WO2023284354A1 - 一种大涵道比涡扇发动机核心舱通风结构及其通风方法 - Google Patents

一种大涵道比涡扇发动机核心舱通风结构及其通风方法 Download PDF

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WO2023284354A1
WO2023284354A1 PCT/CN2022/088460 CN2022088460W WO2023284354A1 WO 2023284354 A1 WO2023284354 A1 WO 2023284354A1 CN 2022088460 W CN2022088460 W CN 2022088460W WO 2023284354 A1 WO2023284354 A1 WO 2023284354A1
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core
air intake
cabin
annular cavity
wall
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PCT/CN2022/088460
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English (en)
French (fr)
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单勇
尹华莉
刘昊
邓明
赵强
谭晓茗
张靖周
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南京航空航天大学
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Publication of WO2023284354A1 publication Critical patent/WO2023284354A1/zh

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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • F02C7/16—Cooling of plants characterised by cooling medium
    • F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

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  • the invention relates to the technical field of engine core compartment ventilation structure, in particular to a ventilation structure and a ventilation method for a core compartment of a turbofan engine with a large bypass ratio.
  • the high-temperature gas transfers heat to the core engine room in the form of heat radiation and convective heat exchange, while the core engine also exchanges heat in the form of heat conduction and heat radiation.
  • Various heat transfer methods are coupled with each other and affect each other, causing the temperature in the engine core engine room to rise. high.
  • the high temperature in the core engine room will cause the core engine components to work abnormally or be damaged, and in severe cases, it will directly affect the flight safety of the aircraft.
  • the ventilation apparatus includes: a fire seal member (for substantially sealing the core nacelle in the gas turbine engine nacelle), and a chimney (the chimney is configured in fluid communication with the fire seal member to enable hot air to pass through the fire seal from the core chamber for the purpose of leading to the outside atmosphere).
  • the device also provides a gas turbine engine and nacelle design including core compartment ventilation.
  • Vedeshkin et al. studied a certain test procedure that can measure hydraulic performance in aeroengine cabin ventilation system components (such as vents and pressure relief doors (PRD)), and conducted experiments using related experimental models. Finally, the obtained The test results and test results were compared and analyzed.
  • Boileau et al. proposed a method for large eddy simulations on unstructured grids to better predict wall heat transfer in turbulent flow environments or typical aerospace applications with complex geometries . This method is used to study the coupling between periodic vortex shedding and wall heat transfer by phase-averaged analysis.
  • Mohammadinia et al. proposed a genetic algorithm-based ventilation and cooling optimization method for aircraft piston engines. They established an engine cooling model, analyzed the ventilation and cooling principles of aircraft piston engines, and then used genetic algorithms to optimize the ventilation and cooling procedures of aircraft piston engines. The results show that the method can effectively optimize the ventilation and cooling schemes of aviation piston engines. During the optimization process, the average convergence value is always greater than 95.00%. Under different working conditions, the maximum failure rate of the optimized ventilation and cooling scheme is 0.03. And its maximum failure rate is lower than other similar methods.
  • Ma Wenchang and others established a numerical calculation model for the spatial temperature distribution in the engine compartment of a certain type of aircraft, and calculated the cooling of the engine compartment of a certain type of aircraft through two calculation methods of cold air flow and heat transfer in the nacelle, and compared the calculation results Through the analysis, conclusions with engineering application value are obtained, which provide a theoretical basis for the design of aircraft engine compartments.
  • the cooling structure of the core cabin of a turbofan engine with a large bypass ratio is basically formed by opening holes on the outer wall of the engine core cabin, that is, openings on the inner wall of the engine culvert channel, and introducing the culvert airflow;
  • the heat in the core nacelle is carried away and flows out from the ring seam at the end of the core nacelle.
  • the structure must not have too many air-inducing openings on the wall, which will lead to unfavorable factors such as uneven cooling inside the core nacelle;
  • the air intake effect which inevitably affects the incoming air flow, does not even reach the minimum cooling air volume.
  • a large bypass ratio turbofan engine core cabin ventilation structure proposed by the patent of the present invention adds structures such as intake ring cavity and exhaust grille, which is expected to make up for the cooling of the current high bypass ratio turbofan engine core cabin. Insufficient structure.
  • the purpose of the present invention is to provide a ventilation structure and ventilation method for the core cabin of a turbofan engine with a large bypass ratio.
