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研究生： 鄧先浩
Teng, Hsien-Hao
論文名稱： F16翼前緣延伸設計空氣動力學特性之數值研究
Design and Numerical Study on Aerodynamics Performance for F16 Leading Edge Extension
指導教授： 林三益
Lin, San-Yih
學位類別： 碩士
Master
系所名稱： 工學院 - 航空太空工程學系
Department of Aeronautics & Astronautics
論文出版年： 2018
畢業學年度： 106
語文別： 中文
論文頁數： 110
中文關鍵詞： 翼前緣延伸 、升阻力係數 、俯仰力矩 、渦流破裂
外文關鍵詞： LEX, lift/drag ratio, pitch moment, vortex burst
相關次數： 點閱：144  下載：24 
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    • 本文探討翼前緣延伸(Leading Edge Extension,LEX)對現代多用途戰鬥機氣動特性之影響。它可以提升戰鬥機整體空氣動力特性和機動性，使得戰機在高攻角方面具有挑戰性的氣動力控制。翼前緣延伸空氣動力學特性，近年用於軍事戰機上的效果非常顯著。當戰鬥機在高攻角狀態下，翼前緣延伸的前緣形成渦流，流過機翼上產生可用氣流來使機翼延遲失速和減少升力損失。本研究使用Dassault Systemes的高階電腦輔助設計軟體CATIA來做翼前緣延伸設計，並運用商業套裝軟體ANSYS CFX進行F16氣動力外型的外部流場之數值模擬。數值研究上使用ANSYS CFX高階數值方法，求解非穩態可壓縮流那威爾-史托客方程式(Navier-Stokes equations)來探討F16空氣動力流場特性。在紊流模型上，採用剪應力傳輸(Shear Stress Transport k-ω)紊流模型。計算網格由ANSYS ICEM產生非結構性網格，採用混合型網格，於流道壁面周圍建立棱柱型網格(Prism Mesh)來模擬邊界層黏性流場，其他計算領域則採用四面體網格(Tetrahedron Mesh )，其模擬結果分析升力係數、阻力係數、俯仰力矩係數之關係與探討F16在不穩定的流場中強渦流結構與機翼上部分相互作用、流場分離現象對機翼的升阻比、渦旋破裂位置、速度分布等之影響。改變翼前緣延伸的設計，對整個F16的空氣動力特性比原本來的佳，升阻比及渦流強度的增加，渦流破裂位置延遲。在穿音速巡航時，機翼後方將產生震波及紊流結構複雜其影響空氣動力特性，但在機翼前方設置小翼片，可破壞震波區邊界層結構，增加速度，並在高攻角時產生強勁的渦流使得整體升力增加。

    In this research, we use commercial software, ANSYS CFX, to carry out the simulations of the F16 aerodynamics flow fields. The flight with a modified Leading Edge Extension (LEX) is proposed to increase the lift/drag ratio. The Shear Stress Transport k-ω turbulent model is used. The unstructured grid system is generated by the ICEM CFD. The prism grid system around the wall surface is generated to simulate boundary layer viscosity flow field and Tetrahedron Mesh is used for the other computation domain. The lift, drag, and pitch moment are computed. The strong vortex structures upper the wing and vortex bursts under different sweep angle of LEX are investigated.

