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Characteristics of a coherent jet enshrouded in a supersonic fuel gas 预览
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作者 Fei Zhao Rong Zhu Wen-rui Wang 《矿物冶金与材料学报:英文版》 SCIE EI CAS CSCD 2020年第2期173-180,共8页
Based on a current coherent jet,this study proposes a supersonic combustion(SC)coherent jet in which the main oxygen jet is surrounded by a supersonic fuel gas.The characteristics of the proposed coherent jet are anal... Based on a current coherent jet,this study proposes a supersonic combustion(SC)coherent jet in which the main oxygen jet is surrounded by a supersonic fuel gas.The characteristics of the proposed coherent jet are analyzed using experimental methods and numerical simulations in the high-temperature environment of electric arc furnace(EAF)steelmaking.The SC coherent jet achieved stable combustion in the EAF steelmaking environment.The simulated combustion temperature of the supersonic shrouding methane gas was 2930 K,slightly below the theoretical combustion temperature of methane–oxygen gas.The high speed and temperature of the supersonic flame effectively weakened the interaction between the main oxygen jet and the external ambient gas,inhibiting the radial expansion of the main oxygen jet and maintaining its high speed and low turbulence over a long distance.These features improved the impact capacity of the coherent jet and strengthened the stirring intensity in the EAF bath. 展开更多
关键词 EAF steelmaking coherent jet supersonic shrouding fuel gas supersonic combustion field characteristics
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Integrated supersonic wind tunnel nozzle
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作者 Junmou SHEN Jingang DONG +4 位作者 Ruiqu LI Jiang ZHANG Xing CHEN Yongming QIN Handong MA 《中国航空学报:英文版》 SCIE EI CAS CSCD 2019年第11期2422-2432,共11页
In supersonic wind tunnels, the airflow at the exit of a convergent-divergent nozzle is affected by the connection between the nozzle and test section, because the connection is a source of disturbance for supersonic ... In supersonic wind tunnels, the airflow at the exit of a convergent-divergent nozzle is affected by the connection between the nozzle and test section, because the connection is a source of disturbance for supersonic flow and the source of disturbance generated by this disturbance propagates downstream. In order to avoid the disturbance, the test can only be carried out in the rhombus area. However, for the supersonic nozzle, the rhombus region is small, limiting the size and attitude angle of the test model. An integrated supersonic nozzle is a nozzle and a test section as a whole, which is designed to weaken or eliminate the disturbance. The inviscid contour of the supersonic nozzle is based on the method of characteristics. A new curve is formed by the smooth connection between the inviscid contour and test section, and the boundary layer is corrected for the overall curve. Integrated supersonic nozzles with Mach number 1.5 and 2 are designed, which are based on this method. The flow field is validated by numerical and experimental results. The results of the study highlight the importance of the connection about the nozzle outlet and test section. They clearly show that the wave system does not exist at the exit of the supersonic nozzle, and the flow field is uniform throughout the test section. 展开更多
