CN108182319A - A kind of supersonic speed integration Nozzle Design method - Google Patents
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Abstract
一种超声速一体化喷管设计方法，涉及风洞试验领域；包括如下步骤：步骤(一)、建立喷管模型,包括收缩段、膨胀段、试验段和边界层；步骤(二)、根据TG段曲线的特征线方程求得TG段曲线；根据AD段曲线的特征线方程求得AD段曲线,得到膨胀段外壁的完整曲线；步骤(三)、计算边界层的位移厚度；得到边界层曲线；步骤(四)、获得收缩段的完整曲线；步骤(五)、获得试验段的完整曲线；本发明使得试验区域增大，且试验段马赫数均方根偏差、轴向马赫数梯度和气流偏转角等参数满足国军标先进指标。
A supersonic integrated nozzle design method, related to the field of wind tunnel testing; including the following steps: step (1), establish a nozzle model, including contraction section, expansion section, test section and boundary layer; step (2), according to TG The characteristic line equation of the section curve obtains the TG section curve; obtains the AD section curve according to the characteristic line equation of the AD section curve, and obtains the complete curve of the outer wall of the expansion section; step (3), calculates the displacement thickness of the boundary layer; obtains the boundary layer curve Step (4), obtain the complete curve of contraction section; Step (5), obtain the complete curve of test section; The present invention makes test area increase, and test section Mach number root mean square deviation, axial Mach number gradient and airflow Parameters such as deflection angle meet the advanced indicators of the national military standard.
Description
技术领域technical field
本发明涉及一种风洞试验领域，特别是一种超声速一体化喷管设计方法。The invention relates to the field of wind tunnel tests, in particular to a supersonic integrated nozzle design method.
背景技术Background technique
目前在超声速风洞试验中，由于喷管出口马赫角较小，导致有效的试验均匀区面积较小，为了避免试验模型暴露在试验均匀区外，只能采用较小试验模型。较小的试验模型在风洞中获得的试验数据与实际飞行器模型所获得试验数据存在一定的偏差，影响试验精度。同时，较小的试验均匀区，限制模型连续变换姿态角，提高风洞运行成本。At present, in the supersonic wind tunnel test, due to the small Mach angle at the exit of the nozzle, the effective test uniform area is small. In order to avoid the test model being exposed outside the test uniform area, only a small test model can be used. There is a certain deviation between the test data obtained in the wind tunnel of the small test model and the test data obtained by the actual aircraft model, which affects the test accuracy. At the same time, the small uniform area of the test limits the continuous change of the attitude angle of the model and increases the operating cost of the wind tunnel.
发明内容Contents of the invention
本发明的目的在于克服现有技术的上述不足，提供一种超声速一体化喷管设计方法，使得试验区域增大，且试验段马赫数均方根偏差、轴向马赫数梯度和气流偏转角等参数满足国军标先进指标。The purpose of the present invention is to overcome the above-mentioned deficiencies in the prior art, and provide a supersonic integrated nozzle design method, so that the test area is enlarged, and the Mach number root mean square deviation of the test section, the axial Mach number gradient and the airflow deflection angle, etc. The parameters meet the advanced indicators of the national military standard.
