CN113987694A - A prediction method of flow field parameter distribution of rotary detonation engine based on space propulsion algorithm - Google Patents

A prediction method of flow field parameter distribution of rotary detonation engine based on space propulsion algorithm Download PDF

Info

Publication number
CN113987694A
CN113987694A CN202111092248.1A CN202111092248A CN113987694A CN 113987694 A CN113987694 A CN 113987694A CN 202111092248 A CN202111092248 A CN 202111092248A CN 113987694 A CN113987694 A CN 113987694A
Authority
CN
China
Prior art keywords
detonation
point
representing
flow
detonation wave
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Pending
Application number
CN202111092248.1A
Other languages
Chinese (zh)
Inventor
李锐
徐惊雷
黄帅
陈匡世
俞凯凯
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Nanjing University of Aeronautics and Astronautics
Original Assignee
Nanjing University of Aeronautics and Astronautics
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Nanjing University of Aeronautics and Astronautics filed Critical Nanjing University of Aeronautics and Astronautics
Priority to CN202111092248.1A priority Critical patent/CN113987694A/en
Publication of CN113987694A publication Critical patent/CN113987694A/en
Pending legal-status Critical Current

Links

Images

Classifications

    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/17—Mechanical parametric or variational design
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/20—Design optimisation, verification or simulation
    • G06F30/28—Design optimisation, verification or simulation using fluid dynamics, e.g. using Navier-Stokes equations or computational fluid dynamics [CFD]
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
    • G06F2119/08—Thermal analysis or thermal optimisation

Landscapes

  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Theoretical Computer Science (AREA)
  • General Physics & Mathematics (AREA)
  • Geometry (AREA)
  • Evolutionary Computation (AREA)
  • General Engineering & Computer Science (AREA)
  • Mathematical Analysis (AREA)
  • Mathematical Optimization (AREA)
  • Pure & Applied Mathematics (AREA)
  • Computer Hardware Design (AREA)
  • Fluid Mechanics (AREA)
  • Mathematical Physics (AREA)
  • Computing Systems (AREA)
  • Algebra (AREA)
  • Computational Mathematics (AREA)
  • Combined Controls Of Internal Combustion Engines (AREA)

Abstract

本发明公开了一种基于空间推进算法的旋转爆震发动机流场参数分布预测方法，包括步骤：预估爆震波波前流动参数；根据单一比热比的CJ爆震模型，确定爆震波后流动参数；将爆震波后流动参数和远下游流动参数赋值给初值线；通过空间推进算法确定激波坐标系下的流动参数；确定可燃混合气喷注起始位置和终了位置；迭代收敛旋转爆震发动机流场参数，并通过坐标转换获得实验室坐标系下的流动参数。本发明通过耦合空间推进算法和CJ爆震波模型，通过坐标转换获得实验室坐标系下的流动参数，能够快速获取旋转爆震发动机流场结构和出口流动参数分布，克服了现有算法获取旋转爆震发动机流场耗时长等缺点，降低了旋转爆震发动机构型设计周期。

Figure 202111092248

The invention discloses a method for predicting the distribution of flow field parameters of a rotary detonation engine based on a space propulsion algorithm. parameters; assign the flow parameters after the detonation wave and the flow parameters far downstream to the initial value line; determine the flow parameters in the shock wave coordinate system through the space propulsion algorithm; determine the starting position and end position of the combustible mixture injection; iteratively convergent rotary explosion The flow field parameters of the vibration engine were obtained, and the flow parameters in the laboratory coordinate system were obtained through coordinate transformation. By coupling the space propulsion algorithm and the CJ detonation wave model, the invention obtains the flow parameters in the laboratory coordinate system through coordinate transformation, can quickly obtain the flow field structure of the rotary detonation engine and the distribution of the flow parameters at the outlet, and overcomes the problem of obtaining the rotary detonation by the existing algorithm. The shortcoming of detonation engine flow field, such as time-consuming, reduces the design cycle of rotary detonation engine.

Figure 202111092248

Description

一种基于空间推进算法的旋转爆震发动机流场参数分布预测 方法A method for predicting flow field parameter distribution of rotary detonation engine based on space propulsion algorithm

技术领域technical field

本发明属于旋转爆震发动机流场数值计算领域，具体涉及一种基于空间推进算法的旋转爆震发动机流场参数分布预测方法。The invention belongs to the field of numerical calculation of the flow field of a rotary detonation engine, in particular to a method for predicting the distribution of flow field parameters of a rotary detonation engine based on a space propulsion algorithm.

背景技术Background technique

旋转爆震发动机作为一种新型推进装置，如图2所示，常见的结构为同心环柱，环柱底端开有可燃混合气喷注槽口。环腔内存在单个或多个爆震波朝喷注平面倾斜，且沿周向传播。紧邻爆震波后的气体经过迅速释热，温度和压力得到显著提升，而后一系列膨胀波使气流膨胀加速。当靠近喷注平面的燃烧产物的压力低于可燃混合气的喷注总压时，可燃混合气得以注入环腔。爆震波和可燃混合气/燃烧产物间断面，围成一个充满可燃混合气三角区，确保爆震燃烧的可持续性。在流场结构上，爆震波、剪切层以及诱导激波相交于三角区上顶点。大部分气流沿着旋转爆震发动机出口轴向排出，但流动参数沿周向的变化仍显著。Rotary detonation engine is a new type of propulsion device, as shown in Figure 2, the common structure is a concentric ring column, and the bottom end of the ring column is provided with a combustible mixture injection slot. There are single or multiple detonation waves in the annular cavity that are inclined towards the injection plane and propagate in the circumferential direction. The gas immediately after the detonation wave undergoes rapid heat release, the temperature and pressure are significantly increased, and then a series of expansion waves accelerate the expansion of the gas flow. When the pressure of the combustion products near the injection plane is lower than the total injection pressure of the combustible mixture, the combustible mixture is injected into the annular cavity. The detonation wave and the combustible mixture/combustion product discontinuity form a triangle area full of combustible mixture to ensure the sustainability of detonation combustion. In the flow field structure, the detonation wave, the shear layer and the induced shock wave intersect at the upper vertex of the triangle area. Most of the airflow is discharged axially along the outlet of the rotary detonation engine, but the variation of flow parameters along the circumferential direction is still significant.

相较于燃气涡轮发动机和冲压发动机，旋转爆震发动机在工作中熵增较小，热效率更高，受到各研究机构的广泛关注。为了获取旋转爆震发动机流场的完整信息，高精度数值仿真技术必不可少。爆震燃烧过程中运动激波与化学放热反应强烈耦合，导致传统的基于时间推进的数值模拟网格尺度和离散时间步过小。为了获取旋转爆震发动机的气动性能，往往耗费较多的时间和计算资源。Compared with gas turbine engines and ramjets, rotary detonation engines have less entropy increase and higher thermal efficiency during operation, and have received extensive attention from various research institutions. In order to obtain complete information on the flow field of a rotary detonation engine, high-precision numerical simulation technology is essential. The kinematic shock wave is strongly coupled with the chemical exothermic reaction in the process of detonation combustion, which leads to too small grid scale and discrete time step for traditional numerical simulation based on time advancement. In order to obtain the aerodynamic performance of the rotary detonation engine, it often consumes a lot of time and computing resources.

发明内容SUMMARY OF THE INVENTION

技术目的：为解决上述技术问题，本发明提供了一种基于空间推进算法的旋转爆震发动机流场参数分布预测方法，通过耦合空间推进算法和单一γ的CJ爆震波模型，并利用空间推进算法所对应的流线网格的随流特性，确定激波坐标系下爆震波后流动参数以及爆震波远下游流动参数，并通过坐标转换获得实验室坐标系下的流动参数，有效降低旋转爆震发动机流场解算耗时，实现旋转爆震发动机流动参数分布的快速预测。Technical purpose: In order to solve the above technical problems, the present invention provides a method for predicting the flow field parameter distribution of a rotary detonation engine based on a space propulsion algorithm. The flow characteristics of the corresponding streamline grid are used to determine the flow parameters after the detonation wave and the flow parameters in the far downstream of the detonation wave in the shock wave coordinate system, and obtain the flow parameters in the laboratory coordinate system through coordinate transformation, which effectively reduces the rotational knock The engine flow field calculation is time-consuming, and the rapid prediction of the flow parameter distribution of the rotary detonation engine is realized.