关键词 Boundary layer DISTURBANCE Flow field Integrated supersonic nozzle Supersonic wind tunnels The rhombus region
基于RBF神经网络和遗传算法的超声速Licher双翼优化设计研究 预览
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作者 马博平 王刚 +1 位作者 叶坤 叶正寅 《航空科学技术》 2019年第9期73-80,共8页
基于Busemann双翼的设计方法,采用径向基函数神经网络(Radial-Basis Function Neural Network,RBFNN)和基于遗传算法(Genetic Algorithm,GA)的优化技术对Licher双翼进行了优化设计以提高设计马赫数情况下的升阻比。通过计算流体力学(Com... 基于Busemann双翼的设计方法,采用径向基函数神经网络(Radial-Basis Function Neural Network,RBFNN)和基于遗传算法(Genetic Algorithm,GA)的优化技术对Licher双翼进行了优化设计以提高设计马赫数情况下的升阻比。通过计算流体力学(Computational Fluid Dynamics,CFD)方法在无黏性和黏性模式下对优化设计结果进行了验证。结果表明,与典型的Busemann双翼相比,优化后的双翼构型在无黏模拟情况下的升力和升阻比分别提高了27.3%和27.4%,黏性模拟情况下则提升了近60%和40%,表明本文采用的方法对于将双翼构型应用于未来超声速运输机领域具有很大的潜力。 展开更多
关键词 超声速 双翼 升阻比 Busemann双翼 Licher双翼
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Experimental study on nanosecond pulsed pin-to-plate discharge in supersonic air flow
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作者 李益文 庄重 +3 位作者 庞磊 段朋振 丁志文 张百灵 《等离子体科学与技术:英文版》 SCIE EI CAS CSCD 2019年第6期122-131,共10页
Development of magnetohydrodynamic acceleration technology is expected to improve wind tunnel simulation capability and testing capability.The underlying premise is to produce uniform and stable plasma in supersonic a... Development of magnetohydrodynamic acceleration technology is expected to improve wind tunnel simulation capability and testing capability.The underlying premise is to produce uniform and stable plasma in supersonic air flow,and gas discharge is an effective way to achieve this.A nanosecond pulsed discharge experimental system under supersonic conditions was established,and a pin-to-plate nanosecond pulsed discharge experiment in Mach 2 air flow was performed to verify that the proposed method produced uniform and stable plasma under supersonic conditions.The results show that the discharge under supersonic conditions was stable overall,but uniformity was not as good as that under static conditions.Increasing the number of pins improved discharge uniformity,but reduced discharge intensity and hence plasma density.Under multi-pin conditions at 1000Hz,the discharge was almost completely corona discharge,with the main current component being the displacement current,which was smaller than that under static conditions. 展开更多
关键词 nanosecond pulse DISCHARGE SUPERSONIC air flow pin-to-plate DISCHARGE plasma UNIFORM and stable
内置隔板和前缘喷流对机翼油箱热防护的影响 预览
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作者 赵璇 孙智 孙建红 《南京理工大学学报》 CAS CSCD 北大核心 2019年第6期784-792,共9页
