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RESEARCH MEMORANDUM -

EFFECT OF NOSE SHAPE AND TRAILING-EDGE BLUNTNESS ON

THE AERODYNAMIC CHARACTERISTICS OF AN UNSWEPT

WING OF ASPECT RATIO 3.1, TAPER RATIO 0.4,

AND 3-PERCENT THICKNESS

By- John C. Heitmeyer

Ames AeronauticalMoffett Field,

Mborator ycam.

*M - -wiE&tM----- AX .-. . . . —.

NATIONAL AD~lSOFOR AERONAUTICS

WASHINGTONMarch 15, 1954

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NACA RM A54A04

NA’ITONALADVISORY cowm FOR

RESEARCH MEMORANDUM

AERONAWITCS

EFFECT OF NOSE SHAPE AND TRAILING-EDGE BIUWlltESSON

THE AERODYNAMIC CHARACTERISTICS OF AN UNSWEPT

WING OF ASPECT RATIO 3.1, TAPER RATIO ok,

By

S-mcm THIcmzss

John C. Eeitmeyer

SUMMARY

The effects of blunting the trailing edge and/or rounding the lead-ing edge upon the lift, drag, and pitching-moment characteristics of aplane tapered wing in combination with a body have been experimen~lyinvestigated at Mach numbers ranging from 0.61 to 0.93 and from 1.20to 1.$)0. Results indicate that blunting the trailing edge to 0.3 of themaximum airfoil thickness reduced the forward movement of the aerodynamiccenter, noted for the sharp-trailing-edge sections at high subsonicspeeds without, in.general, a significant reduction in the maximum lift-drag ratio. At all Mach numbers, blunting the trailing edge increasedthe minimum drag. Rounding the leading edges of the wings having eitherthe sharp or the blunt trailing edges decreased the minimum drag andjtith one exception, increased the maximum lift-drag ratios at subsonicspeeds. At supersonic speeds the opposite effects were noted.

INTRODUCTION

It was shown in reference 1 that for unswept wings with sharp trail-ing edges the forward movement of the aerodynamic center with increasingMach number at high subsonic speeds was reduced or eliminated if thetrailing-edge thickness was made equal to or greater than one-half themaximum thickness of the section. Unfortunately, this improvement instability characteristics was accompanied by an increase in minimum drag .and a decrease in msximum lift-drag ratio. It was_suggested, therefore,that in light of the results of reference 2, it might be possible torealize the improved pitching-moment characteristics without the delete-rious effects on drag by thickening the trailing e@e to less than

2 NACA RM A54A04

one-half the maximum thickness of the section. The present investigationwas undertaken, therefore, to determine experimentally the aerodynamiccharacteristics of the wing-body combination of re~erence 1, employingthe same plane tapered wing of aspect ratio 3.1 with 3-percent-thick,circular-arc, biconvex sections, but modified to have a trailing-edge-thickness equal to 0.3 of the maximum airfoil thickness. The presentinvestigation was extended to include also wings having the more con-ventional round-nose airfoil sections to determine if the aforementionedbeneficial effects due to blunting the trailing edge would be realizedon wings having round-nose airfoil sections.

The experimental data for the models with wings having sharp traili-ng edges were obtained from the tabulated-results of reference 3. Theresults for the models with the blunt-trailing-edge wings were obtainedduring the present investigation and are presented herein.

NOTATTON

a.c.

b

CD

‘%nin

CL

cm

c

E

dCL

-Z

aerodynamic center position, percent 5 from leading edge of @

wing span

dragdrag coefficient, —

qs

minimum drag coefficient

liftlift coefficient, — .

qs

pitching-moment coefficient,pitching moment

-qsE(Pitching moments were referred to a horizontal axis throughthe quarter point of the wing mean aerodynamic chord.)

local wing chord

p’ C2W

mean aerodynamic chord of wing, o

/:i2C dy

rate of change of lift coefficient with angle of attack, per deg(The slope is measured between lift coefficients of -0.2 –and +0.2.]

~

—

m

—

●

.—

.=

.

,—

.

.— .-

F

,

J

NACA RM A54A04 3

-.

.

.