  • the structural design of the cavity and a set of exhaust grilles enables the outer culvert cooling airflow introduced by the ram air intake effect to enter the core engine room evenly in the circumferential direction, and at the same time, it can effectively adjust the cooling airflow after the cooling airflow is introduced from the outer culvert.
  • the flow direction of the flow in the core cabin can achieve better flow and heat exchange effect in the cabin; and the suction effect of the external air flow can be used to make the gas in the core cabin easier to discharge and achieve a better exhaust effect.
  • a core compartment ventilation structure of a turbofan engine with a large bypass ratio comprising:
  • the outer duct wall (1), the core engine room wall (4), the core machine casing (7), and the outer duct flow channel (2) is formed between the outer duct wall (1) and the core engine room wall (4) ), forming a core cabin (6) between the core cabin wall (4) and the core machine casing (7);
  • an air intake annular chamber (3) is provided at a position inside and upstream of the core nacelle, and the air intake annular chamber (3) is used to communicate with the outer stent flow channel (2) and the core nacelle (6), and an exhaust grille (5) is also arranged at the core nacelle wall surface (4) and the tail position, and the exhaust grille (5) is used to drain the inside of the core nacelle (6) The cooling gas is discharged into the outer bypass channel (2).
  • the inside of the air intake annular cavity (3) is an annular hollow cavity, the central axis of which coincides with the central axis (8) of the core cabin (6), and the air intake annular cavity (3) ) is coincident with the core cabin wall surface (4), and a plurality of small air inlet holes (9) are evenly arranged on the side wall;
  • a continuous and inwardly concave inclined surface is formed, and the inclined surface is the end surface (10) of the annular cavity, and a plurality of exhaust small holes (11) are evenly arranged on the end surface (10) of the annular cavity.
  • the number of the air inlet holes (9) is 4-6.
  • a plurality of exhaust grilles (5) are evenly arranged on the circumferential surface of the core cabin wall (4), and the exhaust grills (5) are multi-layer sheet structures with a plurality of blades;
  • a method for ventilating a core compartment ventilation structure of a turbofan engine with a large bypass ratio comprising the following steps:
  • Step S1 Utilizing the stamping effect, the culvert airflow in the culvert flow channel (2) flows into the interior of the air intake annular cavity (3) through the air intake holes (9) on the core cabin wall (4);
  • Step S2 after the cooling air flow enters the interior of the air intake annular cavity (3), it first flows in the circumferential direction, and then flows through the exhaust holes (11) on the end face (10) of the air intake annular cavity (3). ), uniformly circumferentially and at a certain angle into the core cabin (6);
  • Step S3 the cooling airflow flows backward in the core cabin (6), and is discharged from the exhaust grille (5) into the outer culvert flow channel (2).
  • the present invention introduces an air intake annular cavity structure, so that the outer culvert cooling air flows into the air intake annular cavity radially through the air intake small holes on the wall surface of the core cabin, First, it flows in the circumferential direction inside the air intake annular cavity, and then enters the core engine room evenly in the circumferential direction from the small exhaust hole on the end face of the annular cavity.
  • the end face of the air intake ring cavity is designed as a ring surface with a certain angle between the normal direction and the central axis of the core cabin, and the included angle is less than 90 degrees, so that the cooling air flow can be effectively adjusted
  • the flow direction of the airflow enables the outer culvert cooling airflow to flow into the core engine room at a certain angle, so as to better achieve the purpose of cooling the surface of the core engine casing and obtain a better flow and heat exchange effect in the cabin.
  • the design of the exhaust grille in the present invention increases the use of external air flowing through the exhaust grille to produce a suction effect on the basis of relying on the intake pressure of the core cabin to drive the exhaust, making it easier to discharge the gas in the core cabin , and mix with the outer air flow in the outer channel to achieve a better exhaust effect.
  • Fig. 1 is a two-dimensional schematic diagram of the ventilation structure of the engine core compartment provided in Embodiment 1;
  • Fig. 2 is the schematic diagram of the principle of the ventilation structure of the engine core compartment provided in embodiment 1;
  • Fig. 3 is a structural schematic diagram of a core engine room with an outer culvert
  • Fig. 4 is a structural schematic diagram of a core engine room with an outer culvert
  • Fig. 5 is a schematic structural diagram of a core engine room without external connotations
  • Fig. 6 is a structural schematic diagram of a core engine room without external connotations
  • Fig. 7 is a front view of the air intake annular cavity
  • Fig. 8 is a side view of the air intake annular cavity
  • Fig. 9 is a schematic diagram of the structure of the air inlet annular cavity
  • Fig. 10 is a schematic diagram of the structure of the air inlet annular cavity.