为了研究飞机在超声速巡航状态下的气动热效果及对机翼进行相应的热防护设计,该文分析了内置隔板和前缘反向喷流2种不同热防护的防热效果。采用Spalart-Allmaras(SA)湍流模型求解Navier-Stokes方程,对NACA23012翼型在不同马赫数下的气... 为了研究飞机在超声速巡航状态下的气动热效果及对机翼进行相应的热防护设计,该文分析了内置隔板和前缘反向喷流2种不同热防护的防热效果。采用Spalart-Allmaras(SA)湍流模型求解Navier-Stokes方程,对NACA23012翼型在不同马赫数下的气动加热进行了数值分析,并对机翼油箱的热防护进行了研究。结果表明:2种热防护的降温及防热效果均会随来流马赫数的增加而增强;翼型的不对称性会导致上下翼面流场特性的不同,从而产生上下翼面热防护特性的明显差异;当来流马赫数较大时,在降低机翼表面温度及油箱热负荷方面,喷流作为主动式热防护的防热效果要明显优于被动式隔板热防护的防热效果;当来流马赫数为3.0,喷流流量为0.0018 m 3/s时,机翼下翼面平均温度可降低约46.7%。 展开更多
关键词 内置隔板 前缘喷流 机翼 油箱 热防护 超声速 气动加热
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Boundary-layer transition of advanced fighter wings at high-speed cruise conditions
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作者 Yiming DU Zhenghong GAO +1 位作者 Chao WANG Qianhuan HUANG 《中国航空学报:英文版》 SCIE EI CAS CSCD 2019年第4期799-814,共16页
The achievement of laminar flow in the boundary layer at high-speed cruise conditions may further, in addition to shock-wave control, reduce the drag and extend the range of military fighter aircraft. To this end, a f... The achievement of laminar flow in the boundary layer at high-speed cruise conditions may further, in addition to shock-wave control, reduce the drag and extend the range of military fighter aircraft. To this end, a further investigation on transitional boundary-layer flow of fighter wings is needed due to different configurations from the wings used on conventional transport aircraft. In this paper, wind tunnel experiments and numerical simulations were conducted on three-dimensional transition of thin diamond-shaped wings used on advanced fighter aircraft at tran/supersonic design points. A newly proposed correlation of crossflow transition which includes the effect of surface roughness was introduced into the c-Rehttransition model. Predicted results were in good agreement with flow visualizations. Results showed that the strength of the crossflow component grew rapidly around the leading edge because of the severe flow acceleration, just as the same as wings with a large aspect ratio. However, there seemed no regular pattern of instabilitydominance variation in span-wise for a diamond configuration. The dominance of different instability mechanisms strongly depended on the local pressure distribution. Hereby, the research recommended a ‘‘roof- like ' shape of pressure distribution to suppress both crossflow and Tollmien-Schlichting(T-S) instabilities. Besides, a sharp suction peak with a serious pressure rise should be cut off to avoid stronger instabilities. Further discussions also revealed an independence of the unit Reynolds number when transition was triggered by T-S instabilities. Aerodynamic force comparisons indicated that further benefit on drag reduction could be expected by including the three-dimensional transition effect into a wing design process. 展开更多
关键词 Boundary layer TRANSITION FIGHTER AIRCRAFT design Supersonic AIRCRAFT WINGS TRANSONIC wing AERODYNAMICS Wind tunnel measurements
三组分气体超声速凝结过程数值模拟与实验研究 预览 被引量:2
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作者 曹学文 边江 +2 位作者 靳学堂 尹鹏博 杨文 《石油学报(石油加工)》 EI CAS CSCD 北大核心 2019年第1期99-110,共12页
