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World War Ii Fighter Aerodynamics

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World War Ii Fighter Aerodynamics WORLD WAR II FIGHTER AERODYNAMICS BY DAVID LEDNICER EAA 135815 Previously, we have explored reat strides air-in foil is a NACA 2213, transitioning craft design were to a NACA 2209.4 at the tip rib. the aerodynamics of modern of era madethe in The Fw 190, which was designed G 1935-1945, 1930se andth f o this , d usee en dth e ath t is most evident in the NACA 23000 series of airfoils. homebuilt aircraft. Here, we design of fighter aircraft of this wine Th g root airfoi NACa s i l A period. thisFor reason, evalu-an 23015.3 and the tip airfoil a NACA will instead look at a different ation of three prominent fighter 23009. The P-51's wing, designed aircraft of this era, the North earle inth y 1940s, use earln sa y class aircraftof - WorldII War American P-51 Mustang,Su- the laminar flow airfoil which is a permarine Spitfire and the Focke NACA/NAA hybrid- calle45 e dth Wulf Fw 190 is presented here. 100. The wing root airfoil (of the fighters. As time progresses, As so much misinformation has basic trapezoidal wing, excluding appeared theseon aircraft, refer- the inboard leading edge exten- many valuable ofthe lessons ences will citedbe supportto the sion) is 16% thick, while the airfoil data discussed here. at the tip rib is 11.4% thick. With learned in the original design the inboard leading edge exten- Wing Geometry sion wine th , g root e airfoith n o l of vintage aircraft are being P-51B is 15.2% thick and on the P-51D 13.8% thick. The later I nsensea , these three aircraft model P-51H use dNACa A 66,2- lost. It is the purpose of this types represent three stages within (1.8)15.5 a=.6 at the wing root and a single generation of fighter de- a NACA 66,2-(1.8)12 a=.6 at the modernstudyuse to aerody- velopment mose . b Thi n t seasilca y tip and has no inboard leading seen in the wing airfoils used on edge extension. namic analysis tools recoverto the aircraft Spitfiree Th . , designed It is interesting to note that ap- 1930sd mi e i,n th use NACe dth A proximatel degreey2 f washouso t 2200 serie f airfoilsso , whics hwa l thre s useal n wa edo aircraft. some of this lost knowledge. new at the time. The wing root air- However, the distribution of twist SPORT AVIATION 85 upper surface extends fairl bacr yfa k wingie th n o s chord. This indicates thawine th t g shoul capable db f eo supporting a fairly large amount of laminar flow e P-5Th .1 Mustans gi renowned for being one the first air- craft to make use of airfoils designed to be capable of having extensive runs f laminao r flow. Bot Spitfire hth d ean Fw 190 use airfoils that do not support substantial amount f laminaso r flow. Figur- e1 A two-dimensional cut through the Wing twist distrib- wing pressure and skin friction distri- utions for the P- butions calculated by VSAERO on 6 0. 5 0. 1 4 0. 3 0. 0.7 0.8 0.9 51B, P-51 D, Semispan Fraction Spitfire and Fw Mustane th g sho) (Fig5 w. thata t a , 190. representative cruise condition, the win capabls gwa f sustainineo g long varie r eacdfo h aircraft Spitfire Th . e low. This depth was necessary to laminar boundary layer runs, with wing has a constant incidence (2 de- hous outware eth d retracting landing transition occurrin t roughlf ga o % y47 grees) to the dihedral break, where gear and wing gun ammunition boxes. chord. However, this calculation is for the twist starts. This aircraft actually an ideal case, for a wing without fas- has 2.25 degrees of washout, distrib- P-51 Mustang Analysis teners, gaps, misalignments or surface 0 ute19 d w linearlF e Th y. (Fig1) . waviness. During World War II, a win unusuas gi than i l degree2 t f so originae Th l North American Avia- Mustang was flight tested by NACA washout exists betweea rooe d nth an t tion drawing set for the Mustang are wit wakha e rake behin wine dth t ga point at 81.5% semispan. Outboard of available fro Nationae md th an r Ai l roughly 66.7% semispae Th n. (Ref1) . this location there is no more Space Museum frienA . f mindo e liv- result f thiso s test show that sern i , - washout incidence th , e holding fixed ing in England, Arthur Bentley, had vice the aircraft was unlikely to have a at zero degrees. The basic trapezoida kins lwa d d enougobtainean t se h e dth substantial laminar flow on the wing P-51e win th P-51 d f g2 Bo an s