CN118911838A - Active jet control method for forced transition of boundary layer of ultra-high speed air inlet channel - Google Patents
Active jet control method for forced transition of boundary layer of ultra-high speed air inlet channel Download PDFInfo
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- 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
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Abstract
本发明提供一种超高速进气道边界层强制转捩主动喷气控制方法，在吸气式超高速飞行器前体或进气道压缩面上沿垂直流动方向布置一排等间距喷孔，气源经过输送管路从喷孔射出，喷射气体与来流空气相互作用，产生流向涡结构，诱导超高速边界层强制转捩为湍流；控制方法包括确定喷孔安装位置、确定喷射介质、确定喷孔直径、确定喷前总压、确定喷孔间距和喷孔数目，该方法与现有的被动式强制转捩控制方法相比，可以显著拓宽边界层强制转捩控制范围，降低控制措施导致的进气道性能损失，且避免了转捩带与高速气流的直接碰撞，基本不需要额外的热防护设计，对于保障吸气式超高速飞行器宽范围稳定工作具有重要意义。
The present invention provides an active jet control method for forced transition of boundary layer in an ultra-high-speed inlet, wherein a row of equally spaced nozzles are arranged on a front body or an inlet compression surface of an air-breathing ultra-high-speed aircraft in a direction perpendicular to the flow, a gas source is ejected from the nozzles through a delivery pipeline, the ejected gas interacts with the incoming air to generate a flow vortex structure, and the ultra-high-speed boundary layer is induced to be forced to transition to turbulence; the control method comprises determining the installation position of the nozzles, determining the ejection medium, determining the nozzle diameter, determining the total pressure before ejection, determining the nozzle spacing and the number of nozzles; compared with the existing passive forced transition control method, the method can significantly widen the control range of forced transition of the boundary layer, reduce the loss of inlet performance caused by control measures, and avoid direct collision between the transition zone and the high-speed airflow, and basically does not require additional thermal protection design, which is of great significance for ensuring the wide range and stable operation of the air-breathing ultra-high-speed aircraft.
Description
技术领域Technical Field
本发明吸气式超高速飞行器及发动机领域，特别是进气道流动控制设计方面。The present invention relates to the field of air-breathing hypersonic aircraft and engines, and particularly to the design of air inlet flow control.
背景技术Background Art
对于吸气式超高速飞行器，如果进气道边界层内的流体为层流，则容易导致压缩拐角和唇口激波反射区的流动发生分离，这种分离具有极大的不确定度，可能导致进气道不起动和飞行器失控，从而给飞行器带来严重的安全隐患；但若进入进气道的流体转捩为湍流，则可大大增强边界层的抗逆压梯度能力，从而有效抑制这种分离，便于进气道起动。实际上，尺寸不大的吸气式超高速飞行器在高空中飞行时，由于来流湍流度和噪声均较低（低于0.1%），进气道内的边界层往往很难发生自然转捩，导致飞行存在安全隐患，因此，需要在前体加装控制装置，促进边界层强制转捩为湍流，增加边界层抗逆压梯度能力，减少流动分离，保障进气道起动。For air-breathing hypersonic aircraft, if the fluid in the boundary layer of the inlet is laminar, it is easy to cause the flow separation in the compression corner and the lip shock wave reflection area. This separation has great uncertainty and may cause the inlet to fail to start and the aircraft to lose control, thus posing a serious safety hazard to the aircraft; but if the fluid entering the inlet turns into turbulence, the boundary layer's ability to resist adverse pressure gradients can be greatly enhanced, thereby effectively suppressing this separation and facilitating the start of the inlet. In fact, when a small-sized air-breathing hypersonic aircraft flies at high altitudes, due to the low turbulence and noise of the incoming flow (less than 0.1%), the boundary layer in the inlet is often difficult to naturally transition, resulting in safety hazards in flight. Therefore, it is necessary to install a control device on the forebody to promote the forced transition of the boundary layer to turbulence, increase the boundary layer's ability to resist adverse pressure gradients, reduce flow separation, and ensure the start of the inlet.
