个互不干扰的冷却单元，避免了由于单个冲击腔导致的外环块冷却前贫后富的问题，形成了具有针对性的冷却结构，提高了冷气的利用率，进而在相同涡轮前温度的前提下提高了发动机性能。The invention adds a layer of throttling cavity between the air supply cavity and the impact cavity of the outer ring block of the turbine, and divides the throttling cavity into a plurality of independent composite cooling throttling sub-cavities along the axial direction of the turbine. Corresponding to the cavity, the impact cavity of composite cooling is also divided into multiple independent impact sub-cavities along the axial direction of the turbine, so that the entire outer ring block is divided into multiple cooling units that do not interfere with each other along the axial direction of the turbine, avoiding the need for a single impact cavity. The resulting problem that the outer ring block is lean before cooling and then rich, forms a targeted cooling structure, improves the utilization rate of cold air, and improves the engine performance under the premise of the same temperature before the turbine.

下面结合附图和实施例进一步描述本发明，应该理解，以下所述实施例旨在便于对本发明的理解，而对其不起任何限定作用。The present invention will be further described below with reference to the accompanying drawings and embodiments, and it should be understood that the following embodiments are intended to facilitate the understanding of the present invention, but do not have any limiting effect on it.

如图3所示，本实施例的分腔供气的涡轮外环块供气结构，沿冷气流动方向(图3中显示从上至下)依次包括节流板1、节流腔2、冲击板3、冲击腔4和气膜板5。As shown in FIG. 3 , the air supply structure of the turbine outer ring block for air supply by dividing the cavity of this embodiment includes a throttle plate 1, a throttle cavity 2, an impingement plate 1, a throttle cavity 2, an impingement plate 1, a throttling cavity 2, a shock absorbing Plate 3, impact chamber 4 and air film plate 5.

节流板1与冲击板3间隔设置且两者之间形成节流腔2，节流板1上设置有供流体流入节流腔2的多个阵列排布的节流孔11。节流腔2中成排间隔设置有多个第一柱肋6，将节流腔2沿涡轮轴向(图4箭头所指方向)顺排分割成多个相互独立的节流分腔21。The throttle plate 1 and the impingement plate 3 are arranged at intervals and a throttle cavity 2 is formed therebetween. The throttle plate 1 is provided with a plurality of orifices 11 arranged in an array for the fluid to flow into the throttle cavity 2 . The throttle chamber 2 is provided with a plurality of first column ribs 6 at intervals in a row, and divides the throttle chamber 2 into a plurality of independent throttle sub-chambers 21 in a row along the turbine axial direction (the direction indicated by the arrow in FIG. 4 ).

冲击板3和气膜板5间隔设置且两者之间形成冲击腔4，冲击板3上设置有供流体从各节流分腔21流入冲击腔4的多个阵列排布的冲击孔31。特别地，由于在冲击腔内可能会减弱沿涡轮轴向上的流量分配，本发明在冲击腔对应节流腔的位置设置同样的分腔结构，以保证流量的供给，进而保证充足的冷却。具体地，在冲击腔4中成排间隔设置多个第二柱肋7，将冲击腔4分割成多个相互独立的冲击分腔41，各冲击分腔41内包括多个扰流柱42，并且使节流分腔21和冲击分腔41的数量相等且顺排对应。气膜板5上设置有供流体从各冲击分腔41流出的多个阵列排布的气膜孔51。第一柱肋6和第二柱肋7的长度方向与涡轮轴向方向垂直。The impingement plate 3 and the gas film plate 5 are spaced apart and an impingement cavity 4 is formed therebetween. The impingement plate 3 is provided with a plurality of arrayed impingement holes 31 for allowing fluid to flow into the impingement cavity 4 from each throttling sub-chamber 21 . In particular, since the flow distribution in the axial direction of the turbine may be weakened in the impingement cavity, the present invention provides the same sub-chamber structure at the position of the impingement cavity corresponding to the throttle cavity to ensure the supply of flow, thereby ensuring sufficient cooling. Specifically, a plurality of second column ribs 7 are arranged in a row in the impact chamber 4 at intervals, and the impact chamber 4 is divided into a plurality of mutually independent impact sub-cavities 41. Each impact sub-cavity 41 includes a plurality of spoiler columns 42. And the throttle sub-cavities 21 and the impact sub-cavities 41 are equal in number and correspond in sequence. The gas film plate 5 is provided with a plurality of gas film holes 51 arranged in an array for the fluid to flow out from each impact sub-chamber 41 . The longitudinal directions of the first column rib 6 and the second column rib 7 are perpendicular to the turbine axial direction.