    中文摘要 i Extended Abstract iii 誌謝 viii 目錄 ix 符號說明 xvi 第一章 緒論1 1.1前言1 1.2研究動機與目的1 1.3文獻回顧2 1.4內容大鋼4 第二章 空氣動力學理論基礎6 2.1基礎理論6 2.1.1基本氣動力參數6 2.2 有限翼展7 2.2.1後掠翼8 2.3 穿音速流體8 2.4 靜態穩定與控制9 2.4.1 氣動力中心(aerodynamic center)10 2.4.中性點(Neutral point)與靜態限度(Static Margin)10 2.4.3重心位置(center of gravity)11 第三章數值方法12 3.1統御方程式12 3.2紊流模型(Turbulence Modeling) 14 3.2.1 S-A(Spalart-Allmaras)model 14 3.2.2 SST (Shear-Stress Transport)k-ω model 15 3.2.3 DES(Detached Eddy Simulation)model 18 3.3 壁面函數19 3.4 快速傅立葉轉換19 3.5 幾何外形20 3.6 網格生成20 3.7 Leading edge extension(LEX)設計20 第四章程式與物理模型驗證22 4.1 NACA0012翼剖面之流場模擬驗證22 4.1.1不可壓縮流低雷諾數流場模擬22 4.1.2不可壓縮流高雷諾數流場模擬23 4.1.3可壓縮流流場模擬24 4.1.4 NACA0012與NACA64a206穿音速流場模擬比對25 4.2 ONERA M6機翼流場之模擬驗證26 4.3DLR-F6 Wing-Body全機流場之模擬驗證27 4.4F16全機流場之模擬驗證28 4.5網格獨立性分析29 4.6時間步獨立性分析30 第五章結果與討論31 5.1低馬赫數下F16與F16翼前緣設計後之空氣動力特性比較31 5.1.1壓力變化之比較31 5.1.2升力係數變化之比較32 5.1.3LEX與機翼前緣之流場分佈33 5.1.4渦流核心切向速度之比較33 5.1.5渦流破裂位置34 5.2高馬赫數下F16與F16翼前緣設計後之空氣動力特性比較34 5.2.1升力、阻力、俯仰力矩係數變化之比較35 5.2.2渦流結構(Q-criterion)流場分佈影響36 5.3 SST k-ω與DES紊流模型之比較36 第六章結論與建議38 6.1結論38 6.2建議39 參考文獻 40 表3.1 自動體網格類型與生成方法概述45 表3.2 F16模型幾何外型數據47 表4.1 ONERA M6 機翼幾何外型數據47 表4.2 ONERA M6機翼與不同紊流模型和網格方法比對升力係數和阻力係數48 表4.3 DLR-F6 Wing-Body幾何外型數據48 表4.4 DLR-F6網格設置比較49 圖 2-1有限翼展對升力性能的影響52 圖 2-2垂直自由流影響翼面壓力分佈52 圖 2-3流體通過不同升力面所形成之漩渦形狀53 圖 2-4各種不同形狀機翼之C1及C2值53 圖 2-5穿音速機翼周圍流體分佈54 圖 3-1 F16幾何模型55 圖 3-2 F16加裝後掠角65度Canard幾何模型55 圖 3-3 F16加裝後掠角55度Canard幾何模型56 圖 4-1 NACA0012流場(a)邊界條件;(b)網格配置;(c)近壁面處網格58 圖 4-2 NACA0012雷諾數100升力係數之比對58 圖 4-3 NACA0012雷諾數100升力係數之比對59 圖 4-4 NACA0012雷諾數3E+06升力係數之比對59 圖 4-5 NACA0012雷諾數3E+06阻力係數之比對60 圖 4-6 M∞=0.8, α = 0 º, NACA0012無因次量壓力分佈60 圖 4-7 M∞=0.8, α = 0 º, NACA0012無因次量速率分佈61 圖 4-8 M∞=0.8, α = 0 º,NACA0012壓力係數Cp分佈61 圖 4-9 M∞=0.8, α = 0 º, NACA0012與NACA64a206無因次量壓力分佈 62 圖 4-10 M∞=0.8, α = 0 º,NACA0012與NACA64a206壓力係數Cp分佈 62 圖 4-11 ONERA M6 機翼幾何外型[36]63 圖 4-12 ONERA M6 機翼流體計算域幾何63 圖 4-13 ONERA M6 機翼(a)網格設置(b)近壁面處網格64 圖 4-14 ONERA M6 機翼邊界條件設定65 圖 4-15 ONERA M6 機翼壓力分佈65 圖 4-16 (a-f) ONERA M6 機翼表面上不同位置的壓力係數分佈68 