本发明的上述目的是通过如下技术方案予以实现的：Above-mentioned purpose of the present invention is achieved by following technical scheme:
一种超声速一体化喷管设计方法，包括如下步骤：A supersonic integrated nozzle design method, comprising the steps of:
步骤(一)、建立喷管模型；包括收缩段、膨胀段、试验段和边界层；其中，收缩段、膨胀段和试验段沿轴向依次首尾相连；边界层包覆在膨胀段和试验段的外壁；收缩段与膨胀段的连接处为喉道；喉道对应喷管外壁的位置为T点；膨胀段和试验段的连接处对应喷管外壁的位置为D点；收缩段轴向远离膨胀段的一端为入口端；膨胀段轴向远离收缩段的一端为出口端；Step (1), establishing a nozzle model; including a contraction section, an expansion section, a test section and a boundary layer; wherein, the contraction section, the expansion section and the test section are connected end-to-end along the axial direction; the boundary layer covers the expansion section and the test section The outer wall of the contraction section and the expansion section is the throat; the position of the throat corresponding to the outer wall of the nozzle is point T; the connection between the expansion section and the test section is the point D corresponding to the outer wall of the nozzle; the axial distance of the contraction section is One end of the expansion section is the inlet end; the end of the expansion section axially away from the contraction section is the outlet end;
步骤(二)、预先设定膨胀段的最大膨胀角θ、马赫数M和膨胀段的出口端直径；选取膨胀段中最大膨胀角θ对应点，将该点定义为G点；以G点为起始点，沿指向试验段的方向在膨胀段的外壁取点A；GA段为圆锥曲线段；建立坐标系oxy；根据坐标系oxy分别建立AD段曲线的特征线方程和TG段曲线的特征线方程；根据TG段曲线的特征线方程求得TG段曲线；根据AD段曲线的特征线方程求得AD段曲线；Step (2), preset the maximum expansion angle θ of the expansion section, the Mach number M and the outlet diameter of the expansion section; select the point corresponding to the maximum expansion angle θ in the expansion section, and define this point as point G; Starting point, take point A on the outer wall of the expansion section along the direction pointing to the test section; the GA section is a conic section; establish the coordinate system oxy; according to the coordinate system oxy, respectively establish the characteristic line equation of the curve of the AD section and the characteristic line of the curve of the TG section Equation; obtain the TG section curve according to the characteristic line equation of the TG section curve; obtain the AD section curve according to the characteristic line equation of the AD section curve;
将TG段曲线的G点和AD段曲线的A点直线连接；得到膨胀段外壁的完整曲线；Connect the point G of the curve of the TG section with the point A of the curve of the AD section in a straight line; obtain the complete curve of the outer wall of the expansion section;
步骤(三)、建立卡门动量积分方程计算边界层的动量厚度δ′；并根据动量厚度δ′计算边界层的位移厚度δ；即为边界层曲线；Step (3), establish the Karman momentum integral equation to calculate the momentum thickness δ' of the boundary layer; and calculate the displacement thickness δ of the boundary layer according to the momentum thickness δ'; it is the boundary layer curve;
步骤(四)、根据步骤(二)中的TG段曲线的特征线方程，得到T点下游的斜率；另T点下游的斜率与T点上游的斜率相同；得到收缩段的完整曲线；Step (4), according to the characteristic line equation of the TG section curve in step (2), obtain the slope of the downstream of the T point; Another slope of the downstream of the T point is identical with the slope of the upstream of the T point; obtain the complete curve of the contraction section;
步骤(五)、预先设定欠膨胀修正因子λ；试验段的外壁曲线为直线段；根据步骤(二)中AD段曲线的特征线方程，得到D点上游的斜率；试验段的斜率为D点上游的斜率乘以欠膨胀修正因子λ；得到试验段的完整曲线。Step (5), pre-setting under-expansion correction factor λ; the outer wall curve of the test section is a straight line segment; according to the characteristic line equation of the AD section curve in step (2), the slope upstream of point D is obtained; the slope of the test section is D The slope upstream of the point is multiplied by the underexpansion correction factor λ; the complete curve for the test section is obtained.
在上述的一种超声速一体化喷管设计方法，所述的步骤(二)中，坐标系oxy的建立方法为：In above-mentioned a kind of supersonic integrated nozzle design method, in described step (two), the establishment method of coordinate system oxy is:
以T点为坐标原点，x轴正方向为沿轴向指向出口端方向，y轴正方形为竖直向上方向。Take point T as the coordinate origin, the positive direction of the x-axis is the direction along the axial direction to the outlet end, and the square of the y-axis is the vertical upward direction.