目前已有的边界层强制转捩控制主要采用被动式的强制转捩装置，可分为分布式转捩带和离散式的转捩带。分布式的转捩带多采用很小的固体颗粒（如金刚砂等）制成，颗粒之间的尺寸很小难以区分；而离散的转捩带多采用金属颗粒，颗粒大小一般在几毫米左右，颗粒形状一般为圆柱形、钻石型、三角形、后掠斜坡型等。这些被动式的控制装置具有结构简单，控制效果明显，稳定可靠的优点，但其缺点也非常明显：往往只能针对较窄的飞行包线范围，当来流条件偏离设计点后，其要么会因粗糙颗粒高度过低导致强制转捩效果失效，给飞行器带来安全隐患，要么会因粗糙颗粒高度过高而引起不必要的进气道性能损失，且带来额外的热防护需求。The existing boundary layer forced transition control mainly adopts passive forced transition devices, which can be divided into distributed transition zones and discrete transition zones. Distributed transition zones are mostly made of very small solid particles (such as corundum, etc.), and the size of the particles is very small and difficult to distinguish; while discrete transition zones are mostly made of metal particles, the particle size is generally about a few millimeters, and the particle shape is generally cylindrical, diamond-shaped, triangular, swept slope type, etc. These passive control devices have the advantages of simple structure, obvious control effect, stability and reliability, but their disadvantages are also very obvious: they can often only be used for a narrow range of flight envelopes. When the incoming flow conditions deviate from the design point, the forced transition effect will either fail due to the low height of the coarse particles, which will bring safety hazards to the aircraft, or the coarse particles will cause unnecessary loss of inlet performance due to the high height, and bring additional thermal protection requirements.
本发明为了克服被动式强制转捩带的缺点，提供了一种主动喷气的边界层强制转捩控制方法，可以在较宽飞行马赫数和高度范围内实现对边界层强制转捩的宽范围低损耗控制。In order to overcome the shortcomings of the passive forced transition zone, the present invention provides a boundary layer forced transition control method for active jet, which can achieve wide-range low-loss control of boundary layer forced transition within a wider range of flight Mach numbers and altitudes.
发明内容Summary of the invention
本发明的目的在于提供一种采用壁面喷气的超高速进气道边界层强制转捩主动喷气控制方法，该方法与现有的被动式强制转捩控制方法相比，可以显著拓宽边界层强制转捩控制范围，降低控制措施导致的进气道性能损失，且避免了转捩带与高速气流的直接碰撞，基本不需要额外的热防护设计，对于保障吸气式超高速飞行器宽范围稳定工作具有重要意义。The purpose of the present invention is to provide an active jet control method for forced transition of the boundary layer of an ultra-high-speed inlet using wall jets. Compared with the existing passive forced transition control method, this method can significantly widen the control range of the boundary layer forced transition, reduce the performance loss of the inlet caused by control measures, and avoid the direct collision between the transition zone and the high-speed airflow. Basically, no additional thermal protection design is required, which is of great significance for ensuring the wide range and stable operation of air-breathing ultra-high-speed aircraft.
本发明的上述技术目的是通过以下技术方案得以实现的：The above technical objectives of the present invention are achieved through the following technical solutions:
一种超高速进气道边界层强制转捩主动喷气控制方法，在吸气式超高速飞行器前体或进气道压缩面上沿垂直流动方向布置一排等间距喷孔，气源经过输送管路从喷孔射出，喷射气体与来流空气相互作用，产生流向涡结构，诱导超高速边界层强制转捩为湍流；An active jet control method for forced transition of a boundary layer in a hyperspeed inlet is provided, wherein a row of equally spaced spray holes are arranged on the front body or the compression surface of the inlet of an air-breathing hyperspeed aircraft in a direction perpendicular to the flow direction, and a gas source is ejected from the spray holes through a delivery pipeline, and the ejected gas interacts with the incoming air to generate a flow vortex structure, thereby inducing forced transition of the hyperspeed boundary layer into turbulence;
控制方法包括确定喷孔安装位置、确定喷射介质、确定喷孔直径、确定喷前总压、确定喷孔间距和喷孔数目，包括以下步骤：The control method includes determining the installation position of the spray hole, determining the spray medium, determining the diameter of the spray hole, determining the total pressure before spraying, determining the spray hole spacing and the number of spray holes, and includes the following steps:
（1）确定喷孔安装位置，喷孔安装于吸气式超高速飞行器的前体或进气道压缩面上；(1) Determine the nozzle installation position. The nozzle is installed on the front body or the compression surface of the air intake of the air-breathing hypersonic vehicle;
对于多波系压缩进气道，喷孔安装于第一道压缩面上，For multi-wave compression inlet, the nozzle is installed on the first compression surface.
