2、畢業(yè)設計（論文）外文參考資料及譯文封面

1、<p>　　畢 業(yè) 設 計（論 文）外 文 參 考 資 料 及 譯 文</p><p>　　譯文題目： 非定常翼型繞流的間歇前緣渦輪脫落標準離散渦方法 </p><p>　　學生姓名： 學　　號： </p><p>　　?！　I(yè)：

2、 </p><p>　　所在學院： </p><p>　　指導教師： </p><p>　　職　　稱： </p&g

3、t;<p><b>　　年 月 日</b></p><p><b>　　說明：</b></p><p>　　要求學生結合畢業(yè)設計（論文）課題參閱一篇以上的外文資料，并翻譯至少一萬印刷符（或譯出3千漢字）以上的譯文。譯文原則上要求打印（如手寫，一律用400字方格稿紙書寫），連同學校提供的統(tǒng)一封面及英文原文裝訂，于畢業(yè)設計（論文

4、）工作開始后2周內完成，作為成績考核的一部分。</p><p>　　Discrete-vortex method with novel shedding</p><p>　　criterion for unsteady aerofoil flows with</p><p>　　intermittent leading-edge vortex shedding<

5、;/p><p>　　Kiran Ramesh 1, ?, Ashok Gopalarathnam 1 , Kenneth Granlund 2 ,</p><p>　　Michael V. Ol 2 and Jack R. Edwards 1</p><p>　　1 Department of Mechanical and Aerospace Engineering,

6、North Carolina State University, Raleigh,NC 27695-7910, USA2 US Air Force Research Laboratory, Air Vehicles Directorate, AFRL/RBAL, Building 45,2130 8th Street, WPAFB, OH 45433, USA(Received 18 November 2013; revised 15

7、May 2014; accepted 21 May 2014;first published online 23 June 2014)</p><p>　　Unsteady aerofoil flows are often characterized by leading-edge vortex (LEV)shedding. While experiments and high-order computation

8、s have contributed to ourunderstanding of these flows, fast low-order methods are needed for engineeringtasks. Classical unsteady aerofoil theories are limited to small amplitudes andattached leading-edge flows. Discrete

9、-vortex methods that model vortex sheddingfrom leading edges assume continuous shedding, valid only for sharp leading edges,</p><p>　　or shedding governed by ad-hoc criteria such as a critical angle of attac

10、k, validonly for a restricted set of kinematics. We present a criterion for intermittent vortexshedding from rounded leading edges that is governed by a maximum allowableleading-edge suction. We show that, when using uns

11、teady thin aerofoil theory, thisleading-edge suction parameter (LESP) is related to the A 0 term in the Fourier seriesrepresenting the chordwise variation of bound vorticity. Furthermore, for any aerofoiland</p>&

12、lt;p>　　of the motion kinematics. When the instantaneous LESP value exceeds the criticalvalue, vortex shedding occurs at the leading edge. We have augmented a discrete-time,arbitrary-motion, unsteady thin aerofoil theo

13、ry with discrete-vortex shedding from theleading edge governed by the instantaneous LESP. Thus, the use of a single empiricalparameter, the critical-LESP value, allows us to determine the onset, growth, and</p>&l

14、t;p>　　termination of LEVs. We show, by comparison with experimental and computationalresults for several aerofoils, motions and Reynolds numbers, that this computationallyinexpensive method is successful in predicting

15、 the complex flows and forces resultingfrom intermittent LEV shedding, thus validating the LESP concept.</p><p>　　Key words: computational methods, separated flows, vortex interactions</p><p>　　

16、1. Introduction</p><p>　　Unsteady flow phenomena are prevalent in a wide range of problems in nature andengineering. These include, but are not limited to, dynamic stall in rotorcraft andwind turbines, leadi

17、ng-edge vortices in delta-wings, micro air vehicle (MAV) design,?Email address for correspondence: kramesh2@ncsu.eduDiscrete-vortex method with novel shedding criterion for aerofoil flow 501 gust handling and flow contro

18、l. Unsteady flows are characterised by rapid changes inthe circulation of the aerofoil, apparent-m</p><p>　　Theodorsen (1935) developed a potential flow solution for the unsteady airloads on a flat plate und

19、ergoing harmonic, small-amplitude oscillations in pitch and plunge.</p><p>　　Wagner (1925) solved for the lift response of an aerofoil undergoing a step change in angle of attack (indicial response). The uns