技术方案：为实现上述目的，本发明采用的技术方案为：Technical scheme: In order to realize the above-mentioned purpose, the technical scheme adopted in the present invention is:

一种基于空间推进算法的旋转爆震发动机流场参数分布预测方法，其特征在于：包括以下步骤：A method for predicting the flow field parameter distribution of a rotary detonation engine based on a space propulsion algorithm, characterized in that it comprises the following steps:

(1)、预先估计喷注面进口平均速度，根据可压缩流动关系式，确定爆震波波前流动参数；(1) Pre-estimate the average velocity at the inlet of the injection surface, and determine the detonation wave front flow parameters according to the compressible flow relationship;

(2)、将爆震波前流动参数代入单一比热比的CJ爆震模型中，确定爆震波后流动参数；(2) Substitute the flow parameters before the detonation wave into the CJ detonation model of a single specific heat ratio, and determine the flow parameters after the detonation wave;

(3)、将爆震波后流动参数以及远下游流动参数赋值给初值线；第一次迭代时爆震波远下游流动参数由爆震波后压力膨胀至爆震波前压力获得，此后迭代中的爆震波远下游的流动参数则根据前一次计算结果；(3) Assign the flow parameters after the detonation wave and the far downstream flow parameters to the initial value line; in the first iteration, the far downstream flow parameters of the detonation wave are obtained from the pressure after the detonation wave to the pressure before the detonation wave. The flow parameters in the far downstream of the shock wave are based on the previous calculation results;

(4)、将初值线代入空间推进算法中，确定激波坐标系下的旋转爆震发动机流场结构及流动参数分布；(4) Substitute the initial value line into the space propulsion algorithm to determine the flow field structure and flow parameter distribution of the rotary detonation engine in the shock coordinate system;

(5)、将喷注面解点的压力值与喷注总压相比较，当解点压力值小于喷注总压时，对喷注流场和爆震波下游流场进行耦合求解，当解点压力值大于喷注总压时，仅求解爆震波下游流场；(5) Compare the pressure value of the solution point on the injection surface with the total injection pressure. When the pressure value of the solution point is less than the total injection pressure, the coupling solution of the injection flow field and the downstream flow field of the detonation wave is carried out. When the point pressure value is greater than the total injection pressure, only the downstream flow field of the detonation wave is solved;

(6)、确定爆震波远下游位置的爆震波前的流动参数的面积平均值，返回步骤(2)-步骤(5)进行迭代，直至远下游的平均气流角度为0，跳出循环，执行步骤(7)；(6), determine the area average value of the flow parameters of the detonation wave front at the far downstream position of the detonation wave, return to step (2)-step (5) for iteration, until the average airflow angle in the far downstream is 0, jump out of the cycle, and execute the steps (7);

(7)、根据爆震波倾角大小，确定激波坐标系的旋转角度，并根据速度三角形确定实验室坐标系下的旋转爆震发动机流动参数分布。(7) Determine the rotation angle of the shock wave coordinate system according to the inclination angle of the detonation wave, and determine the flow parameter distribution of the rotary detonation engine in the laboratory coordinate system according to the velocity triangle.

进一步的，所述步骤(1)包括如下步骤：Further, described step (1) comprises the steps:

(11)、根据旋转爆震发动机飞行工况，确定旋转爆震进口可燃混合气的总压Pt和总温Tt，并估计旋转爆震发动机进口平均速度

Figure BDA0003267983790000021
(11) According to the flight conditions of the rotary detonation engine, determine the total pressure Pt and total temperature Tt of the combustible mixture at the rotary detonation inlet, and estimate the average velocity of the rotary detonation engine inlet
Figure BDA0003267983790000021

(12)、根据可压缩流关系式，确定爆震波前平均静温

Figure BDA0003267983790000022
平均静压
Figure BDA0003267983790000023
以及平均密度
Figure BDA0003267983790000024
(12), according to the compressible flow relationship, determine the average static temperature before the detonation wave
Figure BDA0003267983790000022
Average static pressure
Figure BDA0003267983790000023
and the average density
Figure BDA0003267983790000024

进一步的，所述步骤(2)包括如下步骤：Further, described step (2) comprises the steps:

(21)、构造单一比热比的CJ爆震模型，模型方程为如下形式：(21) Construct a CJ knock model with a single specific heat ratio, and the model equation is as follows:

Figure BDA0003267983790000025
Figure BDA0003267983790000025

Figure BDA0003267983790000026
Figure BDA0003267983790000026

其中，MCJ表示爆震波马赫数，H表示无量纲释热量，Q表示化学反应热，Rg表示气体常数，γ为可燃混合气和燃烧产物的比热比；Among them, M CJ represents the detonation wave Mach number, H represents the dimensionless heat release, Q represents the chemical reaction heat, R g represents the gas constant, and γ represents the specific heat ratio of the combustible mixture and combustion products;

(22)、根据跨爆震波前后质量守恒定律、动量守恒定律以及能量守恒定律，确定爆震波后流动参数，包括爆震波后静压

Figure BDA0003267983790000031
爆震波后密度
Figure BDA0003267983790000032
和爆震波后温度
Figure BDA0003267983790000033
(22) According to the law of conservation of mass, the law of conservation of momentum and the law of energy conservation before and after the detonation wave, determine the flow parameters after the detonation wave, including the static pressure after the detonation wave
Figure BDA0003267983790000031
Density after detonation wave
Figure BDA0003267983790000032
and post-detonation temperature
Figure BDA0003267983790000033

进一步的，所述步骤(4)包括如下步骤：Further, described step (4) comprises the steps:

(41)、通过主流方向无粘通量E确定中间变量χ，进而获得初值线上的状态量Q，中间变量χ的表达式为如下形式：(41), the intermediate variable χ is determined by the inviscid flux E in the main flow direction, and then the state quantity Q on the initial value line is obtained. The expression of the intermediate variable χ is as follows:

Figure BDA0003267983790000034
Figure BDA0003267983790000034

(42)、根据初值线上速度分布，确定下游解点坐标

Figure BDA0003267983790000035
建立流线网格，
Figure BDA0003267983790000036
表示x方向上第j个下游解点的x坐标，
Figure BDA0003267983790000037
表示x方向上第j个初值线点的x坐标；(42) Determine the coordinates of the downstream solution point according to the velocity distribution on the initial value line
Figure BDA0003267983790000035
Create a streamline grid,
Figure BDA0003267983790000036
represents the x coordinate of the jth downstream solution point in the x direction,
Figure BDA0003267983790000037
Indicates the x-coordinate of the j-th initial value line point in the x-direction;

(43)、根据本质无振荡插值方法和限制函数，获得步进半步的沿x方向的

Figure BDA0003267983790000038
和
Figure BDA0003267983790000039
的数值解，进而获取步进长度为1/2位置的状态量
Figure BDA00032679837900000310
和
Figure BDA00032679837900000311
(43), according to the intrinsically non-oscillating interpolation method and the limit function, obtain the half-step along the x direction
Figure BDA0003267983790000038
and
Figure BDA0003267983790000039
The numerical solution of , and then obtain the state quantity of the position where the step length is 1/2
Figure BDA00032679837900000310
and
Figure BDA00032679837900000311

(44)、根据激波极曲线理论，通过空间x方向上步进长度为1/2设定值的相邻两点，确定空间推进算法所对应的步进长度为1/2位置的解点压力和气流角，进而获得黎曼自相似解，相邻两点状态曲线的表达式为如下形式：(44) According to the shock pole curve theory, through two adjacent points in the x-direction of space where the step length is 1/2 the set value, determine the solution point where the step length corresponding to the space propulsion algorithm is 1/2 pressure and airflow angle, and then obtain the Riemann self-similar solution, the expression of the state curve of two adjacent points is as follows:

Figure BDA00032679837900000312
Figure BDA00032679837900000312

Figure BDA00032679837900000313
Figure BDA00032679837900000313

其中，ΦB表示初值线下侧点的状态曲线，θB表示初值线下侧点的气流角，f(MB,αB)表示初值线下侧点的兰金许贡关系式，αB表示下游空间中任一点压力与初值线下侧点压力之比，ν(MB)表示初值线的下侧点的普朗特-迈耶函数，ΦT表示初值线上侧点的状态曲线，θT表示初值线上侧点的气流角，f(MT,αT)表示初值线上侧点的兰金许贡关系式，αT表示下游空间任一点压力与初值线上侧点压力之比，ν(MT)表示初值线的上侧点的普朗特-迈耶函数；Among them, Φ B represents the state curve of the point on the lower side of the initial value line, θ B represents the airflow angle of the point on the lower side of the initial value line, and f( MB , α B ) represents the Rankin Xugong relation of the point on the lower side of the initial value line , α B represents the ratio of the pressure at any point in the downstream space to the pressure at the lower point of the initial value line, ν(M B ) represents the Prandtl-Meyer function of the lower point of the initial value line, Φ T represents the initial value line The state curve of the side point, θ T represents the airflow angle of the side point on the initial value line, f(M T ,α T ) represents the Rankin-Huge relationship of the side point on the initial value line, α T represents the pressure at any point in the downstream space The ratio of the pressure at the upper point on the initial value line, ν(M T ) represents the Prandtl-Meyer function of the upper point on the initial value line;

(45)、通过对可压缩欧拉方程进行差分离散，获得单个空间递进步长下的流动参数，下游沿流动x方向的差分方程为如下形式：(45) Through differential discretization of the compressible Euler equation, the flow parameters under a single spatial progressive step are obtained, and the downstream differential equation along the flow x direction is in the following form:

Figure BDA0003267983790000041
Figure BDA0003267983790000041

其中，

Figure BDA0003267983790000042
表示初值线相邻两点的y坐标差值，
Figure BDA0003267983790000043
表示下游相邻两解点的y坐标差值，
Figure BDA0003267983790000044
表示空间步进1/2长度位置的上侧格点的沿y方向的无粘通量，
Figure BDA0003267983790000045
表示空间步进1/2长度位置的上侧格边的斜率，
Figure BDA0003267983790000046
表示空间步进1/2长度位置的上侧格点的沿x方向上的无粘通量，
Figure BDA0003267983790000047
表示空间步进1/2长度位置的下侧格点的沿y方向上的无粘通量，
Figure BDA0003267983790000048
表示空间步进1/2长度位置的下侧格边的斜率，
Figure BDA0003267983790000049
表示空间步进1/2长度位置的下侧格点的沿x方向上的无粘通量。in,
Figure BDA0003267983790000042
Represents the y-coordinate difference between two adjacent points on the initial value line,
Figure BDA0003267983790000043
Represents the y-coordinate difference between the two adjacent solution points downstream,
Figure BDA0003267983790000044
represents the inviscid flux along the y-direction of the upper lattice at the 1/2-length position of the spatial step,
Figure BDA0003267983790000045
Indicates the slope of the upper grid edge at the 1/2 length position of the space step,
Figure BDA0003267983790000046
represents the inviscid flux along the x-direction of the upper lattice point at the 1/2-length position of the space step,
Figure BDA0003267983790000047
represents the inviscid flux along the y-direction of the lower lattice at the 1/2-length position of the space step,
Figure BDA0003267983790000048
Indicates the slope of the lower grid edge at the 1/2 length position of the space step,
Figure BDA0003267983790000049
Represents the inviscid flux along the x-direction of the lower grid at the 1/2-length position of the spatial step.