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lirt-iwg ratio

maximum lift-drag ratio

thickness of the trailing edge

free-stresm Mach number

free-stream @mmic pressure

Reynolds number based on the mean aerodynamic chord of the wing

radius of body

wing area formed by extending the leading ~d trailing edges tothe plane of symmetry

msximum wing thickness

longitudinal distance from nose of body

lateral distance measured perpendicular to plane of symmetry

angle of attack of body axis, deg

APPARATUS

The experimental.investigation was conducted in the Ames 6- by 6-footsupersonic wind tunnel. In this wind tunnel, the test Mach number canbe varied continuously and the stagnation pressure can be regulated tomaintain a given test Reynolds number.

The models were mounted on a straight sting in the wind tunnel, thediameter of the sting being about 93 percent of the diameter of the bodybase. A k-inch-diameter, four-component, strain-gage balance enclosedwithin the body of the model measured the aerodynamic forces and momentsexperienced by the model.

The model and pertinent dimensions are shown in figure 1. The.air-foil section of each of the four wings considered in the present inves-tigation is illustrated in figure 2. The sectionsing edge were derived by thickening the section atto 0.3 of the maximum thickness and fairing to the

a..........

having a blunt trail-the trailing edgebiconvex surface by

means of straight lines. The forward half of the round-nose sectionsWZLSa semiell.ipsewhich for the thickness ratios considered closelyresembled the NACA 66-003 airfoil section.1

.-.

The solid steel wing of reference 1, having a 3-percent-thickbicon-vex section, was modified, by the addition of bismuth-tin alloy, to

2

obtain the desired contours for the remaining three wings. With theexception of the aluminum nosepiece, the body was constructed of steel.

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All exposed surfaces of the wing and body were smooth and polished.

REDUCTTON OF DATA

A complete discussion of the methods used to reduce the wind-tunneldata to coefficient form and of the various corrections applied to thedata can be found in reference 1. These corrections, which were of thesame magnitude for all models, account for the following factors:

1. Induced effects of the tunnel walls at subsonic speeds resultingfrom lift on the model. The magnitude of these corrections which wereadded to the uncorrected data were as follows:

● -”

Aa = 0.57 CL, deg .

ACD = 0.0100 CL2

2. Change in the~velocity of the air streem in the vicinity of themodel at subsonic speeds due to constriction of the flow by the tunnelwalls. At a Mach nuniberof 0.93 (the maximum subsonic Mach number atwhich data up to kh” angle of attack could be obtained without chokingthe wind tunnel) this correction amounted to about a ~-percent increasein the Mach number over that determined from a calibration of the windtunnel without a model in place.

3. The longitudinal force experienced by the body of the model dueto the streamwlse variation of static pressure measured in the test sec-tion at subsonic and supersonic speeds without a model in place. Thiscorrection varied from -0.0007 (at a Nkch number of 1.3) to +0.0010 (ata Mach number of 0.93).

It should be noted that the drag coefficients presented herein havebeen adjusted to represent drag coefficients for which the fuselage basepressure would be equal to the free-stresm static pressure. .-

%Fhe ordinates of the NACA 66-003 section were obtained by halving theordinates of the NACA 66-006 section.

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NACA RM A74A04 5 ““-

Tests of the models having the round-nose airfoils in both an uprightand inverted position have indicated that an effective stream angle ofabout -O.1° exists in the pitch plane of the models at both subsonic andsupersonic speeds. The data presented herein have not been corrected forthis effect.

RESULTS AND DISCUSSION

The data obtained during the present investigation of two tq,peredunswept tings with blunt-trailing-edge sections have been tabulated intables I and II. Comparable data for the wings having a sharp-trailing-edge section ha’vebeen ~resented in the tables of reference 3. Analysisof the data from reference 3 showed that the drag data obtained at thelower test Reynolds numbers (1.4 million and 2.4 million) at subsonicspeeds exhibited considerable asynmetry between data obtained at positiveand negative lift coefficients. At supersonic speeds, the slight asym-metry observed was considerably less than at subsonic speeds and wasconsidered to be within the accuracy of the drag measurements.

