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RESEARCH MEMORANDUM

THE EFFECTS AT TRANSONIC SPEEDS OF THICKENING THE

TRAILING EDGE OF A WING WITII A 4-PERCENT-

THICK CIRCULAR-ARC AIRFOIL

By Joseph W. Cleary and George L. Stevens

Ames Aeronautical LaboratoryMoff ettField, Calif, -

?Nassificationcane.’l~d(IXchanudto~~~~~@f~.4@............)

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NATIONAL ADVISORY COMMITTEEFOR AERONAUTICS

WASHINGTONDecember 11, 1951

.

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TECH LIBRARY KAFB, NM.

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NATIO~ ADVIEX2RYCOM@ZMEE TOR AERONAUTICS

RESEARCH MEMORANDUM

-1

.

.

EEE’EECTSAT TRANSONIC SPEEDS CD?THICKENING THE

TRKILIRGEX+EOFAW~W~A~~- ‘

THICK C3RCUIAR=KBC AIRFOIL

By Joseph W. C!learyand George L. Stevens

SUMMARY

.—

The effects of a systematic variation of trailing~dge thicknessof a symmetrical circ~ c airfoil on the aer&Qnamic forcej momentjbase-pressure, and wake fIuctuations have been tivestigated using thetransonic-bump testing technique. An investigation of the effects ofone boattail modification was also made. The atifoils were 4 percentthick, of rectsmgular plan form} and of aspect ratio 4. The test-covered a Mach nunber range from 0.60 to 1.10 with a correspondingReynolds znmtxx range from about “1.’7to 2.2 million,

At subsonic Mach nunbers, the results show a beneficial effect onthe lift-drag ratios with no measurable increase in minimum drag coef-ficient for a trailing+dge thickness equal to 0.3 of the airfoil.thick-,ness. Higher li-urve slopes were observed in the transonic Machnunber range for all the blunt-trailing~dge airfoils as compared tothe sharpedged airfoil.and higher maximum lift coefficients were notedat 0.6 Mach nunibere Surface ro@ness appeared to have a significanteffect on the pitching+noment characteristics of the circuLar+uc andboattailed airfoils, particularly at high s~sonic lkch nunibers, Base-pressure coefficients for the blunt-trailing+Xge airfoils increasedfrom root to tip. Ecreasing the trailing-edge thickness generaldy

.-.

caused a decrease of basepressure coefficient-

Surveys in the wakes of the airfoils indicated apwoximatel.y thesame level of wake fluctuaticms for trailing+dge thicknesses of O and0.3 of the airfoil thickness> but indicated marked increases in wakefluctuations for trailing-edge thicknesses of 0.6 and 1.0.

4

IMWWIIENT‘“:$iwmui -. ‘

E==@_—

2 —— .,,.. NACA R.MA51Jll

INTRODUCTION .

At supersonic Mach ?mubers, the airfoil section having minimumdrag for a prescribed structural strength or stiffness may have a blunttrailing edge, as has been shown through theoretical considerationsbyChapman in reference 1. The twtiimeasionaJ characteristics of suchairfoils of moderate thickness as co~ared with a more conventionalairfoil having a sharp trailing edge have been investigated experimen-tally at subsonic Mach nunbers (reference 2). Although the results ofthat investigation indicated higher minimum drag for the blunt-trailing-edge airfoil} gains in maximum lift were observed and the lift--curveslope increased at Mach nuniberswhere the Lift-curve slope decreasedfor the sharp-trailing+dge airfoil. Thus, it would seem that blunt-trailing~dg~ airfoil; nm~be of some practical value at.both subsonic _. ..and supersonic Mach nrmbers. —.

The present investigationwas undertaken to evaluate the effect , .of’an increase in trailing+dge thickness on the aerodynamic character-istics of a thin three-dimensionalwing in the transonic I@ch numberrange. For this investigation} the thick trailing+dge airfoils wereformed by building up the trailhg edge to the desired thickness andthen fairing to the origtil a@foil by straight lines. The forwardportion of the ctrcular—arc section remained intact and the resulting -airfoil was not one of the optimum sections.derived in reference 1.

.

—

.

--

drag c.oefficknt

lift coefficient

NOTATION

(twtbe semisnan aqs &‘)

(twice semisnan ift

qs ‘)

cm pitchi~ment coefficient, referred to

.

m

(twice semispan pitching momentqsE )“

amplitude of total-pressure,flucttition,square foot

lff-ag ratio .-

0.25 75 ,..

pounds per

L.