进一步的，所述步骤(6)包括如下步骤：Further, described step (6) comprises the steps:

(61)、根据远下游沿y方向的流动参数分布，识别爆震波前可燃混合气所在区间；(61), according to the flow parameter distribution along the y direction in the far downstream, identify the interval where the combustible gas mixture is located before the detonation wave;

(62)、确定爆震波前可燃混气所在区间的流动参数的面积平均值，面积平均公式为如下形式：(62), determine the area average value of the flow parameters of the area where the combustible air mixture before the detonation wave is located, and the area average formula is as follows:

Figure BDA00032679837900000410
Figure BDA00032679837900000410

Figure BDA00032679837900000411
Figure BDA00032679837900000411

Figure BDA00032679837900000412
Figure BDA00032679837900000412

Figure BDA00032679837900000413
Figure BDA00032679837900000413

其中，

Figure BDA00032679837900000414
表示爆震波前平均速度的x分量，
Figure BDA00032679837900000415
表示爆震波前平均速度的y分量，nx表示爆震波面法矢的x分量，ny表示爆震波面法矢的y分量，
Figure BDA00032679837900000416
表示单位质量所含内能，根据牛顿-拉弗森多元迭代法，联立求解面积平均公式的四个方程，获得爆震波前流动参数
Figure BDA0003267983790000051
以及
Figure BDA0003267983790000052
in,
Figure BDA00032679837900000414
represents the x-component of the mean velocity of the detonation front,
Figure BDA00032679837900000415
represents the y-component of the average velocity of the detonation wavefront, nx represents the x-component of the normal vector of the detonation wavefront, ny represents the y-component of the normal vector of the detonation wavefront,
Figure BDA00032679837900000416
Represents the internal energy contained in the unit mass. According to the Newton-Raphson multivariate iterative method, the four equations of the area average formula are solved simultaneously to obtain the detonation wavefront flow parameters
Figure BDA0003267983790000051
as well as
Figure BDA0003267983790000052

进一步的，所述步骤(7)包括如下步骤：Further, described step (7) comprises the following steps:

(71)、根据爆震波前气流角，修正爆震波相对于喷注平面的倾角，修正公式为如下形式：(71), according to the airflow angle in front of the detonation wave, correct the inclination angle of the detonation wave relative to the injection plane, and the correction formula is as follows:

Figure BDA0003267983790000053
Figure BDA0003267983790000053

其中，

Figure BDA0003267983790000054
表示经过修正的爆震波相对于喷注平面的倾角，
Figure BDA0003267983790000055
表示未经过修正的爆震波相对于喷注平面的倾角；in,
Figure BDA0003267983790000054
represents the inclination of the corrected detonation wave relative to the injection plane,
Figure BDA0003267983790000055
Represents the inclination of the uncorrected detonation wave relative to the injection plane;

(72)、根据多次迭代修正后的爆震波相对于喷注平面的倾角，确定激波坐标系的旋转角度，通过二维旋转矩阵，将出口方向转为水平，旋转角度以及二维旋转矩阵为如下形式：(72), determine the rotation angle of the shock wave coordinate system according to the inclination of the detonation wave corrected by multiple iterations relative to the injection plane, and convert the exit direction to the horizontal through the two-dimensional rotation matrix, the rotation angle and the two-dimensional rotation matrix in the following form:

Figure BDA0003267983790000056
Figure BDA0003267983790000056

Figure BDA0003267983790000057
Figure BDA0003267983790000057

Figure BDA0003267983790000058
Figure BDA0003267983790000058

其中，

Figure BDA0003267983790000059
表示激波坐标系旋转角度，xnew表示激波坐标系旋转后的流场中任意一点的x坐标，ynew表示激波坐标系旋转后的流场中任意一点的y坐标，xold表示激波坐标系未旋转的流场中任意一点的x坐标，yold表示激波坐标系未旋转的流场中任意一点的y坐标，unew表示激波坐标系旋转后的流场中任意一点的速度x分量，vnew表示激波坐标系旋转后的流场中任意一点的速度y分量，uold表示激波坐标系未旋转的流场中任意一点的速度x分量，vold表示激波坐标系未旋转的流场中任意一点的速度y分量；in,
Figure BDA0003267983790000059
represents the rotation angle of the shock coordinate system, x new represents the x coordinate of any point in the flow field after the shock coordinate system is rotated, y new represents the y coordinate of any point in the flow field after the shock coordinate system is rotated, and x old represents the shock coordinate system. The x coordinate of any point in the flow field when the wave coordinate system is not rotated, y old is the y coordinate of any point in the flow field when the shock coordinate system is not rotated, and u new is the value of any point in the flow field after the shock coordinate system is rotated. The velocity x component, v new is the velocity y component of any point in the flow field after the shock coordinate system is rotated, u old is the velocity x component of any point in the flow field when the shock coordinate system is not rotated, v old is the shock coordinate is the y-component of the velocity at any point in the unrotated flow field;

(73)、根据矢量加减原则，通过速度三角形法，将激波坐标系内的速度转换为实验室坐标系内的速度，转换关系式为如下形式：(73) According to the principle of vector addition and subtraction, the velocity in the shock wave coordinate system is converted into the velocity in the laboratory coordinate system through the velocity triangle method, and the conversion relationship is as follows:

Figure BDA0003267983790000061
Figure BDA0003267983790000061

vlab＝vnew v lab = v new

其中，ulab表示实验室坐标系下任意一点速度的x分量，vlab表示实验室坐标系下任意一点速度的y分量。Among them, u lab represents the x component of the velocity at any point in the laboratory coordinate system, and v lab represents the y component of the velocity at any point in the laboratory coordinate system.

进一步的，所述步骤(6)中爆震波前可燃混气所在区间的上侧区间的气流角，在新一轮迭代的激波坐标系里修正为0。Further, in the step (6), the airflow angle of the upper section of the section where the combustible air-fuel mixture before the detonation wave is located is corrected to 0 in the shock wave coordinate system of a new round of iteration.

有益效果：由于采用了上述技术方案，本发明具有如下技术效果：Beneficial effect: Owing to adopting above-mentioned technical scheme, the present invention has following technical effect:

本发明提供的一种基于空间推进算法的旋转爆震发动机流场参数分布预测方法，通过耦合空间推进算法和单一γ的CJ爆震波模型，利用空间推进算法所对应的流线网格的随流特性，确定激波坐标系下爆震波后流动参数以及爆震波远下游流动参数，并通过坐标转换获得实验室坐标系下的流动参数，可以快速解算旋转爆震发动机流场，快速评估旋转爆震发动机的气动性能，同时精确捕捉流场中的爆震波、剪切层以及诱导激波等流动结构，极大地缩短了旋转爆震发动机构型设计周期。The present invention provides a method for predicting the flow field parameter distribution of a rotary detonation engine based on a space propulsion algorithm. By coupling the space propulsion algorithm and the CJ detonation wave model of a single γ, the following flow of the streamline grid corresponding to the space propulsion algorithm is utilized. It can determine the flow parameters after the detonation wave and the flow parameters in the far downstream of the detonation wave in the shock wave coordinate system, and obtain the flow parameters in the laboratory coordinate system through coordinate transformation. The aerodynamic performance of the detonation engine is improved, and the flow structures such as the detonation wave, shear layer and induced shock wave in the flow field are accurately captured, which greatly shortens the design cycle of the rotary detonation engine.

附图说明Description of drawings

图1为本发明基于空间推进算法的旋转爆震发动机流场参数分布预测方法中的步骤(1)到步骤(6)确定的激波坐标系下的计算域；1 is a computational domain under the shock coordinate system determined by steps (1) to (6) in the method for predicting the flow field parameter distribution of a rotary detonation engine based on a space propulsion algorithm of the present invention;

图2为旋转爆震发动机的实物结构示意图；Fig. 2 is the physical structure schematic diagram of rotary detonation engine;

图3为初值线上相邻两点的激波极曲线的相交过程；Figure 3 shows the intersection process of the shock pole curves of two adjacent points on the initial value line;

图4为空间推进算法从初值线到解点的步进过程，E-P表示格边点，C-P表示格心点，Const X表示初值线以及解点所在直线，P-M表示普朗特-迈耶膨胀波，Sh表示激波，Sl表示无粘滑移线；Figure 4 shows the step-by-step process of the space advance algorithm from the initial value line to the solution point. E-P represents the grid edge point, C-P represents the grid center point, Const X represents the initial value line and the straight line where the solution point is located, and P-M represents the Prandtl-Meyer Expansion wave, Sh means shock wave, Sl means no stick-slip line;

图5为旋转爆震发动机流场中爆震波参数收敛历程；Fig. 5 is the detonation wave parameter convergence process in the flow field of the rotary detonation engine;

图6为激波坐标系下的旋转爆震发动机流线网格；Fig. 6 is the streamline grid of the rotary detonation engine under the shock coordinate system;

图7为空间推进算法所获得的旋转爆震发动机温度场；Fig. 7 is the temperature field of the rotary detonation engine obtained by the space propulsion algorithm;

图8为空间推进算法所获得的实验室坐标系下的旋转爆震发动机速度场；Fig. 8 is the velocity field of the rotary detonation engine in the laboratory coordinate system obtained by the space propulsion algorithm;

图中，11-上游爆震波对应初值线，12-爆震波远下游流场对应初值线，13-爆震波远下游流场冗余度段对应初值线，14-倾角与冗余段气流折转角一致的假想无粘壁面，15-倾角等于爆震波倾角的喷注平面阻塞段，16-可燃混合气模化进气段，17-爆震波远下游流场对应解点所在直线，18-下游爆震波对应解点所在位置，19-倾角与可燃混合气模化进气段的气流折转角一致的假想无粘壁面。In the figure, 11 - the upstream detonation wave corresponds to the initial value line, 12 - the detonation wave far downstream flow field corresponds to the initial value line, 13 - the detonation wave far downstream flow field redundancy section corresponds to the initial value line, 14 - the dip angle and the redundant section The imaginary non-stick wall surface with the same airflow turning angle, 15- the injection plane blocking section with the inclination angle equal to the inclination angle of the detonation wave, 16- the modeled intake section of the combustible mixture, 17- the straight line corresponding to the solution point of the far downstream flow field of the detonation wave, 18 - The position of the solution point corresponding to the downstream detonation wave, 19 - an imaginary non-stick wall surface whose inclination angle is consistent with the airflow turning angle of the modeled intake section of the combustible mixture.

