C.P. No. 1261

MINISTRY OF DEFENCE (PROCUREMENT EXECUTIVE)

AERONAUTICAL RFSEARCH COUNCll

CURRENT ‘PAPERS

NPL 9615 and NACA 0012

A Comparison of Aerodynamic Data

BY . ’ N. Gregory and P. G. Wilby

Aerodynamics Division, NPL

LONDON: HER MAJESTY’S STATIONERY OFFICE

1973

PRICE 8Op NET

CP No. 126;~. Rowember, 7968

NPL 9615 AND rucA co12 - A COMPARISON OF AERODYNAMIC Wl!A - by -

N. Gregory aad P. G. Wllby

Aerodynamics Division, NPL

SDMIURY

Ordinates, surface slopes and curvatures are listed for the two eerofoils together with a detailed tabulation of lift, drag and pitching moment data obtained at Mach numbers between 0.3 and 0.85 ln the NPL 36 IL x 14 ln. trensonx tunnel. The aerodynamic characteristics and all the pressure distribution are plotted, with some comparisons.

1. Introduction

NACA 0012 is a standard section frequently used for helicopter rotors and NPL 9615 is a derivative of it having a 6.20 extension to the chord and a drooped leading edge with larger radius of curvature. The position of NPL 9615 in the current programme of aerofoil section development for helicopter use will be reported on separately, but as this section has been found to possess advantages over NACA 0012, its description and measured aerodynamic characteristics are given herewith and some comparisons are drawn.

2. Section Shapes

in 1932 , The NACA four-digit series of wing sections was first reported on

, and the following formulae for the thickness distribution and leading edge radius are taken from Ref. 2.

f y/c = t/c- 0.29690 v’x/c - 0.12600 x/c - 0.35160 (x/c)’ 0.20 (

+ 0.28430 wc)3 - 0.10150 wc)4 >

R,/" = 1.1019 it/c12

NACA 0012 is obtained by putting t/c = 0.12 in the above equations, and the ordinates are listed in 'IYLBLR .l together with surface slopes and curvatures.

The new section, NPL 9615, was obtained by taking the rear portion NACA 0012 and modifying ordinates forward of the position of maximum thickness, extending and drooping the nose. The ordinates for NPL gb15 are non-d~mens~onal~seci with respect to the extended chord and are listed in 'l.kBLR 2, also with surface

slopes/

*Replaces A.R.C.30 657 - NPL Aero Special Report 017.

slopes and ciirv3tures. The nose portions of the serofo1ls are com:jnrcd. in Fig. : and the upper surf2ce ",wvature aistr5ut1r;ns 3hC'd cf ~osltiorl of liV?x1mum thickness ars compared in Fig. 2. The tbickness/chorl ratio of NPL 9615 is 0.413 commpsrei ~th. 0.12 of INCA 0012, and the eev: profile blends smoothly into the NLCh 03 2 rem i;ortion st ';/c (b ased 0.1 the neb+ chord end nose posltioa as orq,in ) aqxl to 0.28333 for the nppcr surfzoe end eq~,cl to 0.3$09, t% poc,ition for ma~murn thickness, on tho lover.

'J&e leading-edge radius of N?L 9615 is O.Ol88j, non-d~mansiocclised in term of the actual cho?d. Th3s should be compared >:ith O.@Y? for the standard. NACA OOi2, (a figure ,rhich is reduced to O.r)M for a NACA four-digit thickness distribution with the II.% thickness/chord ratio of NPL 9615).

Cne of the &ms of the design features of NPL 7615 W.SS to reduce the curvature of the upper surface, relative to that for NACA 0012, in the region rihere supersonic floz develops at incidenoe. This helps to reduce the maximum velocities and hence the strength of the shock wave that tarmlnates the supersonic region. The reduction of curvature was made possible by the extended leahng edge which also allowed the ixorpcxation of leading-edge droop and 311 Increase of leading-edge radius, both of mhich help to insre%se ths lc;: qeed CL,,, . Thhe droop and increased radius also have an effect on t;le development CI? the 10~31 supsrsonic firm.

