本研究探討具有NACA 0012剖面的有限長度後掠懸臂翼之流場與氣動力性能，主要物理參數包括：雷諾數、攻角及後掠角。透過表面油膜流技術觀察機翼升力面之表面流場模態，並輔以煙線流場可視化技術以了解低雷諾速時之流場特徵，再應用質點影像速度儀(Particle Image Velocimetry)，取得量化之表面流場結構之全域速度場。並使用六力平衡儀量測後掠翼之氣動力性能。在低雷諾數區域，當後掠角45o時，翼表面之流場行為有五種特徵模態產生，分別為貼附表面流、分離尾流，分離渦旋尾流、翼前緣分離尾流、鈍體尾流模態，在攻角0o時，翼面上之煙線維持平滑地通過上、下翼面，且於翼後緣處相接，隨著攻角增加，層流邊界層在機翼吸力側分離，機翼吸力側及壓力側分離的邊界層形成剪流層不穩定波。攻角持續增加，吸力側所形成之剪流層和從壓力側分離往下游發展的剪流層兩者交亙作用，產生類似Kármán形式的渦旋流逸結構，由於高攻角的關係，使得機翼迎風面尺寸增加，則機翼前緣與後緣之剪流層不穩定波不再具有交互作用，各自形成渦旋流逸的結構。在高雷諾數區域，後掠角30o、38o及45o時，發現七種模態，分別為貼附表面流、層流分離、分離泡、翼前緣分離泡、分離泡延展、分離泡破裂及鈍體尾流，隨著攻角的增加，流體開始從升力面產生分離，而層流邊界層過渡成為紊流邊界層，使其在升力面表面再接觸，而形成分離泡，分離泡隨著攻角的增加往翼前緣前進，且分離泡尺寸隨著前進而縮小，攻角繼續的增加，分離泡尺寸開始增加，最後分離泡破裂消失，且紊流邊界層在升力面產生分離。使用拓樸分析的結果發現與煙線的特徵相似。將六力平衡儀所量測的氣動力特性結果與表面流場比對，發現在機翼升力面之表面流場情形與其氣動力性能有著密切的關係。在後掠角30o、38o及45o時，機翼會在表面流場在當分離泡延展模態時產生偽失速(false stall)的氣動力特性，而當攻角持續增加，當模態進入分離泡破裂時升力則持續上升，直到攻角在30o ~ 35o間才開始失速。在升力與阻力係數的特性方面，直翼的升阻比來的比後掠翼來得高，而後掠翼之最高升阻比處並不如同後掠角在0o、15o一樣，在失速處前具有最高的升阻比，而是在偽失速前具有較高的升阻比，當後掠角大於30o時，較晚發生失速外，後掠翼45o較直翼的最大升力係數約增加了16.4%，但失速阻力係數增加到直翼的4.5倍之多，而力矩係數也是隨著後掠角的增加而往負值增加。最大升阻力係數比原來直翼減少約55%。

Effects of the back-swept angle, chord Renolds number, and root angle-of-attack on the surface flow and the aerodynamic performance of a finite back-swept wing are experimentally studied. The experiments are conducted in a wind tunnel. The cross-sectional profile of the wing model is NACA 0012 and the back-swept angle is varied from 0o to 45o. The aspect ratio of the wing is 5. The surface flow field is visualized by using the techniques of the smoke-wire and surface-oil flow. Five characteristic flow modes are found in low Renolds number regime according to the smoke-streak flow patterns and seven are categorized at high Renolds numbers by observing the patterns of the surface-oil flow. Flow topologies of the surface flow modes in various characteristic flow regimes and the time-dependent wake evolution processes are analyzed by the separatrices, alley ways, and critical points. A six component balance is employed to measure the aerodynamic performance. It is found that the aerodynamic performance of the wing and the surface-oil flow characteristic modes are closely related. When the back-swept angle is great than 30o, the angle of attack at stall would be delayed from 10o~15o of straight wing to 30o~35o of present cases. This may be induced by the increased rotation energy of the secondary flows which are generated by the back swept of the wing when the back-swept angle is greater than 30 o. Because the secondary flows create rotation energy and increase the turbulence intensity, the surface boundary layer thus has ability to resist the negative effect against the adverse pressure occurred at high angle-of-attack. The stall AOA is therefore deferred.

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