具体实施方式Detailed ways

下面结合附图及实施例对本发明作更进一步的说明。The present invention will be further described below in conjunction with the accompanying drawings and embodiments.

本发明公开了一种基于空间推进算法的旋转爆震发动机流场参数分布预测方法，包括步骤：预先估计爆震波波前流动参数；根据单一比热比的CJ爆震模型，确定爆震波后流动参数；将爆震波后流动参数和远下游流动参数赋值给初值线；通过空间推进算法确定激波坐标系下的流动参数；确定可燃混合气喷注起始位置和终了位置；迭代收敛旋转爆震发动机流场参数，并通过坐标转换获得实验室坐标系下的流动参数。The invention discloses a method for predicting the distribution of flow field parameters of a rotary detonation engine based on a space propulsion algorithm. The method includes the steps of: pre-estimating the detonation wave front flow parameters; parameters; assign the flow parameters after the detonation wave and the far downstream flow parameters to the initial value line; determine the flow parameters in the shock wave coordinate system through the space propulsion algorithm; determine the starting position and end position of the combustible mixture injection; iterative convergence of the rotary explosion The flow field parameters of the vibration engine were obtained, and the flow parameters in the laboratory coordinate system were obtained through coordinate transformation.

结合图1所示，11为上游爆震波对应初值线，12为爆震波远下游流场对应初值线，13对应爆震波远下游流场冗余度段对应初值线。14对应倾角与冗余段气流折转角一致的假想无粘壁面，15对应倾角等于爆震波倾角的喷注平面阻塞段，16对应可燃混合气模化进气段，17对应爆震波远下游流场对应解点所在直线，18对应下游爆震波对应解点所在位置，19对应倾角与可燃混合气模化进气段的气流折转角一致的假想无粘壁面。1, 11 is the initial value line corresponding to the upstream detonation wave, 12 is the initial value line corresponding to the far downstream flow field of the detonation wave, and 13 corresponds to the initial value line corresponding to the redundancy segment of the far downstream flow field of the detonation wave. 14 corresponds to the imaginary inviscid wall with the same inclination angle as the airflow turning angle of the redundant section, 15 corresponds to the injection plane blocking section with the inclination angle equal to the inclination angle of the detonation wave, 16 corresponds to the modeled intake section of the combustible mixture, and 17 corresponds to the flow field far downstream of the detonation wave The straight line corresponding to the solution point, 18 corresponds to the position of the downstream knock wave corresponding to the solution point, and 19 corresponds to the imaginary non-stick wall surface whose inclination angle is consistent with the airflow turning angle of the combustible mixture modeling intake section.

激波坐标系是假设运动的爆震波线段AB保持不动，以线段AB方向作为y坐标方向，以与线段AB垂直的方向作为x坐标方向，形成的坐标系。实验室坐标系是以喷注平面的线段AM方向组为x坐标方向，以与线段AM垂直的方向作为y坐标方向，形成的坐标系。The shock wave coordinate system is a coordinate system formed by assuming that the moving detonation wave line segment AB remains stationary, the direction of the line segment AB is taken as the y coordinate direction, and the direction perpendicular to the line segment AB is taken as the x coordinate direction. The laboratory coordinate system is a coordinate system formed by taking the line segment AM direction group of the injection plane as the x coordinate direction and the direction perpendicular to the line segment AM as the y coordinate direction.

θave角为平均气流角，位于线段IJ稍上游的位置，θ角为喷注平面(激波坐标系时倾斜的线段AJ)和爆震波AB夹角的余角。The angle θ ave is the average airflow angle, which is located slightly upstream of the line segment IJ, and the angle θ is the residual angle between the injection plane (the inclined line segment AJ in the shock wave coordinate system) and the angle of the detonation wave AB.

预测方法具体包括以下步骤：The prediction method specifically includes the following steps:

步骤(1)，预先估计喷注面进口平均速度，根据可压缩流动关系式，确定爆震波波前流动参数；Step (1), pre-estimating the average velocity of the injection surface inlet, and determining the detonation wave front flow parameters according to the compressible flow relationship;

步骤(2)，将爆震波前流动参数代入单一比热比γ的CJ爆震模型中，确定爆震波后流动参数；In step (2), the flow parameters before the detonation wave are substituted into the CJ detonation model of a single specific heat ratio γ, and the flow parameters after the detonation wave are determined;

步骤(3)，第一次迭代中的爆震波远下游的流动参数由爆震波后压力膨胀至爆震波前压力获得，此后迭代中的爆震波远下游的流动参数则根据前一次计算结果，将爆震波后流动参数以及远下游流动参数赋值给初值线；如图1所示，初值线由线段AB和BD共同构成，而爆震波AB后的流动参数与前一次计算的线段IJ爆震波后的流动参数一致，同时爆震波上方的线段BD的流动参数分布与前一次计算的线段IE的流动参数分布一致，具体数值则通过线性插值获得。In step (3), the flow parameters in the far downstream of the detonation wave in the first iteration are obtained by expanding the pressure after the detonation wave to the pressure in the front of the detonation wave, and the flow parameters in the far downstream of the detonation wave in the subsequent iterations are based on the previous calculation results. The flow parameters after the detonation wave and the far downstream flow parameters are assigned to the initial value line; as shown in Figure 1, the initial value line is composed of line segments AB and BD, and the flow parameters after the detonation wave AB are the same as the line segment IJ detonation wave calculated in the previous calculation. The latter flow parameters are consistent, and the flow parameter distribution of the line segment BD above the detonation wave is consistent with the flow parameter distribution of the line segment IE calculated in the previous calculation, and the specific values are obtained by linear interpolation.

步骤(4)，将初值线代入空间推进算法中，确定激波坐标系下的旋转爆震发动机流场结构及流动参数分布；Step (4), substitute the initial value line into the space propulsion algorithm, and determine the flow field structure and flow parameter distribution of the rotary detonation engine under the shock coordinate system;

步骤(5)，将喷注面解点的压力值与喷注总压相比较，当解点压力值小于喷注总压时，对喷注流场和爆震波下游流场进行耦合求解；若解点压力值大于等于总压时，喷注流场不存在，此时仅进行爆震波下游流场的求解。Step (5), compare the pressure value of the solution point of the injection surface with the total injection pressure, and when the pressure value of the solution point is less than the total injection pressure, the coupling solution of the injection flow field and the downstream flow field of the detonation wave is carried out; When the pressure value of the solution point is greater than or equal to the total pressure, the injection flow field does not exist, and only the downstream flow field of the detonation wave is solved.

步骤(6)，确定爆震波远下游位置的爆震波前的流动参数的面积平均值，返回步骤(2)进行迭代，直至远下游的平均气流角度为0，跳出循环；Step (6), determine the area average value of the flow parameters of the detonation wave front at the far downstream position of the detonation wave, return to step (2) for iteration, until the average airflow angle far downstream is 0, jump out of the cycle;

步骤(7)，根据爆震波倾角大小，确定激波坐标系的旋转角度，并根据速度三角形确定实验室坐标系下的旋转爆震发动机流动参数分布；Step (7), according to the size of the detonation wave inclination, determine the rotation angle of the shock wave coordinate system, and determine the flow parameter distribution of the rotary detonation engine under the laboratory coordinate system according to the velocity triangle;

步骤(8)，将实验室坐标系下的旋转爆震发动机流动参数分布输出成数据文件。Step (8), output the flow parameter distribution of the rotary detonation engine in the laboratory coordinate system into a data file.

结合图3，本发明中，相邻两点的状态曲线的交点参数通过牛顿-拉弗森迭代方法获得。结合图4，本发明中，由初值线向下游解点推进过程中，通过直接计算激波和膨胀波之间的相互作用获得流动参数，不需额外构造单元点过程，通过自适应的改变格边的斜率，形成流线网格。Referring to FIG. 3 , in the present invention, the intersection parameter of the state curves of two adjacent points is obtained by the Newton-Raphson iteration method. Referring to Fig. 4, in the present invention, in the process of advancing from the initial value line to the downstream solution point, the flow parameters are obtained by directly calculating the interaction between the shock wave and the expansion wave. The slope of the grid edges, forming a streamline grid.