3. Test Conditions

The aero&;namic data for the tiiro section- ner: obtcined under i,&nticel xnditions SO that the coupar?son should no? be i:?flucrced by these conditions. The tests were carried out ~1t.h IO in. (0.254 m.) chord model3 spacnlng tC I& In. (0.356 m.) dimension of the NF'L 36 in. x 14 in. (0.72 n. x 0.36 m.) transonic wind-tunnel, which operates at atmospheric stngnsttlqn pressare. The floor and ceiling of tie tunnel were slotted (& slots, overell open-area rztlo = 0.33) and were 31 in. (0.79 m.) apart throughout the length of the xrorkxnG section. These conditions s-e close to those giving blockage-free anti loft-lcterference-free resulC,s, and no corrections for wall constraint have been ..pplie:.

tither tests are now in hand to determine optimum test conditions 2nd to calibratz fhe tunnel precisely. Lift iderf'erence 2nd blockage ar? affected by both wall divergence and open-area ratio and it is not clear whet!xr a rial.l configuration can be found to give zero values 01 n lift interference and blockage simultaneously, or Thsther such a mall vould remain entirely interference-free at the high values of CL and Mach numbtir which nzcZ to be covered in the present tests.

For the present, it should suffice that any nsglected cor,straint correction; should be small, and may be equivalent to a small change in frsa-stream Mach number. Furthermore, the oq~r,rison hatvrcen the 5.~0 sections sho!GS not be influenced by mhether the idsntxal test conditions art; entirely interfarenoc-fret. In one respect they are not: the YllCk en&--:~sll 3OUrldq~ layer.5 lead to a SerLotis departure from two-dimensional testing conditions at high angles of incidence and hence to a reduction in CL -

max

Companion tests hsve been carFled out in the 13 ft x 9 ft loti speed wind tunnel on a model of NAG OOl2. llthough the end effects 3re much further removed from m-La-span wzth a planform aspect ratio of 3.6 compared sith 1.4 for the 36 m. x 14 in. iflnd tunnel, the tests revealed e gain in CL ImY of over 0.1 when prematurs flow sspnrstion at the ends yrss inhibited by- boundary-layer control by suction.

-3-

All results wero obtained v~lth a roughness band of 730-270 mesh carborundus? presentbet-ween the leading edge and 0.02 chord on both surfaces. Sufficient roughness was required to prcxiucc b:Jndary-layer transition ahead of strong shocks in order to avoid optimistic values for CL max at high Mach number. On the other hand, too much roughness Was likely tc prcduce 10,~ values of CL ma:: at low speed and c hi& d-zerall level cf drag. T!le ban: that iiab chosen ~rO~lded

a compromise roughess that could be used over the whole range of the tests and gave a reasonable simulation of the conditions on a full-scale helicopter blade. However, the most important point to note is tha t the same roughness band was used on both models so as to ensure a valid comparison of one set of results with the other. DLrect-shadow photographs revealed +&t with the band present, transition oc-cd between 0.10 and 0.40 chord ;o"mstream of the band., depending on Mach number and pressure gradient. Without the band, trcnsition mould have occurred in an ursealistic position much further aft.

44 Results

Lift and pitching moment =ere found by integration from the distribution of pressure round the centre portion of the aarofoil measured at about 43 static pressure hole stations in the surface of the model. Profile drag was ohtaiqed by make traverse.

The measurements were taken in order to construct tables of aerodynsmic characteristics at 0.05 intervals in Mach number and $' intervals in incidence requxed as input for the comparative machine computation of rotor performance and it was therefore necessary to double-smooth ths expermental observetions. This has therefore J.reedy been done and the tables here presented contain the smoothed values, limited to the (M, a) regions actually covered by the tests, and obtained by Interpolation where necessary: Only a few of the $d settings wers actually tested. The graphs on the other hand make use of vsZ.ues measured at 0.025 intervals in Idach number in regions where the values are varying rapidly. TABLIZ 3 lists values of CL, CD and Cm

C/I 1 for WL 9615 at Mach numbers betneen

0.3 and 0.85 with incidences between -2' and the stall, and TA%?? 4 contains corresponding information at positive values of inridence for the symmetrical section NACA 0012. The test Reynolds number varied fran 1.7 x 10s at M = 0.3 to 3.75 x 10s at M = 0.85. Curves shoving the variation at CL, CD, md %I& with incidence are plotted in Figs. 3, l+ and 5 forNPL 9615, and m Figs. 6, 7 and 8 for NACA 0012.