对于三维曲面压缩进气道，喷孔安装于前体上，确保由喷射气流诱导出的流向涡有足够的失稳距离；For the three-dimensional curved compression inlet, the nozzle is installed on the front body to ensure that the streamwise vortex induced by the jet flow has enough instability distance;
同时，喷孔安装位置还需考虑飞行器内有足够的空间可以放置输气管路；确定安装位置后，通过层流状态的流体力学计算，获取安装位置处的层流边界层厚度δ，此处边界层厚度δ定义为总焓边界层厚度，边界层外缘的总焓为来流总焓的99 %；At the same time, the nozzle installation position also needs to consider whether there is enough space in the aircraft to place the gas pipeline. After determining the installation position, the laminar boundary layer thickness δ at the installation position is obtained through fluid mechanics calculation in the laminar state. Here, the boundary layer thickness δ is defined as the total enthalpy boundary layer thickness. The total enthalpy at the outer edge of the boundary layer is 99% of the total enthalpy of the incoming flow.
（2）确定喷射介质，喷射介质采用空气或氮气，储气压力不小于1个大气压；(2) Determine the injection medium. The injection medium shall be air or nitrogen, and the gas storage pressure shall not be less than 1 atmosphere;
（3）确定喷孔直径0.3-1mm，确定喷前总压P j ，喷孔形状为圆形，喷孔与压缩面垂直，使得喷射气流与边界层内气流相互垂直；气流穿透深度h取为当地边界层厚度的0.5倍，气流穿透深度定义为：获取垂直喷孔中心线上的流场马赫数分布，定义沿该垂直喷孔中心线上马赫数第一次达到局部极大值时对应的点距壁面距离为穿透深度h，由于气流穿透深度满足关系式： ，其中p 2为采用声速喷射气流静压，p为无喷射时当地流体静压，参数a、b通过流体力学计算进行标定，标定获得参数a、b后，代入关系式，得到喷射气流静压p 2，再根据等熵关系式，其中γ=1.4为比热比，M=1为喷射马赫数，得喷前总压P j= 1.893p 2 ；(3) Determine the nozzle diameter of 0.3-1mm, determine the total pressure before spraying Pj , the nozzle shape is circular, the nozzle is perpendicular to the compression surface, so that the injection airflow and the airflow in the boundary layer are perpendicular to each other; the airflow penetration depth h is taken as 0.5 times the local boundary layer thickness, and the airflow penetration depth is defined as: obtain the flow field Mach number distribution on the center line of the vertical nozzle, and define the distance from the point on the center line of the vertical nozzle to the wall when the Mach number first reaches the local maximum as the penetration depth h . Since the airflow penetration depth satisfies the relationship: , where p2 is the static pressure of the jet airflow using sonic velocity, p is the static pressure of the local fluid when there is no jet, and the parameters a and b are calibrated by fluid mechanics calculation. After the parameters a and b are obtained by calibration, they are substituted into the relationship to obtain the static pressure p2 of the jet airflow, and then according to the isentropic relationship , where γ=1.4 is the specific heat ratio, M=1 is the injection Mach number, and the total pressure before injection is P j= 1.893p 2 ;
（4）确定喷孔间距d和喷孔数目，喷孔间距d按照穿透深度h的3.33~4.66倍设置，喷孔安装数目根据喷孔作用宽度范围来确定，喷孔作用宽度范围需对飞行器开展流体力学计算，通过壁面极限流线获取进入进气道内的流体在展向的范围，在喷孔安装位置处的宽度w即为喷孔作用宽度范围，将该宽度w除以喷孔间距d，喷孔安装数目为不小于w/d的整数。(4) Determine the nozzle spacing d and the number of nozzles. The nozzle spacing d is set at 3.33 to 4.66 times the penetration depth h . The number of installed nozzles is determined based on the effective width range of the nozzles. The effective width range of the nozzles requires fluid mechanics calculations for the aircraft. The spanwise range of the fluid entering the inlet duct is obtained through the wall limiting streamlines. The width w at the nozzle installation position is the effective width range of the nozzles. The width w is divided by the nozzle spacing d. The number of installed nozzles is an integer not less than w/d.