20、teady lift coefficient due to arbitrary changes in angle of attack could hence be calculated by superposition using the Duhamel integral (Leishman 2002, chapter 8). McCune, Lam & Scott (1990) have presented a nonline

21、ar theory of unsteady potential flow which allows determination of lift and moment coefficients for large-amplitude motions.</p><p>　　Although these and other classical theories (Garrick 1937; von Kárm&

22、#225;n & Sears 1938) have proven invaluable in offering insight into unsteady aerodynamics and in fields such as aeroelasticity, their applicability in many problems is hindered by their inherent assumptions (small a

23、mplitudes, planar wake, fully attached flow). Advances in computational fluid dynamics (CFD) and experimental techniques have facilitated the</p><p>　　detailed study and analysis of unsteady phenomena. Ol et

24、 al. (2009a) and McGowan et al. (2011) have analysed the forces and flow fields for unsteady motions over a broad parameter space using both experimental and computational methods. Garmann & Visbal (2011) and Granlun

25、d, Ol & Bernal (2013) have investigated pitching flat plates in detail through computational and experimental methods respectively. Baiket al. (2012), Pitt Ford & Babinsky (2013) and Rival et al. (2014) have stud

26、ied the</p><p>　　effects and influence of leading-edge vortices using experimental techniques. However,these methods are not suitable for applications such as real-time simulation, rapid analysis, control an

27、d design, because of cost and time considerations. Since closed-form solutions from theory are incapable of capturing the various nonlinear effects,we may suitably augment theory with numerical procedures to expand its r

28、ange of</p><p>　　applicability. Brunton, Rowley & Williams (2013) and Wang & Eldredge (2013) have developed such phenomenologically augmented theoretical methods. With computing power and technology

29、advancing rapidly, low-order numerical models (constructed by augmenting classical theory) may provide the perfect balance between fidelity and cost.</p><p>　　In this paper, a discrete-vortex method with a n

30、ovel shedding criterion is proposed for modelling of massively separated, vortex-dominated flows.</p><p>　　Significant investigation of unsteady flow phenomena has been carried out by researchers interested

31、in studying and finding methods to suppress dynamic stall.</p><p>　　Dynamic stall refers to unsteady separation and stall phenomena on aerofoils that execute time-dependent motion, where the effective angle

32、of attack exceeds the static stall angle (McCroskey 1981, 1982). This process results in a delayed onset of flow separation/stall, followed by the shedding of a vortex from the leading edge of the aerofoil which traverse

33、s the aerofoil chord (Leishman 2002, chapter 9). Although this</p><p>　　vortex enhances the lift when it stays over the surface of the aerofoil, it also creates large nose-down pitching moments and flow sepa

34、ration over the entire aerofoil when it convects off the trailing edge. Hence dynamic stall can lead to violent vibrations and dangerously high airloads, resulting in material fatigue and structural failure. A good revie

35、w of experimental and numerical approaches toward understanding and predicting dynamic stall in given by Carr (1988) and Carr et al. (1990). A la</p><p>　　with a demonstration of their capabilities is given

36、in Leishman (2002, chapter 9).These models however, rely on several empirical parameters and can only be used in conditions that are bounded by validation with experimental data.</p><p>　　502 K. Ramesh, A. G

37、opalarathnam, K. Granlund, M. Ol and J. Edwards</p><p>　　Discrete-vortex methods have also been used extensively to model unsteady</p><p>　　separated flows. These methods are typically based on

38、potential-flow theory, and the shear layers representing separated flow emanate from the surface in the form of discrete vortices. Clements & Maull (1975) and Saffman & Baker (1979) provide detailed background on

39、 the historical development of the discrete-vortex method.A review of more recent progress on the application of vortex methods for flow simulation is given by Leonard (1980). Clements (1973), Sarpkaya (1975), Kiya &

40、 Arie (1977), and </p><p>　　to model flow past inclined plates and bluff bodies. Katz (1981) has developed a method for partially separated flow past an aerofoil, where the location of separation on the aero

41、foil has to be known through experiment or other means.</p><p>　　More recently, low-order methods based on discrete vortices have been developed by Ansari</p><p>　　Zbikowski & Knowles (2006a

42、), Wang & Eldredge (2013), Xia & Mohseni</p><p>　　(2013), and Hammer, Altman & Eastep (2014) to model leading-edge vortices in unsteady flows, with applications to insect flight and MAV aerodynam