.NACA RM A!51XU mgg@?I!I!-. 3

---Irax

Mach number

local.hlachzumher

base-pressure coefficient(%%9

Reynolds nurdxsrbased on mean aero@mmnic chord

total wing area (twice wing area of semispan model),square feet

velocity, feet per second

twice span of semis- mcdel, feet

local wing chord, feet

/’/”/2 c2d.\mean aerodynamic chord (Jo -\

\ f-’b/2 j, feet

\J Cw’trailing+dge thiclmess, f~et /

free+ tream static pressure, pounds per square foot

base pressure, pounds per square foot

free+ tream dynamic pressurefoot

maximum wing thictiess, feet

spanwise distance from plane

airfoil thickness ratio

angle of attack, degrees

air density, slugs per cubic

().,~PQ2

2p Pounds per square

of symmetry, feet

foot

—.-@NFngm.y. NACA RM A51JI-1

.4

dc~TH7

a.c.

slope of lift curm, per degree

aerd.ynamic-center position, percent

M’PmTus AND mlEIS

The tests were conducted in the Ames 16-foottunnel. The bump testing technique, as described

-.

of=” ..

high-speed windin reference a, was—

employed to extend the test Wch nmiber into the superscmic range.- Themodels were camtilever=muunted on an electrical strain-gage halmcepermitting simultaneous measurements of lift, drag, and pitching moment.Figure 1 shbws typical model installations on the transonic bump.

The m&lels were rectangular airfoils ham an effective aspectratio of 4, as illustrated in figure 2. The models were mule of steel-.End plates were mounted near the root to reduce the flow leakagethrough the bump. The sections of the blunt-trailing-edge airfoilswere derived from a symmetrical circular-arc section by building up th6trailing edge to the desired thictiess and then fairing to the originalairfoil section by straight lines as illustrated in,figure 3. Thus theblunt-trailing-edge airfoils retained the &percent maximum thiclmessof the basic circular-arc section and the forward”portionremainedintact. The airfoils kd trailing-edge thicknesses of 0, 0.3, 0.6, and1.0 times the mximum thiclmess.-

One airfoil wastested which had a boattail trailing edge. Itssection was formed by modifying the airfoil with a 1.0 trailing-edge-thickness ratio by chamfering the trailing edge to 0.6 t~c~e~s fi”~ “.O.~ chord to the tmiling edge. (See fig. 3.}

fi Base pressures were measured by means of an orifice at each offour spanwise stations along the blmt trailing edges. These spanwisestations were at 25, 3’7.5,62.5, and 87.5 percent of the wing semi-span. A mercury manometer was used to measure the pressures.. Waketotal-pressure fluctuations were measured by means of quick-responsepressure cells mpunted on probes one chord length ”behindthe wingtraili~ edge. These probes measured fluctuating pressures at three

— —

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—

—- .—.- —

.

-. ---

. —

-.

.. - :-

. ..

..

points & t% wakes as-shown infrom these cells were amplifiedrecorded & an oscillograph.

figure 1. The electrical impulses.-

with a carrier-currentamplifier andz

-.—.... l..—

.

NACA RM A5UU.

----

.

5

TESTS

Measurements nwle during the tests consisted of lift, drag, pitc&ing moment, base pressures, ad fluctuations of totil pressures in thewakes of the airfoils.

Tests with surface roughuess on each airfoil were made to evaluateits effect on the force sad moment characteristics. The surface rough–ness consisted of No. 60 Carborundum grains lightly sprayed on a bondingagent on ths upper and lower surfaces from the leading edge to 0.10chord.

The tests covered a &chnuniber raage from O.60 to 1.10. Since theflow deflection angle at the leading edge of the airfoils was abut4.5° the ’leading+dge shockwave never became attached. Thus the flawfield about the model was transonic for the highest Machnuniber of thetest. For the Mach number range of t~ tests, the Reynolds nunibervariedfrom about 1.7 million to 2.2of+attack r-e extended from?nmibersIt was limited by the

millionas sk in figure 4. The angle-about -6° to +16° but at the higher Machnmiel strength to lower values.

CORKU!TIONS

A tare of 0.0022 was stitracted from the drag coefficient to takeinto account the effects of the end plate and the drag ~s corrected forthe interaction of the balance force and moment components. Blockageand tunnel-all interference effects-were assumed negligible since themodels were small with respect to the flow field. The indicated testMach numbers.were evaluated from the tunnel calibration for the bumpand represent an average of the Mach numbers over the region occupiedby the model. Typical Mach number contoms are illustrated infigure 5. The data have been corrected for flow inclination which wasfound to exist over the bump. The-effects of flow curvature were not - ~investigated.

Effects

FUZSULTSAND DISCUSSION

of a Systematic Variation of theTrailing=dge Thiclmess

.The lift, drag, and pitching+noment characteristics of the air-

foils are presented in figures 6, 7, 8, and 9. The data are presentedfor the models with and without surface roug%ness. For the Reynolds.