在上述的一种超声速一体化喷管设计方法，所述的步骤(二)中，AD段曲线的特征线方程为：In above-mentioned a kind of supersonic integrated nozzle design method, in described step (two), the characteristic line equation of AD section curve is:
TG段曲线的特征线方程为：The characteristic line equation of the TG segment curve is:
式中，x为坐标系oxy的横坐标；In the formula, x is the abscissa of the coordinate system oxy;
y为坐标系oxy的纵坐标；y is the ordinate of the coordinate system oxy;
γ喷管中的气体比热比。Gas specific heat ratio in the gamma nozzle.
在上述的一种超声速一体化喷管设计方法，所述的步骤(三)中，通过卡门动量积分方程计算边界层的动量厚度δ′的方法为：In the above-mentioned a kind of supersonic integrated nozzle design method, in the described step (3), the method for calculating the momentum thickness δ' of the boundary layer by the Karman momentum integral equation is:
其中，H为附面层形状因子；Among them, H is the boundary layer shape factor;
Cf为摩擦系数。C f is the coefficient of friction.
在上述的一种超声速一体化喷管设计方法，所述的步骤(三)中，位移厚度δ的计算方法为：In the above-mentioned a kind of supersonic integrated nozzle design method, in the described step (3), the calculation method of the displacement thickness δ is:
δ＝Hδ′ (4)。δ=Hδ' (4).
本发明与现有技术相比具有如下优点：Compared with the prior art, the present invention has the following advantages:
(1)本发明采用步骤1特征线方法设计的无粘型线，无粘型线斜率连续，气流在喷管内流动时流动不会产生波系；(1) The present invention adopts the non-viscous profile line designed by the characteristic line method of step 1, the non-viscosity profile line slope is continuous, and the flow of the air flow will not produce a wave system when flowing in the nozzle pipe;
(2)本发明采用步骤2的边界层修正方法，符合气流的流动规律，喷管出口的流场品质优良；(2) The present invention adopts the boundary layer correction method of step 2, which conforms to the flow law of the air flow, and the flow field at the outlet of the nozzle is of good quality;
(3)本发明采用步骤3的连接方式，使得喷管收缩段和膨胀段斜率连续，气流流过喉道T点时，流动不会产生分离；(3) The present invention adopts the connection method of step 3, so that the slope of the nozzle shrinkage section and the expansion section is continuous, and when the air flow passes through the throat T point, the flow will not be separated;
(4)本发明采用步骤4和步骤5的优化方法，使得喷管出口DQ没有明显的膨胀波系和激波，使得试验区域由CDF增大到CDSN。试验段马赫数均方根偏差、轴向马赫数梯度和气流偏转角等参数满足国军标先进指标。(4) The present invention adopts the optimization method of step 4 and step 5, so that the nozzle outlet DQ has no obvious expansion wave system and shock wave, so that the test area is increased from CDF to CDSN. Parameters such as Mach number root mean square deviation, axial Mach number gradient and airflow deflection angle in the test section meet the advanced indicators of the national military standard.
附图说明Description of drawings
图1为本发明超声速一体化喷管示意图；Fig. 1 is the schematic diagram of supersonic integrated nozzle of the present invention;
图2为本发明超声速一体化喷管设计流程示意图。Fig. 2 is a schematic diagram of the design process of the supersonic integrated nozzle of the present invention.