43、ics. Although these methods are based on potential theory, they capture the essential physics in flows of interest by augmentation of inviscid theory with discrete-vortex shedding.</p><p>　　Apart from provid

44、ing a means to calculate the force coefficients on the aerofoil, these methods also enable study of the flow features and identification of dominant unsteady effects which require further modelling. These are significant

45、 advantages of this class of methods over semi-empirical methods, which only allow determination of the force coefficients through empirical fitting. However, the computational complexity increases as O(n 2 ) (when fast

46、summation methods are not used), where n is</p><p>　　Moreover, the methods cited above assume some ad-hoc start and stop criteria for vortex shedding, such as continuous shedding from a given location (valid

47、 only for sharp edges) or shedding that starts and stops depending on whether the local angle of attack exceeds a critical value (valid only for a small range of motions). A more general vortex shedding criterion is need

48、ed to make discrete-vortex methods broadly applicable to a wide range of geometries (including aerofoils with rounded leading</p><p>　　edges) and arbitrary unsteady motions.</p><p>　　In previous

49、 research (Ramesh et al. 2013b), the authors have developed an unsteady aerofoil theory based on potential flow, which holds uniformly regardless of amplitude and reduced frequency of motion, and shape of trailing wake.

50、This method was applied to a pitch-up, hold, pitch-down motion which was characterised by the shedding of a strong leading-edge vortex (LEV) during a part of the motion. The method was seen to predict well even under con

51、ditions of large amplitude and high reduced frequ</p><p>　　LEV formation was required. With this objective, the leading-edge suction parameter(LESP) was developed by the authors (Ramesh et al. 2011). This pa

52、rameter is a measure of the suction at the leading edge and it was shown that initiation of LEV formation always occurred at the same critical value of LESP, regardless of motion kinematics so long as the aerofoil and Re

53、ynolds number of operation were the same.</p><p>　　These methods are reviewed in §2.1. In this research, a discrete-vortex method is proposed in which the LESP criterion is used to modulate the initiati

54、on, growth, and termination of intermittent LEV shedding. The extension of the LESP criterion to handle LEV shedding and termination is discussed in §2.2. Details of the vortex method are discussed in §2.3. The

55、 LESP-modulated discrete-vortex method (LDVM) from the current work is validated for use in LEV-dominated flows against results from CFD an</p><p>　　新脫落的非定常翼型繞流的間歇前緣渦脫落標準離散渦方法</p><p>　　不穩(wěn)定機翼流動常常

56、為前沿(LEV)脫落。而實驗和高階計算造成了我們了解這些流，快速低工程所需的方法任務。經(jīng)典不穩(wěn)定僅限于小振幅和機翼理論附加的前沿流動。Discrete-vortex渦旋脫落的方法模型只從前緣假設不斷脫落,有效尖銳前緣或減少由特別標準如臨界迎角、有效只有一組限制的運動學。我們提出一個標準間歇從圓的前緣，脫落是由最大允許尖端的吸入。我們表明,當使用不穩(wěn)定的薄機翼理論，這一點尖端吸參數(shù)(LESP)項有關傅里葉級數(shù)代表綁定渦度弦向變化。此外，對

57、于任何機翼和雷諾數(shù)，LESP的臨界值，這是獨立的運動的運動學。當瞬時LESP值超過臨價值，渦旋脫落發(fā)生在前緣。我們增強了離散時間，arbitrary-motion,不穩(wěn)定的薄機翼理論的discrete-vortex脫落前緣由瞬時LESP。因此,使用一個單一的經(jīng)驗參數(shù),critical-LESP值，使我們能夠確定發(fā)病,增長,列弗的終止。我們顯示通過對比實驗和計算結果幾個翼面，動作和雷諾茲數(shù),計算便宜的方法成功地預測復雜的流和力產生的從間歇

58、列脫落，從而驗證LES概念。</p><p>　　關鍵詞:計算方法，分離流，渦輪相互作用</p><p><b>　　1、介紹</b></p><p>　　不穩(wěn)定流動現(xiàn)象是普遍在性質和范圍廣泛的問題工程。這包括,但不限于,在旋翼和動態(tài)失速風力渦輪機,尖端在三角翼的漩渦,微型飛行器(微型飛行器設計,?通信的電子郵件地址:kramesh2@ncsu