本发明步骤(1)，具体包括如下步骤：Step (1) of the present invention specifically comprises the following steps:

(11)、根据旋转爆震发动机飞行工况，确定旋转爆震进口可燃混合气的总压Pt和总温Tt，并估计旋转爆震发动机进口平均速度

Figure BDA0003267983790000081
(11) According to the flight conditions of the rotary detonation engine, determine the total pressure Pt and total temperature Tt of the combustible mixture at the rotary detonation inlet, and estimate the average velocity of the rotary detonation engine inlet
Figure BDA0003267983790000081

(12)、根据可压缩流关系式，确定爆震波前流动参数，包括平均静温

Figure BDA0003267983790000082
平均静压
Figure BDA0003267983790000083
以及平均密度
Figure BDA0003267983790000084
(12) According to the compressible flow relationship, determine the detonation wave front flow parameters, including the average static temperature
Figure BDA0003267983790000082
Average static pressure
Figure BDA0003267983790000083
and the average density
Figure BDA0003267983790000084

存在以下关系：The following relationships exist:

Figure BDA0003267983790000085
Figure BDA0003267983790000085

Figure BDA0003267983790000091
Figure BDA0003267983790000091

Figure BDA0003267983790000092
Figure BDA0003267983790000092

其中，所述Cp和γ分别为可燃混合气和燃烧产物的定压比热和比热比，Rg表示气体常数。Wherein, the C p and γ are the constant pressure specific heat and the specific heat ratio of the combustible mixture and the combustion product, respectively, and R g represents the gas constant.

步骤(2)具体包括如下步骤：Step (2) specifically includes the following steps:

(21)、根据双γ的CJ爆震模型，简化可得单一比热比γ的CJ爆震模型：(21) According to the CJ knocking model of double γ, the CJ knocking model of single specific heat ratio γ can be obtained by simplification:

Figure BDA0003267983790000093
Figure BDA0003267983790000093

Figure BDA0003267983790000094
Figure BDA0003267983790000094

(22)、根据跨爆震波前后质量守恒定律、动量守恒定律以及能量守恒定律，确定爆震波后流动参数的表达式为如下形式：(22) According to the law of conservation of mass, the law of conservation of momentum and the law of energy conservation before and after the detonation wave, the expression of the flow parameters after the detonation wave is determined as follows:

Figure BDA0003267983790000095
Figure BDA0003267983790000095

Figure BDA0003267983790000096
Figure BDA0003267983790000096

Figure BDA0003267983790000097
Figure BDA0003267983790000097

其中，MCJ表示爆震波马赫数，H表示无量纲释热量，Q表示化学反应热，Rg表示气体常数，

Figure BDA0003267983790000098
表示爆震波前平均静温，
Figure BDA0003267983790000099
表示爆震波前平均密度，
Figure BDA00032679837900000910
表示爆震波前平均静压，
Figure BDA00032679837900000911
表示爆震波后静压，
Figure BDA00032679837900000912
表示爆震波后密度，
Figure BDA00032679837900000913
表示爆震波后温度。Among them, M CJ represents the detonation wave Mach number, H represents the dimensionless heat release, Q represents the chemical reaction heat, R g represents the gas constant,
Figure BDA0003267983790000098
is the average static temperature before the detonation wave,
Figure BDA0003267983790000099
is the average density of the detonation wavefront,
Figure BDA00032679837900000910
is the average static pressure before the detonation wave,
Figure BDA00032679837900000911
represents the static pressure after the detonation wave,
Figure BDA00032679837900000912
is the density after the detonation wave,
Figure BDA00032679837900000913
Indicates the post-detonation temperature.

步骤(4)具体包括如下步骤：Step (4) specifically includes the following steps:

(41)、通过主流方向无粘通量E确定中间变量χ，进而获得初值线上的状态量Q，中间变量χ的表达式为如下形式：(41), the intermediate variable χ is determined by the inviscid flux E in the main flow direction, and then the state quantity Q on the initial value line is obtained. The expression of the intermediate variable χ is as follows:

Figure BDA00032679837900000914
Figure BDA00032679837900000914

E1、E2、E3、E4、是指主流方向无粘通量；E 1 , E 2 , E 3 , E 4 , refer to the non-viscous flux in the mainstream direction;

(42)、根据初值线上速度分布，确定下游解点坐标，建立流线网格，下游解点的表达式为如下形式：(42) According to the velocity distribution on the initial value line, determine the coordinates of the downstream solution point, establish a streamline grid, and the expression of the downstream solution point is as follows:

Figure BDA0003267983790000101
Figure BDA0003267983790000101

Figure BDA0003267983790000102
Figure BDA0003267983790000102

Figure BDA0003267983790000103
表示x方向上第j个下游解点的x坐标，
Figure BDA0003267983790000104
表示x方向上第j个初值线点的x坐标，Δx表示网格沿x方向步进长度，
Figure BDA0003267983790000105
表示y方向上第j个下游解点的y坐标，
Figure BDA0003267983790000106
表示y方向上第j个初值线点的x坐标，
Figure BDA0003267983790000107
表示y方向上第j初值线点的y向速度分量，
Figure BDA0003267983790000108
表示y方向上第j初值线点的x方向速度分量；
Figure BDA0003267983790000103
represents the x coordinate of the jth downstream solution point in the x direction,
Figure BDA0003267983790000104
Represents the x coordinate of the jth initial value line point in the x direction, Δx represents the step length of the grid along the x direction,
Figure BDA0003267983790000105
represents the y coordinate of the jth downstream solution point in the y direction,
Figure BDA0003267983790000106
Represents the x-coordinate of the j-th initial value line point in the y-direction,
Figure BDA0003267983790000107
represents the y-direction velocity component of the j-th initial value line point in the y-direction,
Figure BDA0003267983790000108
Indicates the x-direction velocity component of the j-th initial value line point in the y-direction;

(43)、根据本质无振荡插值方法和限制函数，获得步进半步的沿x方向的

Figure BDA0003267983790000109
和
Figure BDA00032679837900001010
的数值解，进而获取步进长度为1/2位置的状态量
Figure BDA00032679837900001011
和
Figure BDA00032679837900001012
(43), according to the intrinsically non-oscillating interpolation method and the limit function, obtain the half-step along the x direction
Figure BDA0003267983790000109
and
Figure BDA00032679837900001010
The numerical solution of , and then obtain the state quantity of the position where the step length is 1/2
Figure BDA00032679837900001011
and
Figure BDA00032679837900001012

(44)、根据激波极曲线理论，通过空间x方向上步进长度为1/2设定值的相邻两点，确定空间推进算法所对应的步进长度为1/2位置的解点压力和气流角，进而获得黎曼自相似解，采用的相邻两点的状态曲线包括如下激波极曲线方程：(44) According to the shock pole curve theory, through two adjacent points in the x-direction of space where the step length is 1/2 the set value, determine the solution point where the step length corresponding to the space propulsion algorithm is 1/2 pressure and airflow angle, and then obtain the Riemann self-similar solution. The state curve of two adjacent points used includes the following shock pole curve equation:

Figure BDA00032679837900001013
Figure BDA00032679837900001013

Figure BDA00032679837900001014
Figure BDA00032679837900001014

其中，ΦB表示初值线下侧点的状态曲线，θB表示初值线下侧点的气流角，f(MB,αB)表示初值线下侧点的兰金许贡关系式，αB表示下游空间中任一点压力与初值线下侧点压力之比，ν(MB)表示初值线的下侧点的普朗特-迈耶函数，ΦT表示初值线上侧点的状态曲线，θT表示初值线上侧点的气流角，f(MT,αT)表示初值线上侧点的兰金许贡关系式，αT表示下游空间任一点压力与初值线上侧点压力之比，ν(MT)表示初值线的上侧点的普朗特-迈耶函数。Among them, Φ B represents the state curve of the point on the lower side of the initial value line, θ B represents the airflow angle of the point on the lower side of the initial value line, and f( MB , α B ) represents the Rankin Xugong relation of the point on the lower side of the initial value line , α B represents the ratio of the pressure at any point in the downstream space to the pressure at the lower point of the initial value line, ν(M B ) represents the Prandtl-Meyer function of the lower point of the initial value line, Φ T represents the initial value line The state curve of the side point, θ T represents the airflow angle of the side point on the initial value line, f(M T ,α T ) represents the Rankin-Huge relationship of the side point on the initial value line, α T represents the pressure at any point in the downstream space The ratio of the pressure at the upper point on the initial value line, ν(M T ) represents the Prandtl-Meyer function of the upper point on the initial value line.

(45)、通过对可压缩欧拉方程进行差分离散，获得单个空间递进步长下的流动参数，下游沿流动x方向的差分方程为如下形式：(45) Through differential discretization of the compressible Euler equation, the flow parameters under a single spatial progressive step are obtained, and the downstream differential equation along the flow x direction is in the following form:

Figure BDA0003267983790000111
Figure BDA0003267983790000111

其中，

Figure BDA0003267983790000112
表示初值线相邻两点的y坐标差值，
Figure BDA0003267983790000113
表示下游相邻两解点的y坐标差值，
Figure BDA0003267983790000114
表示空间步进1/2长度位置的上侧格点的沿y方向的无粘通量，
Figure BDA0003267983790000115
表示空间步进1/2长度位置的上侧格边的斜率，
Figure BDA0003267983790000116
表示空间步进1/2长度位置的上侧格点的沿x方向上的无粘通量，
Figure BDA0003267983790000117
表示空间步进1/2长度位置的下侧格点的沿y方向上的无粘通量，
Figure BDA0003267983790000118
表示空间步进1/2长度位置的下侧格边的斜率，
Figure BDA0003267983790000119
表示空间步进1/2长度位置的下侧格点的沿x方向上的无粘通量。in,
Figure BDA0003267983790000112
Represents the y-coordinate difference between two adjacent points on the initial value line,
Figure BDA0003267983790000113
Represents the y-coordinate difference between the two adjacent solution points downstream,
Figure BDA0003267983790000114
represents the inviscid flux along the y-direction of the upper lattice at the 1/2-length position of the spatial step,
Figure BDA0003267983790000115
Indicates the slope of the upper grid edge at the 1/2 length position of the space step,
Figure BDA0003267983790000116
represents the inviscid flux along the x-direction of the upper lattice point at the 1/2-length position of the space step,
Figure BDA0003267983790000117
represents the inviscid flux along the y-direction of the lower lattice at the 1/2-length position of the space step,
Figure BDA0003267983790000118
Indicates the slope of the lower grid edge at the 1/2 length position of the space step,
Figure BDA0003267983790000119
Represents the inviscid flux along the x-direction of the lower grid at the 1/2-length position of the spatial step.