圖 4-17 DLR-F6 Wing-Body幾何外型69 圖 4-18 DLR-F6 Wing-Body流體計算域幾何69 圖 4-19 DLR-F6 Wing-Body(a)網格設置(b)近壁面處網格70 圖 4-20 DLR-F6邊界條件設定71 圖 4-21 DLR-F6 Wing-Body 機翼壓力分佈71 圖 4-22 (a-g) DLR-F6 Wing-Body機翼表面上不同位置的壓力係數分佈 75 圖 4-23 DLR-F6 Wing-Body升力係數之比對75 圖 4-24 DLR-F6 Wing-Body阻力係數之比對76 圖 4-25 DLR-F6 Wing-Body俯仰力矩係數之比對76 圖 4-26 DLR-F6 Wing-Body表面流線77 圖 4-27 F16流體計算域幾何77 圖 4-28 F16(a)網格設置(b)近壁面處網格78 圖 4-29 F16邊界條件設定79 圖 4-30 F16在M∞=0.8,Re=2.5e+06，SST k-ω與S-A紊流模型與實驗數據[15]之比對(a)升力係數CL;(b)阻力係數CD;(c)俯仰力矩係數80 圖 4-31 F16,M∞=0.8, α = 6 º (a)升力係數(b)阻力係數網格獨立性分析81 圖 4- 32 F16,M∞=0.8, α = 6 º升力係數之時間步獨立性分析82 圖 5-1 F16Λ65˚非結構網格設置83 圖 5-2 F16Λ55˚非結構網格設置83 圖 5-3 F16M∞=0.2無因次量壓力分佈圖84 圖 5-4 F16Λ65˚M∞=0.2 無因次量壓力分佈圖85 圖 5-5 F16Λ55˚M∞=0.2 無因次量壓力分佈圖86 圖 5-6 F16LEX表面壓力係數(a)FS 5(m);(b)FS 5.5(m);(c)FS 6(m) 87 圖 5-7 F16Λ65˚LEX表面壓力係數(a)FS 4(m);(b)FS 4.5(m);(c)FS 5(m)88 圖 5-8 F16Λ55˚LEX表面壓力係數(a)FS 4(m);(b)FS 4.5(m);(c)FS 5(m)89 圖 5-9 F16 M∞=0.2,升力係數與相對攻角比較90 圖 5-10F16、F16Λ65˚、F16(Λ55˚), M∞=0.2,升力係數與相對攻角比較 90 圖 5-11 F16LEX與機翼前緣瞬時流線91 圖 5-12 F16Λ65˚LEX與機翼前緣瞬時流線92 圖 5-13 F16Λ55˚LEX與機翼前緣瞬時流線93 圖 5-14 F16FS7.5(m)橫切面上無因次量總壓力分佈圖94 圖 5-15 F16Λ65˚FS 7.5(m)橫切面上無因次量總壓力分佈圖95 圖 5-16 F16Λ55˚FS 7.5(m)橫切面上無因次量總壓力分佈圖96 圖 5-17 F16 LEX渦流分佈圖(x=0.43l)97 圖 5-18 F16 LEX渦流核心破裂之前切向速度(x=0.43l)97 圖 5-19 F16Λ65˚渦流分佈圖(x=0.33l)98 圖 5-20 F16Λ65˚渦流核心破裂之前切向速度(x=0.43l)98 圖 5- 21F16Λ55˚渦流分佈圖(x=0.26l)99 圖 5-22 F16Λ55˚渦流核心破裂之前切向速度(x=0.26l)99 圖 5-23 F16、F16Λ65˚、F16Λ55˚之LEX渦流破裂位置比較100 圖 5-24 F16、F16Λ65˚、F16Λ55˚, M∞=0.8,升力係數與相對攻角比較 101 圖 5-25 F16、F16Λ65˚、F16Λ55˚,M∞=0.8,阻力係數與相對攻角比較 101 圖 5-26 F16、F16Λ65˚、F16Λ55˚,M∞=0.8,俯仰力矩係數與相對攻角比較102 圖 5-27 F16周圍流動之紊流結構(Q-criterion)103 圖 5-28 F16Λ65˚周圍流動之紊流結構(Q-criterion)104 圖 5-29 F16Λ55˚周圍流動之紊流結構(Q-criterion)105 圖 5-30 M∞=0.8,F16表面上渦流分佈106 圖 5-31 M∞=0.8,F16Λ65˚表面上渦流分佈107 圖 5-32 M∞=0.8,F16Λ55˚表面上渦流分佈108 圖 5-33 SST k-ω turbulence model和DES turbulence model振幅與頻率關係圖109 圖 5-34 SST k-ω turbulence model 機翼上渦流分離分佈圖110 圖 5-35 DES turbulence model 機翼上渦流分離分佈圖110

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