  • the present embodiment provides a core compartment ventilation structure of a turbofan engine with a large bypass ratio, including:
  • Outer duct wall 1, core engine room wall 4, core machine casing 7, outer culvert flow channel 2 is formed between outer duct wall 1 and core engine room wall 4, and core is formed between core engine room wall 4 and core engine case 7 cabin 6;
  • an air intake annular chamber 3 is arranged at the position upstream of the core cabin, and the air intake annular chamber 3 is used to communicate with the outer bypass flow channel 2 and the core cabin 6, and at the wall surface 4 of the core cabin and at the rear position
  • An exhaust grill 5 is provided, and the exhaust grill 5 is used to discharge the cooling gas inside the core nacelle 6 into the outer culvert flow channel 2 .
  • the inside of the air intake annular cavity 3 is an annular hollow cavity, and its central axis coincides with the central axis 8 of the core cabin 6 of the core cabin 6, and the side wall of the air intake annular cavity 3 is in line with the The walls 4 of the core cabin are overlapped, and 4-6 small air intake holes 9 are evenly arranged on the side walls; a continuous and inwardly concave slope is provided on the exhaust side of the air intake annular cavity 3, and the slope is The end face 10 of the annular cavity, on which a plurality of exhaust holes 11 are evenly arranged, the angle between the end face 10 of the annular cavity and the central axis 8 of the core nacelle is less than 90 degrees, more specifically, in In this embodiment, the end face of the air inlet annular cavity is designed as a circular surface with a certain angle between the normal direction and the central axis of the core nacelle, and the included angle is less than 90 degrees, so that the cooling airflow can be effectively adjusted.
  • a plurality of exhaust grilles 5 are evenly arranged on the peripheral surface of the core cabin wall 4, and the exhaust grills 5 are multi-layer sheet structures with multiple blades; wherein, in the radial direction In the direction, there is no overlap between the blades and there is a gap; in the axial direction, the tail and the head overlap between the blades, and the tail end of the upstream blade is located outside the front end of the downstream blade to realize the core nacelle 6 The connection between the fluid area and the fluid area of the external channel 2.
  • the exhaust grille provided in this embodiment on the basis of relying on the intake pressure of the core cabin to drive the exhaust, increases the use of external air flowing through the exhaust grille to generate a suction effect, so that the core cabin
  • the inner air is easier to discharge and mixes with the outer air flow in the outer channel to achieve better exhaust effect.
  • this embodiment provides a core compartment ventilation method for a turbofan engine with a large bypass ratio based on the ventilation structure for the core compartment of a turbofan engine with a large bypass ratio provided in Embodiment 1, which specifically includes:
  • Step 1 Utilizing the stamping effect, the culvert airflow in the culvert flow channel 2 flows into the cavity of the air intake annular cavity 3 connected to the core cabin wall 4 through the air intake holes 9 on the core cabin wall 4;
  • Step 2 After the cooling air enters the interior of the air intake annular chamber 3, it flows in the circumferential direction first, and then is evenly discharged in the circumferential direction from the exhaust holes 11 on the end face 10 of the annular chamber of the air intake annular chamber 3 In the core engine room 6; because there is an inclination angle at the end surface 10 of the annular cavity of the air intake annular cavity 3, the cooling air flow can flow into the core engine room 6 at a certain angle, so as to better achieve the purpose of cooling the surface of the core machine casing , to obtain a better flow and heat transfer effect in the cabin;
  • Step 3 The cooling air flow flows backward in the core cabin 6, and is discharged from the exhaust grille 5 into the outer culvert flow channel 2; at the same time, the suction effect of the outer culvert airflow makes it easier to discharge the gas in the core cabin 6 and connect with the outer culvert.
  • the external airflow in runner 2 mixes.

