助手标题  
全文文献 工具书 数字 学术定义 翻译助手 学术趋势 更多
查询帮助
意见反馈
   三角翼 在 航空航天科学与工程 分类中 的翻译结果: 查询用时:0.041秒
图标索引 在分类学科中查询
所有学科
航空航天科学与工程
力学
仪器仪表工业
数学
更多类别查询

图标索引 历史查询
 

三角翼
相关语句
  delta wing
    The Experimental Investigation of Stall Characteristics for 75° Sweep Slender Delta Wing
    75°前缘后掠角细长三角翼失速特性实验研究
短句来源
    The effects of vectored jet on leading edge vortex breakdown of 65° delta wing
    矢量喷流对65°三角翼前缘涡破裂的影响
短句来源
    Experimental Investigation on the Leading edge Vortex Breakdown and Their Control for Flow over a 60° Delta Wing
    60°三角翼前缘涡破裂及其控制实验研究
短句来源
    Experimental Studies on Leading-Edge Vortex Breakdown and Control of Flow over 76°/40° Double Delta Wing
    76°/40°双三角翼前缘涡破裂及其控制实验研究
短句来源
    Effect of the apex flap on the leading edge vortex breakdownover a 70° delta wing
    尖顶襟翼对70°三角翼前缘涡破裂的影响
短句来源
更多       
  delta wings
    Experimental Investigation on Unsteady Aerodynamic Characteristics of Rapidly Pitching Delta Wings
    快速俯仰三角翼的非定常空气动力实验研究
短句来源
    An experimental investigation on unsteady aerodynamic characteristics was carried out for two flat--plate delta wings of aspect ratio 1. 5 and 2. 0 undergoing large amplitude pitching motions inthe 3 × 2.5 low -- speed wind tunnel at NUAA.
    展弦比为1.5和2.0的两个平板三角翼作快速俯仰振荡运动的非定常空气动力实验研究在南京航空航天大学3×2.5米低速风洞中进行。
短句来源
    A fully implicit Lower-Upper-Symmetric-Gauss-Seidel(LU-SGS) with pseudo time sub-iteration and the modified Jameson's central scheme are applied to solve thin layer Navier-Stokes(N-S) equations with laminar hypothesis and Baldwin-Lomax(B-L) model for the vortical flows around delta wings at high angles of attack(AOA).
    采用由伪时间子迭代格式实现的二阶精度LU SGS方法进行时间推进,并以Jameson中心加人工粘性格式进行空间离散,应用层流假设和Baldwin Lomax(B L)模式,求解雷诺平均的薄层Navier Stokes(N S)方程组以模拟细长三角翼大迎角流动.
短句来源
    A CALCULATING METHOD OF THE BREAKDOWN FEATURE OF THE LEADING-EDGE SEPARATING VORTEX FOR SLENDER DELTA WINGS
    细长三角翼前缘分离涡破裂特性计算方法
短句来源
    AN EXPERIMENTAL INVESTIGATION OF DELTA WINGS WITH LEADING-EDGE VORTEX SEPARATION
    三角翼前缘分离涡场的实验研究
短句来源
更多       
  a delta wing
    Study of crossflow topology of unsteady vortex motion over a delta wing
    三角翼非定常旋涡运动横截面拓扑结构研究
短句来源
    Experimental investigation of Gurney flaps on the lift enhancement of a delta wing
    三角翼Gurney襟翼增升实验研究
短句来源
    Numerical simulation of a delta wing at high angles of attack
    三角翼大迎角绕流的数值模拟
短句来源
    Vortex structure on a delta wing in unsteady free stream via Particle Image Velocimetry
    PIV测量非定常自由来流中的三角翼前缘涡
短句来源
    Study of aerodynamic characteristics of coupled-forced pitching and rolling of a delta wing
    三角翼受迫俯仰滚转耦合运动的气动特性研究
短句来源
更多       
  “三角翼”译为未确定词的双语例句
    The Effect of Outboard Leading-Edge Shape of 75°/45°Double-Delta Wing on Separated Flow Characteristics at Large Incidence
    75°/45°双三角翼外翼前缘形状对大迎角分离流动特性影响
短句来源
    An experimental study of the aerodynamic forces on double-deltawings with sideslips is presented,a =-3°~42 ,B=-20°~20 , Re=1.3×106°.
    本文对不同展弦比双三角翼有侧滑时的气动力进行了实验研究,α=--3°~42°,β=--20°~20°,雷诺数为1.3×10~6。
短句来源
    The rolling-up of the leading-edge vortex sheet and interaction with the trailing-edge vortex are simulated numerically by using the two-dimensional unsteady analogue and vortex-in-cell method.
    本文应用二维非定常比拟和恪子涡(Vortex-in-Cell)方法,数值模拟了三角翼前缘涡层的卷起以及与尾涡的相互作用。
短句来源
    Good agreement was found between the calculations and the experiments.
    对三角翼及双三角翼气动特性的计算表明,计算值与实验值符合得相当好。
短句来源
    The Reynolds number was fixed at 0. 72×106 based on the root chord length.
    以三角翼根弦为参考长度的雷诺数为0.72×106。
短句来源
更多       
查询“三角翼”译词为用户自定义的双语例句