结合液滴成核与生长模型,以及气、液流动控制方程建立了超声速凝结流动数学模型,对空气+水+乙醇三组分(双可凝)气体超声速流动条件下凝结特性进行了数值计算,研究了三组分气体超声速凝结特性影响因素,通过与空气+水双组分(单可凝)气体对... 结合液滴成核与生长模型,以及气、液流动控制方程建立了超声速凝结流动数学模型,对空气+水+乙醇三组分(双可凝)气体超声速流动条件下凝结特性进行了数值计算,研究了三组分气体超声速凝结特性影响因素,通过与空气+水双组分(单可凝)气体对比,分析了第二种可凝组分对凝结成核的影响,并开展了实验验证与对比分析。结果表明:随着三组分气体中乙醇含量的升高,Laval喷管内成核率、液滴数均增大,但成核区收窄,液滴生长区向前移动;在入口可凝气体为饱和状态下,升高入口温度与压力均能促进凝结的发生,使Wilson点向喉部移动,进而提高出口气体湿度;与双组分气体相比,三组分气体发生凝结的Wilson点更靠近喉部,出口湿度更大,说明三组分气体发生凝结时,两种可凝气体的凝结过程是相互促进的;Laval喷管沿程压力及Wilson点测试结果与数值计算结果吻合较好,说明所建立的数学模型具有较高的准确性。 展开更多
关键词 三组分气体 超声速 凝结 数值模拟 实验
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纵向沟槽面超声速阻力特性大涡模拟 预览
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作者 余奕甫 王兵 王强 《航空计算技术》 2019年第5期49-53,共5页
针对超声速壁面边界层流动问题,提出一种基于纵向梯形沟槽薄膜的减阻方法。利用基于有限差分的程序对纵向梯形沟槽壁面超声速流场进行数值模拟,对于湍流的模拟采用大涡模拟(LES)方法。通过对比超声速平面壁面流动结构与超声速梯形沟槽... 针对超声速壁面边界层流动问题,提出一种基于纵向梯形沟槽薄膜的减阻方法。利用基于有限差分的程序对纵向梯形沟槽壁面超声速流场进行数值模拟,对于湍流的模拟采用大涡模拟(LES)方法。通过对比超声速平面壁面流动结构与超声速梯形沟槽壁面边界层流动结构特征的区别,分析减阻特性,并给出了来流Mach数对给定沟槽壁面减阻效率的影响情况。结果表明,超声速流动情况下,对于一种特定构型的纵向梯形沟槽面,存在一个具有良好减阻效率的来流Mach数区间,通过分析湍流边界层流场结构特征,总结出该梯形沟槽面的减阻规律。特别地,当来流Mach数M=2.0时,该梯形沟槽能够产生2.18%的最大减阻效率。 展开更多
关键词 大涡模拟 超声速 纵向沟槽面 阻力 减阻效率
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高超声速气动热数值模拟的网格模式相关性研究 预览
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作者 吕水燕 张传侠 +1 位作者 叶坤 徐健 《兵器装备工程学报》 CAS 北大核心 2019年第3期82-86,共5页
采用双椭球体的经典模型为研究对象,在8. 04Ma和10. 02Ma的工况下进行了气动热数值模拟,计算采用结构网格和非结构网格模式,分别获得了相应工况下双椭球体中心子午线上的热流密度,得到了网格模式相关性规律。计算结果表明:两种网格均与... 采用双椭球体的经典模型为研究对象,在8. 04Ma和10. 02Ma的工况下进行了气动热数值模拟,计算采用结构网格和非结构网格模式,分别获得了相应工况下双椭球体中心子午线上的热流密度,得到了网格模式相关性规律。计算结果表明:两种网格均与试验值热流分布趋势较为一致;非结构网格的Y+优于结构网格时,在表面大范围区域内可以获得比结构网格更优的计算结果;非结构网格在驻点处热流计算误差超过100%。该计算对高超声速飞行器的热流计算起到了重要的参考作用。 展开更多
关键词 高超声速气动热 双椭球体 热流密度 网格模式相关性 CFD
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Inverse design of low boom configurations using proper orthogonal decomposition and augmented Burgers equation
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作者 Yidian ZHANG Jiangtao HUANG +2 位作者 Zhenghong GAO Chao WANG Bowen SHU 《中国航空学报:英文版》 SCIE EI CAS CSCD 2019年第6期1380-1389,共10页
Mitigation of sonic boom to an acceptable stage is a key point for the next generation of supersonic transports. Meanwhile, designing a supersonic aircraft with an ideal ground signature is always the focus of researc... Mitigation of sonic boom to an acceptable stage is a key point for the next generation of supersonic transports. Meanwhile, designing a supersonic aircraft with an ideal ground signature is always the focus of research on sonic boom reduction. This paper presents an inverse design approach to optimize the near-field signature of an aircraft, making it close to the shaped ideal ground signature after the propagation in the atmosphere. Using the Proper Orthogonal Decomposition(POD) method, a guessed input of augmented Burgers equation is inversely achieved. By multiple POD iterations, the guessed ground signatures successively approach the target ground signature until the convergence criteria is reached. Finally, the corresponding equivalent area distribution is calculated from the optimal near-field signature through the classical Whitham F-function theory. To validate this method, an optimization example of Lockheed Martin 1021 is demonstrated. The modified configuration has a fully shaped ground signature and achieves a drop of perceived loudness by 7.94 PLdB. This improvement is achieved via shaping the original near-field signature into wiggles and damping it by atmospheric attenuation. At last, a nonphysical ground signature is set as the target to test the robustness of this inverse design method and shows that this method is robust enough for various inputs. 展开更多