6 .-. N.ACARM A~lJll -..

number ramge of these tests it is believed that laminar flow was ~in-tained over a large part of the smooth wing surfaces. When surface .

roughness was employed, it is probable that most of the boundary layerwas turbulent. The force amd.moment data are summarized in figures 10,U., and 12 for the smooth models. Base-pressure data are presented infigures 13, 14, and 15.

Lift.-An examination of figure 10 shows that for a trailing+d.getbicb~s= of O the lift+urve slope increasec”gradually with increasingMach number and then decreased as supersonic Mach nuniberswere attained. .-

Tbe effect on the lift-curve slope of increasing the trailing-edge thick-ness was small at low subsonic Wch nunibersbut as the Mach number wasincreased, -thelift+urve slopes of the blunt=trailing+dge airfoilsincreased more rapidly than for the c$rcular+rc airfoil. For trailing-edge thicknesses of 0.3 and 0.6, maximum slopes were reached at 0.94 —.Mach n~ber, but with 1.0 trailing-edge thicl.messa maximum was attainedat about 0.98 Mach nuuiber. At high subsonic Mach nunibers,relativelylarge lift-curve slopes have been attained by 10-percent-thick circular-arc airfoils kting” blunt trailing edges (reference2) and by aK&percent-thick wedge (reference 4). The results of reference 4 alsoshow that at high subsonic Mach numbers the lift-curve sloye of a con-ventional airfoil decreased; whereas tlat of the wedge airfoil of thesame thickness continued to increase. In the present case, a comparisonof the Mft-curve s~opes of an NACA 63A.004airfoil of the same phn form .

and aspect ratio (reference ~) with those of the 0.3 arid0,6 blunt-trailin~dge airfoils of circular-arc origin shows similar values oflift-curve slope in the transonfc Mach number range. Thus it appears w

that while increases in lift-curve slope can be expected by increasingthe trailing-edge thickness of circular+rc airfoils, the markedimprovement in lift+urve slope of blunt-trailing-edge airfoils over

—

conventional airfoils indicated by references 2 and 4 would not occurin the transonic range for airfoil thicknesses of the order of 4 percent.The effect of increasing the trailing-edge thickness of conventionalairfoils was not considered in the present Investigation but the resultsof reference 6 show tkt for a 10-percentAhick conventional airfoil,increasing the trailing-edge thickness increases the li~ve slope.

,

Althougha sufficiently high angle of attack was not reached at.=

the higher Mach numbers to show the effect of trailing+dge thicknesson maximum lift coefficient, the data do indicate progressively higher —

maximum lift coefficients as the trailingAge thictiess was increasedat 0.60,Mach number. The effect of surface rouglmess on the lift char-

.—

acteristics appeared”practically negligible except for slightly lowermaximum lift coefficients for the various airfoils at 0.60 Mach number, A—

w

1

. (

NACA RM A51Jll — 7

PitChin&!111Oment.- At a lift coefficient of 0.1, increasing thetrailing+dge kness appeared to cause a slight rearward movement ofthe aerodynamic c~nter but the variation of aerodynamic center withMach nuniberwas similar for all trailing+dge thicknesses (fig. 10).The abrupt rearward movement of the aerodynamic center as superscmicMach numbers were approached was of the order of 10 to li?yercent ofthe mem aerodynamic chord.for all trailing+dge thiclmesses. Ingeneral, the effect of surface roughness on the pitch$ng+noment charac–teristics was of minor importance for all trailing+dge thiclmessesexcept O (circulaRrc section). For this exception, surface roug+nessappeared to reduce the nonlinearities for pitching+uoment coefficientthat occurred near zero lift at s~sonic Mach nuribers. Thus it seemsthat surface roughness and, therefore, the:type of boundary-layer flowaltered the pressure distribution in such a way as to have little effecton the lift but a noticeable effect on the pitching moment. It iSknown from reference 7 that for a slightly thicker circular+rc airfoilat zero angle of attack the t~ of boundary layer had a significanteffect on the local Mach nuniberdistribution at Mach nunibersabove thecritical, but had an unimportant effect for ~ch nunibersless than thecritical. Similarly, the results of the present investigation showthat the effect of surface roughness on the pitching moment was largestat high subsonic Mach nunibers.

* 222%3”-The variation of drag coefficient with lkch nunibershownin figure 11 illustrates the relative changes in drag that can beattributed to increasing the trailing-edge thickness. The data indicate

. that the minimum drag (at zero lift) of the 0.3 blunt-trailing-dgeairfoil was about the same as that of the circular+rc airfoil. However,at lift coefficients of 0.2 and 0.4 lower drags were measured with the0.3 blun-trailiqdge airfoil. Similarly, increasing the tiailing-edge thiclmess reduced the drag rise due to lift as compared with theshar@zailing+edge airfoil.