As a consequericeof the above analysis, only data obtained at thehighest test Reynolds’number (3.6 million) are presented graphically at .

both subsonic and supersonic Mach numbers. At supersonic speeds, how-ever, data obtained at a Reynolds number of 2*4 million have also beenpresented @aphically for all models in order to extend or supplementthe range of Machsonic MacB number

nunibersjsince power limitation restricted the super-range at the highest test Reynolds nwiber.

Effect of Trailing-E&e Bluntness

Lift and pitching-moment characteristics.- The variations of liftcoefficient with angle of attack and of pitching-moment coefficient tithlift coefficient for the models under consideration are presented infigures 3 and 4, respectively. The variation with Mach number of thelift-curve slope and position of the aerodynamic center are presentedin figures 5 and 6, respectively. The results show that blunting thetrailing edge of a sharp-nose airfoil increased the value of lift-curveslope throughout the lift-coefficient range, as shown in figure 3, andthe Mach number range, as shown in figure 5. This increase was mostevident at subsonic speeds and in the range of lift coefficients, -0.2to 0.2 (for which the results of fig. 5(a) are applicable), wherein thewing with the biconvex section exhibited low values of lift-cum?e slope.These small values of lift-curve slope near zero lift have %een obse&vedin two-dimensional investigations at subsonic speeds of 6- and 10-percent-thick biconvex sections at Reynolds numbers comparable to those of the

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6 NACA RM A54A04

present investigation (refs. 4 and 5). Reference 4 indicates, however,that at Reynolds numbers approaching full scale (18 million), the non-linear variation near zero lift is eliminated and the value of the lift-curve slope is substantially increased. It is believed,therefore, thatthe benefits indicated in figure 5(s.)that result from blunting thetrailing edge of the sharp-nose airfoil are optimistic, amd that at full-scale Reynolds numbers such benefits at subsonic speeds would resemblemore closely those shown for the round-nose section (fig. 5(b)). Theincrease in lift-curve slope, due to blunting the trailing edge of theWing with the round-nose section, at the high subsonic Mach numbers isattributed to the greater lifting pressures developed over the rear por-tion of the airfoil, due to the rearward shift of the terminal shocks.This increase in the loading (lifting pressure) near the trailing edgeat the high subsonic Mach numbers is indicated also by the more negativepitching-moment coefficients shown in figure 4. At supersonic speedsthe observed increase in the vaJue of the lift-curve slope due to blunt-ing the trailing edge of the wing with the sharp-nose airfoil is greaterthan would be predicted by two-dimensional second-order theory. Thereason for this effect is not known at present.

Blunting the wing trailing edge to 0.3 of the maximum thickness ofthe airfoil for the wings having either a rouud-nose or sharp-nose air-foil reduced the over-all travel of the aerodynamic center with variationin Mach number (fig. 6). This result was in agreement with that shownby reference 1 for wings with trailing-edge thickness equal to or greaterthan one-half the maximum thickness.

Drag characteristics.-The variation of the drag coefficient at var-ious lift coefficients with Mach number for the models under considera-tion is presented in figure 7. Blunting the trailing edge increased theminimum drag of the models with either the sharp-nose or round-nose air-foils, at all test Mach numbers. At subsonic speeds the increment inminimum drag, due to blunting the trailing edge, was about constant forboth the sharp-nose airfoil and the round-nose airfoil. However, atsupersonic speeds, the variation with Mach number of the increment inminimum drag for the sharp-nose airfoil and the round-nose airfoil wasnoticeably different. In the former case the increment remained aboutconstant, whereas in the latter case the increment decreased with increas-ing Mach number. Results of a free-flight investigation (ref. 6) of awing-body combination, having a plan form similar to that of the modelof the present investigation, indicate that at a Reynolds number of about8.o million the increment in drag due to blunting the trailing edge of awing with a k-percent-thickbiconvex airfoil decreases with increasingsupersonic Mqch number. Thus, it is possible that at higher Reynoldsnumbers the increment in drag noted for the sharp-nose airfoils woulddecrease with Mach number in a manner similar to that noted for theround-nose airfoils.