关键词 Aeroacoustics AUGMENTED BURGERS equation LOW BOOM configuration Optimization Supersonic AERODYNAMICS
“音爆云”现象与“音爆”有关系吗? 预览
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作者 张华 《力学与实践》 北大核心 2019年第2期239-243,共5页
'音爆云'实质是飞机飞过高湿度空气时因局部流场加速使温度降低至露点以下而形成的水汽凝结云团,将其称为凝结云更为恰当。将凝结云按飞行速度和不同特征分成3类:低亚声速不规则凝结云、高亚声速锥形凝结云和超声速凝结云。本... '音爆云'实质是飞机飞过高湿度空气时因局部流场加速使温度降低至露点以下而形成的水汽凝结云团,将其称为凝结云更为恰当。将凝结云按飞行速度和不同特征分成3类:低亚声速不规则凝结云、高亚声速锥形凝结云和超声速凝结云。本文分别讨论了3类凝结云的形成机制、不同特点及其与激波、突破'声障'和'声爆'的关系。第1类与'声爆'无关;第2类伴随局部弱激波、未突破'声障',与'声爆'基本无关;第3类则与超声速激波及其'声爆'有关。 展开更多
关键词 音爆云 声障 声爆 激波 超声速
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超音速雾化喷枪的设计与分析 预览
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作者 焦峥辉 王美妍 +2 位作者 任雪娇 张小全 薄文露 《现代制造技术与装备》 2019年第12期79-80,95,共3页
空气喷枪作为一种重要的喷涂工具,在各行各业得到了广泛应用。为了提高液体的雾化效果,结合目前喷枪的结构和工作原理,设计了一种新型超音速雾化喷枪。该喷枪重新设计结构,并对设计的喷枪内部的流场进行仿真模拟,明确了结构设计的合理... 空气喷枪作为一种重要的喷涂工具,在各行各业得到了广泛应用。为了提高液体的雾化效果,结合目前喷枪的结构和工作原理,设计了一种新型超音速雾化喷枪。该喷枪重新设计结构,并对设计的喷枪内部的流场进行仿真模拟,明确了结构设计的合理性及其雾化效果,可为雾化喷枪的设计与研究提供重要的参考作用。 展开更多
关键词 超音速 雾化 拉法尔喷管 喷枪
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Effects of dimensional wall temperature on velocity-temperature correlations in supersonic turbulent channel flow of thermally perfect gas
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作者 XiaoPing Chen XinLiang Li ZuChao Zhu 《中国科学:物理学、力学、天文学英文版》 SCIE EI CAS CSCD 2019年第6期64-74,共11页
Direct numerical simulations of temporally evolving supersonic turbulent channel flows of thermally perfect gas are conducted at Mach number 3.0 and Reynolds number 4800 for various values of the dimensional wall temp... Direct numerical simulations of temporally evolving supersonic turbulent channel flows of thermally perfect gas are conducted at Mach number 3.0 and Reynolds number 4800 for various values of the dimensional wall temperature to study the influence of the latter on the velocity-temperature correlations. The results show that in a fully developed turbulent channel flow, as the dimensional wall temperature increases, there is little change in the mean velocity, but the mean temperature decreases. The mean temperature is found to be a quadratic function of the mean velocity, the curvature of which increases with increasing dimensional wall temperature. The concept of 'recovery enthalpy' provides a connection between the mean velocity and the mean temperature, and is independent of dimensional wall temperature. The right tails of probability density function of the streamwise velocity fluctuation grows with increasing dimensional wall temperature. The dimensional wall temperature does not have a significant influence on the Reynolds analogy factor or strong Reynolds analogy(SRA). The modifications of SRA by Huang et al. and Zhang et al. provide reasonably good results, which are better than those of the modifications by Cebeci and Smith and by Rubesin. 展开更多