特别地，使节流分腔21和冲击分腔41的数量相等且顺排对应，可以将整个供气结构分成多个互相独立的冷却单元，从而避免由于单个冲击腔导致的外环块冷却前贫后富的问题，形成具有针对性的冷却结构。In particular, by making the throttle sub-chambers 21 and the impact sub-chambers 41 equal in number and corresponding in sequence, the entire air supply structure can be divided into a plurality of mutually independent cooling units, thereby preventing the outer ring block from being depleted before cooling due to a single impact chamber After the rich problem, a targeted cooling structure is formed.

特别地，为了保证充足的冷却，各冷却单元尺寸越小越好，则各个孔越小越好。基于现有的常规加工技术，本实施例将所有孔型均设为圆形孔，孔径为0.4mm。In particular, in order to ensure sufficient cooling, the smaller the size of each cooling unit, the better, and the smaller the size of each hole, the better. Based on the existing conventional processing technology, in this embodiment, all hole types are set as circular holes with a diameter of 0.4 mm.

此外，为了保证外环块的有效冷却，即，将外环块有限的冷气进行合理分配，本发明在将外环块沿涡轮轴向分为多个互不干扰的冷却单位的基础上，对各冷却单元所包括的节流孔的数量以及各节流孔的大小进行设计，以使节流孔沿涡轮轴向具有不同的节流能力，实现对空气系统引气进行分配，然后通过冲击扰流气膜的复合冷却结构对外环块。有利地，设置各冷却单元的节流孔的数量，使得进入各冷却单元的流体的流量沿涡轮轴向方向成梯度变化，实现整个分腔的流阻调节，从而实现冷气流经各个分腔后具有不同的冷气供给，进而保证了沿着涡轮轴线方向上具有一定的流量梯度，保证了外环块各个位置的充足冷却。然后根据沿涡轮轴向上分割出的冷却单元的出口压力及出口温度不同，通过设计各冷却单元进气节流孔总面积来调整得到不同的进气节流面积，具体确定过程如下：In addition, in order to ensure the effective cooling of the outer ring block, that is, to reasonably distribute the limited cooling air of the outer ring block, the present invention divides the outer ring block into a plurality of cooling units that do not interfere with each other along the turbine axis. The number of orifices included in each cooling unit and the size of each orifice are designed so that the orifices have different throttling capabilities along the turbine axis, so as to distribute the bleed air of the air system, and then pass the impacting and turbulent air. The composite cooling structure of the membrane is an outer ring block. Advantageously, the number of orifices of each cooling unit is set, so that the flow rate of the fluid entering each cooling unit changes in a gradient along the axial direction of the turbine, so as to realize the flow resistance adjustment of the entire sub-chamber, so as to realize the flow of cold air after passing through each sub-chamber. It has different cold air supply, thus ensuring a certain flow gradient along the axis of the turbine, and ensuring sufficient cooling of each position of the outer ring block. Then, according to the different outlet pressures and outlet temperatures of the cooling units divided up along the turbine axis, different intake throttle areas are obtained by designing the total area of the intake throttle holes of each cooling unit. The specific determination process is as follows:

1)计算获取涡轮主流与外环块的交界面沿涡轮轴向的压力及温度分布；1) Calculate and obtain the pressure and temperature distribution of the interface between the main flow of the turbine and the outer ring block along the axial direction of the turbine;