    我想查看译文中含有:的双语例句
例句
为了更好的帮助您理解掌握查询词或其译词在地道英语中的实际用法,我们为您准备了出自英文原文的大量英语例句,供您参考。
  delta wing
Effect of mach number on hypersonic flow past a delta wing with blunt edges
      
Supersonic three-dimensional flow around a delta wing with blunted leading edges
      
We consider the problem of steady flow of an inviscid, non-heat-conducting gas about a delta wing which is spherically blunted at the nose and cylindrlcally blunded on the leading edges, at an angle of attack.
      
Numerical solution of the problem of supersonic gas flow past an arbitrary delta wing surface in the compression region
      
We study the hypersonic flow of an inviscid ideal gas past a delta wing of small aspect ratio at a finite angle of attack.
      
更多          
  delta wings
Experimental study of flow about flat delta wings at m=5 and angles of attack from 0 to 70°
      
Results are presented of an experimental study of the flow about flat delta wings with sharp and blunt leading edges and sweep angles χ=60, 70, 80° for a freestream Mach number of 5 in the range of angles of attack from 0 to 70°.
      
Hypersonic flow past delta wings of a certain class at angle of attack with an attached shock
      
Supersonic aerodynamic characteristics of delta wings at high angles of attack
      
As examples we consider delta and double-delta wings.
      
更多          
  a delta wing
Effect of mach number on hypersonic flow past a delta wing with blunt edges
      
Supersonic three-dimensional flow around a delta wing with blunted leading edges
      
We consider the problem of steady flow of an inviscid, non-heat-conducting gas about a delta wing which is spherically blunted at the nose and cylindrlcally blunded on the leading edges, at an angle of attack.
      
We study the hypersonic flow of an inviscid ideal gas past a delta wing of small aspect ratio at a finite angle of attack.
      
A delta wing with blunt edges at low angles of attack in hypersonic flow
      
更多          


This paper fully discusses the basic principle and experimental methods of determining aircraft longitudinal dynamic stability derivative during using half-model and various oscillations-free oscillation and fored oscillation. It briefly describes the self-designed and self-menufactured dynamic balance. According to the oomperison of the preliminary experimental results of delta wing model in our 250×200mm2 high speed wind tunnel with FFA62 experimental results, it is proved that this balance is useful. But...

This paper fully discusses the basic principle and experimental methods of determining aircraft longitudinal dynamic stability derivative during using half-model and various oscillations-free oscillation and fored oscillation. It briefly describes the self-designed and self-menufactured dynamic balance. According to the oomperison of the preliminary experimental results of delta wing model in our 250×200mm2 high speed wind tunnel with FFA62 experimental results, it is proved that this balance is useful. But in the course of our experiments, many problems occurred, hence the last part of this paper offers some suggestions as to the solution of them and states how we should endeavour later on.

本文较全面地讨论了当采用不同的振动法,即自由振动法和强迫振动法时,用半模型测定飞行器纵向动导数的基本原理和实验方法。介绍了自行设计、制造的动天平概况。根据在我院250×200mm~2高速风洞中对三角翼的初步测试结果,与FFA62实验结果相比较表明该动天平是适用的。但在实验中发现不少问题,故在文章最后部分提出了改进的意见和今后努力的方向。

The three-dimensional unsteady second-order non-homogeneous differential equation has been derived by superposition of a small disturbance on a given steady three-dimensional flow. Based on the assumption of high Mach numbers this second-order equation for unsteady flow reduces to a form analogous to that for steady flow. This makes it possible to solve the equation by methods used in the second-order theory for steady flows. In the course of solution the flows are constrained and corrected according to the...