步骤(5)中，将喷注面解点的压力值与喷注总压相比较，解点为流场区域中第n个ConstX线上的上游点通过步骤(44)和(45)获得的第n+1个ConstX线的下游点。其中ConstX代表图1中与爆震波AB平行的直线。In step (5), the pressure value of the solution point on the injection surface is compared with the total injection pressure, and the solution point is the upstream point on the nth ConstX line in the flow field region obtained through steps (44) and (45). Downstream point of the n+1th ConstX line. where ConstX represents the straight line parallel to the detonation wave AB in Fig. 1 .

所述步骤(6)包括如下步骤：Described step (6) comprises the steps:

步骤(61)、根据远下游沿y方向的流动参数分布，识别爆震波前可燃混合气所在区间；Step (61), identify the interval where the combustible gas mixture is located before the detonation wave according to the flow parameter distribution along the y direction in the far downstream;

步骤(62)、确定爆震波前可燃混气所在区间的流动参数的面积平均值，面积平均公式为如下形式：Step (62), determine the area average value of the flow parameters in the interval where the combustible gas mixture is located before the detonation wave, and the area average formula is the following form:

Figure BDA00032679837900001110
Figure BDA00032679837900001110

Figure BDA00032679837900001111
Figure BDA00032679837900001111

Figure BDA00032679837900001112
Figure BDA00032679837900001112

Figure BDA0003267983790000121
Figure BDA0003267983790000121

其中，

Figure BDA0003267983790000122
表示爆震波前平均速度的x分量，
Figure BDA0003267983790000123
表示爆震波前平均速度的y分量，nx表示爆震波面法矢的x分量，ny表示爆震波面法矢的y分量，
Figure BDA0003267983790000124
表示单位质量所含内能，A表示爆震波所占面积。根据牛顿-拉弗森多元迭代法，联立求解面积平均公式的四个方程，获得爆震波前流动参数
Figure BDA0003267983790000125
以及
Figure BDA0003267983790000126
in,
Figure BDA0003267983790000122
represents the x-component of the mean velocity of the detonation front,
Figure BDA0003267983790000123
represents the y-component of the average velocity of the detonation wavefront, nx represents the x-component of the normal vector of the detonation wavefront, ny represents the y-component of the normal vector of the detonation wavefront,
Figure BDA0003267983790000124
Represents the internal energy contained in the unit mass, and A represents the area occupied by the detonation wave. According to the Newton-Raphson multivariate iterative method, the four equations of the area average formula are solved simultaneously to obtain the detonation wavefront flow parameters
Figure BDA0003267983790000125
as well as
Figure BDA0003267983790000126

迭代时，在步骤(3)赋值时用到的前一次的计算结果，是指根据步骤(62)得到的爆震波前流动参数进一步通过步骤(2)计算得到的爆震波后流动参数。During the iteration, the previous calculation result used in the assignment of step (3) refers to the flow parameters after the detonation wave further calculated in step (2) according to the flow parameters before the detonation wave obtained in step (62).

所述步骤(7)包括如下步骤：Described step (7) comprises the following steps:

(71)、根据爆震波前气流角，修正爆震波相对于喷注平面的倾角，修正公式为如下形式：(71), according to the airflow angle in front of the detonation wave, correct the inclination angle of the detonation wave relative to the injection plane, and the correction formula is as follows:

Figure BDA0003267983790000127
Figure BDA0003267983790000127

其中，

Figure BDA0003267983790000128
表示经过修正的爆震波相对于喷注平面的倾角，
Figure BDA0003267983790000129
表示未经过修正的爆震波相对于喷注平面的倾角；in,
Figure BDA0003267983790000128
represents the inclination of the corrected detonation wave relative to the injection plane,
Figure BDA0003267983790000129
Represents the inclination of the uncorrected detonation wave relative to the injection plane;

(72)、根据多次迭代修正后的爆震波相对于喷注平面的倾角，确定激波坐标系的旋转角度，通过二维旋转矩阵，将出口方向转为水平，旋转角度以及二维旋转矩阵为如下形式：(72), determine the rotation angle of the shock wave coordinate system according to the inclination of the detonation wave corrected by multiple iterations relative to the injection plane, and convert the exit direction to the horizontal through the two-dimensional rotation matrix, the rotation angle and the two-dimensional rotation matrix in the following form:

Figure BDA00032679837900001210
Figure BDA00032679837900001210

Figure BDA00032679837900001211
Figure BDA00032679837900001211

Figure BDA00032679837900001212
Figure BDA00032679837900001212

其中，

Figure BDA00032679837900001213
表示激波坐标系旋转角度，xnew表示激波坐标系旋转后的流场中任意一点的x坐标，ynew表示激波坐标系旋转后的流场中任意一点的y坐标，xold表示激波坐标系未旋转的流场中任意一点的x坐标，yold表示激波坐标系未旋转的流场中任意一点的y坐标，unew表示激波坐标系旋转后的流场中任意一点的速度x分量，vnew表示激波坐标系旋转后的流场中任意一点的速度y分量，uold表示激波坐标系未旋转的流场中任意一点的速度x分量，vold表示激波坐标系未旋转的流场中任意一点的速度y分量；in,
Figure BDA00032679837900001213
represents the rotation angle of the shock coordinate system, x new represents the x coordinate of any point in the flow field after the shock coordinate system is rotated, y new represents the y coordinate of any point in the flow field after the shock coordinate system is rotated, and x old represents the shock coordinate system. The x coordinate of any point in the flow field when the wave coordinate system is not rotated, y old is the y coordinate of any point in the flow field when the shock coordinate system is not rotated, and u new is the value of any point in the flow field after the shock coordinate system is rotated. The velocity x component, v new is the velocity y component of any point in the flow field after the shock coordinate system is rotated, u old is the velocity x component of any point in the flow field when the shock coordinate system is not rotated, v old is the shock coordinate is the y-component of the velocity at any point in the unrotated flow field;

(73)、根据矢量加减原则，通过速度三角形法，将激波坐标系内的速度转换为实验室坐标系内的速度，转换关系式为如下形式：(73) According to the principle of vector addition and subtraction, the velocity in the shock wave coordinate system is converted into the velocity in the laboratory coordinate system through the velocity triangle method, and the conversion relationship is as follows:

Figure BDA0003267983790000131
Figure BDA0003267983790000131

vlab＝vnew v lab = v new

其中，ulab表示实验室坐标系下任意一点速度的x分量，vlab表示实验室坐标系下任意一点速度的y分量。速度三角形如图1所示，

Figure BDA0003267983790000132
表示激波坐标系下AMLKED区域里的流场速度，速度之间的换算通过矢量加减获得。由于激波坐标系中将激波的空间位置固定不动，而实际流动(实验室坐标系)中激波等流动结构沿着喷注平面切向高速运动，需要根据相对运动的变换规则，因此通过坐标转换获得实际旋转爆震发动机的流场。Among them, u lab represents the x component of the velocity at any point in the laboratory coordinate system, and v lab represents the y component of the velocity at any point in the laboratory coordinate system. The velocity triangle is shown in Figure 1,
Figure BDA0003267983790000132
Indicates the velocity of the flow field in the AMLKED region in the shock coordinate system, and the conversion between the velocities is obtained by vector addition and subtraction. Since the spatial position of the shock wave is fixed in the shock wave coordinate system, and the flow structure such as the shock wave in the actual flow (laboratory coordinate system) moves tangentially and at high speed along the injection plane, it needs to be based on the transformation rules of relative motion, so The flow field of the actual rotary detonation engine is obtained by coordinate transformation.

进一步的，所述步骤(6)中爆震波前可燃混气所在区间的上侧区间的气流角，在新一轮迭代的激波坐标系里修正为0。Further, in the step (6), the airflow angle of the upper section of the section where the combustible air-fuel mixture before the detonation wave is located is corrected to 0 in the shock wave coordinate system of a new round of iteration.

本发明的方法可用于快速评估旋转爆震发动机气动性能，即可以通过建立爆震过程和其下游流场结构的数学模型，避免传统时间推进算法耗费时间和计算资源的特点。The method of the present invention can be used to quickly evaluate the aerodynamic performance of a rotary detonation engine, that is, by establishing a mathematical model of the detonation process and its downstream flow field structure, the traditional time-propulsion algorithm can avoid the time-consuming and computational resource-consuming characteristics.

为了更好的说明本发明，便于理解本发明的技术方案，本发明的典型但非限制性的实施例如下：In order to better illustrate the present invention and facilitate the understanding of the technical solutions of the present invention, typical but non-limiting examples of the present invention are as follows:

旋转爆震发动机进口可燃混合气的喷注总压为5×105Pa，喷注总温为300K，进口平均速度的初始值取450m/s，可燃预混气比热比为1.3961，可燃预混气气体常数为395.75J·kg-1·K-1，燃烧产物比热比为1.1653，燃烧产物气体常数为346.20J·kg-1·K-1，单位质量释热为5.4704×106J·kg-1，可燃混合气活化温度15100K,单步化学反应指前因子1.0×109s-1。旋转爆震发动机中环展开长度314mm，轴向长度75mm。The total injection pressure of the combustible mixture at the inlet of the rotary detonation engine is 5×10 5 Pa, the total injection temperature is 300K, the initial value of the average inlet velocity is 450m/s, the specific heat ratio of the combustible premixed gas is 1.3961, and the combustible premixture is 1.3961. The gas constant of the mixture is 395.75J·kg -1 ·K -1 , the specific heat ratio of the combustion product is 1.1653, the gas constant of the combustion product is 346.20J·kg -1 ·K -1 , and the heat release per unit mass is 5.4704×10 6 J ·kg -1 , the activation temperature of the combustible mixture is 15100K, and the single-step chemical reaction pre-exponential factor is 1.0×10 9 s -1 . The middle ring of the rotary detonation engine has an unfolded length of 314mm and an axial length of 75mm.