具体实施方式Detailed ways
下面结合附图和具体实施例对本发明作进一步详细的描述：Below in conjunction with accompanying drawing and specific embodiment the present invention is described in further detail:
如图2所示为超声速一体化喷管设计流程示意图，由图可知，一种超声速一体化喷管设计方法，包括如下步骤：Figure 2 is a schematic diagram of the supersonic integrated nozzle design process. It can be seen from the figure that a supersonic integrated nozzle design method includes the following steps:
步骤(一)、建立喷管模型；包括收缩段1、膨胀段2、试验段3和边界层4；其中，收缩段1、膨胀段2和试验段3沿轴向依次首尾相连；边界层4包覆在膨胀段2和试验段3的外壁；收缩段1与膨胀段2的连接处为喉道；如图1所示为超声速一体化喷管示意图，由图可知，喉道对应喷管外壁的位置为T点；膨胀段2和试验段3的连接处对应喷管外壁的位置为D点；收缩段1轴向远离膨胀段2的一端为入口端；膨胀段2轴向远离收缩段1的一端为出口端；Step (1), establishing the nozzle model; including contraction section 1, expansion section 2, test section 3 and boundary layer 4; wherein, contraction section 1, expansion section 2 and test section 3 are connected end-to-end in the axial direction; boundary layer 4 Coated on the outer wall of the expansion section 2 and the test section 3; the connection between the contraction section 1 and the expansion section 2 is the throat; as shown in Figure 1 is a schematic diagram of the supersonic integrated nozzle, as can be seen from the figure, the throat corresponds to the outer wall of the nozzle The position of the expansion section 2 and the test section 3 corresponding to the outer wall of the nozzle is point D; the end of the contraction section 1 axially away from the expansion section 2 is the inlet end; the expansion section 2 is axially away from the contraction section 1 One end of is the outlet port;
步骤(二)、预先设定膨胀段2的最大膨胀角θ、马赫数M和膨胀段2的出口端直径；选取膨胀段2中最大膨胀角θ对应点，将该点定义为G点；以G点为起始点，沿指向试验段3的方向在膨胀段2的外壁取点A；GA段为圆锥曲线段；建立坐标系oxy；其中，坐标系oxy的建立方法为：Step (2), preset the maximum expansion angle θ of the expansion section 2, the Mach number M and the outlet end diameter of the expansion section 2; select the point corresponding to the maximum expansion angle θ in the expansion section 2, and define this point as point G; Point G is the starting point, and point A is taken on the outer wall of the expansion section 2 along the direction pointing to the test section 3; section GA is a conic section; coordinate system oxy is established; wherein, the establishment method of coordinate system oxy is:
以T点为坐标原点，x轴正方向为沿轴向指向出口端方向，y轴正方形为竖直向上方向。Take point T as the coordinate origin, the positive direction of the x-axis is the direction along the axial direction to the outlet end, and the square of the y-axis is the vertical upward direction.
根据坐标系oxy分别建立AD段曲线的特征线方程和TG段曲线的特征线方程；根据TG段曲线的特征线方程求得TG段曲线；根据AD段曲线的特征线方程求得AD段曲线；According to the coordinate system oxy, respectively establish the characteristic line equation of the AD section curve and the characteristic line equation of the TG section curve; obtain the TG section curve according to the characteristic line equation of the TG section curve; obtain the AD section curve according to the characteristic line equation of the AD section curve;
AD段曲线的特征线方程为：The characteristic line equation of the AD segment curve is:
TG段曲线的特征线方程为：The characteristic line equation of the TG segment curve is:
式中，x为坐标系oxy的横坐标；In the formula, x is the abscissa of the coordinate system oxy;
y为坐标系oxy的纵坐标；y is the ordinate of the coordinate system oxy;
γ喷管中的气体比热比。Gas specific heat ratio in the gamma nozzle.
将TG段曲线的G点和AD段曲线的A点直线连接；得到膨胀段2外壁的完整曲线；Connect the point G of the curve of the TG section with the point A of the curve of the AD section in a straight line; obtain the complete curve of the outer wall of the expansion section 2;
步骤(三)、建立卡门动量积分方程计算边界层4的动量厚度δ′；并根据动量厚度δ′计算边界层4的位移厚度δ；即为边界层4曲线；Step (3), establish the Karman momentum integral equation to calculate the momentum thickness δ' of the boundary layer 4; and calculate the displacement thickness δ of the boundary layer 4 according to the momentum thickness δ'; it is the boundary layer 4 curve;
通过卡门动量积分方程计算边界层(4)的动量厚度δ′的方法为：The method to calculate the momentum thickness δ′ of the boundary layer (4) through the Karman momentum integral equation is:
其中，H为附面层形状因子；Among them, H is the boundary layer shape factor;
Cf为摩擦系数。C f is the coefficient of friction.