作为优选方式，若飞行器内已有氮气罐，不需要额外的储气罐，只需从氮气罐中引入一条管路与喷孔相连，再通过电磁阀和减压阀进行控制即可。As a preferred method, if there is already a nitrogen tank in the aircraft, no additional gas tank is needed. It is only necessary to introduce a pipeline from the nitrogen tank to connect it to the spray hole, and then control it through the solenoid valve and the pressure reducing valve.
作为优选方式，若飞行器内无现有的氮气气源，则采用空气作为喷射介质，额外配备一个储气罐，储气罐的储气压力不小于1个大气压即可。As a preferred method, if there is no existing nitrogen gas source in the aircraft, air is used as the injection medium, and an additional gas tank is equipped, and the gas storage pressure of the gas tank is not less than 1 atmosphere.
作为优选方式，喷孔直径为0.5mm。As a preferred embodiment, the diameter of the nozzle hole is 0.5 mm.
通过采用上述技术方案，气体以一定喷注压力从喷孔进入压缩面边界层内，与来流相互作用，产生流向涡，流向涡诱导边界层强制转捩为湍流。By adopting the above technical solution, the gas enters the compression surface boundary layer from the nozzle at a certain injection pressure, interacts with the incoming flow, and generates a streamwise vortex, which induces the boundary layer to forcibly transform into turbulence.
本发明相比于现有被动边界层强制转捩控制技术具有如下有益效果：Compared with the existing passive boundary layer forced transition control technology, the present invention has the following beneficial effects:
一、降低强制转捩装置的热防护需求。本发明采用前体或进气道壁面喷射气流的方式实现强制转捩，避免了被动式强制转捩装置中凸出壁面的颗粒结构，也就避免了高速气流与颗粒结构的高速碰撞过程，不会额外产生气动加热，因此可大幅度降低转捩装置的热防护需求。1. Reduce the thermal protection requirements of the forced transition device. The present invention realizes forced transition by jetting airflow from the precursor or the wall of the air inlet, avoiding the particle structure protruding from the wall in the passive forced transition device, thus avoiding the high-speed collision process between the high-speed airflow and the particle structure, and no additional aerodynamic heating is generated, thus significantly reducing the thermal protection requirements of the transition device.
二、有效拓宽工作范围。强制转捩效果的好坏受粗糙颗粒高度或穿透深度与当地边界层厚度的比值影响严重，已有的被动式强制转捩装置的高度是固定不变的，当来流条件偏离设计点后，若当地边界层厚度明显增加，则强制转捩装置有可能失效，若当地边界层厚度明显减小，则因颗粒高度太高会导致粗糙颗粒与主流接触，既会引入较大气动损失，也会引起较大气动加热。本发明采用的边界层强制转捩主动喷气控制方法，在飞行器偏离设计点后，可以通过在输气管路中设置流量调节阀门，从而实现穿透深度的实时调节，结合飞行前的流体力学计算，既可以按既定时序设置喷前压力，也可以通过飞行时的传感器数据判断飞行状态，实时调节喷前压力，从而实现宽范围低损失的强制转捩控制。2. Effectively broaden the working range. The effect of forced transition is seriously affected by the ratio of the height of coarse particles or the penetration depth to the thickness of the local boundary layer. The height of the existing passive forced transition device is fixed. When the incoming flow conditions deviate from the design point, if the thickness of the local boundary layer increases significantly, the forced transition device may fail. If the thickness of the local boundary layer decreases significantly, the coarse particles will contact the mainstream due to the high particle height, which will introduce large aerodynamic losses and cause large aerodynamic heating. The boundary layer forced transition active jet control method adopted by the present invention can achieve real-time adjustment of the penetration depth by setting a flow regulating valve in the air pipeline after the aircraft deviates from the design point. Combined with the fluid mechanics calculation before flight, the pre-spray pressure can be set according to the established timing, and the flight status can be judged by the sensor data during flight, and the pre-spray pressure can be adjusted in real time, thereby realizing a wide range of low-loss forced transition control.
