WORLD WAR IIFIGHTERAERODYNAMICSBY DAVID L E D N I C E REAA 135815

Previously, we have exploredthe aerodynamics of modern

homebuilt aircraft. Here, we

will instead look at a different

class of aircraft - World War II

fighters. As time progresses,

many of the valuable lessonslearned in the original design

of vintage aircraft are being

lost. It is the purpose of this

study to use modern aerody-

namic analysis tools to recover

some of this lost knowledge.

G reat strides in air-craft design weremade in the era of1935-1945, and thisis most evident in the

design of fighter aircraft of thisperiod. For this reason, an evalu-ation of three prominent fighteraircraft of this era, the NorthAmerican P-51 Mustang, the Su-permarine Spitfire and the FockeWulf Fw 190 is presented here.As so much misinformation hasappeared on these aircraft, refer-ences will be cited to support thedata discussed here.

Wing Geometry

In a sense, these three aircrafttypes represent three stages withina single generation of fighter de-velopment. This can be most easilyseen in the wing airfoils used onthe aircraft. The Spitfire, designedin the mid 1930s, used the NACA2200 series of airfoils, which wasnew at the time. The wing root air-

foil is a NACA 2213, transitioningto a NACA 2209.4 at the tip rib.The Fw 190, which was designedat the end of the 1930s, used theNACA 23000 series of airfoils.The wing root airfoil is a NACA23015.3 and the tip airfoil a NACA23009. The P-51's wing, designedin the early 1940s, uses an earlylaminar flow airfoil which is aNACA/NAA hybrid called the 45-100. The wing root airfoil (of thebasic trapezoidal wing, excludingthe inboard leading edge exten-sion) is 16% thick, while the airfoilat the tip rib is 11.4% thick. Withthe inboard leading edge exten-sion, the wing root airfoil on theP-51B is 15.2% thick and on theP-51D 13.8% thick. The latermodel P-51H used a NACA 66,2-(1.8)15.5 a=.6 at the wing root anda NACA 66,2-(1.8)12 a=.6 at thetip and has no inboard leadingedge extension.

It is interesting to note that ap-proximately 2 degrees of washoutwas used on all three aircraf t .However, the distribution of twist

SPORT AVIATION 85

0.1 0.3 0.4 0.5 0.6

Semispan Fraction0.7 0.8 0.9

Figure 1-Wing twist distrib-utions for the P-51B, P-51 D,Spitfire and Fw190.

varied for each aircraft. The Spitfirewing has a constant incidence (2 de-grees) to the dihedral break, wherethe twist starts. This aircraft actuallyhas 2.25 degrees of washout, distrib-uted linearly (Fig. 1). The Fw 190wing is unusual in that 2 degrees ofwashout exists between the root and apoint at 81.5% semispan. Outboard ofthis location there is no morewashout, the incidence holding fixedat zero degrees. The basic trapezoidalwing of the P-51B and P-51D has 2degrees of washout, with the tip rib at-.85 degrees of incidence. However,addition of the drooped inboard lead-ing edge extension modifies theappearance of the twist distribution.Lift distributions for the three aircraftshow the results of these twist distrib-utions (Fig. 2). These lift distributionswere calculated, using VSAERO,with the aircraft trimmed at 360 ktsand 15,000 feet of altitude to repre-sentative Gross Weights and CGlocations.

The Spitfire wing is famous forhaving an elliptic planform. Indeed,the chord distribution is elliptical. Anexamination of the resulting circula-t ion d is t r ibut ion for a tr immedcondition mentioned above, showsthat the loading distribution is not el-liptical, though it is probably the mostoptimum of the three aircraft from theinduced drag standpoint. The reasonfor deviation from elliptical is the 2degrees of washout that have beenadded to the e l l ip t ica l p lanform,which shifts the loading inboard. Theelliptical wing planform appears tohave been chosen primarily to pro-vide greater wing depth in the inboardportion of the wing, whi le keepingthe airfoil thickness-to-chord ratios

86 JANUARY 1999

low. This depth was necessary tohouse the outward retracting landinggear and wing gun ammunition boxes.