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NACA RM A54A04 7

Blunting the trailing edge of the wings, having either the sharp orthe round leading edges, in general, reduced the maximum lift-drag ratioonly slightly at subsonic speeds (see fig. 8), in spite of the largeincrease in the minimum drag noted previously. The favorable character-istics of drag tie to lift indicated at subsonic speeds are believed tobe related to an increase in the base pressure (negative drag) at theblunt trailing edge of the wings with amgle of attack. (See ref. 1.)At supersonic speeds, the results of figure 7 indicate that the effectsof blunting the trailing edge upon the characteristics of drag due tolift of these wings were influencedby the shape of the leading edge.‘Ihecharacteristics of drag due to lift of the wing with the round lead-ing edge were improved by blunting the trailing edge; as a result, themsximum lift-drag ratios for the round-nose airfoil were not affected(fig. 8). The characteristics of drag due to lift of the wing with thesharp leading edge were not influenced significantly by blunting thetrailing edge, with the result that the maximum lift-drag ratios werereduced due to the increase in minimum drag.

Effect of Variation in Trailing-Edge ‘Thickness

The variation of several aerodynamic parameters with trailing-edgethickness parameter h/t for the model having a sharp-nose airfoil, atfive selected Mach numbers, is presented in figures 9 and 10. Theresults presented for thicknesses of one-half and greater were obtainedfrom reference 1.2 ‘l%edata indicate that the value of the lift-curveslope at each Mach number is relatively unaffected by thickening thetrailing edge beyond h/t = 0.3. IMcept at a Mach number of 0.91, theposition of the aero@xamic center was not influenced by variations inthe trailing-edge thickness. At all Mach numibers,the minimum dragincreased with increasing value of h/t. The maximum lift-drag ratiosare reduced slightly as the trailing edge is thickened from O to 0.3 ofthe maximum section thickness and then are decreased appreciably withfurther increases in trailing-edge thickness.

Effect of Wing Nose Shape

Lift and pitching-moment characteristics.- The variations of liftcoefficient with angle of attack and of pitching-moment coefficient withlift coefficient for the models are presented in figures U_ and E?,respectively. The variation with Mach number of the lift-curve slopeand position of the aerodynamic center are presented in figures 13and 14. With the exception of the lift-curve-slopes for the wings with

2Data obtained at subsonic speeds were corrected in the manner describedin reference .3.

8 NACARMA54A04

sharp trailing edges at subsonic speeds, little difference was noted inthe lift and pitching-moment characteristics near zero lift for wingswith either sharp or round leading edges (see figs. 13 and 14). Asmentioned previously, however, the small lift-curve slopes of the wingwith a biconvex section at subsonic speeds are believed to be a Reynoldsnumber phenomenon; at full-scale Reynolds number, it would be expectedthat the lift-curve slope for this wing would be larger and approachthat of the wing with the round-nose sharp-trailing-edge section.

.

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At lift coefficients outside the range for which the results offigures 13 and 14 are applicable, -0.2 to 0.2, the most significant effectof rounding the leading edges of these wings occurred at a Mach number of0.91 (see figs. 11 and 12). For wings with either sharp or blunt trailingedges, rounding the leading edges reduced the lift coefficient at whichthe wings became neutrally stable by approximately 0.1.

Drag characteristics.- The results of figure 15 indicate that theeffects on the minimum drag due to rounding the leading edges of thewings were influenced by Mach number. At Mach numbers above 1.2, theminimum drag coefficients were greater for the models with round leadingedges while at subsonic speeds, the opposite effect was obtained. Thereduction in minimum drag at subsonic speeds is believed to be relatedto the larger regions of lsminar boundary layer present on the wings with

.

rounded leading edges (see ref. 3). The increase in minimum drag observedat Mach numbers greater than 1.2 is related directly to the higher wavedrag associated with the bow wave being detached from the round leading

.

edges.

A comparison of the ~tation of msximum lift-drag ratio with Machnumber (fig. 16) shows that rounding the leading edges of the wingsIncreased the (L/D)u significantly only at the lower subsonic Machnumbers and decreased the (L/D)H at supersonic speeds. me increase

noted at the lower subsonic Mach numbers was due to both a reduction inminimum drag and in the bag due to lift (see fig. 15). Examination ofthe results indicates that when the critical Mach nuniberis exceeded,the value of the drag due to lift for the wings tith round leading edgesapproaches that for the wings with sharp leading edges. The effects ofrounding the leading edges of the wings on the maximum lift-drag ratiothen results primarily from changes in the miriimumdrag coefficient.