三、为宽域飞行的吸气式超高速飞行器提供了切实可行的边界层强制转捩控制方案。本发明可以实现边界层强制转捩的宽范围低损失控制，可以采用燃油供应系统的氮气作为气源，通过输气管路连接储气罐与喷孔，通过电磁阀进行流量控制，符合我国现有工业水平及技术基础，可以在实际飞行中使用。3. It provides a feasible boundary layer forced transition control scheme for wide-range air-breathing ultra-high-speed aircraft. The present invention can achieve wide-range low-loss control of boundary layer forced transition, can use nitrogen from the fuel supply system as the gas source, connect the gas storage tank and the spray hole through the gas pipeline, and control the flow through the solenoid valve, which is in line with my country's current industrial level and technical foundation and can be used in actual flight.
附图说明BRIEF DESCRIPTION OF THE DRAWINGS
图1为本发明多波系压缩进气道示意图。FIG1 is a schematic diagram of a multi-wave compression air inlet according to the present invention.
图2为本发明喷射气体与来流相互作用产生流向涡并逐渐失稳破碎的过程。FIG. 2 shows the process of the interaction between the injected gas and the incoming flow to generate a streamwise vortex and gradually become unstable and break up.
图3为本发明进气道压缩面上壁面热流对比及转捩位置示意图。FIG3 is a schematic diagram of the heat flux comparison and transition position on the compression surface of the air inlet duct of the present invention.
图中，1为喷孔。In the figure, 1 is a nozzle.
具体实施方式DETAILED DESCRIPTION
以下通过特定的具体实例说明本发明的实施方式，本领域技术人员可由本说明书所揭露的内容轻易地了解本发明的其他优点与功效。本发明还可以通过另外不同的具体实施方式加以实施或应用，本说明书中的各项细节也可以基于不同观点与应用，在没有背离本发明的精神下进行各种修饰或改变。The following describes the embodiments of the present invention through specific examples, and those skilled in the art can easily understand other advantages and effects of the present invention from the contents disclosed in this specification. The present invention can also be implemented or applied through other different specific embodiments, and the details in this specification can also be modified or changed in various ways based on different viewpoints and applications without departing from the spirit of the present invention.
实施例1Example 1
实施例提供一种超高速进气道边界层强制转捩主动喷气控制方法，在吸气式超高速飞行器前体或进气道压缩面上沿垂直流动方向布置一排等间距喷孔，气源经过输送管路从喷孔射出，喷射气体与来流空气相互作用，产生流向涡结构，诱导超高速边界层强制转捩为湍流；The embodiment provides an active jet control method for forced transition of a boundary layer in a hyperspeed inlet, wherein a row of equally spaced spray holes are arranged on a front body or a compression surface of an air-breathing hyperspeed aircraft in a direction perpendicular to the flow, and a gas source is ejected from the spray holes through a delivery pipeline, and the ejected gas interacts with the incoming air to generate a flow vortex structure, thereby inducing a forced transition of the hyperspeed boundary layer into turbulence;
控制方法包括确定喷孔安装位置、确定喷射介质、确定喷孔直径、确定喷前总压、确定喷孔间距和喷孔数目，包括以下步骤：The control method includes determining the installation position of the spray hole, determining the spray medium, determining the diameter of the spray hole, determining the total pressure before spraying, determining the spray hole spacing and the number of spray holes, and includes the following steps:
（1）确定喷孔安装位置，喷孔安装于吸气式超高速飞行器的前体或进气道压缩面上；(1) Determine the nozzle installation position. The nozzle is installed on the front body or the compression surface of the air intake of the air-breathing hypersonic vehicle;
对于多波系压缩进气道，喷孔安装于第一道压缩面上，For multi-wave compression inlet, the nozzle is installed on the first compression surface.