关键词 direct numerical simulation velocity-temperature correlation SUPERSONIC FLOW channel FLOW thermally PERFECT GAS strong Reynolds analogy
超声速喷流混合流场大涡模拟 被引量:1
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作者 朱志斌 程晓丽 潘宏禄 《航空动力学报》 EI CAS CSCD 北大核心 2019年第1期210-216,共7页
以光学窗口外冷喷流为研究背景,采用大涡模拟方法对后台阶外形切向喷流混合流场进行了研究。数值方法基于隐式亚格子模型,采用高精度WENO格式进行空间离散,并通过超声速平面混合层流动对数值方法进行了考核验证。喷流混合流场计算模型... 以光学窗口外冷喷流为研究背景,采用大涡模拟方法对后台阶外形切向喷流混合流场进行了研究。数值方法基于隐式亚格子模型,采用高精度WENO格式进行空间离散,并通过超声速平面混合层流动对数值方法进行了考核验证。喷流混合流场计算模型与试验一致,来流和喷流马赫数分别为3.4和2.5。数值模拟清晰地捕捉到了流场波系以及混合剪切层、壁面边界层等典型流场结构,并精细预测了混合层发生失稳、转捩及发展为充分发展湍流的时空发展过程。数值模拟得到的湍流大尺度结构的位置和形态与实验图像一致。通过对瞬时流场、统计平均流场和脉动参数的分析,揭示了流场结构特征及其时空演化规律,并获得了流场密度脉动特性。 展开更多
关键词 超声速 冷却喷流 剪切层 转捩 湍流 大涡模拟
钝头双锥喷流致冷流场结构及密度脉动特性
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作者 朱志斌 潘宏禄 程晓丽 《航空动力学报》 EI CAS CSCD 北大核心 2019年第7期1425-1432,共8页
采用大涡模拟方法对钝头双锥喷流致冷流场开展了数值模拟,研究了超声速喷流混合流场结构特征及密度脉动特性。大涡模拟方法基于隐式亚格子模型,空间离散采用高精度通量限制型紧致格式,时间推进采用显式Runger-Kutta方法。数值模拟清晰... 采用大涡模拟方法对钝头双锥喷流致冷流场开展了数值模拟,研究了超声速喷流混合流场结构特征及密度脉动特性。大涡模拟方法基于隐式亚格子模型,空间离散采用高精度通量限制型紧致格式,时间推进采用显式Runger-Kutta方法。数值模拟清晰地捕捉到了流场波系结构,精细地预测了流动发生失稳、转捩以及发展为充分发展湍流的物理过程,直接获得了流场密度脉动特性。通过有、无喷流状态对称面流场的对比,发现超声速喷流能够有效冷却光学窗口;喷流与主流形成的混合层不稳定,很快发生失稳和转捩,形成大尺度湍流结构,进而引起强烈的密度脉动。此外,获得了钝头双锥整体模型喷流致冷流场的空间发展形态特征。 展开更多
关键词 超声速 喷流致冷 混合层 转捩/湍流 大涡模拟
天然气超声速脱二氧化碳技术研究 预览
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作者 傅健 袁汝华 +4 位作者 李大全 王军 郑松贤 李友行 赵西廓 《当代化工》 CAS 2019年第10期2240-2244,共5页
进行了Laval喷管的结构设计.基于真实气体状态方程和湍流方程,结合凝结成核与液滴生长理论,建立了描述喷管内超声速气体凝结流动的数学模型,进行了CO2-CH4气体凝结流动规律研究.研究结果表明:在特定的入口温度与压力条件下可以实现CO2... 进行了Laval喷管的结构设计.基于真实气体状态方程和湍流方程,结合凝结成核与液滴生长理论,建立了描述喷管内超声速气体凝结流动的数学模型,进行了CO2-CH4气体凝结流动规律研究.研究结果表明:在特定的入口温度与压力条件下可以实现CO2气体的凝结与脱除,当气体发生凝结后,喷管内形成气、液两相流动,产生的亚微米级微小液滴可随气流运动至喷管出口;CO2气体成核过程在时间和空间上表现出急剧性,凝结核心形成后,液滴生长过程可维持较长时间和距离,直至液滴到达喷管出口;由于凝结的发生和液滴生长过程释放了大量潜热,喷管内表现出明显的凝结冲波现象,压力下降减缓,温度出现回升. 展开更多
关键词 天然气 二氧化碳 喷管 超声速 凝结
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匀速俯仰运动及角速率对横向喷流的影响
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作者 赖江 赵忠良 +2 位作者 王晓冰 李浩 李玉平 《航空学报》 EI CAS CSCD 北大核心 2019年第10期65-74,共10页
为研究运动对横向喷流干扰特性的影响,数值模拟了导弹模型匀速俯仰运动过程的超声速横向喷流,获取了运动状态下的横向喷流干扰量,并对比分析了俯仰运动和角速率对喷口附近流场结构、模型表面极限流线、表面压力分布和子午线压力变化及... 为研究运动对横向喷流干扰特性的影响,数值模拟了导弹模型匀速俯仰运动过程的超声速横向喷流,获取了运动状态下的横向喷流干扰量,并对比分析了俯仰运动和角速率对喷口附近流场结构、模型表面极限流线、表面压力分布和子午线压力变化及气动特性和干扰放大因子造成的影响。结果表明:模拟参数范围内,动态及角速率影响随运动方向及迎角范围而发生变化;中小迎角时主要影响上游分离区和尾部偏折效应,大迎角时弓形激波位置变化显著;俯仰运动的气动特性和横向喷流干扰特性出现动态迟滞,且随角速率增加而增强;动态大迎角下由于压力平台效应减弱,其力矩放大因子受俯仰运动影响更为明显,出现偏离静态的不利结果。 展开更多
关键词 导弹模型 横向喷流 超声速 动态气动特性 放大因子
子弹稳定装置翼片动态张开过程研究 预览
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作者 陈佩银 杨永刚 +1 位作者 薛再清 张韩宇 《导弹与航天运载技术》 CSCD 北大核心 2019年第1期24-29,35共7页
为解决子母战斗部抛撒后子弹在超声速飞行时其稳定装置翼片动态张开过程的工程计算问题,基于气体斜激波理论和普朗特-梅耶膨胀波理论,分析在高速气流条件下翼片张开过程中的受力情况,建立了翼片动态张开过程的动力学方程,并用MATLAB编... 为解决子母战斗部抛撒后子弹在超声速飞行时其稳定装置翼片动态张开过程的工程计算问题,基于气体斜激波理论和普朗特-梅耶膨胀波理论,分析在高速气流条件下翼片张开过程中的受力情况,建立了翼片动态张开过程的动力学方程,并用MATLAB编程实现。用此方法计算了某子弹稳定装置在一定初始条件下翼片的运动,分析了不同子弹攻角和角速度对子弹翼片张开过程的影响,并就其对工程设计的影响进行了定性讨论。计算结果可为子弹稳定装置的飞行强度计算提供依据. 展开更多