图5是本发明快速解算迭代时爆震波各参数的收敛历程，θinc表示爆震波倾角，HD表示爆震波长度，P1/P2表示爆震波前波后静压比，Ulab表示爆震波速。图6是本发明快速解算时在激波坐标系下的流线网格。图7是本发明所得旋转爆震发动机的温度场，准确的捕捉到激波增压过程和爆震波后气流膨胀加速过程。图8是本发明所得旋转爆震发动机在实验室坐标系下的速度场，准确的捕捉到旋转爆震发动机的气流流向变化过程，实验室坐标系的定义是以喷注平面的线段AM方向组为x坐标方向，以与线段AM垂直的方向作为y坐标方向，形成的坐标系。Fig. 5 is the convergence history of each parameter of the detonation wave during the rapid solution iteration of the present invention, θ inc represents the inclination angle of the detonation wave, H D represents the length of the detonation wave, P 1 /P 2 represents the static pressure ratio before and after the detonation wave, and U lab represents the detonation wave. detonation wave velocity. Fig. 6 is the streamline grid in the shock coordinate system during the fast solution of the present invention. 7 is the temperature field of the rotary detonation engine obtained by the present invention, which accurately captures the shock wave supercharging process and the airflow expansion acceleration process after the detonation wave. Fig. 8 is the velocity field of the rotary detonation engine obtained by the present invention under the laboratory coordinate system, and accurately captures the change process of the airflow direction of the rotary detonation engine. The definition of the laboratory coordinate system is the line segment AM direction group of the injection plane is the x-coordinate direction, and the direction perpendicular to the line segment AM is used as the y-coordinate direction to form a coordinate system.

表1是本发明快速解算所得的爆震波参数与传统时间推进算法解算所得爆震波参数的对比数据。Table 1 is the comparison data of the detonation wave parameters obtained by the rapid solution of the present invention and the detonation wave parameters obtained by the solution of the traditional time advancing algorithm.

表1Table 1

Figure BDA0003267983790000141
Figure BDA0003267983790000141

与传统时间推进算法所得旋转爆震发动机的爆震波参数相比，本发明所得爆震波高度相对偏差为1.38％，爆震波倾角相对偏差为1.35％，爆震波前后静压相对偏差为1.05％，爆震波波速相对偏差为0.97％，可见空间推进算法具有时间推进算法解算旋转爆震流场同等精度。Compared with the detonation wave parameters of the rotary detonation engine obtained by the traditional time advance algorithm, the relative deviation of the detonation wave height obtained by the present invention is 1.38%, the relative deviation of the inclination angle of the detonation wave is 1.35%, and the relative deviation of the static pressure before and after the detonation wave is 1.05%. The relative deviation of the shock wave velocity is 0.97%. It can be seen that the space propulsion algorithm has the same accuracy as the time propulsion algorithm to solve the rotating detonation flow field.

表2是在4核8线程i7CPU环境下本发明解算旋转爆震发动机流场耗时与传统时间推进算法解算耗时的对比数据。Table 2 is the comparison data of the time-consuming solution of the flow field of the rotary detonation engine of the present invention and the time-consuming solution of the traditional time-propulsion algorithm under the environment of 4-core 8-thread i7CPU.

表2Table 2

算法algorithm CPU数目及型号Number and model of CPU 耗时time consuming 空间推进算法space propulsion algorithm 4核8线程，i74 cores 8 threads, i7 5h5h 传统时间推进算法Traditional Time Advancement Algorithm 4核8线程，i74 cores 8 threads, i7 48h48h

本发明是一种基于空间推进算法的旋转爆震发动机流场参数分布预测方法，使用4核8线程i7的CPU进行求解运算，总耗时5h。而传统的时间推进算法使用4核8线程进行求解运算，总耗时48h。本发明比传统时间推进算法求解速度快8.6倍，能有效的降低旋转爆震发动机构型设计和选取周期。The present invention is a method for predicting the flow field parameter distribution of a rotary detonation engine based on a space propulsion algorithm, which uses a CPU with 4 cores and 8 threads i7 to perform the calculation, which takes 5 hours in total. The traditional time advancement algorithm uses 4 cores and 8 threads to solve the operation, which takes 48h in total. The invention is 8.6 times faster than the traditional time advancing algorithm, and can effectively reduce the design and selection period of the rotary detonation engine.

以上所述仅是本发明的优选实施方式，应当指出：对于本技术领域的普通人员来说，在不脱离本发明原理的前提下，还可以做出若干改进和润饰，这些改进和润饰也应视为本发明的保护范围。The above are only the preferred embodiments of the present invention, and it should be pointed out that: for those skilled in the art, without departing from the principles of the present invention, several improvements and modifications can be made, and these improvements and modifications should also be It is regarded as the protection scope of the present invention.

Claims (7)