CONCLUSIONS

The present report presents results of a wind-tunnel investigationto determine the effects of blunting the trailing edge and of roundingthe leading edge upon the aerodynamic characteristics of a plane taperedwing of aspect ratio 3.1 in combination with a body. For the Mach

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:YNACA RM A54A04

numbers considered in the present investigation, 0.61 to 0.93 and 1.20to 1.90, the following results were obtained:.

1. The investigation of the effects of blunting the trailing edge. of wings, having either round- or sharp-leading-edge airfoils, to 0.3 of

the maximum airfoil thidsness showed that:

(a) At high subsonic speeds and at a Reynolds number of3.8 million, blunting the trailing edge increased the lift-curve slope and reduced the forWard movement of the aero-_ic center with Mach number without, in general, anappreciable decrease in the vslue of the maximum lift-dragratio.

(b) At all test Mach numbers, blunting the trailingedge increased the minimum drag.

2. The investigation of the effects of variation in the trailing-edge thickness of the wing having & sharp-leading-edge airfoil showedthat for the range of Mach numbers considered, there were no aerodynamicbenefits to be derived by thickening the trailing edge beyond 0.3 of the

. maximum airfoil thickness.

3, The investigation of the effects of rounding the leading edge. of wings having either a sharp or a blunt trailing edge of 0.3 the maxi-

mum section thickness showed that:

(a) Rounding the leading edge reduced the minimum dragand, in general, increased the ~imum lift-drag ratio atsubsonic speeds while having the opposite effect at super-sonic speeds.

(b) Only the lift characteristics of the sharp-trailing-edge airfoil were sffected by rounding the leading edge.These effects were, however, confined to data obtained atsubsonic speeds and in the vicinity of zero lift.

(c) At all Mach numbers, the pitching-moment charac-teristics of either the blunt- or sharp-trailing-edgemodels in the region of zero lift were not influenced byrounding the leading edge. At high subsonic speeds (M=O.91)rounding the leading edge reduced the lift coefficient atwhich the longitudinal stability became neutral.

.Ames Aeronautical Laboratory

National Advisory Committee for Aeronautics. Moffett Field, Calif., Jan. h, 1954

10 NACA RM A54A04

REFERENCES

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1. Dugaa, Duane W.: Effects of Three ~es of Blunt Trailing Edges onthe Aerodynamic Characteristics of a Plane Tapered Wing of AspectRatio 3.1, With a 3-Percent-Thick Biconvex Section. NACA RM A52E01, -1952.

2. Cleary, Joseph W., and Stevens, G=rge L.: The Effects at Transoni.cSpeeds of Thickening the Trailing Edge of a Wing With a &PercentThick Circular-Arc Airfoil. NACA RMA51Jll, 1951.

3. Hall$ Charles F.: Lift, Drag, and Pitching-Moment of Low-Aspect-RatioWings at Subsonic and Supersonic Speeds. NACA RM A53A30, 1953.

4. Underwood, William J.,Tunnel InvestigationCircular-Arc Airfoil1946.

5. Summers, James L., and

and Nuber, Robert J.: Two-Dimensional Wind-at High Reynolds Numbers of Two SymmetricalSections With High-Lift Devices. NACA RM L6K22,

Page, Wiliiam A.: Lift and Moment Character-istics at Subsonic Mach Numbers of Four 10-Percent-Thick AirfoilSections of Varying l%ailing-Edge Thickness. NACA RM A50J09, 1950. -

6. Morrow, John D.: Measurements of the Effect of Trailing-Edge Thick-ness on the Zero-Lift Drag of Thin Low-Aspect-Ratio Wings.

.NACA

RM L50F26, 1950.—

NACA RM A54A04 11

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TABLE I.- WIND-TUNNXL DATA FOR AN UNSWEPT WING OF ASPECT3-PERCENT-TRICK,ROUNDED-NOSE, BLUNT-’IRAILING-EDGE

EwITo3.1 WITHSECTION

.

.