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Abstract

一种大涵道比涡扇发动机核心舱通风结构，包括：外涵道壁面(1)，核心机舱壁面(4)，核心机机匣(7)，外涵道壁面(1)与核心机舱壁面(4)之间构成外涵流道(2)，核心机舱壁面(4)与核心机机匣(7)之间构成核心机舱(6)；其中，在核心机舱内部且上游的位置处设置有进气圆环腔(3)，进气圆环腔(3)用于连通外涵流道(2)与核心机舱(6)，并且在核心机舱壁面(4)且尾部位置处还设置有排气格栅(5)，排气格栅(5)用于将核心机舱(6)内部的冷却气体排入外涵流道(2)中。该结构更好的达到为核心机机匣表面降温目的，获得更好的舱内流动换热效果。还提供了一种通风方法。

Description

一种大涵道比涡扇发动机核心舱通风结构及其通风方法 技术领域
本发明涉及发动机核心舱通风结构技术领域，特别是涉及一种大涵道比涡扇发动机核心舱通风结构及其通风方法。
背景技术
燃气涡轮发动机核心机正常工作时要向外散发大量的热量。高温燃气以热辐射及对流换热的形式向核心机舱传递热量，而核心机内部又以导热、热辐射等形式换热，各种换热方式相互耦合，相互影响，造成发动机核心机舱内温度升高。核心机舱内的高温会引起核心机部件工作异常或损坏，严重时会直接影响飞机的飞行安全。因此考虑到核心机舱内附件和结构的寿命，设计合适的核心机舱通风冷却系统方案提供核心机舱冷却所需的气流来保证核心机舱和核心机机匣的温度在规定的限制范围内十分重要。
Sawyers-abbott等人研究了一种发动机核心舱通风装置。该通风装置包括：防火密封构件(用于基本密封燃气涡轮发动机机舱中的核心机舱)、以及烟囱(烟囱构造成与防火密封构件流体连通的形式，以实现将热空气通过防火密封件从核心室引导至外部大气的目的)。该装置还提供了一种燃气涡轮发动机和包括核心舱通风装置的机舱的设计方案。Vedeshkin等人研究了可以在航空发动机机舱通风系统组件(如通风口和泄压门(PRD))中测得液压性能的某种测试程序，并利用相关的实验模型进行了试验，最后对得到的测试结果和试验结果进行了对比和分析。结果表明，该测试程序的计算结果与实验模型的试验结果吻合良好。Boileau等人则提出了一种在非结构化网格上进行大涡模拟的方法，其目的是为了更好地预测在湍流环境下或是具有复杂几何形状的典型航空应用中的壁面传热情况。该方法用于通过相平均分析研究周期性涡旋脱落与壁传热之间的耦合。近些年来，Mohammadinia等人又提出了一种基于遗传算法的飞机活塞发动机通风冷却优化方法。他们通过建立发动机冷却模型，分析飞机活塞发动机的通风和冷却原理，然后采用遗传算法对飞机活塞发动机的通风和冷却程序进行优化。结果表明，该方法可以有效地优化航空活塞发动机的通风和冷却方案。在优化过程中，平均收敛值始终大于95.00％。在不同的工作条件下，优化的通风和冷却方案的最大故障率为0.03。且其最大故障率低于其他类似方法。
国内学者同样对发动机的通风冷却系统进行了大量的实验研究。王杏涛等人对某型发动机舱进行二维结构简化，通过数值计算的方法，研究了发动机舱通风冷却、辐射遮挡和隔热 层等对发动机舱蒙皮的冷却降温效果。缪国君等人研究了民用飞机涡扇发动机短舱内的通风冷却系统，介绍了通风冷却系统的设计方案，提出了通风冷却系统的设计要求和设计内容，给出了通风冷却系统的验证思路。张亚海等人针对某型直升机动力舱通风冷却系统，提出了一种简化的基于旋翼下洗流的动力舱通风冷却性能计算方法，利用商业CFD软件，计算了悬停状态下动力舱通风冷却系统性能，分析了湍流模型、旋翼下洗流以及发动机散热率等因素对动力舱温度场和排气引射器性能的影响。王涛等人总结、介绍了几种通风冷却系统进气设计方案及其特点，并对试飞结果进行了整理分析。王玉梅等人对某飞机的发动机舱温超限的故障进行分析，针对发动机舱温超限现象对发动机舱通风冷却系统的进气方式提出了多种方案。马文昌等人建立了某型飞机发动机舱内空间温度分布的数值计算模型，通过两种短舱冷气流量和换热的计算方法对某型飞机发动机舱冷却进行了计算，并对计算结果进行对比分析，得到了具有工程应用价值的结论，为飞机发动机舱设计提供了理论依据。