2)根据使各冷却单元宽度，将所述供气结构划分成n个冷却单元；同时将涡轮主流与外环块的交界面分割成n个单元，n个单元与n个冷却单元的气膜孔出口位置一一对应，基于步骤1)中获取的涡轮主流与外环块的交界面沿涡轮轴向的压力及温度分布，确定每个单元下的平均温度和平均压力，即确定每个冷却单元出口处的平均温度和平均压力；2) According to the width of each cooling unit, the air supply structure is divided into n cooling units; at the same time, the interface between the main flow of the turbine and the outer ring block is divided into n units, and the air film between the n units and the n cooling units The positions of the holes are in one-to-one correspondence. Based on the pressure and temperature distribution along the turbine axis at the interface between the turbine main flow and the outer ring block obtained in step 1), the average temperature and average pressure under each unit are determined, that is, each cooling unit is determined. Average temperature and average pressure at the outlet of the unit;

3)分析单个冷却单元的流量特性，获取流经单个冷却单元不同出口温度下的进出口压比及流经流量的关系，基于步骤2)中确定的每个冷却单元出口处的平均温度和平均压力，确定每个冷却单元的进出口压力边界以及温度边界；3) Analyze the flow characteristics of a single cooling unit, and obtain the relationship between the inlet and outlet pressure ratios and flow through a single cooling unit at different outlet temperatures, based on the average temperature and average flow rate at the outlet of each cooling unit determined in step 2). Pressure, determine the inlet and outlet pressure boundaries and temperature boundaries of each cooling unit;

4)根据材料许用温度需求确定流经各冷却单元的冷气流量，步骤3)中获取的流经单个冷却单元不同出口温度下的进出口压比及流经流量的关系，确定各冷却单元进口处的节流孔总面积。4) Determine the flow of cold air flowing through each cooling unit according to the allowable temperature requirements of the material, and determine the relationship between the inlet and outlet pressure ratios and flow through a single cooling unit at different outlet temperatures obtained in step 3), and determine the inlet of each cooling unit total orifice area.

假设，将涡轮主流与外环块交界面沿涡轮轴向分成n个单元，求得第i个单元的平均压力为P(i)(i＝1～n)，平均温度为T(i)(i＝1～n)。通过计算，单个冷却单元不同温度、不同进口面积下的流量特性为m＝f1(Tout，Pin/Pout，Ain)，Tout表示冷却单元的出口温度，Pin表示冷却单元的入口压力，Pout表示冷却单元的出口压力，Ain表示冷却单元的入口面积，f1表示m与Tou、Pin/Pout、Ain的映射关系；此时外环块单元最高温度为Tmax＝f2(Tout，Pin/Pout，m)，f2表示Tmax与Tout、Pin/Pout、m的映射关系，Tout表示冷却单元的出口温度；则冷却单元的入口面积Ain＝g1(Tout，Pin/Pout，m)，Pin，Tin已知，g1表示Ain与Tout、Pin/Pout、m的映射关系。当材料需用温度已知时，其所需的流量为mneed＝g2(Tmax，Tout，Pin/Pout)，g2表示mneed与Tmax、Tout、Pin/Pout的映射关系，则第i个冷却单元的进口面积为Ain(i)＝g1(T(i)，Pin/P(i)，mneed)(i＝1～n)。Assuming that the interface between the turbine main flow and the outer ring block is divided into n units along the turbine axis, the average pressure of the i-th unit is obtained as P(i)(i=1～n), and the average temperature is T(i)( i=1～n). Through calculation, the flow characteristics of a single cooling unit at different temperatures and different inlet areas are m=f 1 (T out , P in /P out , A in ), where T out represents the outlet temperature of the cooling unit , and Pin represents the cooling unit’s Inlet pressure, P out represents the outlet pressure of the cooling unit, A in represents the inlet area of the cooling unit, f 1 represents the mapping relationship between m and T ou , P in /P out , and A in ; at this time, the maximum temperature of the outer ring block unit is T max =f 2 (T out , P in /P out , m), f 2 represents the mapping relationship between T max and T out , Pin /P out , m , and T out represents the outlet temperature of the cooling unit; then the cooling unit The inlet area of A in =g 1 (T out , P in /P out , m), P in , T in are known, and g 1 represents the mapping relationship between A in and T out , P in /P out , and m. When the required temperature of the material is known, the required flow rate is m need =g 2 (T max , T out , P in /P out ), and g 2 represents m need and T max , T out , P in /P The mapping relationship of out , the inlet area of the ith cooling unit is A in (i)=g 1 (T(i), P in /P(i), m need ) (i=1～n).