Dha to sort through it for the drawings that and transition occurred in the first degree washoutf so t a b ri p , ti wit e hth were of relevance to my endeavor. It 15% of the chord. Testing in an as- -.85 degree f incidenceso . However, founs wa d that model P-51B/e th f so C manufactured condition showed additio droopee th f no d inboard lead- and P-51D/K were relatively easo yt slightly lower drag and further, when g edgin e extension modifiee th s prepare, as the North American Avia- wine refines th gwa removo dt e wavi- appearanc twise th f te o distribution . tion drawings contained surface nes surfacd san e imperfections draa , g Lift distribution three th r e sfo aircraf t coordinates familiaa n i , r Fuselage leve s measurewa l d indicativa f eo show the results of these twist distrib- Station/Buttline/Waterline system. substantial regio f laminano r flow. utions (Fig. 2). These lift distributions However NASe th , M drawind di t gse Wartime windtunne lMuse testth - f so were calculated, using VSAERO, not appea contaio rt wine nth g defini- tang's wing airfoil in Germany gave wit aircrafe hth ts trimmekt 0 36 t da tion. Afte f searchingr o quitt bi ea I , similar results. (Ref2) . and 15,000 feet of altitude to repre- was put in touch with the Ed Horkey, Early model P-5e th 1f so experi - sentative Gross Weights and CG who had been the Chief Aerodynami- enced boundary layer separation ni locations. cist on the P-51 at North American. Ed the radiator inlet duct. Pilots reported e SpitfirTh e win famous gi r fo s was kind enough to supply the wing rumblina g noise emanating froe mth having an elliptic planform. Indeed, definition drawings the chord distribution is elliptical. An for both the P-5 IB Figure 2 - Calculated wing loading comparison with the air- examination of the resulting circula- and P-5 ID. craft trimmed at 360 kt and 15,000 feet altitude to repre- tion distributio a trimme r fo n d The pressure dis- sentative gross weightlocationsG C d san . condition mentioned above, shows tributions calculated thaloadine th t g distributio- el t no s ni by VSAERO on the liptical, though it is probably the most P-51d P-51an D B optimum of the three aircraft from the showe ar Fign 3 ni . induced drag standpoint. The reason S an . Particularld4 y o for deviation from elliptical is the 2 noteworthy is the re- degree f washouso t that have been gion of strong suction added to the elliptical planform, on the P-5 ID bubble which shifts the loading inboard. The canopy. This region (J elliptical wing planform appears to is not present on the have been chosen primarily to pro- less bulged P-51B vide greater wing inboardepte th n hi d canopy botn O . h air- portio winge th f no , while keeping craft the suction the airfoil thickness-to-chord ratios regiowine th gn no Semispan Fraction 86 JANUARY 1999 ductwork behind and beneath the base of the windscreen. the cockpit on early model The computation indicates Mustangs investigato T . e this thaboundare th t y layer sepa- phenomena, a complete Mus- rates approximately 6 inches tang fuselage was installed in i nwindscreene fronth f o t , a wind tunnenewle th t a ly increasine th o t e du g pres- opened NACA Ame- Re s sur thin ei s region (Fig. .8) search Center. It was found The boundary layer traces tharesule rumble th th t s t ewa that sto t separatiopa n have of the separated flow in the been restarte winde th n -do cooling inlet duct strikine gth shielpoine th t dta where eth radiator (Ref. 3). Changes, static pressure is the same as both in duct shape and the t separationa t ha i . Suc hsepa - additio deea f no p boundary t iratiopresenno s i n o t layer splitter on the inlet othecithee airo th f tw r- o r eliminate rumble dth d ean craft reviewed here. improve aircraft'e dth s cool- However, thifeatura s i s e e resultingTh .f theso s e Figure 3 - Pressure distribution calculated on the P-51B quite commo automon no - changes can be seen in the Mustang. biles and is related to the VSAERO boundary layer calculation, were presen drawingse th n o t t bu , slope of the windscreen. The Spit- which shows that boundary layer on preparation of the fuselage proved to fire's windscreen is at a 35-degree the upper surfac cooline th f eo g sys- be difficult as a global coordinate sys- forware anglth o et d deck, while eth doe m t separatte sno e unti bacr fa l n ki tem was not used.
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