工程上，大涵道比涡扇发动机核心机舱的冷却结构基本是由发动机核心机舱外侧壁面，即发动机外涵通道内侧壁面开孔，引入外涵气流；引入的气流通过对流换热的方式将发动机核心机舱内的热量带走，从核心机舱末端环缝流出。该结构因为考虑到外涵流动损失，一方面，壁面引气开孔数目不能过多，从而会带来核心机舱内部周向冷却不均匀的不利因素；另一方面，完全依靠外涵气流的冲压进气效应，从而不可避免的对引入外涵空气流量产生影响，甚至达不到最低的冷却空气量。为了增加引气量，只有增加壁面引气开孔数，从而增加了外涵的流动损失。以上两点相互影响，不可调和。因此，本发明专利提出的“一种大涵道比涡扇发动机核心舱通风结构”增加进气圆环腔和排气格栅等结构，有望弥补当前大涵道比涡扇发动机核心机舱的冷却结构的不足。
发明内容
有鉴于此，本发明的目的在于提供一种大涵道比涡扇发动机核心舱通风结构及其通风方法，其在现有小孔进、排气的基础上，通过引入一种进气圆环腔和一组排气格栅的结构设计，使利用冲压进气效应引入的外涵冷却气流能在周向上均匀的进入核心机舱内，同时能够有效的调整自外涵引入冷却气流后，气冷却流在核心机舱内的流动方向，以达到更好的舱内流动换热效果；并利用外涵气流的抽吸效应使得核心机舱内气体更容易排出，达到更好的排气效果。
为了达到上述目的，本发明采用如下技术方案：
一种大涵道比涡扇发动机核心舱通风结构，包括：
外涵道壁面(1)，核心机舱壁面(4)，核心机机匣(7)，所述外涵道壁面(1)与所 述核心机舱壁面(4)之间构成外涵流道(2)，所述核心机舱壁面(4)与所述核心机机匣(7)之间构成核心机舱(6)；
其中，在所述核心机舱内部且上游的位置处设置有进气圆环腔(3)，所述进气圆环腔(3)用于连通所述外涵流道(2)与所述核心机舱(6)，并且在所述核心机舱壁面(4)且尾部位置处还设置有排气格栅(5)，所述排气格栅(5)用于将所述核心机舱(6)内部的冷却气体排入所述外涵流道(2)中。
进一步的，所述进气圆环腔(3)的内部为环状中空腔体，其中心轴与核心机舱(6)的核心机舱中心轴线(8)重合，所述进气圆环腔(3)的侧壁与所述核心机舱壁面(4)重合，并且该侧壁上均匀设置有多个进气小孔(9)；在所述进气圆环腔(3)的排气侧设有一圈连续的且向内凹陷的斜面，该斜面为圆环腔端面(10)，在该圆环腔端面(10)上还均匀布置了多个排气小孔(11)。
进一步的，所述进气小孔(9)的数量为4-6个。
进一步的，所述核心机舱壁面(4)上周向均匀设置有多个排气格栅(5)，所述排气格栅(5)为多层片状结构，具有多个叶片；
其中，在径向方向上，叶片与叶片之间不搭接并且留有空隙；在轴向上，叶片与叶片之间尾、首重叠，上游的叶片的尾端位于下游叶片前端的外侧，用以实现所述核心机舱(6)流体区域与外涵流道(2)流体区域的贯通。
一种大涵道比涡扇发动机核心舱通风结构的通风方法，包括如下步骤：
步骤S1、利用冲压效应，外涵流道(2)中的外涵气流通过核心机舱壁面(4)上的进气小孔(9)流入与进气圆环腔(3)的腔体内部；
步骤S2、冷却气流进入进气圆环腔(3)腔体内部后先沿周向流动，然后从进气圆环腔(3)的圆环腔端面(10)上的排气小孔(11)，周向均匀地且以一定角度的流入核心机舱(6)内部；
步骤S3、冷却气流在核心机舱(6)内向后流动，从排气格栅(5)排入外涵流道(2)。
本发明的有益效果是：