The three-dimensional unsteady second-order non-homogeneous differential equation has been derived by superposition of a small disturbance on a given steady three-dimensional flow. Based on the assumption of high Mach numbers this second-order equation for unsteady flow reduces to a form analogous to that for steady flow. This makes it possible to solve the equation by methods used in the second-order theory for steady flows. In the course of solution the flows are constrained and corrected according to the PLK method, and singularities are thus eliminated. The crucial point in this procedure is to find the correct particular solutions. Two particular solutions are used. One is the approximate three-dimensional particular solution. The other is obtained under the assumption of local two-dimensionality. In addition, the uniform particular solution is given, from which the uniform second-order solutions may be obtained. Then, we have treated the unsteady problem for delta wings with low aspect ratio and supersonic leading edges. The Mach number range for application of the present theory is from supersonic to low hypersonic values with reduced frequencies up to near unity. The theoretical results derived in this work can be used to calculate the unsteady aerodynamic characteristics of wings having arbitrary airfoil sections.As experimental information for similar conditions is not yet available, we can only compare our results with those of Liu D. D. . For this reason, only the derivation for a flat delta wing oscillating at low frequencies has been carried out and an analytical expression is obtained for the first order expansion of the unsteady velocity potential. In the range of Mach numbers 4 to 8, our results are in agreement with those of Lui D. D. .It is also shown that under conditions of three-dimensional thin wings the second-order theory is valid up to Mδ=1.0, while according to application of the second-order theory to bodies of revolution by Van Dyke, the useful upper limit of M5 is only 0.7. Hence, with Mδ=0.7-1 .0, the principal non-linear effects can be calculated by our second-order theory, while for thin wings the third-order terms connected with heat transfer and entropy change can be ignored.

本文处理了超音速三元薄翼非定常问题,通过PLK法使二次解均匀有效。首先考虑零攻角或初始攻角时,已知基本定常绕流叠加高-量级的非定常小扰动流,把它线性化。本方法从健全的基本方程出发,使用高马赫数近似,将非定常二次方程化简,其形式与定常二次方程类似,因而有可能利用定常二次理论的方法求解。特解是求解的关键。鉴于精确特解的复杂性,本报告采用了一种近似特解。 本方法适于一般超音速和完全高超音速之间的马赫数区域(约3~8),折合频率可达至1左右。可较精确地估计厚度,初始攻角对非定常气动力,力矩的影响。 目前据我们所知,还没有有关实验数据,只能和一些理论结果进行比较。为此对低频有初始攻角的超音速前缘平板三角翼进行了计算,在马赫数3~8,与D.D.Liu~[6]比较吻合。计算结果表明,三元薄翼二次理论可用到高超音速相似参数Mδ=1.0。

Velocity gradient,pressure gradient and circulation gradient are used to model the separated vortex flow field over a leading edge delta wing.The deceleration or adversed pressure of out-flow in the direction of x-axis promotes the breakdown of the vortex core,so does the larger circulation gradient in the direction of the axis.The induced larger adversed pressure gradient of the vortex core in the direction of axis or larger radial pressure difference results in the vortex breakdown,so the idea is more completely...

Velocity gradient,pressure gradient and circulation gradient are used to model the separated vortex flow field over a leading edge delta wing.The deceleration or adversed pressure of out-flow in the direction of x-axis promotes the breakdown of the vortex core,so does the larger circulation gradient in the direction of the axis.The induced larger adversed pressure gradient of the vortex core in the direction of axis or larger radial pressure difference results in the vortex breakdown,so the idea is more completely obtained about the vortex breakdown mechanism.The axisymmetric vortex breakdown can not take place until the Reynolds number becomes larger.

用速度梯度、压力梯度和环量梯度模拟大攻角尖前缘三角翼流场。外流场轴向减速和轴向逆压促进涡核的破裂,大的轴向环量梯度也促进涡核的破裂。分离涡的破裂是由于涡核内部诱导过大的轴向逆压或径向压差引起的,因而对涡破裂机理得到更完整的认识。在较大雷诺数条件下,才有可能出现轴对称型破裂涡。

 
<< 更多相关文摘    
图标索引 相关查询

 


 
CNKI小工具
在英文学术搜索中查有关三角翼的内容
在知识搜索中查有关三角翼的内容
在数字搜索中查有关三角翼的内容
在概念知识元中查有关三角翼的内容
在学术趋势中查有关三角翼的内容
 
 

CNKI主页设CNKI翻译助手为主页 | 收藏CNKI翻译助手 | 广告服务 | 英文学术搜索
版权图标  2008 CNKI-中国知网
京ICP证040431号 互联网出版许可证 新出网证(京)字008号
北京市公安局海淀分局 备案号:110 1081725
版权图标 2008中国知网(cnki) 中国学术期刊(光盘版)电子杂志社