TABLE I.- WIND-TUNNEL

NACA

DATA FOR AN UNSWEPT WING OF ASPjCT RATIO

RMA54A04

3.1 mm3-PERCENT-!HIICK,ROUNDED-NOSE, BLUNT-’IRAILING-WE-SEC~ ON-Concluded

.

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.,

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HACA RM A54A04 \ 13

TABLE II.- WIND-TUNNEL DATA FOR AN UNSWEPT WING OF ASPECT RA’ITO3.1 WITH3-PERCENT-THICK, SHARP-NOSE, BLUNT-TRAILING-EDGE SECTTON

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All dimensions

shown in inches

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n Equafr’bn of

Moment Center

— 24.731

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Figure 1.- Plan view and front view of model.G

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r f =0.03c for 0// Wings

~E/lipticai contour

I’&’E 0.50C

t Point of tangency

Note z Sketches not

to scale h=L!3 t

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Point of ttmgency

Points of tangency

1

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II(a) Sketches of airfoil section of varioue wings. I

Bicawex contour __ -., -——— -—-. —-- -

-- —//””

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---- ----- — / s

Effipticai contour--—- . ..— *

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(b) Detail of leading ewe (true ficale].~

Sectionsof variuus wings*g

R@llre 2.- 4=

, . .

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.8

.6

.4

-.2

-.4

Angle of attack, c, deg

Figure 3.- l&kCt Of blunt t?.’at~~

angle of.

(a) Sharp-nose airfoils.

edge (h/t = 0.3) on %he vexiationof lift coefficientwithattack at various Mach mm?mrs.

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.6

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Angle of ot!ack, U, deg

(b) Round-noee airfoils.

Figuxe 3.- Concluded.

I

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B , 6 . . .

P03

, ,, *

(a) Sharp-nose airfoils.

Figure 4.- Effect of blunt trailing edge (h/t = O. 3) on the vaxiati.onof pitching-momentcoeffi-cient with lift coei’ficieritat various Mach mmbers.

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Pitching- munent coefffcikmt, Cm

(b) Round-nose airfoils.

Figure 4.- concluded.

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3.8x106 24x106o ●

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m

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Mach number, M

(b) Round-nose airfoils.

FWLre 5.- Effect of blunt trailing edge (h/t = 0.3) on the variation of lift-me slop with

Mach number.

.8 /.4 1.6 /.8 2.0Moch%rnberfff

(a) Sharp-rime airfoils.

.8 10 /.2 L4 16 1.8 2.0

0 .2 4 .6 .0 LO 12 /.4 L6 18 2.0

Mach number, M

(b) Round-nose airfoils,

Figure 6.- Effect of blunt trailing edge (h/t = O. 3) on the variation of’ aerodynamic center with

Mmh mmiber.

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NACA RM A54A04 ~! 23

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R38x1062.4 x106 -

~’ n m

o .2 .4 .6 .8 1.0 1.2 1.4 !.6 1.8 2.oMach number, M

(a) Sharp-nose airfoils.

Figure 7.- Effect of blunt trailing edge (h/t = 0.3) on the vsriationdrag coefficient at various lift coefficients with Mach number.

of

.10R

3.8 X /0624X 106

< > 0 ●

c >❑ 9

HACA RM A54A04

o 2 .4 .6 .8 f.o IJ? 1.4 1.6 \e8 P.c3Mach number, M

(b) Round-nose air~oils.

Figure 7.- Concluded,

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NACA RM A54A04 25

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10 ‘

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3.8 X 1062.4X 106

6 ~o ●

~ng

40

Figure 8.-

.2 .4 .6 .8 10 /!2 M I@ IBMach number, M

(b) Round-nose airfoils.

Effect ofbhnt trailing edge (h/t) = 0.3) on theof maximum lift-drag ratio with Mach number.

20

variation

/)/t

MR

Z8X106 2.4x106

0.61 0 ●

0.91i.20 : :/.40 A A

/.70 A A

o .2 .4 .6 .8 1.0

fl/ t

Figure 9.- Variation of lit’t-curve slope and aerodynamic center of the sharp-nose airfoil with

trailing-edge thickness pszameter, h/t.

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L70 : A

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Figure 10. - Variation of mimlmum drag coefficient and maxlmm LMt-d.rau ratio of the sham-noseairfoil with tAfng-edge thickness parameter, hit.