59、.eduDiscrete-vortex方法機翼流小說脫落標準501陣風處理和流控制。不穩(wěn)定流動的特點是快速變化機翼的循環(huán),表觀質量的影響,流動分離和旋渦流場。理論工作在這些主題可以追溯到1930年代和1920年代。Theodorsen(1935)開發(fā)潛在的流不穩(wěn)定飛行時的載重量的解決方案平板進行諧波,小振幅振蕩在音高和暴跌。瓦格納(1925)解決電梯響應機翼發(fā)生階躍變在攻角(指標的響應)。由于任意的非定常升力系數(shù)攻角的變化因此可以計算疊

60、加使用杜哈梅積分(全新2002年,第8章)。麥克卡尼,林&斯科特(1990)提出了一個非線性非定常勢流理論允許的決心的升力和力矩系數(shù)大幅度動作。雖然這些和其他經(jīng)典理論(灰呂1937;·馮·卡門和西爾斯1938)已被證明在提供洞察非定常空氣動力學和無價的空氣彈性等領域,他們在許多問題阻礙了其適用性固有的假設(小振幅,平面后，完全附流)。進步計算流體動力學(CFD)和實驗</p><p>

61、　　促進了詳細研究和分析的不穩(wěn)定現(xiàn)象。Ol et al .(2009)和麥高文et al。(2011)分析了部隊和流場的非定常運動廣泛的參數(shù)空間用實驗和計算方法。Garmann& Visbal(2011)和Granlund Ol &伯納爾(2013)已經(jīng)調查了俯仰平面詳細板分別通過計算和實驗方法。Baiket al。(2012年),皮特福特&巴賓斯基(2013)和競爭對手et al。(2014)的研究效果和使用實

62、驗技術前沿漩渦的影響。然而,這些方法不適合應用,如實時仿真、快速分析、控制和設計,因為成本和時間考慮。自閉,形成解決方案從理論不能捕獲各種非線性效應,我們可以適當增加理論和數(shù)值程序,擴大其范圍適用性。勃氏,羅利和威廉姆斯(2013)和王&·(2013)發(fā)達,進而增強理論方法等。與計算力量和技術發(fā)展迅速,低階數(shù)值模型(由增加經(jīng)典理論)可以提供完美的保真度和成本之間的平衡。摘要discrete-vortex方法提出了新穎的

63、脫落標準造型大大分開,vortex-dominated流動。重大的不穩(wěn)定流動現(xiàn)象進行了調查研究人員感興趣的研究,找到方法來抑制動態(tài)失速。動態(tài)失速分離是指非定常在翼面和</p><p>　　502 k·拉梅什,a .Gopalarathnam k . Granlund m 和j·愛德華茲</p><p>　　Discrete-vortex方法也被廣泛地用于模型不穩(wěn)定分離流

64、動。這些方法通常是基于位勢流理論,代表分離流的剪切層是從表面形式離散的漩渦?？巳R門茨&莫爾(1975)和薩夫曼&貝克(1979)提供的詳細的歷史發(fā)展背景discrete-vortex方法。</p><p>　　回顧最近的進展流渦方法的應用仿真是由倫納德(1980年)。克萊門茨(1973),Sarpkaya(1975),琪雅&阿里(1977),和其他研究人員已經(jīng)成功應用這一</p>

65、;<p>　　類的方法模型流過去的斜板和虛張聲勢的身體。Katz(1981)了方法部分分離流過去的機翼,那里的位置分離的機翼必須通過實驗或其他手段。</p><p>　　最近,低階方法開發(fā)了基于離散漩渦安薩里,˙Zbikowski &諾爾斯(2006),夏王&·(2013),和穆赫辛尼經(jīng)常(2013)和錘,奧特曼& Eastep(2014)模型的前沿漩渦不穩(wěn)定流動,

66、應用昆蟲飛行和微型飛機空氣動力學。雖然這些方法都是基于潛在的理論,他們捕獲的基本物理增加的流動感興趣的非粘性的理論與discrete-vortex脫落。除了提供一個意味著計算機翼上的力系數(shù),這些方法也使研究的主導不穩(wěn)定的流動特性和識別效果需要進一步的造型。這些都是這個類的顯著的優(yōu)勢方法在半經(jīng)驗方法,只允許的決心力通過經(jīng)驗擬合系數(shù)。然而,計算復雜度增加O(n - 2)(當不使用快速求和方法),其中n是的旋渦流場,導致可能大量計算時間。此外