P-51 Mustang Analysis

The original North American Avia-tion drawing set for the Mustang areavailable from the National Air andSpace Museum. A friend of mine liv-ing in England, Arthur Bentley, hadobtained the set and was kind enoughto sort through it for the drawings thatwere of relevance to my endeavor. Itwas found that models of the P-51B/Cand P-51D/K were relatively easy toprepare, as the North American Avia-t ion drawings contained surfacecoordinates, in a familiar FuselageStation/Buttline/Waterline system.However, the NASM drawing set didnot appear to contain the wing defini-tion. After quite a bit of searching, Iwas put in touch with the Ed Horkey,who had been the Chief Aerodynami-cist on the P-51 at North American. Edwas kind enough to supply the wingdefinition drawingsfor both the P-5 IBand P-5 ID.

The pressure dis-tributions calculatedby VSAERO on theP-51B and P-51Dare shown in Fig. 3and 4. Particularlynoteworthy is the re-gion of strong suctionon the P-5 ID bubblecanopy. This regionis not present on theless bulged P-51Bcanopy. On both air-craft the suctionregion on the wing

upper surface extends fairly far backon the wingis chord. This indicatesthat the wing should be capable ofsupporting a fairly large amount oflaminar flow. The P-51 Mustang isrenowned for being one the first air-craft to make use of airfoils designedto be capable of having extensive runsof laminar flow. Both the Spitfire andFw 190 use airfoils that do not supportsubstantial amounts of laminar flow.A two-dimensional cut through thewing pressure and skin friction distri-butions calculated by VSAERO onthe Mustang (Fig. 5) show that, at arepresentative cruise condition, thewing was capable of sustaining longlaminar boundary layer runs, withtransition occurring at roughly 47% ofchord. However, this calculation is foran ideal case, for a wing without fas-teners, gaps, misalignments or surfacewaviness. During World War II, aMustang was flight tested by NACAwith a wake rake behind the wing atroughly 66.7% semispan (Ref. 1). Theresults of this test show that, in ser-vice the aircraft was unlikely to have asubstantial laminar flow on the wingand transition occurred in the first15% of the chord. Testing in an as-manufactured condit ion showedslightly lower drag and further, whenthe wing was refined to remove wavi-ness and surface imperfections, a draglevel was measured indicative of asubstantial region of laminar flow.Wartime windtunnel tests of the Mus-tang's wing airfoil in Germany gavesimilar results (Ref. 2).

Early models of the P-51 experi-enced boundary layer separation inthe radiator inlet duct. Pilots reporteda rumbling noise emanating from the

Figure 2 - Calculated wing loading comparison with the air-craft trimmed at 360 kt and 15,000 feet altitude to repre-sentative gross weights and CG locations.

So

(J

Semispan Fraction

ductwork behind and beneaththe cockpit on early modelMustangs. To investigate thisphenomena, a complete Mus-tang fuselage was installed ina wind tunnel at the newlyopened NACA Ames Re-search Center. It was foundthat the rumble was the resultof the separated flow in thecooling inlet duct striking theradiator (Ref. 3). Changes,both in duct shape and theaddition of a deep boundarylayer spli t ter on the i n l e teliminated the rumble andimproved the aircraft's cool-ing. The results of thesechanges can be seen in theVSAERO boundary layer calculation,which shows that boundary layer onthe upper surface of the cooling sys-tem does not separate until far back inthe duct (Fig. 6). The boundary layeron the lower surface of the duct, start-ing fresh behind the oil cooler makes itto within inches of the water radiatorand intercooler before separating. Thelosses in this system are much lowerthan that of the Spitfire. This efficientcooling system arrangement is creditedwith much of the Mustang's superiorperformance over the Spitfire.