At lift coefficients from O to 0.4, the transonic increase indrag coefficient began at or above a Mach nuniberof about 0.90 for alltrailing+dge thicknesses. The minimum drag coefficient (at zero lift)at slightly suprsonic Mach nuniberswas of the order of three timesthe low-peed drag coefficient for the circular+rc airfoil and two andone+alf times-the low-speed drag coefficient for a trailing+dgethickness of 0.60. h e-nation of the drag data (figs. 6(b), 7(b),8(b), and 9(b)) shows that surface roughmss increased the drag forall trailing+dge thicknesses.

Lif&drag ratio.- The results shown in figure 12 indicate higher. lif-g =tios at mibsonic Mach nunbers with a t=iling-edge thick-

ness of 0.3 thm with a thickness of O. hther increase in tmiling-

●edge thickness reduced the lift.&ag ratio except at lift coefficientsabove about 0.65 at 0.60 l&ch nuniber. At supersmic Kch nunibers,the

8“ NACA RM A5LJ11

data indicate approximately the same values of liftArag ratio fortrailing-edge thicknesses of O an.@0.30 and slightly lower values fortrailing-edge thicknesses of 0060 md 1.00. The variation of maximumlift-drag ratio witliMach nuniberis shown in figure 10.

Base pressures.- Base pressures indicate trends which are reflectedin the drag data The increase noted in mini~ drag with increasingtrailing+dge t~ckness and with increasing Mach nuuiberparallels thetrend of decreasing pressure coefficient on the base of the blunt trail-ing edge as shown in figures 13 and 140

At subcritical speeds, the pressure coefficients on the bases ofthe bluntArailing-edge airfo~ls were considerably lower than the pres-sures which would norml.ly occur on the rear portions of a shar~trailing-dge airfoil; hence, the pressure drags of the blunt airfoilswould be expected to be higher than the pressure drag of the circular-arc airfoil. .

As the fre=tream Mach nzmiberand also the Reynolds nuuiberwasincreased through the transonic rmge, a decrease in base=pressurecoefficient occurred at speeds approximately corresponding to the drag-divergence Mach nwiber. This trend is shown in figure 14 for the inner-most base=pressure measuring station. This decrease prohbly resultswhen a supersonic e~nsion occurs around the sharp corner of ths %lunttrailing edge. The magnitude of this expansion is detemnined by theshape of the wake and Is sufficient to result in a low pressure whichwas about @ perc”mt of the free-stream static pressure for the blunttrailing edge with a thiclmess ratio of 1.0. With increasing Machnuniber,the lwse-pressure coefficients increased and the base pressurewas approxi~tely 40 percent of the free+ tream static pressure up tothe highest speeds of the test for this airfoil.

A spanwise gradient of base pressure was found to exist asindicated in figure 15. This gradient, with increasing pressure fromroot to tip, could be partially due to the velocity gradient over thebump normal to the bump surface; however, this variation was not aslarge as the s~nwise gradient of base pressure. . —

The variation of base-pressure coefficient with angle of attack forthe smooth airfoils with various.trailing-edge thiclmesses and at thedifferent spanwise stations is shown in,figure 13(a). At low subsonicspeeds the minimum base-pressure coefficient occurred at zero angle ofattack. At transonic speeds the trendsbase~ressure coefficient with changing

The effect of surface roughness oncomparison of figures is(a) and 13(bj.face roughness caused the base-pressure

indicated essentially a constantangl.e.ofattack.

the base pressme is shown by aIn general, the addition of sur-trends to be more consistent.

C-ring the rough condition with the smooth condition indicates that ‘. roughness increased the base yressure and delayed the negative peak of

base–pressure coefficients at transonic speeds to a higher Mach nmiber.These trends are shown in figure 14. The minimum drag was higher forthe airfoils with surface roughness, indicating that any decrease ofbase dragby the additicm of surface roughness was more tti canceledby increased friction &ag.

Investi@tion of a Boattailed.Trailing Edge

Figure 16 presents the lift, -g, and pitching+xmen t character-istics of a @attailed airfoil formed by chamfering the trailing edge ofthe airfoil hati~ 1.0 trailin~dge thickness. (See fig. 3.) Theresults without surface roughness are ~rized in figures 17 and 18and are co-red with those of the airfoil having 1.0 trailing~dgethickness.

It is ap~rent from figure 17 tlwt boattailing the airfoilincreased the Mftiurve slope in the tremsonic Mach nwiber range. Fora lift coefficient of 0.1, the change in the aerodynamic center of the ~boattailed airfoil in the transonic l&ch ntier range was about the same

. as t~t for the airfoil having 1.0 trailing+dge thickness but the vari-ation with Mach number was less abrupt as supersonic Mach nuibers wereapproached. A comparison of figure 16 with figure 9 indicates $hat beat+

.tailing the airfoil nmde the variation of pitching+oment coefficientwith lift coefficient more nonlinear for Mach nunibersfrom shout 0.70 toO.go. However, adding surface roughness to the boattailed airfoil made -.the pitchingaoment curves more linear for these Mach nunibers.

Figure 18 shuws that bcatbiling the airfoil reduced the tinimumdrag coefficient significantly but at lift coefficients of 0.2 amd 0.4a smaller reduction in drag coefficient was realized. Thus, onlyslightly higher maximum lift-%~ag ratios were obtained for the @at-tailed airfoil as shown in figure 17.

wakewill

Wake Fluctuation Characteristics ‘

Measurements were made of the total~essure fluctuations in theof the test models since it is possible tkt these measurementsgive an indication of the buffeting characteristics of the blunt-

trailing+dge airfoils. Unpublished da~ obtained by this methd yielda“buffet boundary which agrees well with the buffet boundary obtainedby the accepted methd using an accelerometer at the airplane center ofgravity.

—.. -. —--”

10 NACA RM A51Jll

The wake fluctuations are presented in figure 19 which shows theHimum total=pressure fluctuation ditided by free-stream dynamic pres-

()‘we *as a function of Mch nw?iberfor the various trailing-edge

thicknesses. The maxim total-pressure fluctuation shown in this fig-ure is the mximum picked up by the three probes between the angles ofattack of -1° and +6°. This methcd of presentation is used since thepressure fluctuations were random and showed no consistent variationwith angle of attack. Limitations of this method are: (1) the pickup ‘probes were stationary and were not necessarily in the position to pickup the maximum wake total-pressure fluctuation; and (2) insufficientdata were obtained to make a statistical analysis o? the variation ofwake fluctuation with angle of attack.

At hw S~dS (M = 0.60)the wake total-pressure fluctuationsincreased markedly with increasing trailing-edge thickness. The dis-turbeaces set w in the wake were probably sufficiently strong to causean ~teady circulation to be eskblished around the airfoil, resultingin unsteady forces and, hence, some buffeting.

lnvesti~tion of figure 19 shaws that the wake fluctuationsbehindthe airfoils having O and 0.3 trailing-edge thiclmesses were relativelysmll at luw speeds, increased gradually to a peak of 0.25q at a Mechnu?iberofl0.90, and then decreased slightly. The wake fluctuationsbehind the airfoil with 0.6 trailing~dge thickness renmined.practicallyconstant up to 0.90 Mach nuniberand then increased sharply. The airfoilhavinga trailing~dge thickness of 1.0 hd a wake fluctuation of 0.69qat 0.60 Mach ?nmiber,decreased to a minimum of 0.31q at 0.92 Machnumber, and then increased sharply again as the speed was furtherincreased.

—

In general, t“heblunt=trailing-edge airfoils had larger wake fluc-tuations tkn the basic circular-e.rcairfoil and it also appears fromthis test (see fig. 20) that increasing the trailing-edge thictiessabove O.~0 of the maximum thickness results in relatively large wake

—

totil=pressure fluctuations even at low speeds and hence increases thepossibility of buffeting.

The effectof boattailing the blunt tmiling edge was to decreasethe amplitude of the wake total=pressure fluctuation as compared withthe amplitude behind the 1.0 blunt-trailing-edge airfoil at most speedsand also indicated lower amplitudes at Mach nunbers greater than 0.90as compared with the 0.6 blunt-trailing+dge airfoil. Boattailingappears to offer a practical means of reducing the total-pressure fluc-tuations in the wake and hence the possibility of buffeting of blunt-trailing-edge airfoils.

.

L

..—

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=-

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. .

.

.

-+

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. .

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—

..—

.

b -- --- - .

NACA RM A51Jll.

11

CONCLUDING KWARICS.

T&nsonic wind-tunnel tests of’a series of k-percent-thick biconvexairfoils with var~ng amounts of trailing-edge thickness showed anincrease of the lift-drag ratios with no measurable increase in minimumdrag coefficient for a trailing+dge thickness of 0.3 of the maximumairfoil thiclmess. Higher ”lift-curveslopes were obse~ed In the t~sotic Mach number range for all the blunt-trailing-edge airfoils ascompared to the circula~c aitioil and higher msximum lift coeffi-cients were noted at O.6 Mach number. Surface roughness had a signifi-cant effect on the pitching+noment characteristics of the circular-arcand Imattailed airfofls, particularly at high subsonic hhch ntiers.Increasing base pressure from rmt to tip was observed for all theblunt=trailing-edge airfoils, with a progressive decrease in basepressures as the traillng-dge thickness was increased. Surveys inwake of the airfoils indicated approximateely the same level of wakefluctuations for trailing-dge thiclmesses of O and 0.3 but markedincreases in wake fluctuations for trailing-edge thicknesses of 0.6

‘the

1.0.