1. A method for predicting the distribution of flow field parameters of a rotary detonation engine based on a space propulsion algorithm is characterized by comprising the following steps of: the method comprises the following steps:
(1) estimating the average speed of an inlet of an injection surface in advance, and determining the flow parameters of the detonation wave front according to a compressible flow relation;
(2) substituting the flow parameters of the detonation wave front into a CJ detonation model with a single specific heat ratio to determine the flow parameters of the detonation wave front;
(3) assigning the flow parameters after the detonation wave and the flow parameters far downstream to an initial value line; the flow parameters far downstream of the detonation wave in the first iteration are obtained by expanding the pressure after the detonation wave to the pressure before the detonation wave, and the flow parameters far downstream of the detonation wave in the later iteration are obtained according to the calculation result of the previous iteration;
(4) substituting the initial value line into a space propulsion algorithm, and determining a flow field structure and flow parameter distribution of the rotary detonation engine under a shock wave coordinate system;
(5) comparing the pressure value of the solution point of the injection surface with the total injection pressure, when the pressure value of the solution point is less than the total injection pressure, performing coupled solution on the injection flow field and the detonation wave downstream flow field, and when the pressure value of the solution point is more than or equal to the total injection pressure, only performing solution on the detonation wave downstream flow field;
(6) determining the area average value of the flow parameters of the detonation wave front at the far downstream position of the detonation wave, returning to the step (2) to the step (5) for iteration, jumping out of a loop until the average airflow angle at the far downstream is 0, and executing the step (7);
(7) and determining the rotation angle of the shock wave coordinate system according to the inclination angle of the detonation wave, and determining the flow parameter distribution of the rotary detonation engine under the laboratory coordinate system according to the velocity triangle.
2. A method for predicting distribution of flow field parameters of a rotary detonation engine based on a spatial propelling algorithm according to claim 1, wherein the step (1) comprises the following steps:
(11) determining the total pressure Pt and the total temperature Tt of the combustible mixture at the rotary detonation inlet according to the flight condition of the rotary detonation engine, and estimating the average speed at the inlet of the rotary detonation engine
Figure FDA0003267983780000011
(12) Determining the average static temperature of the detonation wave front according to the compressible flow relation
Figure FDA0003267983780000012
Mean static pressure
Figure FDA0003267983780000013
And average density
Figure FDA0003267983780000014
3. The method for predicting the distribution of the flow field parameters of the rotary detonation engine based on the spatial propelling algorithm is characterized in that the step (2) comprises the following steps:
(21) constructing a CJ knock model with a single specific heat ratio, wherein a model equation is in the form of:
Figure FDA0003267983780000021
Figure FDA0003267983780000022
wherein M isCJRepresenting the Mach number of the detonation wave, H representing the dimensionless heat release amount, Q representing the heat of chemical reaction, RgRepresenting a gas constant, gamma is the specific heat ratio of the combustible mixture and the combustion products;
(22) determining post-detonation wave flow parameters including post-detonation wave static pressure parameters according to the mass conservation law, momentum conservation law and energy conservation law before and after the detonation wave
Figure FDA0003267983780000023
Density after detonation wave
Figure FDA0003267983780000024
And post detonation wave temperature
Figure FDA0003267983780000025
4. A method for predicting distribution of flow field parameters of a rotary detonation engine based on a spatial propelling algorithm according to claim 3, wherein the step (4) comprises the following steps:
(41) determining an intermediate variable chi through the non-viscous flux E in the main flow direction, and further obtaining a state quantity Q on an initial value line, wherein the expression of the intermediate variable chi is in the following form:
Figure FDA0003267983780000026
(42) determining downstream solution point coordinates according to the velocity distribution on the initial value line
Figure FDA0003267983780000027
A streamline grid is established, and the streamline grid is established,
Figure FDA0003267983780000028
representing the x coordinate of the jth downstream solution point in the x direction,
Figure FDA0003267983780000029
representing the x coordinate of the jth initial value line point in the x direction;
(43) obtaining step by step half-steps along the x-direction according to an essentially oscillationless interpolation method and a limiting function
Figure FDA00032679837800000210
And
Figure FDA00032679837800000211
to obtain the state quantity of 1/2 position of step length
Figure FDA00032679837800000212
And
Figure FDA00032679837800000213
(44) according to the shock wave polar curve theory, the solution point pressure and the airflow angle of the space propulsion algorithm at the position 1/2 in the stepping length are determined through two adjacent points of which the stepping length in the space x direction is 1/2 set values, and then a Riemann self-similarity solution is obtained, wherein the expression of the state curve of the two adjacent points is as follows:
Figure FDA00032679837800000214
Figure FDA00032679837800000215
wherein phiBA state curve representing the lower side point of the initial value line, thetaBAir flow angle, f (M), representing the lower side point of the initial value lineB,αB) Rankine allowances relation, alpha, representing the lower side point of the initial value lineBDenotes the ratio of pressure at any point in the downstream space to pressure at the lower side of the initial value line, v (M)B) Prandtl-Meyer function, phi, representing the lower side point of the initial lineTA state curve representing the upper point of the initial value line, thetaTAirflow angle, f (M), representing the upper point of the initial value lineT,αT) Rankine allowances relation, alpha, representing upper side point of initial value lineTDenotes the ratio of pressure at any point in the downstream space to pressure at a point on the initial line, v (M)T) A planter-meier function representing an upper point of the initial value line;
(45) obtaining flow parameters under a single space progressive step length by carrying out differential dispersion on a compressible Euler equation, wherein the downstream differential equation along the flow x direction is in the following form:
Figure FDA0003267983780000031
wherein,
Figure FDA0003267983780000032
represents the difference value of y coordinates of two adjacent points of the initial value line,
Figure FDA0003267983780000033
the y coordinate difference of two adjacent solution points at the downstream is shown,
Figure FDA0003267983780000034
the inviscid flux in the y-direction of the upper grid point representing the length position of spatial step 1/2,
Figure FDA0003267983780000035
the slope of the upper grid edge representing the length position of spatial step 1/2,
Figure FDA0003267983780000036
the unbounded flux in the x-direction of the upper grid point representing the length position of spatial step 1/2,
Figure FDA0003267983780000037
the inviscid flux in the y-direction of the lower grid point representing the length position of spatial step 1/2,
Figure FDA0003267983780000038
the slope of the lower lattice edge representing the length position of spatial step 1/2,
Figure FDA0003267983780000039
the non-viscous flux in the x-direction of the lower grid point representing the length position of spatial step 1/2.
5. A method for predicting distribution of flow field parameters of a rotary detonation engine based on a spatial propelling algorithm according to claim 4, characterized in that the step (6) comprises the following steps:
(61) identifying the interval of the combustible mixed gas in the detonation wave front according to the flow parameter distribution of the far downstream along the y direction;
(62) determining the area average value of the flow parameters of the interval where the detonation wave front combustible mixed gas is located, wherein the area average formula is as follows:
Figure FDA00032679837800000310
Figure FDA00032679837800000311
Figure FDA0003267983780000041
Figure FDA0003267983780000042
wherein,
Figure FDA0003267983780000043
an x component representing the mean velocity of the detonation wave front,
Figure FDA0003267983780000044
y component representing mean velocity of detonation wave front, nxX component, n, representing the normal vector of the detonation wave frontyRepresents the y-component of the normal vector of the knock wave front,
Figure FDA0003267983780000045
expressing the internal energy contained in unit mass, simultaneously solving four equations of an area average formula according to a Newton-Laverson multivariate iterative method to obtain a detonation wave front flow parameter
Figure FDA0003267983780000046
And
Figure FDA0003267983780000047
a represents the area occupied by the detonation wave.
6. A method for predicting distribution of flow field parameters of a rotary detonation engine based on a spatial propelling algorithm according to claim 5, characterized in that the step (7) comprises the following steps:
(71) and correcting the inclination angle of the detonation wave relative to the injection plane according to the airflow angle of the detonation wave front, wherein the correction formula is as follows:
Figure FDA0003267983780000048
wherein,
Figure FDA0003267983780000049
indicating the inclination of the modified detonation wave relative to the injection plane,
Figure FDA00032679837800000410
representing the inclination of the uncorrected detonation wave relative to the injection plane;
(72) determining the rotation angle of a shock wave coordinate system according to the inclination angle of the detonation wave corrected by multiple iterations relative to the injection plane, and converting the outlet direction into the horizontal direction through a two-dimensional rotation matrix, wherein the rotation angle and the two-dimensional rotation matrix are in the following forms:
Figure FDA00032679837800000411
Figure FDA00032679837800000412
Figure FDA00032679837800000413
wherein,
Figure FDA00032679837800000414
representing the rotation angle, x, of the shock coordinate systemnewRepresenting the x coordinate and y coordinate of any point in the flow field after the rotation of the shock wave coordinate systemnewThe y coordinate and x of any point in the flow field after the shock wave coordinate system rotates are expressedoldRepresenting the x coordinate, y, of any point in the non-rotating flow field of the shock coordinate systemoldY-coordinate, u, representing any point in the non-rotating flow field of the shock coordinate systemnewRepresenting the velocity x component, v, of any point in the flow field after the rotation of the shock wave coordinate systemnewRepresenting the velocity y component u of any point in the flow field after the rotation of the shock wave coordinate systemoldRepresenting the x component, v, of the velocity at any point in the non-rotating flow field of the shock coordinate systemoldRepresenting the velocity y component of any point in the non-rotating flow field of the shock wave coordinate system;
(73) converting the speed in the shock wave coordinate system into the speed in the laboratory coordinate system by a speed triangle method according to the vector addition and subtraction principle, wherein the conversion relation is as follows:
Figure FDA0003267983780000051
vlab＝vnew
wherein u islabX component, v, representing the velocity at any point in the laboratory coordinate systemlabThe y component represents the velocity at any point in the laboratory coordinate system.
7. The method for predicting the distribution of the flow field parameters of the rotary detonation engine based on the spatial propelling algorithm according to claim 1, wherein the method comprises the following steps: and (4) correcting the airflow angle of the upper side interval of the detonation wave front combustible mixed gas in the step (6) to be 0 in a new iteration shock wave coordinate system.
CN202111092248.1A 2021-09-17 2021-09-17 A prediction method of flow field parameter distribution of rotary detonation engine based on space propulsion algorithm Pending CN113987694A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN202111092248.1A CN113987694A (en) 2021-09-17 2021-09-17 A prediction method of flow field parameter distribution of rotary detonation engine based on space propulsion algorithm

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN202111092248.1A CN113987694A (en) 2021-09-17 2021-09-17 A prediction method of flow field parameter distribution of rotary detonation engine based on space propulsion algorithm

Publications (1)

Publication Number Publication Date
CN113987694A true CN113987694A (en) 2022-01-28

Family

ID=79736023

Family Applications (1)

Application Number Title Priority Date Filing Date
CN202111092248.1A Pending CN113987694A (en) 2021-09-17 2021-09-17 A prediction method of flow field parameter distribution of rotary detonation engine based on space propulsion algorithm

Country Status (1)

Country Link
CN (1) CN113987694A (en)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN117688697A (en) * 2024-02-02 2024-03-12 中国人民解放军空军工程大学 Rotating detonation engine inlet design method
CN118395639A (en) * 2024-06-20 2024-07-26 中国人民解放军空军工程大学 Design method of rotary detonation engine spray pipe

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN117688697A (en) * 2024-02-02 2024-03-12 中国人民解放军空军工程大学 Rotating detonation engine inlet design method
CN117688697B (en) * 2024-02-02 2024-04-26 中国人民解放军空军工程大学 Design method of rotary detonation engine air inlet channel
CN118395639A (en) * 2024-06-20 2024-07-26 中国人民解放军空军工程大学 Design method of rotary detonation engine spray pipe

Similar Documents

Publication Publication Date Title
CN104298805B (en) Hypersonic aircraft CFD aerodynamic modeling methods
CN106508028B (en) A kind of determination complex appearance aircraft supersonic speed, the hypersonic method for having angle of attack tremor secure border
CN112417596B (en) A parallel mesh simulation method for aero-engine combustion chamber flow model
CN113987694A (en) A prediction method of flow field parameter distribution of rotary detonation engine based on space propulsion algorithm
CN107679319A (en) A kind of Algebra modeling method of circumferential pulsating stress item in through-flow model of turbine
CN104346499B (en) A kind of windy turbofan engines design method based on computer platform
Galindo et al. Development and validation of a radial variable geometry turbine model for transient pulsating flow applications
CN104834785B (en) The modeling method of aero-engine steady-state model based on simplex spline function
CN112818573A (en) Method for acquiring boundary layer non-local variable information for unstructured grid
Xianhong et al. Aerodynamic design and numerical simulation of over-under turbine-based combined-cycle (TBCC) inlet mode transition
CN106777642B (en) A kind of Forecasting Methodology of film cooling structure discharge coefficient
Lengyel et al. Design of a counter rotating fan-an aircraft engine technology to reduce noise and CO2-emissions
Godard et al. Efficient design investigation of a turbofan in distorted inlet conditions
CN109299497B (en) Simplification and Equivalent Method for Tight Exhaust Membrane Holes in Ni-based Single Crystal Turbine Blades
CN117933138A (en) Design method, device, equipment and medium of continuous wave-absorbing Mach nozzle
Ma et al. Dynamic multi-objective optimization of scramjet inlet based on small-sample Kriging model
Ishikawa et al. Near Field Sonic Boom Simulations for C608 Airplane of the Third AIAA SPW by Unstructured/Structured Overset Grid Method
Blechschmidt et al. A machine learning approach for the prediction of time-averaged unsteady flows in turbomachinery
Xie et al. Rapid supersonic performance prediction for 2D ramjet inlets
CN115470726A (en) A fast prediction method for hypersonic inlet flow field based on deep learning
Childs et al. Overflow Simulation Guidelines for Orion Launch Abort Vehicle Aerodynamic Analyses
Robles Vega et al. Part I: Validation and Verification of CFD Analysis for NASA SDT Transonic Fan Stage Test Rig
Sanders et al. CFD performance predictions of a serpentine diffuser configuration in an annular cascade facility
Lin et al. Effective boundary conditions and numerical method for flow characteristics of aeroengine compressor at high Mach flight
Horvath et al. Application of CFD numerical simulation for intake port shape design of a diesel engine

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination