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 aircraft 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. For instance, Fw 22-degrea 190' t a s si e angld ean the duct (Fig. 6). The boundary layer bulkheads could onl locatee y- b ac y db the P-51'3a 1t -degrea s si e angle- Ev . lowee th n ro surfac ducte th f ,e o start - cumulating ing fresh behind the oil cooler makes it distances from a to within inches of the water radiator known reference, and intercooler before separating. The insystea m more losse thin si s syste muce mar h lower aki thao nt t usen di than that of the Spitfire. This efficient the design of ships. cooling system arrangement is credited The surface pres- with Mustang'e mucth f ho s superior sure distribution performance ove Spitfiree rth . calculatee th r dfo Mustane Th repu a lon s d gha g -ha Spitfir shows i X eI n tation for being longitudinally in Fig. 7. Unlike the locationunstablG C t af t ea s resulting Mustang chorde th , - from the addition of a long-range wise extent of suc- fuel tank added behind the pilotis e winth tio gn o n seat. Results of a wind tunnel test of upper surfacn eca a P-51 B (Ref. 18) place the aircraftis be seen to be rela- power-off stick-fixed Neutral Point tively small, limit- at 39.11% MAC, which agrees quite inamoune gth f o t Figure 4 - Pressure distribution calculated on the P-51 D Mustang. well with the VSAERO results, laminar flow the wing can support. It which places this point at 38.97% -1.00 MAC. P-51Bs could be flown at CGs is interesting thae tth greatest suction on /Upper Surface Transition s 31.55a (Reft C af . %r MA .4) fa s a —\——— ———I——— Stick-fixe stick-freo dt e effectd an s the entire aircraft ap- / Lower Surface Transition power effects accoun r roughlfo t y pears on the bulged difference7.5C %MA . canopy. Other strong suctions appeat a r Supermarine Spitfire cornere e th th f so Analysis windshield, which f s o mad wa p u e Arthur Bentley also was able to sup- panels of flat armor wite originae m hth y pl l Supermarine shard glasha pd san drawings for the Spitfire. The Spitfire corners. drawing set contained definition for One of the first various models, ranging fro Spite mth - things to come to 1.00 decides wa Seafire t th I . dfiro t e47 eI light in the VSAERO 0.4 x/c 0.6 0.8 1.0 builo t panee dth l mode represeno t l a t analysi Spite th f -so Spitfir , whiceIX h coul fulle d- b yde regioa fir s i e f no Figure 5 - Calculated Mustang wing airfoil pressure distribu- fined from the drawings. Coordinates separated flow at tion and boundary layer transition locations in cruise, for ideal surface conditions. SPORT AVIATION 87 idently, the Spitfire's the boundary layer separates shortly windscree steepo to s ni . after entering the duct, resulting in a experimentan A l wind- large drag penalty (Fig. Experi9) . - screen, rounded and of mentally, it was determined that the shallower slope, was fitted Spitfire cooling system drag- ex , to a Spitfire IX in 1943 pro- pressed as the ratio of equivalent duced a speed increase of cooling-drag powe totao rt l engine 12 mph, at a Mach number power, was considerably higher than of .79 (Ref. 5). A similar that of other aircraft tested by the windscreen introduced on RAE. Thiattributes swa "tho dt e pres- Seafire th e XVII creds i , - encboundara f eo y layer aheae th f do ited with a speed gain of 7 duct tend precipitato st e separation (Refh . 6) .mp 0 mph40 t a , and makes the ducting problem more Supermarine is often re- difficult" (Ref. Simila8) . r problems gardee th f beins o d a e gon are present on the early model Messer- Figure 6 - Calculated boundary layer separation in the first companies to make schmit f 109througp B tu , E e hth Mustang cooling system breakthroughe th f o e us s model completA . e redesige th f no made by Meredith at RAE cooling system, during development Farnboroug f 109Fa h B f io e n e ,th resultef us o e th n di the design of boundary layer bypass duct, which duct coolinr sfo g significantly improve pressure dth e systems (Ref. 7) . recovery at the radiator face (Ref. 9). In fact, the Spit- The Spitfire has long had a reputa- fire's radiator ducts tio beinf no g longitudinally neutrally were designed us- stable. Results of wartime flight tests ing these guide- Spitfira NACy f b o A AeV (Ref) 10 . lines. However, confirm that the aircraft was indeed the VSAERO cal- longitudinally neutrally stabla t ea culation indicates typica locationG C l NACe Th . - Are the boundary layer port mentions that no change in lowee onth r sur- elevator position was necessary to face of the wing maintain longitudinal trim when is ingested by the changing airspeed, implying thae th t cooling syste positiones mwa locatioG e C th t da f no inlet. Running into stick-fixee th d longitudinal Neutral the severe adverse locatioG PointC e thin Th n.i s test (increasing) pres- was at 31.3% MAC. VSAERO analy- sure gradient aheaSpitfire th df o es placesi power-ofe sth f Figure 7 - Pressure distribution calculated on the Spitfire IX. of the radiator, stick-fixed Neutral Poin t 36.66a t %

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For information, use SPORT AVIATION'S Reader Service Card JANUAR8 8 9 99 Y1 MAC. Standard estimates of power the port wing drop- effects show that the Neutral Point ping violentlyso will shift forward theso 4-5t e %edu thataircraftthe al- effects, which account differe th r -sfo most inverted itself. ence betwee VSAERe nth flighd Oan t In fact, if the Ger- test results. The NACA testing also man fighterwas found there was a stable gradient of pulled into stallg a stick force with increasing airspeed. intighta turn,it This means that the Spitfire was stick- would flickintoout free longitudinally stable. Bobweights the opposite bank in the elevator control circuit helped and an incipient turn the stick-fixed neutrally stable spin was the in- airplane int airplann oa e wit hsmala l evitable outcome if degre f stick-freeo e stabilitye th s A . the pilot did not pilot mostly is aware of stick-free sta- have his wits about bility and low margins of stability are stallhim.The in associated with high maneuverability, landing configura- this was a satisfactory situation. quitetionwas Figur Calculate- e8 d Spitfire windshield boundary layer separation. Separation is calculated to take place at the base of different, there be- windshiele th d wher streamline eth e traces end locatioe .Th n ing intense pre-stall wher separatee eth d flo estimatews i reattacdo t h highep ru Focke Wulf Fw 190 Analysis buffeting before the the windshield is shown by where the streamline traces starboard wing resume. dropped compara- Arthur Bentle oncs ywa e agaie nth tively gently at 102 mph (164 km/h). engine inlinen a , mucs i , h longer than source of my geometrical informa- The results of an USAAF evalua- the BMW engine, giving the D-9 a tion thin I . s case, severale yearh o sag tion of the Fw 190 (Ref. 12 and 13) elongated nose, whic countes hwa r had prepared a set of Fw 190 draw- repor aircrafe th t havo t t gentlea e balanced with a 500mm plug added to a modelin ingr fo s g magazine, stall. However, these reports admi t fuselagetaf e VSAERe th Th . O model working from the original Focke Wulf 19w F 0tha e stalleth t d abruptly when modifies wa represeno y dt b 9 D- a t drawings. Initially, I first modeled a maneuvering. The reason for this re- making these changes and by adding radial engined Fw 190 A-8, but I later ported differenc non-maneuverinn i e g bulgee th d canop0 19 y founw F n do modified this mode represeno t l n a t stall behavio unknowns i r comparA . - D-9s. It was found from the VSAERO inline engineD-90 19 thin ,i w dF s ison of the local wing lift coefficients, results thafuselage th t e stretc- hde case using actual Focke Wulf draw- calculate VSAEROy db t stala , l with signeFocke th y edb Wulf engineers ings. Despite sparse fuselage cross estimatee th d stalling lift coefficients resulted in a slight increase in stick section information, this model wa airfoile sth f o s two-dimensionally fixed stability, with the Neutral Point constructed with relative ease. ) show(Fig12 . s that approximately movin8 gA- froe th m n o 35.8 C %MA The pressure distribution calcu- the inner 40% of the wing reaches to 40.4% MAC on the D-9. It should lated on the Fw 190 A-8 and Fw 190 C]max at the same aircraft angle of at- notee b d these resultt containo o sd n D-9 are shown in Fig. 10 and 11. tack. A wartime Focke Wulf report propeller effects, which were not Here, like on the Spitfire, the chord- indicate) (Ref14 . s tha t highea t r load- modeled. Flight testing of an early wise extent of suction on the wing is ing conditions (i.e., when pulling model Fw 190A indicated that the air- limite choice th y f airfoildb eo d an s more gs) elastic deformation of the craft was "just statically stable; stick not much laminar flow is supported. Fw 190 outer wing shifts the load dis- fixe freed dan , engine offstatid ;an - Also, as on the Spitfire, the bulged tribution outboard. This would cause cally unstable to a slight degree, canopy of the Fw 190 D-9 has a re- even wine morth reaco f gt e o s hit engin " (Refeon . 11). Durin cone gth - gion of strong suction, not present on stalling lift coefficient simultane- tinued0 developmen19 w F e th f o t theFw 190 A-8. ously. Combined wit e sharth h p series, the aircraft's CG moved rear- At the time that the Fw 190 first stalling feature NACe th f so A 230XX ward as fuel tanks and other appeare combatn di s 1941n i , wa t i , airfoils, this would produce the harsh equipmen t fuseaf s addee - th wa t o dt superio contemporare th o rt y fighters stall founCapty b n di . Brown genA . - lage (Ref. 15). This Neutral Point shift on nearly every count. When the RAF tle stall would be evidenced by a during development of the Fw 190D captured the first flyable Fw 190 in more gradual progression of the 2D model would have been quite valuable 1942thorouga , h evaluation revealed stall spanwise. in maintaining the continued growth the Achilles Heal to be a harsh stalling Initial VSAERO calculations were of the design. characteristic, which limited its maneu- made on a model of the Fw 190 A-8. ver margins. Captain Eric Brown states This versio aircrafe th pow s f no wa t - Drag Comparison (Ref. 11): ered by a BMW 80ID radial. The stalling speed 190A-Fw ofthe Naturally questioe th , n aroso t s ea Ther mane ear y conflicting claims 4 in clean configuration was 127 mph how the aerodynamics of this aircraft as to the equivalent flat plate drag (204 stall the km/h) cameand sud- differed fro latere mth , Junkers Jumo area (fthesf o ) e fighter aircraft. Based denly and virtually without warning, 213A powered Fw 190 D-9. The Jumo upon my research, what I believe are

SPORT AVIATION 89 Mustang reenginey b d Rolls-Royce with a Merlin Conclusion ; ; , 65. The P-51B, with an improved cooling system Important design feature f threo s e configuratio s i even n prominent fighteI WorlI r drWa aircraft faster than the Spitfire IX. have been examined by the use of a The difference in perfor- modern Computational Fluid Dynamics mance betweee th n method. It is hoped that the results pre- Mustang and the Spitfire sented here will help demonstrate some is attributed to several fac valuabl-e th f o e lessons learned fron ma tors. These include th e importan fighten i a er t r aircraft design. superior configuration of This information, while historical, still the Mustang's cooling sys- has relevance in today's world of air- tem and the Spitfire's craft design. Important lessons to be relatively high- leveex f o l learned are: Figure 9 - Fw 190 calculated lift coefficient distribution crescence drag, generated Airfoi• l choic surfacd ean e qualite yar at 1g stall. y opeb n wheel wellsa , important in achieving the advantages of nonretractable tail laminar flow. wheel and other • Cooling system duct desigr nfo design details liquid cooled engines must be con- (Ref. 17-19). ducted carefull avoio yt d losses. e populaOn r • Attentio aerodynamio nt c detail, piece of aerody- suc windshiels ha d slope overcomn ca , e namic folklore is the disadvantage of excess wetted area. the low CDswet abrupn A • avoidee t b stal- n at ca lf di value achieved with tention is paid to airfoil selection and the Mustang. Vari- wing twist. ous sources quote • As seen with all three of these air- this valu rangs ea - planes, longitudinal stability and control ing from .003o 8t problems are common, but can be .0043 revieA . f wo avoided by the resourceful designer. available wind tun- r; nel and flight test Author's Note drae g th dat r fo a Mustang demon- This article is dedicated to Edward Figure 10 - Pressure distribution calculated on the Fw 190 A-8. strate neee r sth dfo Horkey and Jeffery Ethell, who both . havin detaill gal f so contributed information vital to this the aircraft present work. Ed died as a result of injuries sus- the most accurate value showe sar n ni if the drag is to be accurately mea- tained in traffic accident in July 1996. Table 1. sured. Subscale wind tunnel tests of killes Jefcrase Locka th wa f f n hdo i - The wetted aircrafe areath f e so ar t the P-51 P-51d Aan B resulte valn di - heed P-38 Lightning in May 1997. calculated by VSAERO, and exclude ues of CDswet, at a representative Fayouno rto havo gt e participaten di ducte th coolinr sfo g systems. cruise lift coefficient, in the range of I hav , WorlII e r londWa g been fasci- Notabl thas Mustani e th t s gha .0046-.0047 (Ref. 20-22). However, nated abou famoue th t w findinsho t gou the largest wetted are f thiao s group these tests usually wer f modeleo s aircraft of this war were designed. The of aircraft, but has the lowest drag. lacking exhaust stacks, surface discon- deeper I have gotten into this pursuit, the Evidence of this is that with the same tinuities, etc. Measurements made in more informatio nhavI e uncovered that version of the Rolls-Royce Merlin full-scale wind tunne- P le testth f so has proven to be valuable in my daily propelled an r installed Mustane th , g 5flighd IBan ) (Refte test23 th . f so work as an aerodynamicist. I have be- fasteh mp measures r3 Xwa 2 e b o dt P-51 P-d 5an A I B) (Ref (Ref24 . . 21) come convinced tha a studt f o y than the Spitfire IX (Ref. 16). The resulted in a value of CDswet of ap- ihistorical aerodynamics importann a s i t Allison Mustana s wa n gpowereX d proximately .0053. part of an aerodynamicistis ongoing edu-

TABLE 1

Aircraft f Wetted Area ^Dswet Ref. SpitfirX eI 5.40 ft2 831.2ft2 .0065 16 P-51B Mustang 4.61 ft2 874.0 ft2 .0053 21 P-51D Mustang 4.65 ft2 882.2 ft2 .0053 27 8 A- 0 Fw19 5.22 ft2 735.0 ft2 .0071 26 9 D- 0 19 w F 4.77 ft2 761.6ft2 .0063 26

JANUAR0 9 Y 1999 cation. To this end, one of my goals has nautical Society, August 1944. High-SpeeE RA e ) Stafth 22 f o df become to try and disseminate the ) Phillips10 , W.H Venseld .an , J.R., Wind Tunnel, "High Speed Wind- knowledge I have unearthed, this article "Measurement Flyine th f so g Qualities Tunnel Tests of Models of Four Single being an effort towards this end. For of a Supermarine Spitfire VA Airplane," Engined Fighters (Spitfire, Spiteful, those seeking further informatio thin ni s NASA WR L-334, September 1942. Attacke Mustang)d an r , Parts 1-5," regard, I recommend taking a look at my 11) Brown, Capt. Eric, "Viewed Aeronautical Research CounciM R& l ilncomplete Guide to Airfoil Usage! at: Fro Cockpite mth ; Tank's Second No. 2535, edited by W.A. Mair, 1951. http://amber.aae.uiuc.edu/~m Iron," Air International, Vol. 