Ames Aeronautical I.aboratory,Mtional Advisory Committee for Aeronautics,

Moffett Field, Calif.

JWZERENCES

1. Chapman, Dean R. : Airfoil l?rofilesf’orMinimum I?ressureDrag atSupersonic Velocities - Genemal Analysis With Application toLinearized Supersonic Flow. NACA TN 2264, 1~1.

2. Summers, James L., and Page, William A.: Lift and Moment Chamcter-istics at fhibsonicMach Numbers of Four l&Percent#I!hickAirfoilSections of Var@ng Trailing~dge Thickness. N/MA RMA50J09, 1~0.

3. Axelson, John A., and Taylor, Robert A.: Preliminary Investig3.tionof the Transonic Characteristics of an NACA Ehihnergedklet.NACA RMA50C13, 1~.

4. Eggers, A. J., Jr.: Aerodynamic Characteristics at Subcritical andSupercritical Mach Numbers of Two Airfoil Sections fifing Sharp.Leading Edges and Extreme Rearward Positions of Maximum Thick-ness. NACARMA7C1O, 1947.

u

12 GONF131ENTIALq—— N/MARM A51Jll

5. Nelson,Warren H., sxd McDevitt, John B. : The Transonic Character-istics of 17 Rectangular, Symmetrical Wing Models of Var@ug

● Aspect Ratio and Thickness. N/WA HM A51A1.2,1951.

6. Hemenover, Albert D., and Graham, Donald J.: Influence of AirfoilTrailing~dge Angle and Trailing~dg~bickness Variation on theEffectiveness of a Plain Flap at High @bsonic Mach Nu?ibers.NACA RM A51C12a, 1951.

7. Li.e??mm,IkneWolfgang,Askenas, Harry, and Cole, Julian D.:Experiments in Tramonic Flow. U.S. Air Force Technical Report

-.

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.

No; 5667, I?eb. 9, 1948.—

●

.

.-

.

.- —

NACA RM A~Ull.

—_J, : —

->

mNmEmpL4flF-

13

Figure l.- Typical model mounted on the transonic bump.

.

14 .

. -1-, Fence

NACA RM A51Jll.

Figure 2.-Dimensionsof theredungukv

wings.

““, -

.5

.

.

NACA RM A51%U -a 15.

.

.

“,

Leuding

edge f, 0.04cforoU

y—’ h, 0.3 tI ,

#

po.5oc-4 ‘=Bodtoiled airfoil

~-4cJF”

Figure 3.- Airfoil profiles.

0.6~

/.ot

0.6~

@mm??’iz?’i-”’

2.2#

2./

.ffl

/.8

/.7

/.6

.5 .6 .7 .8 .9 /.0 /./ 1.2

Mach number, M“v

Figure 4.- Variation of Reynold% number with Mach numbe~.,

1“”:’ I

i,b

NACARM A51%U.

20.

/0

.

/0

o

17

.

70 80 90 /00 //0 /20Bump station, inch es

=29=Figure 5.- Typical Mach number contours of the Ames

16-foot wino’-tupn el fransonic bump.

. .——.+!!oti~ :

—

.

PCn

1.0

.8

.6

-.4

-.6

I A I d II I

-4 0 4 8 12 16 for M, 0.60

Angle of attack, a, deg

{a) c’ b% u.

Figure 6. –Aerodynamic characteristics at various MOCh numbers. CkUfO$-OrC Pro fife, # O.

#

1.0

.8

.6

-.4

-,6

I t

I 1 I

I 1 -’1 Iw=i I hl I I 1. I I I I

I I

o .04 .08 ,12 .16 .20 .24 for M, 0.60

Drag coefficient. co

(b) CL VS CD

Figure 6,- Continued

1 ,“

\ I

It)o

1.0

.8

.6

so4

.2

-.4

-.6

t

-Ho Without surface roughness

A With surface roughnessI I

—. tll El

.04 0 .04 -.08 for M, 0.60

Pitching-moment coefficient,

(c) GLVS Cm

Figure 6.– Concluded.

,* !,, ,, f: 1

c.

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i

c!i0’f

El1;

!!’j

i.