1、本发明在小孔进气的基础上，通过引入一种进气圆环腔结构，使外涵冷却气流通过核心机舱壁面上的进气小孔沿径向流入进气圆环腔后，先在进气圆环腔内部沿周向流动，再从圆环腔端面上的排气小孔周向均匀地进入核心机舱中。
2、本发明将进气圆环腔的圆环腔端面设计为一个法线方向与核心机舱中心轴线存在一定角度，且夹角小于90度的圆环面，这样能够有效的调整引入冷却气流后气流的流动方向，使 外涵冷却气流能以一定的角度流入核心机舱内，更好的达到为核心机机匣表面降温的目的，获得更好的舱内流动换热效果。
3、本发明中排气格栅的设计在原本依靠核心机舱进气压力驱使排气的基础上增加了利用外涵空气流经排气格栅产生抽吸效应，使得核心机舱内气体更容易排出，并与外涵流道中的外涵气流混合，以达到更好的排气效果。
附图说明
图1、为实施例1中提供的发动机核心舱通风结构二维示意图；
图2、为实施例1中提供的发动机核心舱通风结构的原理示意图；
图3、为带外涵的核心机舱的结构示意图；
图4、为带外涵的核心机舱的结构示意图；
图5、为不带外涵的核心机舱的结构示意图；
图6、为不带外涵的核心机舱的结构示意图；
图7、为进气圆环腔的正视图；
图8、为进气圆环腔的侧视图；
图9、为进气圆环腔的的结构示意图；
图10、为进气圆环腔的的结构示意图。
附图中：
1-外涵道壁面；2-外涵流道；3-进气圆环腔；4-核心机舱壁面；5-排气格栅；6-核心机舱；7-核心机机匣；8-核心机舱中心轴线；9-进气小孔；10-圆环腔端面；11-排气小孔。
具体实施方式
为使本发明实施例的目的、技术方案和优点更加清楚，下面将结合本发明实施例中的附图，对本发明实施例中的技术方案进行清楚、完整地描述，显然，所描述的实施例是本发明一部分实施例，而不是全部的实施例。基于本发明中的实施例，本领域普通技术人员在没有做出创造性劳动前提下所获得的所有其他实施例，都属于本发明保护的范围。
实施例1
参见图1-图10，本实施例提供一种大涵道比涡扇发动机核心舱通风结构，包括：
外涵道壁面1，核心机舱壁面4，核心机机匣7，外涵道壁面1与核心机舱壁面4之间构成外涵流道2，核心机舱壁面4与核心机机匣7之间构成核心机舱6；
其中，在核心机舱内部且上游的位置处设置有进气圆环腔3，进气圆环腔3用于连通外涵流道2与核心机舱6，并且在核心机舱壁面4且尾部位置处还设置有排气格栅5，排气格栅 5用于将核心机舱6内部的冷却气体排入外涵流道2中。
具体的说，在本实施例中，进气圆环腔3的内部为环状中空腔体，其中心轴与核心机舱6的核心机舱中心轴线8重合，进气圆环腔3的侧壁与核心机舱壁面4重合，并且该侧壁上均匀设置有4-6个进气小孔9；在进气圆环腔3的排气侧设有一圈连续的且向内凹陷的斜面，该斜面为圆环腔端面10，在该圆环腔端面10上还均匀布置了多个排气小孔11，圆环腔端面10与核心机舱中心轴线8的夹角小于90度，更具体的说，在本实施例中将进气圆环腔的圆环腔端面设计为一个法线方向与核心机舱中心轴线存在一定角度，且夹角小于90度的圆环面，这样能够有效的调整引入冷却气流后气流的流动方向，使外涵冷却气流能以一定的角度流入核心机舱内，更好的达到为核心机机匣表面降温的目的，获得更好的舱内流动换热效果。
具体的说，在本实施例中，核心机舱壁面4上周向均匀设置有多个排气格栅5，排气格栅5为多层片状结构，具有多个叶片；其中，在径向方向上，叶片与叶片之间不搭接并且留有空隙；在轴向上，叶片与叶片之间尾、首重叠，上游的叶片的尾端位于下游叶片前端的外侧，用以实现核心机舱6流体区域与外涵流道2流体区域的贯通。更具体的说，本实施例中提供的排气格栅在原本依靠核心机舱进气压力驱使排气的基础上，增加了利用外涵空气流经排气格栅产生抽吸效应，使得核心机舱内气体更容易排出，并与外涵流道中的外涵气流混合，以达到更好的排气效果。
实施例2