The Mustang has long had a repu-tation for being long i tud ina l lyunstable at aft CG locations resultingfrom the addition of a long-rangefuel tank added behind the pilotisseat. Results of a wind tunnel test ofa P-51 B (Ref. 18) place the aircraftispower-off stick-fixed Neutral Pointat 39.11% MAC, which agrees quitewell with the VSAERO resul ts ,which places this point at 38.97%MAC. P-51Bs could be flown at CGsas far aft as 31.55% MAC (Ref. 4).Stick-fixed to stick-free effects andpower effects account for roughly7.5% MAC difference.

Supermarine SpitfireAnalysis

Arthur Bentley also was able to sup-ply me with the original Supermarinedrawings for the Spitfire. The Spitfiredrawing set contained definition forvarious models, ranging from the Spit-fire I to the Seafire 47. It was decidedto build the panel model to represent aSpitfire IX, which could be fully de-fined from the drawings. Coordinates

Figure 3 - Pressure distribution calculated on the P-51BMustang.

were present on the drawings, butpreparation of the fuselage proved tobe difficult as a global coordinate sys-tem was not used. For instance,bulkheads could only be located by ac-cumulatingdistances from aknown reference,in a system moreakin to that used inthe design of ships.

The surface pres-sure dis t r ibut ioncalculated for theSpitfire IX is shownin Fig. 7. Unlike theMustang, the chord-wise extent of suc-t ion on the wingupper surface canbe seen to be rela-tively small, limit-ing the amount oflaminar flow thewing can support. Itis interesting that thegreatest suction onthe entire aircraft ap-pears on the bulgedcanopy. Other strongsuctions appear atthe corners of thewindshield, whichwas made up ofpanels of flat armorglass and had sharpcorners.

One of the firstthings to come tolight in the VSAEROanalysis of the Spit-fire is a region ofseparated flow at

the base of the windscreen.The computation indicatesthat the boundary layer sepa-rates approximately 6 inchesin front of the windscreen,due to the increasing pres-sure in this region (Fig. 8).The boundary layer traces

that stop at separation havebeen restarted on the wind-shield at the point where thestatic pressure is the same asi hat at separation. Such a sep-i r a t i o n is not present onci ther of the other two air-craft reviewed here.However, this is a featurequite common on automo-bi les and is related to the

slope of the windscreen. The Spit-fire's windscreen is at a 35-degreeangle to the forward deck, while theFw 190's is at a 22-degree angle andthe P-51's is at a 31 -degree angle. Ev-

Figure 4 - Pressure distribution calculated on the P-51 D Mustang.

-1.00

1.00

/Upper Surface Transition—\——— ———I———

/ Lower Surface Transition

0.4 x/c 0.6 0.8 1.0

Figure 5 - Calculated Mustang wing airfoil pressure distribu-tion and boundary layer transition locations in cruise, for idealsurface conditions.

SPORT AVIATION 87

Figure 6 - Calculated boundary layer separation in theMustang cooling system

Figure 7 - Pressure distribution calculated on the Spitfire IX.

idently, the Spitfire'swindscreen is too steep.An experimental wind-screen, rounded and ofshallower slope, was fittedto a Spitfire IX in 1943 pro-duced a speed increase of12 mph, at a Mach numberof .79 (Ref. 5). A similarwindscreen introduced onthe Seafire XVII, is cred-ited with a speed gain of 7mph, at 400 mph (Ref. 6).

Supermarine is often re-garded as being one of thefirst companies to makeuse of the breakthroughsmade by Meredith at RAE

Farnborough inthe design ofducts for coolingsystems (Ref. 7).In fact, the Spit-fire's radiator ductswere designed us-ing these guide-lines. However,the VSAERO cal-culation indicatesthe boundary layeron the lower sur-face of the wingis ingested by thecooling systeminlet. Running intothe severe adverse(increasing) pres-sure gradient aheadof the radiator,

the boundary layer separates shortlyafter entering the duct, resulting in alarge drag penalty (Fig. 9). Experi-mentally, it was determined that theSpitfire cooling system drag, ex-pressed as the ratio of equivalentcooling-drag power to total enginepower, was considerably higher thanthat of other aircraft tested by theRAE. This was attributed to "the pres-ence of a boundary layer ahead of theduct tends to precipitate separationand makes the ducting problem moredifficult" (Ref. 8). Similar problemsare present on the early model Messer-schmitt Bf 109, up through the Emodel. A complete redesign of thecooling system, during developmentof the Bf 109F, resulted in the use of aboundary layer bypass duct, whichsignificantly improved the pressurerecovery at the radiator face (Ref. 9).