对于三维曲面压缩进气道，喷孔安装于前体上，确保由喷射气流诱导出的流向涡有足够的失稳距离；For the three-dimensional curved compression inlet, the nozzle is installed on the front body to ensure that the streamwise vortex induced by the jet flow has enough instability distance;
同时，喷孔安装位置还需考虑飞行器内有足够的空间可以放置输气管路；确定安装位置后，通过层流状态的流体力学计算，获取安装位置处的层流边界层厚度δ，此处边界层厚度δ定义为总焓边界层厚度，边界层外缘的总焓为来流总焓的99 %；At the same time, the nozzle installation position also needs to consider whether there is enough space in the aircraft to place the gas pipeline. After determining the installation position, the laminar boundary layer thickness δ at the installation position is obtained through fluid mechanics calculation in the laminar state. Here, the boundary layer thickness δ is defined as the total enthalpy boundary layer thickness. The total enthalpy at the outer edge of the boundary layer is 99% of the total enthalpy of the incoming flow.
（2）确定喷射介质，喷射介质采用空气或氮气，储气压力不小于1个大气压；(2) Determine the injection medium. The injection medium shall be air or nitrogen, and the gas storage pressure shall not be less than 1 atmosphere;
（3）确定喷孔直径0.3-1mm，确定喷前总压P j ，喷孔形状为圆形，喷孔与压缩面垂直，使得喷射气流与边界层内气流相互垂直；气流穿透深度h取为当地边界层厚度的0.5倍，气流穿透深度定义为：获取垂直喷孔中心线上的流场马赫数分布，定义沿该垂直喷孔中心线上马赫数第一次达到局部极大值时对应的点距壁面距离为穿透深度h，由于气流穿透深度满足关系式： ，其中p 2为采用声速喷射气流静压，p为无喷射时当地流体静压，参数a、b通过流体力学计算进行标定，标定获得参数a、b后，代入关系式，得到喷射气流静压p 2，再根据等熵关系式，其中γ=1.4为比热比，M=1为喷射马赫数，得喷前总压P j= 1.893p 2 ；(3) Determine the nozzle diameter of 0.3-1mm, determine the total pressure before spraying Pj , the nozzle shape is circular, the nozzle is perpendicular to the compression surface, so that the injection airflow and the airflow in the boundary layer are perpendicular to each other; the airflow penetration depth h is taken as 0.5 times the local boundary layer thickness, and the airflow penetration depth is defined as: obtain the flow field Mach number distribution on the center line of the vertical nozzle, and define the distance from the point on the center line of the vertical nozzle to the wall when the Mach number first reaches the local maximum as the penetration depth h . Since the airflow penetration depth satisfies the relationship: , where p2 is the static pressure of the jet airflow using sonic velocity, p is the static pressure of the local fluid when there is no jet, and the parameters a and b are calibrated by fluid mechanics calculation. After the parameters a and b are obtained by calibration, they are substituted into the relationship to obtain the static pressure p2 of the jet airflow, and then according to the isentropic relationship , where γ=1.4 is the specific heat ratio, M=1 is the injection Mach number, and the total pressure before injection is P j= 1.893p 2 ;
（4）确定喷孔间距d和喷孔数目，喷孔间距d按照穿透深度h的3.33~4.66倍设置，喷孔安装数目根据喷孔作用宽度范围来确定，喷孔作用宽度范围需对飞行器开展流体力学计算，通过壁面极限流线获取进入进气道内的流体在展向的范围，在喷孔安装位置处的宽度w即为喷孔作用宽度范围，将该宽度w除以喷孔间距d，喷孔安装数目为不小于w/d的整数。(4) Determine the nozzle spacing d and the number of nozzles. The nozzle spacing d is set at 3.33 to 4.66 times the penetration depth h . The number of installed nozzles is determined based on the effective width range of the nozzles. The effective width range of the nozzles requires fluid mechanics calculations for the aircraft. The spanwise range of the fluid entering the inlet duct is obtained through the wall limiting streamlines. The width w at the nozzle installation position is the effective width range of the nozzles. The width w is divided by the nozzle spacing d. The number of installed nozzles is an integer not less than w/d.
在一些实施例中，若飞行器内已有氮气罐，不需要额外的储气罐，只需从氮气罐中引入一条管路与喷孔相连，再通过电磁阀和减压阀进行控制即可。In some embodiments, if there is already a nitrogen tank in the aircraft, no additional gas tank is needed. It is only necessary to introduce a pipeline from the nitrogen tank to connect to the spray hole, and then control it through the solenoid valve and the pressure reducing valve.
