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首页> 外文期刊>Journal of Spacecraft and Rockets >Wake Closure and Afterbody Heating on a Mars Sample Return Orbiter
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Wake Closure and Afterbody Heating on a Mars Sample Return Orbiter

机译：火星样品返回轨道器的尾流关闭和后加热

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摘要

Aeroheating wind-tunnel tests were conducted on a 0.028-scale model of an orbiter concept considered for a possible Mars sample return mission. The primary experimental objectives were to characterize hypersonic near-wake closure and to determine if shear layer impingement would occur on the proposed orbiter afterbody at incidence angles necessary for a Martian aerocapture maneuver. Global heat transfer mappings, surface streamline patterns, and shock shapes were obtained in the NASA Langley Research Center 20-Inch Mach 6 Air and CF_4 Tunnels for postnormal shock Reynolds numbers (based on forebody diameter) ranging from 1.4 X 10~3 to 4.15 X 10~5, angles of attack ranging from —5 to 10 deg at 0-, 3-, and 6-deg sideslip, and normal shock density ratios of 5 and 12. Laminar, transitional, and turbulent shear layer impingement on the cylindrical afterbody was inferred from the measurements and resulted in a localized heating maximum that ranged from 40 to 75% of the reference forebody stagnation point heating.
机译：空气加热风洞试验是在考虑到可能的火星样本返回任务的轨道器概念的0.028比例模型上进行的。主要的实验目的是表征高超声速近尾流闭合并确定在拟议的轨道器后体上是否会以火星航空捕获机动所需的入射角发生剪切层撞击。在NASA兰利研究中心的20英寸马赫6空气和CF_4隧道中获得了总体传热图，表面流线型态和冲击形状，其法线后冲击雷诺数（基于前体直径）范围为1.4 X 10〜3至4.15 X 10〜5，迎角在0、3和6度侧滑时从-5度到10度，正常冲击密度比为5和12。层流，过渡和湍流剪切层撞击圆柱体后车身通过测量可以推断出“最大”误差，并导致局部最大加热，范围为参考前身停滞点加热的40％至75％。

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