1.0

.6

-.2

-.4

-.6

-4 0 4 8 12 16 for M, 0.60

Angie of attack, a, dog

(0) CL w a

Figure Z – Aerodynamic characteristics at various Mach numbers. Biunt -trai/ing-edge profiie, ~ n 0.30.

1.0

.&3

.6

so-1

-.2

-,4

-.6

u ,.04 .08 .12 .16 .20 .24 for U, 0.60

Drag coefficient, Co E

s

(b) C’ M CD

E

Figure Z-Continued.~

E:,

. . * .

. , * , t

1.0

.8

.6

Pitching-moment coefficient CM

(c) CL w cm

Figure Z -Conciuded,

,

,,

I

1.0

.8

.6

so-1

-.2

-.4

-.6

t

v w ./ / / I d b

#+

r —

!/} /– -2 L 7f---t+

d $“I

A I I I I I Iw I I I i I I I I I I I J-4 0 4 8- 12 16 20 for M, 0.60

Angle of ottack, a, deg

(0) CL w Q

Figure 8.- Aerodynamic characteristics at various Mach numbers.

# .

Blunt- iroiiing-edge pro fiie, ~S 0.60;

.

.

. , 6,

1.0

.8

.6

Qj .4

-.2

-.4

0 .04 .08 .f2 .16 .20 .24 .28 for M, 0,60

Drag Coefficient CD

[b) c’ W co

Figure 8.-Continued.

,,, ,,

..

-.08 -.12 -.16 for A4,0.60

Pitching-moment coeff/cien~ C~

(C} C’ VScm

Figure 8.-Concluded.

.:,,. .,,,, ,,-r.

!:

,.. .

J

1.0

.8

.6

Q“

-.2

-,4

,

-4’04 8 12 16 for M, 0.60

Angle ofatfack, a, deg

(a) CL vs u

Figure 9.- Aerodynamic characteristics of various Mach numbers. Blunt-trailing-edge profile, ~S 1.00.

3

n)(n

1.0

.8

,1,

1“

‘Ii

.6

-.2

-.4

0 Without surface roughness.,

a With surface roughness

~.6~o 4I

II I I I I I I I I lb I

, . ,

1.0

.8

.6

u“

-.2

w v M i I IP

.4 I I I I I I I 7 I AI I I 1/ I

r=5=–

.04 0 ,-.04 :08 -. ;2 for M, 0.60

Pitching-moment coefficient, Cm

(c) CL Vs cmFigure 9.-Conciuded.

I

i.,.

30

1

./2

.08

.04

0

60

40

20

0

30

.-. . T.:,=.:: -:.. -NACARM A5LJ11

t

———— :.3—.—-------- ?$ C&, 0./

,,-

i -(

r-- ---.-1 I I I I t

=s9=

.5 .6 .7 .8 .9 /.0 /./ /.2

.

.,

.

,,, , ,–—

Mach number, M

Figure /0.-Voriation of aerodynamic parumefers with Mach numberfor blunt - trui/ing - edge airfoils.

.-. .

.

.

NACA RM A~lJll 31

.08

.04

0

p

Q

.08

.04

0

Al;

—“—— 0.3—.— 0.6-------- /.0 --- ..’” ‘- ‘--

------ ------ --I I —

/ MA ~! ..I I I 1- 1 1

I

-H[G,, O

/-- --- --- --

●

1

.5 .6 .7 .8 .9

Much numbe~ M

Figure /l-Drug coefficient os u function of

/.2

Mach number&

‘ous /if t coefficients.

I’TACARM A51Jll ..

I

/ / ‘~ -

h‘/ “ ~//,-11 \.b‘\\

I

IH+RH-Hh7

———— :.3—.—-------- n

\

o .2 .4 .6 .8

L“ ‘- -

o .2 .4 .6 .8

~if t cot?f ficient, c’

Figure /2. -L if t - drag rutio os a function of Iif t coef ficlent at vc?rious

Mach numbers for different trailing -edge thicknesses.

.—‘coI’?mEmqJ

.-

...

,..

i“

.

--

.-. . .

-—

.,-

--—_

.,--,

.

,

-Lo

-.8 +&&

-.6

-.4 }

-.2 II I 1111 I -.

G -f.o

-.8

-.6

-.4

-.2

n

,100

(260

‘4-20246-4-2024 6-4-20246

A@e of attack, a, deg

+, t230

{0) w-SW+?ci9nW@WWi

fl@we 1.3-B&w pfwsum cwfftifs as a &wttiofa@e ofaftack.

+/-20246

L4Jw

.-

W4=-LO

-.8

-.6

-.4

-.2

Q.. 0

1,1+ 0.60

-.8

-.6

-.4

-.2

n

“-4-20246 4-20246 -4-20246-4-202 46

zlngle of attack, a, deg

+, o#o

(b] Wifh surface roughness.