-.

1IIr

,8

.6

-.2

-.4-4 0 4 8 L? for M=O.61

mg)e ofaftack, c, deg

(a) sharp-trailing-edge

Figure 11.- ~ect of rounding the section leading ed8e on

angle of attack.

airfoils.

the variation of lift coefficient with

z! I

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.8

.6

.4

I 1 1 I-I I I 1 ( I 1 1 1 I 1 1 I 1 A- 1 1 1 1 1

M=(26I

R=3.8xI@ R%%% R%%O@ R%%b-

-.2

-.4

Angle of aitack, c,deg

(b) Bhnt-trailing-edge airfoils.

Figure 11. - Concluded.

—.

(a) Shin-p-trailing-edge airfoU.s.

Fim I-2.- Effect Of rcunding the section leadlng edge on the mriation of pitching-mnnent coef-

ficient with lift coefficient.

.. , , ,

wo

, B

-..2T l-i m /

y

d4 /

/-.4

.04 0

Pitching-mwrmt coefficient, km

I L#l I I I I~o

I1 I / , , I 1 , I I

~~I I I U-.4 I I

-.04 -,08 -.12 for M=O.61

(b) Blunt-treiling-edge airfoils.

IHgure U.- Concluded.cdP

wn)

I 3.8x106” &xf06 jI

I Iv I I Io

❑

●

m

.06

.04 a I I I I I .

~~ .2 .4 ,6 .8 1.0 L2 !.4 1.6 1.8

1111111

Au I I

G Mach number, M$ (a) Sharp-trailing-edge airfofls.

12.0

.10R

3.8x106 2.4xt06o ●

.08 ~, ❑ ■

.06 . I

II I II II II IF41

.04 I 1 I I I I I I I I I I I I I I.- -Q .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0

Much number, M =5=

(b) Blunt-trailing-edge airfoas.

m 13. - Effect of l’ounding the section leading edge an the variatim of lift-~ slope with

Mach number.

* ● a

I

I 9

E Mach number, M&

~ 60(a) Shem-tmiltng-eUe airfoih.

b?1?

. 3,8x 106 24x106k ~o ●

g 40 ~—~ ~4

8 1P ‘- - “u&E~ 20+Qii~o

O 2 ,4 A .8 1,0 1.2 14 1.6

Mach number, M ‘w

(b) Blunt-trailing-edge airfoils.

Figure 14. - Effect of rounding the eection leading edge

with Mach number.

on the variation of aerodynamic center

1

WIw

34NACA RM A54A04

JO

.09

.08

.07

.06

&~.

s.- .05$LeQ0 .04GpQ

.03

.02

.01

0

R

3.8x fO=2.4 Xd

e . 0“ 1~

1 1 1 I t I

t

I 1 I I –-

I= I I

-1w

I

(

t 1 I I I I I I I !

o .2 .4 .6 .8 f.o 1.2 1.4Mach number, M

‘“6 -

(a) Sharp-trailing-edge airfoils.

Figure 15.- Effect of rounding the section leading edge on the variationof drag coefficient at various lift coefficients with Mach number.

.

.

NACA FM A54A04 35

R

3.8 x 1062.4xI06

n n

a

.020sQQ=- B16*Q=~

~ .012“g

*E .008

2~

z .0040 .2 .4 .6 .8 1.0 1.2 1.4 1.6

Mach number, M=&z+

(b) Blunt-trailing-exe airfoils.

Figure 15.- Concluded.

36 NACARMA54A04

g Much number, M*I (a) Sharp-trailing-edge airfoils..*

16

/4

f,

10

#

R

3.8x 1062.4 X (066 .

< >4

❑ 8v.

o .2 .4 “ .6 .8 /.0 1.2 L4 IS ~8 2.0

Much number, M

‘.’m ‘“ “““ ““(b) Blunt-trailing-edge ai~foils..:.,A ““’’f.>

gurenxf$- Effect of rounding the section leading edge on theo~’-_wimum lift-drag ratio with Mach number.,b+~’”” , .-

J

-# L

@* +. -*Q~at& ,., =..-

variation

.—

.

.

.

.—

.—

.

-M-!= NACA-Langley - 3-15-S4 -325