在本实施例中，在节流腔2中成排间隔设置14个第一柱肋6，冲击腔4中也成排间隔设置14个第二柱肋7，形成位置相互对应且数量均为15的节流分腔21和冲击分腔41。In this embodiment, 14 first column ribs 6 are arranged in a row at intervals in the throttle chamber 2, and 14 second column ribs 7 are also arranged in a row at intervals in the impact chamber 4, and the formation positions correspond to each other and the number is 15. throttling sub-chamber 21 and impact sub-chamber 41.

如图4所示，本实施例沿涡轮轴向的节流孔数量分别为7个，7个，7个，6个，6个，6个，5个，5个，5个，5个，4个，4个，4个，4个，4个。在此设计下，保证了各冷却单元沿涡轮轴向供给的流量按照节流孔面积比例进行了严格的分割，进而实现了外环块的精确冷却。As shown in FIG. 4 , the number of orifices along the axial direction of the turbine in this embodiment are 7, 7, 7, 6, 6, 6, 5, 5, 5, 5, 4, 4, 4, 4, 4. Under this design, it is ensured that the flow supplied by each cooling unit along the axial direction of the turbine is strictly divided according to the area ratio of the orifice, thereby realizing precise cooling of the outer ring block.

特别地，各冲击分腔41包括横向连续排布的多个基元。如图5-6所示，各基元包括由两个第二柱肋7形成的空间以及位于该空间的轴向截面为人字形的扰流柱42、2个冲击孔31和6个气膜孔51。扰流柱42包括顶部421和从顶部421分别向左右两侧且由上向下向外倾斜扩展的两个侧翼422。6个气膜孔51均布于顶部521上方且靠近其中一个第二柱肋7。各侧翼422与两个第二柱肋7形成减缩通道423，2个冲击孔31布置于两个侧翼422的正下方中间位置，使得从2个冲击孔31流入该基元的冷气在经过减缩通道423时，动能不断加速，从而保证较强的横流效应，增强对流换热系数，进而增强了冷却。In particular, each impact sub-chamber 41 includes a plurality of cells arranged in succession in the transverse direction. As shown in Figures 5-6, each element includes a space formed by two second column ribs 7 and a spoiler column 42 with a herringbone-shaped axial cross-section located in the space, 2 impact holes 31 and 6 air film holes 51. The spoiler column 42 includes a top portion 421 and two lateral wings 422 extending from the top portion 421 to the left and right sides and from top to bottom and outwards. The six air film holes 51 are evenly distributed above the top portion 521 and close to one of the second columns. Rib 7. Each side wing 422 and the two second column ribs 7 form a reduction channel 423, and the two impact holes 31 are arranged in the middle position directly below the two side wings 422, so that the cold air flowing into the unit from the two impact holes 31 passes through the reduction channel. At 423, the kinetic energy is continuously accelerated, thereby ensuring a strong cross-flow effect, enhancing the convective heat transfer coefficient, and thus enhancing cooling.

对于本领域的普通技术人员来说，在不脱离本发明创造构思的前提下，还可以对本发明的实施例做出若干变型和改进，这些都属于本发明的保护范围。For those of ordinary skill in the art, without departing from the inventive concept of the present invention, several modifications and improvements can also be made to the embodiments of the present invention, which all belong to the protection scope of the present invention.