在一些实施例中，若飞行器内无现有的氮气气源，则采用空气作为喷射介质，额外配备一个储气罐，储气罐的储气压力不小于1个大气压即可。In some embodiments, if there is no existing nitrogen gas source in the aircraft, air is used as the injection medium, and an additional gas tank is provided, and the gas storage pressure of the gas tank is not less than 1 atmosphere.
在一些实施例中，喷孔直径为0.5mm。In some embodiments, the nozzle hole diameter is 0.5 mm.
实施例2Example 2
为验证本发明中提出的超高速边界层强制转捩主动喷气控制方法的有效性，针对一款4波系压缩的进气道构型（参见图1）设计了一套强制转捩方案，并对该方案进行了数值模拟。In order to verify the effectiveness of the ultra-high-speed boundary layer forced transition active jet control method proposed in the present invention, a forced transition scheme was designed for a 4-wave compression inlet configuration (see Figure 1), and a numerical simulation of the scheme was carried out.
进气道模型长550mm、宽225.7mm、高113mm。进气道的设计马赫数范围为4～7，攻角-1～3度，偏航角低于2度。起动马赫数为4。设计状态为：马赫数6，攻角1°，偏航角0°，来流总压2Mpa，来流总温455K，来流单位雷诺数为18.9*106/m，来流为纯空气，边界层强制转捩主动喷气控制方案确定步骤如下：The inlet model is 550mm long, 225.7mm wide and 113mm high. The design Mach number range of the inlet is 4 to 7, the angle of attack is -1 to 3 degrees, and the yaw angle is less than 2 degrees. The starting Mach number is 4. The design state is: Mach number 6, angle of attack 1°, yaw angle 0°, total pressure of incoming flow 2Mpa, total temperature of incoming flow 455K, unit Reynolds number of incoming flow is 18.9*10 6 /m, the incoming flow is pure air, and the boundary layer forced transition active jet control scheme is determined as follows:
（1）确定喷孔安装位置，该进气道为多波系压缩进气道，喷孔安装位置位于第一道压缩面上，综合考虑机体内部管路布置需求，确定喷孔中心距离进气道前缘88mm。采用层流状态的流体力学计算，获得距前缘88mm处层流边界层厚度为δ=1.2mm，此处静压约为2850Pa。(1) Determine the installation position of the nozzle. The inlet is a multi-wave compression inlet. The nozzle installation position is located on the first compression surface. Considering the internal pipeline layout requirements of the fuselage, the center of the nozzle is determined to be 88mm away from the leading edge of the inlet. Using laminar fluid mechanics calculations, the thickness of the laminar boundary layer at 88mm from the leading edge is δ=1.2mm, and the static pressure here is about 2850Pa.
（2）确定喷射介质，此构型仅为进气道构型，无燃油供应系统，喷射介质选为空气。(2) Determine the injection medium. This configuration is only an intake duct configuration, without a fuel supply system, and the injection medium is air.
（3）确定喷孔直径0.3-1mm，确定喷前总压P j ，喷孔形状为圆形，喷孔与压缩面垂直。喷孔直径为0.5mm。由于气流穿透深度满足关系式： ，其中p 2为采用声速喷射气流静压，p为无喷射时当地流体静压，参数a、b通过流体力学计算进行标定，标定获得参数a、b后，代入关系式，得到喷射气流静压p 2，再根据等熵关系式，其中γ=1.4为比热比，M=1为喷射马赫数，得喷前总压P j= 1.893p 2 ；(3) Determine the nozzle diameter of 0.3-1mm, determine the total pressure before spraying Pj , the nozzle shape is circular, and the nozzle is perpendicular to the compression surface. The nozzle diameter is 0.5mm. Since the air flow penetration depth satisfies the relationship: , where p2 is the static pressure of the jet airflow using sonic velocity, p is the static pressure of the local fluid when there is no jet, and the parameters a and b are calibrated by fluid mechanics calculation. After the parameters a and b are obtained by calibration, they are substituted into the relationship to obtain the static pressure p2 of the jet airflow, and then according to the isentropic relationship , where γ=1.4 is the specific heat ratio, M=1 is the injection Mach number, and the total pressure before injection is P j= 1.893p 2 ;
采用计算流体力学计算三个不同喷射静压下的工况，统计其穿透深度，获得式子,根据该式子，设置=3.0，对应的穿透深度为0.6mm，此时的喷前压力。Computational fluid dynamics was used to calculate the working conditions under three different injection static pressures, and the penetration depth was calculated to obtain the formula , according to this formula, set =3.0, the corresponding penetration depth is 0.6mm, and the pre-spray pressure is .