The Spitfire has long had a reputa-tion of being longitudinally neutrallystable. Results of wartime flight testsof a Spitfire VA by NACA (Ref. 10)confirm that the aircraft was indeedlongitudinally neutrally stable at atypical CG location. The NACA re-port mentions that no change inelevator position was necessary tomain ta in longi tudina l trim whenchanging airspeed, implying that theCG was positioned at the location ofthe stick-fixed longitudinal NeutralPoint. The CG location in this testwas at 31.3% MAC. VSAERO analy-sis of the Spitfire places the power-offstick-fixed Neutral Point at 36.66%

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MAC. Standard estimates of powereffects show that the Neutral Pointwill shift forward 4-5% due to theseeffects, which accounts for the differ-ence between the VSAERO and flighttest results. The NACA testing alsofound there was a stable gradient ofstick force with increasing airspeed.This means that the Spitfire was stick-free longitudinally stable. Bobweightsin the elevator control circuit helpedturn the stick-fixed neutrally stableairplane into an airplane with a smalldegree of stick-free stability. As thepilot mostly is aware of stick-free sta-bility and low margins of stability areassociated with high maneuverability,this was a satisfactory situation.

Focke Wulf Fw 190 Analysis

Arthur Bentley was once again thesource of my geometrical informa-tion. In this case, several years ago hehad prepared a set of Fw 190 draw-ings for a modeling magazine,working from the original Focke Wulfdrawings. Initially, I first modeled aradial engined Fw 190 A-8, but I latermodified this model to represent aninline engined Fw 190 D-9, in thiscase using actual Focke Wulf draw-ings. Despite sparse fuselage crosssection information, this model wasconstructed with relative ease.

The pressure distribution calcu-lated on the Fw 190 A-8 and Fw 190D-9 are shown in Fig. 10 and 11.Here, like on the Spitfire, the chord-wise extent of suction on the wing islimited by the choice of airfoils andnot much laminar flow is supported.Also, as on the Spitfire, the bulgedcanopy of the Fw 190 D-9 has a re-gion of strong suction, not present ontheFw 190 A-8.

At the time that the Fw 190 firstappeared in combat, in 1941, it wassuperior to the contemporary fighterson nearly every count. When the RAFcaptured the first flyable Fw 190 in1942, a thorough evaluation revealedthe Achilles Heal to be a harsh stallingcharacteristic, which limited its maneu-ver margins. Captain Eric Brown states(Ref. 11):

The stalling speed of the Fw 190A-4 in clean configuration was 127 mph(204 km/h) and the stall came sud-denly and virtually without warning,

the port wing drop-ping so violentlythat the aircraft al-most inverted itself.In fact, if the Ger-man fighter waspulled into a g stallin a tight turn, itwould flick out intothe opposite bankand an incipientspin was the in-evitable outcome ifthe pilot did nothave his wits abouthim. The stall inlanding configura-tion was quitedifferent, there be-ing intense pre-stallbuffeting before thestarboard wingdropped compara-tively gently at 102 mph (164 km/h).