位移厚度δ的计算方法为：The calculation method of displacement thickness δ is:
δ＝Hδ′ (4)。δ=Hδ' (4).
步骤(四)、根据步骤(二)中的TG段曲线的特征线方程，得到T点下游的斜率；另T点下游的斜率与T点上游的斜率相同；得到收缩段1的完整曲线；Step (four), according to the characteristic line equation of the TG section curve in step (two), obtain the slope of the downstream of the T point; Another slope of the downstream of the T point is identical with the slope of the upstream of the T point; obtain the complete curve of the contraction section 1;
步骤(五)、预先设定欠膨胀修正因子λ；试验段3的外壁曲线为直线段；根据步骤(二)中AD段曲线的特征线方程，得到D点上游的斜率；试验段3的斜率为D点上游的斜率乘以欠膨胀修正因子λ；得到试验段3的完整曲线。对修正因子λ进行优化，使得喷管出口D和试验段DS无明显的膨胀波和激波，完成超声速一体化喷管的型线设计。Step (5), pre-setting the under-expansion correction factor λ; the outer wall curve of the test section 3 is a straight line segment; according to the characteristic line equation of the AD section curve in the step (2), the slope upstream of the D point is obtained; the slope of the test section 3 is the slope upstream of point D multiplied by the underexpansion correction factor λ; the complete curve for test section 3 is obtained. The correction factor λ is optimized so that there are no obvious expansion waves and shock waves at the nozzle outlet D and the test section DS, and the profile design of the supersonic integrated nozzle is completed.
在喷管轴线处取点C，CD等于膨胀段2出口端半径除以sinβ；β＝arcsin(1/M)；F与C点沿膨胀段2出口径向对称；试验段3的尾端沿径向分别取S、N两个点；本发明将试验区域由CDF增大到CDSN。试验段马赫数均方根偏差、轴向马赫数梯度和气流偏转角等参数满足国军标先进指标。Take point C at the nozzle axis, CD is equal to the radius of the outlet end of expansion section 2 divided by sinβ; β=arcsin (1/M); F and C points are radially symmetrical along the outlet of expansion section 2; Two points S and N are respectively taken in the radial direction; the present invention increases the test area from CDF to CDSN. Parameters such as Mach number root mean square deviation, axial Mach number gradient and airflow deflection angle in the test section meet the advanced indicators of the national military standard.
本发明说明书中未作详细描述的内容属本领域技术人员的公知技术。The content that is not described in detail in the description of the present invention belongs to the well-known technology of those skilled in the art.
Claims (5)
- A kind of 1. supersonic speed integration Nozzle Design method, it is characterised in that：Include the following steps：Step (1) establishes jet pipe model；Including contraction section (1), expansion arc (2), test section (3) and boundary layer (4)；Wherein, Contraction section (1), expansion arc (2) and test section (3) join end to end successively in an axial direction；Boundary layer (4) is coated on expansion arc (2) and examination Test the outer wall of section (3)；The junction of contraction section (1) and expansion arc (2) is venturi；The position that venturi corresponds to jet pipe outer wall is T points； The position that expansion arc (2) corresponds to jet pipe outer wall with the junction of test section (3) is D points；Contraction section (1) is axially away from expansion arc (2) one end is arrival end；Expansion arc (2) is the port of export axially away from one end of contraction section (1)；The port of export diameter of step (2), the maximum swelling angle θ, the Mach number M that preset expansion arc (2) and expansion arc (2)；Choosing Maximum swelling angle θ corresponding points in expansion arc (2) are taken, which is defined as G points；Using G points as starting point, along point test section (3) Direction take point A in the outer wall of expansion arc (2)；GA sections are conic line segment；Establish coordinate system oxy；Distinguished according to coordinate system oxy Establish the characteristic strips equation of AD sections of curves and the characteristic strips equation of TG sections of curves；TG is acquired according to the characteristic strips equation of TG sections of curves Section curve；AD sections of curves are acquired according to the characteristic strips equation of AD sections of curves；The G points of TG sections of curves are connected with the A point straight lines of AD sections of curves；Obtain the complete curve of expansion arc (2) outer wall；Step (3) establishes the momentum thickness δ ' that toll bar momentum integral equation calculates boundary layer (4)；And according to momentum thickness δ ' meters Calculate the displacement thickness δ of boundary layer (4)；As boundary layer (4) curve；The characteristic strips equation of step (4), TG section curves in step (2), obtains the slope in T points downstream；Another T points downstream Slope it is identical with the slope of T points upstream；Obtain the complete curve of contraction section (1)；Step (5) presets and owes expansion modifying factor λ；The outer wall curve of test section (3) is straightway；According to step (2) The characteristic strips equation of middle AD sections of curve obtains the slope of D points upstream；The slope of test section (3) is multiplied by deficient for the slope of D points upstream Expand modifying factor λ；Obtain the complete curve of test section (3).