关键词 子弹稳定装置 超声速 翼片 动态张开
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南钢110t转炉顶吹氧枪喷吹特性模拟与应用研究 预览
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作者 陈兴华 胡志勇 +3 位作者 朱荣 刘福海 姚柳洁 冯强 《河南冶金》 2019年第2期6-8,30共4页
本研究利用Fluent软件对南钢110t转炉顶吹氧枪喷头参数对超音速射流流场分布特性影响进行三维数值模拟,并将研究结果应用到南钢110t转炉常规冶炼过程。研究结果表明,13.5°四孔氧枪在1.7m处(理论枪位)保持较高的射流速度,且射流有... 本研究利用Fluent软件对南钢110t转炉顶吹氧枪喷头参数对超音速射流流场分布特性影响进行三维数值模拟,并将研究结果应用到南钢110t转炉常规冶炼过程。研究结果表明,13.5°四孔氧枪在1.7m处(理论枪位)保持较高的射流速度,且射流有效冲击半径最大。即13.5°四孔氧枪可有效提高氧气与熔池的接触面积,提高氧气利用效率。基于412炉次冶炼数据结果发现,相比于原氧枪,在采用设计流量为24000m^3/h,喷孔夹角13.5°的优化后氧枪时,在相同冶炼条件下,110t钢水的平均冶炼时间及终点碳氧积分别减小1.5min及0.0003,熔池脱磷率提高4.1%,终渣TFe含量下降1.7%。 展开更多
关键词 复吹转炉 氧枪 数值模拟 超音速
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Numerical study of convective heat transfer of a supersonic combustor with varied inlet flow conditions 预览
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作者 W.H.Fan F.Q.Zhong +1 位作者 S.G.Ma X.Y.Zhang 《力学学报:英文版》 SCIE EI CAS CSCD 2019年第5期943-953,共11页
Characteristics of convective heat transfer of a supersonic model combustor with variable inlet flow conditions were studied by numerical simulation in this paper.The three-dimensional flow and wall heat flux at diffe... Characteristics of convective heat transfer of a supersonic model combustor with variable inlet flow conditions were studied by numerical simulation in this paper.The three-dimensional flow and wall heat flux at different air inlet Mach numbers of 2.2,2.8 and 3.2 were studied numerically with Reynolds-averaged Navier-Stokes equations with a shear-stress transport(SST)k-ωturbulence model and a three-step reaction model.Meanwhile,ethylene was chosen as the fuel,and the fixed fuel-to-air equivalence ratio is 0.8 in all cases in this paper.The results of the simulations indicate that wall heat flux distribution of the combustor is very non-uniform with several peaks of wall heat flux at varied locations.For the low inlet Mach number of 2.2,a shock train structure is formed in the isolator,and three peaks of wall heat flux are located respectively on the backward face of the cavity,on the side wall near the fuel injection and on the bottom wall near the injection holes,and a maximum wall heat flux reaches 5.4 MW/m2.For the medium inlet Mach number of 2.8,there exists a much shorter shock structure with three peaks of wall heat flux similar to that of Mach number 2.2.However,as the inlet Mach number increased to 3.2,there is no shock structure upstream of fuel injections,and the combustor flow is in a supersonic mode with different locations and values of wall heat flux peaks.The statistical results of wall heat loading show that the change of total wall heat is not monotonic with the increase of inlet Mach number,and the maximum appears in the case of Mach number being 2.8.Meanwhile,for all the cases,the bottom wall takes up more than 50%of the total heat loading. 展开更多
关键词 Wall heat flux Numerical simulation ETHYLENE Supersonic combustor
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