本实施例在实施例1的基础之上，基于实施例1提供的一种大涵道比涡扇发动机核心舱通风结构，提供一种大涵道比涡扇发动机核心舱通风方法，具体包括：
步骤一：利用冲压效应，外涵流道2中的外涵气流通过核心机舱壁面4上的进气小孔9流入与核心机舱壁面4连接的进气圆环腔3的腔体内部；
步骤二：冷却气流进入进气圆环腔3腔体内部后先沿周向流动，然后从进气圆环腔3的圆环腔端面10上的排气小孔11，周向均匀地排入核心机舱6中；由于进气圆环腔3的圆环腔端面10存在一个倾斜角度，使冷却气流能以一定的角度流向核心机舱6内，更好的达到为核心机机匣表面降温的目的，获得更好的舱内流动换热效果；
步骤三：冷却气流在核心机舱6内向后流动，从排气格栅5排入外涵流道2；同时，外涵气流的抽吸效应使得核心机舱6内气体更容易排出，并与外涵流道2中的外涵气流混合。
本发明未详述之处，均为本领域技术人员的公知技术。
以上详细描述了本发明的较佳具体实施例。应当理解，本领域的普通技术人员无需创造性劳动就可以根据本发明的构思作出诸多修改和变化。因此，凡本技术领域中技术人员依本 发明的构思在现有技术的基础上通过逻辑分析、推理或者有限的实验可以得到的技术方案，皆应在由权利要求书所确定的保护范围内。

Claims (5)

  1. 一种大涵道比涡扇发动机核心舱通风结构，其特征在于，包括：
    外涵道壁面(1)，核心机舱壁面(4)，核心机机匣(7)，所述外涵道壁面(1)与所述核心机舱壁面(4)之间构成外涵流道(2)，所述核心机舱壁面(4)与所述核心机机匣(7)之间构成核心机舱(6)；
    其中，在所述核心机舱内部且上游的位置处设置有进气圆环腔(3)，所述进气圆环腔(3)用于连通所述外涵流道(2)与所述核心机舱(6)，并且在所述核心机舱壁面(4)且尾部位置处还设置有排气格栅(5)，所述排气格栅(5)用于将所述核心机舱(6)内部的冷却气体排入所述外涵流道(2)中。
  2. 根据权利要求1所述的一种大涵道比涡扇发动机核心舱通风结构，其特征在于，所述进气圆环腔(3)的内部为环状中空腔体，其中心轴与核心机舱(6)的核心机舱中心轴线(8)重合，所述进气圆环腔(3)的侧壁与所述核心机舱壁面(4)重合，并且该侧壁上均匀设置有多个进气小孔(9)；在所述进气圆环腔(3)的排气侧设有一圈连续的且向内凹陷的斜面，该斜面为圆环腔端面(10)，在该圆环腔端面(10)上还均匀布置了多个排气小孔(11)。
  3. 根据权利要求2所述的一种大涵道比涡扇发动机核心舱通风结构，其特征在于，所述进气小孔(9)的数量为4-6个。
  4. 根据权利要求1所述的一种大涵道比涡扇发动机核心舱通风结构，其特征在于，所述核心机舱壁面(4)上周向均匀设置有多个排气格栅(5)，所述排气格栅(5)为多层片状结构，具有多个叶片；
    其中，在径向方向上，叶片与叶片之间不搭接并且留有空隙；在轴向上，叶片与叶片之间尾、首重叠，上游的叶片的尾端位于下游叶片前端的外侧，用以实现所述核心机舱(6)流体区域与外涵流道(2)流体区域的贯通。
  5. 一种如权利要求书2-4中任意一项所述的一种大涵道比涡扇发动机核心舱通风结构的通风方法，其特征在于，包括如下步骤：
    步骤S1、利用冲压效应，外涵流道(2)中的外涵气流通过核心机舱壁面(4)上的进气小孔(9)流入与进气圆环腔(3)的腔体内部；
    步骤S2、冷却气流进入进气圆环腔(3)腔体内部后先沿周向流动，然后从进气圆环腔(3)的圆环腔端面(10)上的排气小孔(11)，周向均匀地且以一定角度的流入核心机舱(6)内部；
    步骤S3、冷却气流在核心机舱(6)内向后流动，从排气格栅(5)排入外涵流道(2)。
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