10 No. 2, 23) Lange, R.H., "A Summary of Drag selig/ads/aircraft.html. February 1976. Results From Recent Langley Full-Scale mentiones A previoun di s articlesI , 12) Foster Rickerd , Johan . nJr , Tunnel Test Armf so Navd yan y Airplanes," aeronautican a m a l engineer, specializ- Chester S., "Design Analysis No. 9 NAC L5A3R L-108)AR AC 0(W , 1945. applien i g in d computational fluid The Focke-Wulf 190," Aviation, Oc- 24) Staffs of the High-Speed Tunnel and dynamics. Based in Redmond, Wash- tober 1944. High-Speed Flight Sections, "Research on ington worI , Analyticar kfo l Methods, Wartn 13Va ) , F.D., "Handbook High-Speed Aerodynamic Royae th t sa l Incaerodynamiy M . c (and hydrody- For Fw 190 Airplane," USAAF T-2 Aircraft Establishment from 194 1945,o 2t " namic) consulting projectI haveAM t sa Technical Report F-TR-1 102 ND, Aeronautica l. 2222ResearcNo ,M hR& included submarines, surface vessels, March 1946. edite W.Ay db . Mair, September 1946. automobiles, trains, helicopters, aircraft 14) Gross , "DiP. , e Entwicklunr gde 25) Anon., "Performance Calculations for and space launch vehicles. 1 can be Tragwerkkonstruktio w 190,F n " Model P-51D-5-NA Airplane (NAA Model reached at: [email protected] or: Bericht 176 der Lillenthal-Gesellschaft, NA-109)," North American Aviation Report Analytical Methods, Inc., 2133 2 Teil, January 1944. No. NA-8449, December 1,1944. 152nd AveNE, Redmond, WA 98052 15) Bentley, Arthu , "FockrL. e Wulf 26) Anon., "Widerstandaten von Fighter," Scale Models, July 1978. Flugzeugen" (Drag Data for Aircraft), References 16) Birch, David, Rolls-Royce and Focke Wulf data sheet. * the Mustang, Rolls- 1) Zalovcik, J.A. Profile-DraA ," g Royce Heritage Trust Investigation in Flight of an Experi- Historical Series No. mental Fighter-Type Airplane - The 9, Derby, England, North American XP-51," NACA re- 1987. port, November 1942. 17) Private Commu- 2) Bussmann, K., "Messungem na nication, J. Leland Laminarprofil P-51 Mustang," Aero- Atwood, October 1994. dynamisches Institu r Technischede t n 18) Anon, "Estima- Hochschule Braunschweig, Bericht Increase tioth f no n ei 43/4, January 1943. Performance Obtain- 3) Matthews, H.F., "Eliminatiof no able By Fitting a Rumble From the Cooling Ducts of a Continuously Variable Single-Engine Pursuit Airplane," Radiator Flap," Rolls- NAC A-70R AW , August 1943. Royce Experimental 4) Morgan, Eric B. and Shacklady, Department Report, Edward, Spitfir Historye eTh Puby ,Ke - August 10, 1942. lishing, Stamford Lines, England, 1987. ) Privat19 e Commu- Anon.) 5 , "Aerodynamic Dimensional nication, Ed Horkey, Figure 11 Pressure distribution calculated on the Fw 190 D-9. Data on P-51 B-1 -NA, P-51B-5-NA and November 1994. P-51C-1-NT Airplanes," North American ) Anon.20 , "Wind 2.0 Tunne- l XP Dat r aFo Aviation Repor . NA-5822tNo , August 6,1943. 1.8 - 6) Smith , "ThJ. , e Developmenf to AirplanB 51 e (NAA - 6 1. the Spitfire and Seafire," Journal of Model NA-101)," Mmax Royae th l Aeronautical Society, Vol. North American Avia- Stalled Region

51, April 1947. tion Report No. NA- 1.2 - 7) Meredith, F.W., "Note on the 5548, October 9,1943. 1.0 - Cooling of Aircraft Engines With Spe- 21)Nissen, J.M., Calculated C| Distribution cial Reference to Ethylene Glycol Gadeberg, B.L. and Radiators Enclosed in Ducts," ARC Hamilton, W.T., 1683M R& , August 1935. "Correlatioe th f no 0.4 - ) Hartshorn8 , A.S Nicholsond an . , Drag Characteristics M.A., "The Aerodynamice th f o s of a Typical Pursuit 0.2 - Cooling Aircraft Reciprocating En- Airplane Obtained 0.0 2498 1947M y gines,Ma ,R& . C "AR From High-Speed 0.3 0.4 05 ) Morgan9 , M.B Smeltd an . , R. , Wind-Tunnel and Semispan Fraction "Aerodynamic Features of German Flight Tests," NACA Figure 12 - Boundary layer separation calculated in the Spitfire Aircraft," Journa Royae th f o l Aero- Report 916, 1948. cooling system. SPORT AVIATION 91