- 2. a kind of supersonic speed integration Nozzle Design method according to claim 1, it is characterised in that：The step (2) in, the method for building up of coordinate system oxy is：Using T points as coordinate origin, positive direction of the x-axis is direction straight up to be axially directed to outlet extreme direction, y-axis square.
- 3. a kind of supersonic speed integration Nozzle Design method according to claim 2, it is characterised in that：The step (2) in, the characteristic strips equation of AD sections of curves is：The characteristic strips equation of TG sections of curves is：In formula, x is the abscissa of coordinate system oxy；Y is the ordinate of coordinate system oxy；Specific heats of gases ratio in γ jet pipes.
- 4. a kind of supersonic speed integration Nozzle Design method according to claim 1, it is characterised in that：The step (3) in, the method that the momentum thickness δ ' of boundary layer (4) is calculated by toll bar momentum integral equation is：Wherein, H is boundary-layer form factor；CfFor friction coefficient.
- 5. a kind of supersonic speed integration Nozzle Design method according to claim 1, it is characterised in that：The step (3) in, the computational methods of displacement thickness δ are：δ=H δ ' (4).
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| CN110207934A (en) * | 2019-05-28 | 2019-09-06 | 中国航天空气动力技术研究院 | Effectively extend the method for the high enthalpy impulse wind tunnel flowing time of large scale freedom piston |
| CN111220341A (en) * | 2020-01-21 | 2020-06-02 | 中国空气动力研究与发展中心超高速空气动力研究所 | Design method of wind tunnel high-Mach-number low-Reynolds-number axisymmetric profile spray pipe |
| CN113946904A (en) * | 2021-08-31 | 2022-01-18 | 中国航天空气动力技术研究院 | Design method of large-size low-noise spray pipe |
| CN114969627A (en) * | 2022-05-08 | 2022-08-30 | 中机新材料研究院(郑州)有限公司 | Method for designing molded surface of supersonic circumferential seam spray pipe for gas atomization powder preparation |
| CN115048730A (en) * | 2022-08-15 | 2022-09-13 | 中国人民解放军国防科技大学 | Axial symmetry supersonic velocity spray pipe optimization design method and device based on grid displacement |
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| CN113946904B (en) * | 2021-08-31 | 2024-06-11 | 中国航天空气动力技术研究院 | Design method of large-size low-noise spray pipe |
| CN114969627A (en) * | 2022-05-08 | 2022-08-30 | 中机新材料研究院(郑州)有限公司 | Method for designing molded surface of supersonic circumferential seam spray pipe for gas atomization powder preparation |
| CN115389155A (en) * | 2022-07-29 | 2022-11-25 | 中国航天空气动力技术研究院 | A design method for hypersonic gas-liquid-solid multiphase flow surface nozzle |
| CN115389155B (en) * | 2022-07-29 | 2024-08-06 | 中国航天空气动力技术研究院 | A design method for a hypersonic gas-liquid-solid multiphase flow profile nozzle |
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| CN115048730B (en) * | 2022-08-15 | 2022-10-21 | 中国人民解放军国防科技大学 | Axial symmetry supersonic velocity spray pipe optimization design method and device based on grid displacement |
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