穹顶形扰流柱冲击冷却系统综合换热效率数值模拟

为了分析扰流柱对冲击冷却效率的影响,采用数值模拟方法对穹顶形扰流柱冲击冷却系统进行研究,获得其换热与流动特性,并与平板靶板冲击冷却系统和圆形扰流柱冲击冷却系统进行对比分析。结果表明：穹顶形扰流柱冲击冷却系统可以同时获得良好的换热效果与较小的流动阻力系数。与圆形扰流柱靶板相比,穹顶形扰流柱靶板的Nu 增大了13.8%,而流动阻力却减小了5.3%;其综合换热效率提高了17.9%。从综合换热效率的角度看,穹顶形扰流柱冲击冷却系统优于平板靶板冲击冷却系统和圆形扰流柱冲击冷却系统。     

Experimental research in rotating wedge-shaped cooling channel with multiple non-equant holes lateral inlet

The heat transfer in a novel smooth wedge-shaped cooling channel with lateral ejection of turbine blade trailing edge is experimentally investigated in both non-rotating and rotating cases. Beside the conventional inlet at the bottom of the channel, an extra coolant injection from 8 lateral non-equant holes is introduced to improve the overall heat transfer. The total mass flow rate ratio (lateral mass flow rate/total mass flow rate) varies from 0 to 1.0. The major inlet Reynolds number and rotation number respectively vary from 10000 to 20000 and from 0 to 1.16. Experimental results show that the lateral inlet decreases local bulk temperature and increases local heat transfer at the middle and the top of the static channel. In rotating cases, the lateral inlet notably improves the heat transfer at the high-radius half channel and compensates the negative effects induced by the rotation. Both intensity and uniformity of heat transfer inside the channel are enhanced while flow resistance decreases with proper mass flow rate ratio of coolant from two inlets. The most satisfactory total mass flow rate ratio is around 2/3. This new structural style of cooling channel has huge potential and provides new direction of heat transfer of turbine blade trailing edge.     

一种两管径式脉动热管的流动与传热特性

为提高脉动热管的传热特性,提出了一种两管径式脉动热管结构,并基于质量、动量和能量守恒方程发展了适用的物理和数学模型。这种两管径式脉动热管对蒸发段和冷凝段取不同管径,两者的比值定义为直径比,应用上述理论模型分析了直径比对脉动热管运动规律和传热特性的影响。结果显示:采用两管径结构,可以有效提升脉动热管的自激振荡机制,特别是直径比小于1时的情况。而从传热特性而言,相比于传统等管径式脉动热管(直径比等于1),采用直径比小于1的结构可以使脉动热管的热阻明显减小,采用直径比大于1的结构却反而使传热特性下降。     

Numerical simulation of compression corner flows at Mach number 9

A hypersonic vehicle encounters a wide range of conditions during its complete flight regime.These flight conditions may vary from low to high Mach numbers with varying angles of attack.The near-wall viscous dissipation associated with flows at combined high Mach and Reynolds numbers leads to significant wall heat transfer rates and shear stresses.The shock wave/boundary-layer interaction results in a flow separation region,which commonly augments total pressure losses in the flow and lowers the efficiency of aerodynamic control surfaces such as fins installed on a vehicle.The standard turbulence models,when used to resolve such flows,result in incorrect separation bubble size for large separated flows.Therefore,it results in an inaccurate aerodynamic load,such as the wall pressures,skin friction distribution,and heat transfer rate.In previous studies,the application of the shock-unsteadiness correction to the standard two-equation k-ω turbulence model improved the separation bubble size leading to an accurate pressure prediction and shock definition with the assumption of constant Prandtl number.In the present work,the new shock-unsteadiness modification to the k-ω turbulence model is applied to the hypersonic compression corner flows.This new model with variable Prandtl number is based on the model parameter,which depends upon the local density ratio.The computed wall pressures,heat flux and flow field are compared to the experimental data.A parametric study is carried out by varying compression deflection angles,free stream Reynolds number and wall temperatures to compute the flow field and wall data accurately,particularly in the shock boundary layer interaction region.The new shockunsteadiness modified k-ω model with variable Prandtl number shows an accurate prediction of initial pressure rise location,pressure distribution in the plateau region and heat flux in comparison to the standard k-ω model.     

火星表面大气环境下热球风速仪的对流换热模型及试验验证

火星表面大气环境与一般轨道航天器所处的空间环境存在差异。为了实现极端环境下热模型修正、早期故障筛除、性能测试等目的,一般需要在模拟的低气压有风环境下对火星巡视器进行热试验,试验涉及在1400 Pa左右压力的环境下对0～15 m/s风速进行模拟和测量。文章针对极低气压下的风速测量问题,使用无量纲数分析方法建立恒热流式热球风速传感器表面的换热模型,对其在低气压下的输出、自然对流影响等进行分析,并与低气压下的测试结果进行对比。试验结果显示,在1400 Pa低气压下,热球风速探头表面仍以强制对流换热为主,探头灵敏度约为0.1～0.2 mV/(m·s-1),可以用于极低压力下的风速测量。     

长尾喷管传热及烧蚀特性研究

针对长尾喷管的热防护问题,采用实验和数值仿真结合的方法,研究了采用SQ-2推进剂固体火箭发动机复合结构长尾喷管的传热以及烧蚀特性.仿真程序采用基于格心的有限体积法,对流体域求解Navier-Stokes方程,对固体域求解热传导方程,湍流模型采用k-ωSST模型,通过保证热流密度大小相等、温度连续的方法实现耦合传热计算;...     

超临界压力RP-3壁面结焦对流阻的影响

实验研究了长时间加热条件下航空煤油RP-3在微细不锈钢管内流动过程中结焦对流动及换热的影响规律。实验中系统压力保持为5 MPa,燃油质量流量为3 g/s。在燃油溶解氧达到饱和的条件下,实验段进出口油温分别为130℃和450℃。实验从开始到终止持续36 h。实验结果表明,随着时间的增长,管内的结焦量不断增加。由于壁面结焦现象对管内流动和换热产生严重的影响,管内各段换热在实验前期迅速恶化逐渐趋于稳定,管内流阻随着实验时间的增加持续增长。管内流动阻力随着时间的增长呈现出\"快速增长→平稳增长→急速增长\"的过程。另外,基于实验结果,提出了一种影响系数作为判断结焦对换热器单管影响的工程模型。      

高温燃气流超声速风洞扩压器热防护设计

为实现高热流环境下高温燃气流超声速风洞系统的重复使用,采用数值方法对扩压器热环境进行了分析,采用沸腾换热方法对扩压器进行了热防护设计,并通过试验验证了沸腾换热理论应用的可行性。结果表明,沸腾换热在水流速2m/s的情况下可实现热流2MW/m 2的热防护,应用沸腾换热可以有效降低设计难度,在较低的速度和压力损失下达到更好的换热效果。