The results of an USAAF evalua-tion of the Fw 190 (Ref. 12 and 13)report the aircraft to have a gentlestall. However, these reports admitthat the Fw 190 stalled abruptly whenmaneuvering. The reason for this re-ported difference in non-maneuveringstall behavior is unknown. A compar-ison of the local wing lift coefficients,calculated by VSAERO, at stall withthe estimated stalling lift coefficientsof the a i r fo i l s two-dimensional ly(Fig. 12) shows that approximatelythe inner 40% of the wing reachesC]max at the same aircraft angle of at-tack. A wartime Focke Wulf report(Ref. 14) indicates that at higher load-ing condit ions (i.e., when p u l l i n gmore gs) elastic deformation of theFw 190 outer wing shifts the load dis-tribution outboard. This would causeeven more of the wing to reach itsstal l ing l i f t coefficient simultane-ously . Combined wi th the sharpstalling features of the NACA 230XXairfoils, this would produce the harshstall found in by Capt. Brown. A gen-tle s ta l l would be evidenced by amore gradual progression of the 2Dstall spanwise.

Initial VSAERO calculations weremade on a model of the Fw 190 A-8.This version of the aircraft was pow-ered by a BMW 80ID radial.Naturally, the question arose as tohow the aerodynamics of this aircraftdiffered from the later, Junkers Jumo213A powered Fw 190 D-9. The Jumo

Figure 8 - Calculated Spitfire windshield boundary layer sep-aration. Separation is calculated to take place at the base ofthe windshield where the streamline traces end. The locationwhere the separated flow is estimated to reattach higher upthe windshield is shown by where the streamline tracesresume.

engine, an inline, is much longer thanthe BMW engine, giving the D-9 aelongated nose, which was counterbalanced with a 500mm plug added tothe aft fuselage. The VSAERO modelwas modified to represent a D-9 bymaking these changes and by addingthe bulged canopy found on Fw 190D-9s. It was found from the VSAEROresults that the fuselage stretch de-signed by the Focke Wulf engineersresulted in a slight increase in stickfixed stability, with the Neutral Pointmoving from 35.8% MAC on the A-8to 40.4% MAC on the D-9. It shouldbe noted these results do not containpropeller effects, which were notmodeled. Flight testing of an earlymodel Fw 190A indicated that the air-craft was "just statically stable; stickfixed and free, engine off; and stati-cal ly uns table to a sl ight degree,engine on" (Ref. 11). During the con-tinued development of the Fw 190series, the aircraft's CG moved rear-ward as fuel tanks and otherequipment was added to the aft fuse-lage (Ref. 15). This Neutral Point shiftduring development of the Fw 190Dmodel would have been quite valuablein maintaining the continued growthof the design.

Drag ComparisonThere are many conflicting claims

as to the equivalent flat plate dragarea (f) of these fighter aircraft. Basedupon my research, what I believe are

SPORT AVIATION 89

Figure 9 - Fw 190 calculated lift coefficient distributionat 1g stall.

Figure 10 - Pressure distribution calculated on the Fw 190 A-8.

the most accurate values are shown inTable 1.

The wetted areas of the aircraft arecalculated by VSAERO, and excludethe ducts for cooling systems.

Notable is that the Mustang hasthe largest wetted area of this groupof aircraft, but has the lowest drag.Evidence of this is that with the sameversion of the Rolls-Royce Merlinand propeller installed, the MustangX was measured to be 23 mph fasterthan the Spitfire IX (Ref. 16). TheMustang X was an Allison powered

Mustang reengined byRolls-Royce with a Merlin65. The P-51B, wi th animproved cooling systemconf igura t ion is evenfaster than the Spitfire IX.The difference in perfor-mance between theMustang and the Spitfireis attributed to several fac-tors. These inc lude thesuperior configuration ofthe Mustang's cooling sys-tem and the Spi t f i re ' srelatively high level of ex-crescence drag, generatedby open wheel wells , a

nonretractable tailwheel and otherdesign details(Ref. 17-19).

One popularpiece of aerody-namic folklore isthe low CDswetvalue achieved withthe Mustang. Vari-ous sources quotethis value as rang-ing from .0038 to.0043. A review ofavailable wind tun-nel and flight testdrag data for theMustang demon-strates the need for

. having all details ofthe aircraft present

if the drag is to be accurately mea-sured. Subscale wind tunnel tests ofthe P-51A and P-51B resulted in val-ues of CDswet, at a representativecruise lift coefficient, in the range of.0046-.0047 (Ref. 20-22). However,these tests usually were of modelslacking exhaust stacks, surface discon-tinuities, etc. Measurements made infull-scale wind tunnel tests of the P-5 IB (Ref. 23) and flight tests of theP-51A (Ref. 24) and P-5 IB (Ref. 21)resulted in a value of CDswet of ap-proximately .0053.