67、,上面列舉的方法假設一些特別的啟動和停止的標準渦旋脫落,如連續(xù)流從一個給定的位置(僅為有效銳利的邊緣)或脫落,啟動和停止取決于當?shù)氐慕堑墓舫^一個臨界值(有效的只有一個小范圍的運動)。一個更通用渦旋脫落標準是需要discrete-vortex方法廣泛適用于各種幾何圖形(包括翼面圓領邊緣)和任意的非定常運動。在先前的研究(拉梅什et al . 2013 b),</p><p>　　F IGURE 1。在線(顏色)

68、時域方法的一個例子。2。LESP-modulateddiscrete-vortex方法(LDVM)</p><p><b>　　2.1。背景</b></p><p>　　在這一節(jié)中,作者先前開發(fā)的理論方法總結。感興趣的讀者可以參考拉梅什et al .(2011)和拉梅什et al .(2013 b)為進一步的細節(jié)。</p><p>　　2.1.

69、1。大角度不穩(wěn)定thin-aerofoil理論</p><p>　　大角度不穩(wěn)定thin-aerofoil理論開發(fā)的目的消除了傳統(tǒng)小角度的假設在thin-aerofoil理論無效的活期利息的流動。該方法基于時域方法由卡茨&普羅金(2000)。在圖1中,給出了慣性坐標系OXYZ車身骨架,連接到移動的機翼,Bxyz。在時間t = 0時,兩個坐標系一致和時間t > 0,身體朝著左邊的幀走一個時變的頁面。在

70、每個時間步,一個離散的后緣渦從后緣(TEV)了。類似于古典thin-aerofoil理論,在渦量分布機翼,γ(x)是傅里葉級數(shù),γ(θ,t)= 2 u</p><p>　　”0(t)1 + cosθsinθ+∞Xn = 1一個n(t)的罪(nθ)#(2.1)θ是一個變量的變換與弦向的坐標xx =c2(1?cosθ),(2.2)和0(t),1(t)……,n(t)是依賴于時間的傅里葉系數(shù),c是機翼弦,U是組件的機翼速

71、度- X方向。庫塔條件(0后緣渦度)通過傅里葉級數(shù)的形式執(zhí)行隱式。傅里葉系確定的函數(shù)的瞬時當?shù)卦跈C翼向下運動嗎執(zhí)行流的邊界條件必須保持切向機翼表面上看,0(t)=?1πZπ0W(x,t)Udθ(2.3)n(t)=2πZπ0W(x,t)Ucosnθdθ。(2.4)504 k·拉梅什,a . Gopalarathnam k . Granlund m . Ol和j·愛德華茲zxU交流F IGURE 2。在線(顏色)機翼速度

72、如圖所示(正面)和主的位置。機翼表面誘導速度正常,W(x,t),從此被稱為向下運動,從運動部件運動學計算,描述了圖2,誘導速度的漩渦流場:≡W(x,t)?φB?z=?η?x嗎?U cosα+˙hsinα+?φ列弗?x+?φtev?x。</p><p>　　U sinα??˙α(x?ac)+˙hcosα??φtev?z??φ列弗?z(2.5)φB,φ列弗和φtev速度勢與束縛,前沿和后緣渦度,η(x)是在機翼曲面分

73、布,?φ列弗/?x和?φtev /?x速度誘導切向領先的和弦和后緣離散渦和?φ列弗/?z和?φtev /?z誘導速度正常的和弦。運動參數(shù)包括Z方向的速度下滑,˙h和球場和弦的角度對X方向,α。列弗脫落過程詳細的§2.2。tev是流在每一個時間步如前所述,和他們優(yōu)勢迭代計算,開爾文的循環(huán)條件執(zhí)行:Γb(t)+N tevXm = 1Γtev m +N列弗Xn = 1Γ列弗n = 0(2.6)其中Γb是綁定循環(huán)計算通過整合弦向的分布

74、綁定渦度的機翼弦:Γb = Ucπ嗎?0(t)+1(t)2(2.7)</p><p>　　2.1.2。LESP標準起始列弗的形成已經(jīng)知道了幾十年,主要分離的發(fā)生邊緣是由臨界流參數(shù)的前緣。埃文斯&莫特(1959)表明,尖端分離強直接相關逆壓力梯度,遵循在前緣吸力峰。電子床(1978)展示了一個等價的尖端分離和之間的通信前緣的流速。瓊斯& Platzer爆發(fā)(1997)研究了層流分離的前緣推介NACA