（4）确定喷孔间距d和喷孔数目，喷孔间距d按照穿透深度h的3.33~4.66倍设置，喷孔安装数目根据喷孔作用宽度范围来确定，喷孔作用宽度范围需对飞行器开展流体力学计算，通过壁面极限流线获取进入进气道内的流体在展向的范围，在喷孔安装位置处的宽度w即为喷孔作用宽度范围，将该宽度w除以喷孔间距d，喷孔安装数目为不小于w/d的整数。(4) Determine the nozzle spacing d and the number of nozzles. The nozzle spacing d is set at 3.33 to 4.66 times the penetration depth h . The number of installed nozzles is determined based on the effective width range of the nozzles. The effective width range of the nozzles requires fluid mechanics calculations for the aircraft. The spanwise range of the fluid entering the inlet duct is obtained through the wall limiting streamlines. The width w at the nozzle installation position is the effective width range of the nozzles. The width w is divided by the nozzle spacing d. The number of installed nozzles is an integer not less than w/d.
喷孔间距选为2.5mm，则d/h≈4.2，属于本发明要求的范围内。根据计算流体力学获得的壁面极限流线，得到喷孔作用范围为中心线两侧36mm内，因此作用范围宽度一共为72mm，喷孔间距2.5mm，一共需29个喷孔。The nozzle spacing is selected as 2.5mm, then d/h≈4.2, which is within the range required by the present invention. According to the wall limit streamline obtained by computational fluid dynamics, the nozzle action range is obtained to be within 36mm on both sides of the center line, so the total width of the action range is 72mm, and the nozzle spacing is 2.5mm, so a total of 29 nozzles are required.
采用数值模拟获得了主动喷气下游的流场结构，转捩位置及壁面热流情况，为了简化计算，只对包含三个喷孔的片式模型进行模拟，两侧采用周期性边界条件，此时进气道模型宽度仅为7.5mm。图2展示了喷射气流与来流相互作用产生的流向涡结构，及流向涡结构在下游逐渐失稳、破碎、最终转捩的整个过程。图3展示了壁面热流分布，从壁面热流与全层流计算结果、全湍流计算结果和无控制措施计算结果对比可以看出，本发明得到的强制转捩方法有效促进了边界层转捩位置由第二道压缩拐角下游前移至第一道压缩拐角附近，表明本发明的有效性。Numerical simulation was used to obtain the flow field structure, transition position and wall heat flux downstream of the active jet. In order to simplify the calculation, only the sheet model containing three nozzles was simulated, and periodic boundary conditions were used on both sides. At this time, the width of the inlet model was only 7.5 mm. Figure 2 shows the flow vortex structure generated by the interaction between the jet flow and the incoming flow, and the entire process of the flow vortex structure gradually becoming unstable, breaking, and finally transitioning downstream. Figure 3 shows the wall heat flux distribution. From the comparison of the wall heat flux with the full laminar flow calculation results, the full turbulent flow calculation results and the calculation results without control measures, it can be seen that the forced transition method obtained by the present invention effectively promotes the boundary layer transition position to move forward from the downstream of the second compression corner to near the first compression corner, indicating the effectiveness of the present invention.
上述实施例仅例示性说明本发明的原理及其功效，而非用于限制本发明。任何熟悉此技术的人士皆可在不违背本发明的精神及范畴下，对上述实施例进行修饰或改变。因此，凡所属技术领域中具有通常知识者在未脱离本发明所揭示的精神与技术思想下所完成的一切等效修饰或改变，仍应由本发明的权利要求所涵盖。The above embodiments are merely illustrative of the principles and effects of the present invention, and are not intended to limit the present invention. Anyone familiar with the art may modify or alter the above embodiments without departing from the spirit and scope of the present invention. Therefore, all equivalent modifications or alterations made by a person of ordinary skill in the art without departing from the spirit and technical ideas disclosed by the present invention shall still be covered by the claims of the present invention.
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