Claims (5)

1. The gas supply structure of the turbine outer ring block for supplying gas in a separated cavity is characterized by sequentially comprising a throttle plate, a throttle cavity, an impact plate, an impact cavity and a gas film plate in the flowing direction of a fluid; the throttling plate and the impact plate are arranged at intervals, the throttling cavity is formed between the throttling plate and the impact plate, and a plurality of throttling holes which are arranged in an array mode and used for allowing fluid to flow into the throttling cavity are formed in the throttling plate; a plurality of first column ribs are arranged in the throttling cavity at intervals in rows so as to divide the throttling cavity into a plurality of mutually independent throttling branch cavities; the impact plate and the air film plate are arranged at intervals, the impact cavity is formed between the impact plate and the air film plate, and a plurality of impact holes which are arranged in an array and are used for allowing fluid to flow into the impact cavity from each throttling sub-cavity are formed in the impact plate; a plurality of second column ribs are arranged in the impact cavity at intervals in a row so as to divide the impact cavity into a plurality of independent impact sub-cavities; the air film plate is provided with a plurality of air film holes which are arranged in an array and used for fluid to flow out of each impact sub-cavity; each impact sub-cavity comprises a plurality of turbulence columns; the throttling subchambers and the impact subchambers are equal in number and correspond in a row, so that the gas supply structure is divided into a plurality of cooling units which are independent of each other; the length direction of each first column rib and each second column rib is vertical to the axial direction of the turbine;
the total orifice area of each cooling unit is configured such that the flow rate of the fluid entering each cooling unit changes in a gradient manner in the turbine axial direction.
2. The air supply structure according to claim 1, wherein the total orifice area setting process of each cooling unit is as follows:
1) calculating and obtaining the pressure and temperature distribution of the interface of the turbine main flow and the outer ring block along the axial direction of the turbine;
2) dividing the gas supply structure into n cooling units according to the width of each cooling unit; dividing the interface of the main flow of the turbine and the outer ring block into n units, wherein the n units correspond to the positions of film hole outlets of the n cooling units one by one, and determining the average temperature and the average pressure under each unit based on the pressure and the temperature distribution of the interface of the main flow of the turbine and the outer ring block of the turbine along the axial direction of the turbine, which are obtained in the step 1), namely determining the average temperature and the average pressure at the outlet of each cooling unit;
3) analyzing the flow characteristics of the single cooling unit, acquiring the relation between the inlet-outlet pressure ratio and the flow at different outlet temperatures of the single cooling unit, and determining the inlet-outlet pressure boundary and the temperature boundary of each cooling unit based on the average temperature and the average pressure at the outlet of each cooling unit determined in the step 2);
4) determining the flow rate of cold air flowing through each cooling unit according to the allowable material temperature requirement, and determining the total area of the throttling holes at the inlets of the cooling units according to the relationship between the inlet-outlet pressure ratio and the flow rate of the cold air flowing through the single cooling unit at different outlet temperatures acquired in the step 3).
3. The air supply structure according to claim 2, wherein the total orifice area of each cooling unit is set to gradually decrease in the turbine axial direction.
4. The air supply structure according to any one of claims 1 to 3, wherein the orifice, the impingement holes and the film holes are circular holes having a diameter set according to a cold air demand of the cooling unit.
5. Air supply structure according to any one of claims 1 to 3, characterised in that the turbulence column is a cylindrical or chevron-shaped turbulence column.
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Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN107435563A (en) * 2017-05-05 2017-12-05 西北工业大学 A kind of case structure with tip clearance control and the flowing control of leaf top
CN107559090A (en) * 2016-06-30 2018-01-09 中国航发商用航空发动机有限责任公司 The cooling component of turbine outer ring
CN207813753U (en) * 2017-12-26 2018-09-04 中国航发商用航空发动机有限责任公司 Turbine outer ring cooling structure, turbine structure and aero-engine

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8137056B2 (en) * 2006-03-02 2012-03-20 Ihi Corporation Impingement cooled structure
US8727704B2 (en) * 2010-09-07 2014-05-20 Siemens Energy, Inc. Ring segment with serpentine cooling passages

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN107559090A (en) * 2016-06-30 2018-01-09 中国航发商用航空发动机有限责任公司 The cooling component of turbine outer ring
CN107435563A (en) * 2017-05-05 2017-12-05 西北工业大学 A kind of case structure with tip clearance control and the flowing control of leaf top
CN207813753U (en) * 2017-12-26 2018-09-04 中国航发商用航空发动机有限责任公司 Turbine outer ring cooling structure, turbine structure and aero-engine

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