Conclusion ; ; ,

Important design features of threeprominent World War II fighter aircrafthave been examined by the use of amodern Computational Fluid Dynamicsmethod. It is hoped that the results pre-sented here will help demonstrate someof the valuable lessons learned from animportant era in fighter aircraft design.This information, while historical, stillhas relevance in today's world of air-craft design. Important lessons to belearned are:

• Airfoil choice and surface quality areimportant in achieving the advantages oflaminar flow.

• Cooling system duct design forliquid cooled engines must be con-ducted carefully to avoid losses.

• Attention to aerodynamic detail,such as windshield slope, can overcomethe disadvantage of excess wetted area.

• An abrupt stall can be avoided if at-tention is paid to airfoil selection andwing twist.

• As seen with all three of these air-planes, longitudinal stability and controlproblems are common, but can beavoided by the resourceful designer.

r;

Author's NoteThis article is dedicated to Edward

Horkey and Jeffery Ethell, who bothcontributed information vital to thiswork. Ed died as a result of injuries sus-tained in traffic accident in July 1996.Jeff was killed in the crash of a Lock-heed P-38 Lightning in May 1997.

Far too young to have participated inWorld War II, I have long been fasci-nated about finding out how the famousaircraft of this war were designed. Thedeeper I have gotten into this pursuit, themore information I have uncovered thathas proven to be valuable in my dailywork as an aerodynamicist. I have be-come convinced that a study ofihistorical aerodynamicsi is an importantpart of an aerodynamicistis ongoing edu-

TABLE 1

AircraftSpitfire IXP-51B MustangP-51D MustangFw 190 A-8Fw 190 D-9

f5.40 ft2

4.61 ft2

4.65 ft2

5.22 ft2

4.77 ft2

Wetted Area831.2ft2874.0 ft2882.2 ft2735.0 ft2761.6ft2

^Dswet.0065.0053.0053.0071.0063

Ref.1621272626

90 JANUARY 1999

cation. To this end, one of my goals hasbecome to try and disseminate theknowledge I have unearthed, this articlebeing an effort towards this end. Forthose seeking further information in thisregard, I recommend taking a look at myilncomplete Guide to Airfoil Usage! at:http://amber.aae.uiuc.edu/~mselig/ads/aircraft.html.

As mentioned in previous articles, Iam an aeronautical engineer, specializ-ing in applied computational fluiddynamics. Based in Redmond, Wash-ington, I work for Analytical Methods,Inc. My aerodynamic (and hydrody-namic) consulting projects at AMI haveincluded submarines, surface vessels,automobiles, trains, helicopters, aircraftand space launch vehicles. 1 can bereached at: dave@amiwest.com or:

Analytical Methods, Inc., 2133152nd AveNE, Redmond, WA 98052

References

1) Zalovcik, J.A., "A Profile-DragInvestigation in Flight of an Experi-mental Fighter-Type Airplane - TheNorth American XP-51," NACA re-port, November 1942.

2) Bussmann, K., "Messungen amLaminarprofil P-51 Mustang," Aero-dynamisches Institut der TechnischenHochschule Braunschweig, Bericht43/4, January 1943.

3) Matthews, H.F., "Elimination ofRumble From the Cooling Ducts of aSingle-Engine Pursuit Airplane,"NACA WR A-70, August 1943.

4) Morgan, Eric B. and Shacklady,Edward, Spitfire The History, Key Pub-lishing, Stamford Lines, England, 1987.

5) Anon., "Aerodynamic DimensionalData on P-51 B-1 -NA, P-51B-5-NA andP-51C-1-NT Airplanes," North AmericanAviation Report No. NA-5822, August 6,1943.