Figure 1.3-Gamluded1

,4 I ,,

I

NACA RM A51Jll 35

.

.

-/.2

-.8

-. 4

0

-/.2

-.8

-.4

0

Withoutsurfoce roughness‘---- –-- With surfuce roughness

1 I 1 1 ,

14$ 0.6 1

.5 .6 .7 .8 .9 /.0 L/ 1.2

Moth numbe~ M v

Figure 14.-7,,e Vorjution of Lwse pressure coefficient with

A40eh number of semispon station ~~ 0.25. Angle .of Oftock, o ?

--

.

.Amnmtma-—-NACA RM A51J1136

-.’8

-.4

0

-/.2

-.8

-.4

0

-/.2

IM + 0.6

‘0:98- >- .

‘/.06- T -- ‘- --: -..I

_ .80- >.60- 4

-- ,4 ~ ~-

I $J0.3

-.8

-.4

00

‘“/.06-7 -I

\ . N .

—’.60=

I

.2 .4 .6 .8 /.0

Sen7ispt7n station, ~==z$=

Figure /5.–The variation of buse pressu~e

coefficient with semispun station, 7“Angle of ottock, 0?

.

.

..

.. -----

,---- <,. .—

—

. . ,*, ‘, ___

..—

.-

“‘--..—

.9?.x

,.. i+

-..

.,:“..

..

Wmmmmm7___.... -

, ? .

1.0

.8

.6

-.4

-.6

1 r 1 r 1 r 1 r

I [ I I I I I I1 / I/1 1

I

‘4 O 4 “8 12 16 20 for k?, 0.60

Ang/8ofaltack, a, deg

(a)c’ vs @

Figure 16,-Aerodynamic characteristics ot various Mach numbers.

.

Boot tail trailing edge..

%’

1

,,

I

1.0

.8

.6

so4

-.2

-.4

-.6

t ,’ ~ o Without surface roughness- p

1

A With surface roughness/

-n /,@ bJ ~ -n. />( I I I I I I I

t$k-t--lI (M/ I J I 1/ I I

#11 I

—b .s.

r I

o .04 .08 .12 .16 .20 .24 for bf, O.60

Dreg coefficient, Co ~

(b)CL K coii

Figure 16.- Continued.~

P

P

\

, # ,

Pitching-moment co8fficien/, Cm

(cjCL vs Cm

Figure 16.-Conc/uded.

G

40

./2

o

60

40

20

0

30

o

coNFlmmIii-i- NACA RM A51Jll

. .~. -- ~

/’.@*/~” / “

& ---~“

.-—.

---- - ---

.-CL, 0.1

.. - . . i

Without boat ?ail troiling edge..

------- Wit~” boottail tyailing edge

-- -—.#

. /-- .- _- --- .

6“, 0,/

,

.5 .6 .7 .8

Much numbe~ M

.9 1.0 /./ /.2

Figure Iii- Variution of oerodynomic parameters with Moth

for blunt und boattoifed trailing edges.

~~_~ENTI& ;

—...-—-----.

.

.-

NACARM A51Jll

.

.08

.04

0

.08

.04

0

.08

.

.04

0

Figure

.

-.41

1“——

IWithouf bo attai/ troiling edge

------- wj~~ &oat tai/ trai/ing edget--H

--. --- ---- ---GL, o

I

~-- --- __

.5 .6 .7 .8 .9 /.0 L/ ‘ /.2

Mach number, M v

/8. -Voriufion of drug coefficient with Mach numberfor blunt untf bot?ttuilea’ troiling edges.

—.—

-—

42—.

c@gIn@@iii4 NACARM A51J11 _*

.60 .70 .80 .90 /.00 /./0Moth number, M

/9.-Effect of Mc7chpressure fluctuation at

.

. .,

.

,=

. “+

.,,.-

. ..

number on the wake muximum totol-vurious trulling edge thicknesses.

-—— .-

--’+”

—..

6C

.

.

NACA RM A51JU 43

M.60

------ .80—— .90—-—.98 /i—--—/.06

fl - ~. /‘*

P

./. ,~ / ~

/’// ‘

//

t J‘I ,’

// ,// ~ ~i

-.fl~ - “ //--- v-

/

~

- =%=- ‘

0 .2 .4 .6 .8

T.ruiling edge thickness ~

Wing thickness ‘7

Figure 20.-Effect of troiling edge thickness on themuximum woke total-pressure fluctuutt’on at severuiMuch numbers.

=CONTIDERTIAL>~— =

NACA-Lan@y-12-11-S1 -325