6) Smith, J., "The Development ofthe Spitfire and Seafire," Journal ofthe Royal Aeronautical Society, Vol.51, April 1947.

7) Meredith, F.W., "Note on theCooling of Aircraft Engines With Spe-cial Reference to Ethylene GlycolRadiators Enclosed in Ducts," ARCR&M 1683, August 1935.

8) Hartshorn, A.S. and Nicholson,M.A., "The Aerodynamics of theCooling Aircraft Reciprocating En-gines," ARC R&M 2498, May 1947.

9) Morgan, M.B. and Smelt, R.,"Aerodynamic Features of GermanAircraft," Journal of the Royal Aero-

nautical Society, August 1944.10) Phillips, W.H. and Vensel, J.R.,

"Measurements of the Flying Qualitiesof a Supermarine Spitfire VA Airplane,"NASA WR L-334, September 1942.

11) Brown, Capt. Eric, "ViewedFrom the Cockpit; Tank's SecondIron," Air International, Vol. 10 No. 2,February 1976.

12) Foster, John Jr. and Ricker,Chester S., "Design Analysis No. 9The Focke-Wulf 190," Aviation, Oc-tober 1944.

13) Van Wart , F.D., "HandbookFor Fw 190 Airplane," USAAF T-2Technical Report F-TR-1 102 ND,March 1946.

14) Gross, P., "Die Entwicklung derTragwerkkonstrukt ion Fw 190,"Bericht 176 der Lillenthal-Gesellschaft,2 Teil, January 1944.

15) Bentley, Arthur L., "Focke WulfFighter," Scale Models, July 1978.

16) Birch, David, Rolls-Royce andthe Mustang, Rolls-Royce Heritage TrustHistorical Series No.9, Derby, England,1987.

17) Private Commu-nication, J. LelandAtwood, October 1994.

18) Anon, "Estima-tion of the Increase inPerformance Obtain-able By Fitt ing aContinuously VariableRadiator Flap," Rolls-Royce ExperimentalDepartment Report,August 10, 1942.

19) Private Commu-nication, Ed Horkey,November 1994.

20) Anon., "WindTunnel Data For XP-51B Airplane (NAAModel NA-101),"North American Avia-tion Report No. NA-5548, October 9,1943.

21)Nissen, J.M.,Gadeberg, B.L. andHamil ton, W.T.,"Correlation of theDrag Characteristicsof a Typical PursuitAirplane ObtainedFrom High-SpeedWind-Tunnel andFlight Tests," NACAReport 916, 1948.

22) Staff of the RAE High-SpeedWind Tunnel, "High Speed Wind-Tunnel Tests of Models of Four SingleEngined Fighters (Spitfire, Spiteful,Attacker and Mustang), Parts 1-5,"Aeronautical Research Council R&MNo. 2535, edited by W.A. Mair, 1951.

23) Lange, R.H., "A Summary of DragResults From Recent Langley Full-ScaleTunnel Tests of Army and Navy Airplanes,"NACA ACR L5A30 (WR L-108), 1945.

24) Staffs of the High-Speed Tunnel andHigh-Speed Flight Sections, "Research onHigh-Speed Aerodynamics at the RoyalAircraft Establishment from 1942 to 1945,"Aeronautical Research R&M No. 2222,edited by W.A. Mair, September 1946.

25) Anon., "Performance Calculations forModel P-51D-5-NA Airplane (NAA ModelNA-109)," North American Aviation ReportNo. NA-8449, December 1,1944.

26) Anon., "Widerstandaten vonFlugzeugen" (Drag Data for Aircraft),Focke Wulf data sheet. *

Figure 11 Pressure distribution calculated on the Fw 190 D-9.

2.0

1.8 -

1.6 -

1.2 -

1.0 -

0.4 -

0.2 -

0.0

Stalled RegionMmax

Calculated C| Distribution

0.3 0.4 05

Semispan FractionFigure 12 - Boundary layer separation calculated in the Spitfirecooling system.

SPORT AVIATION 91