The pressure distributions for NPL 9615 are plotted for each incidence in Figs. 9 a - q. Up to 3O incidence the uppe- * and lower surface distributions are plotted separately; above this incidence the 'cue distributions are combined on the same diagram. Correspcnding pressure dlst?zhutions for NACA @Cl2 are shcsvn in Figs. 10 a - cl, for positive mciderces up to 12'.

5. Some Comparisons

The effect of the profile chanss refer-ed to inSection 2 on the aerodynamic characteristics of the 'GJO scctiors is summarised in Fig. 11. This figure shws en improvement in the values of CL max over the vfhole Mech number

range of the tests and also gives boundaries in the (M, CL) plane for the onset of a rapid change in pitching moment cni; for the onset of a rapid rice in drag. The latter boundaries are not easily defined everyWhere, and are only approximate in their location. In particular, at incidences belo>; I', NPL 9615 exhibits a pronounced drag creep preceding the more rapid rise, as can be seen at zero lift in the drag comparison of Fig. 12. This figure serves to emphasise an&her limi'ation of Fig. 11. This is that the values of CD, cm and OL differ ____________________------- - -____________

*2X-2,0 mesh carbmund~m lm~lies grains that were sieved through a %?u%e vrlth 231 wlms to the linear Inch, but which were retained by P ea"ee with 270 wires to the Inch. This ,m,,lles 'inIlls that passed throu&!h P 59"pTe Ppen~ rith side 0.0~7 IL (0.062 mm) but not through one with side 0.0023 In. (O-053 m).

-4-

in general between the twosectloiis ?t any point on thet figure, ad also ilong the boundaries: the values mvst be 1lferr-d from the Tables, but are shown in sore more de%iled ?ompar~sons vhlch follow.

NPL 9615 has dightly larger zero-lift &a& than NJ&A Sl? et Mach numbers below dreg rise, (Fig. 12), though its crlticsl Mxh number for rrpid &rag rise is marginally grextw, an& beyowl t:us point Its ratr of increase of drag is apprecjably less. This latter featxre rriurs also at zL+ze;-o values of CL! hoth vith respect to increase of Mach number and also wrth r?spsct to increases of ST, at constant Mach number, a.s can be SW? from thP comparison of drag polars, Fig. 13.

Drag reductions are odzined in regions of high drag at e.11 Mach nunbers, with both sub-rritxal 2nd super-critical flow. The largest drag redxtionr are obtained et Mach nunbers between 0.55 and 0.65, a ran% that c0ver.s the tip Mach numbers of many hellcopter rotors, and at values of C L 1~ the regw,~ of CL maLY

fcr %4CA OcIl2, the saving in drag ten amount to es much as 3a. In w?er-cr3.ticsl flow, a comparison of t~~cal pressure dlstlzbutions 1s marip in Fig. 14 for & Mach number of 0.6 anI1 a CL wlue of 0.76. (The dlstributfon for WL 9615 was interpolated bet;veen C;lo obzexvatlons). It &llbe seen that the profxle chsn<:e has red.ucedtheveloc~ti?s in the svparsonic region as expected, ard thx has resulted j-n the desired re?.uction =n shock strength, and hence ill .oave drag.

The improvements In CL max and the reduction.? in the hqh-drag level together result in an xnprovement in the maxunum value of the llft/clrag ratio over the whole Mach number range of the comparison, Fig. 15.

The pitching-moment varlatlons w.th Mach n.mber are compared In Fig. 16 for angles of 3' and below. On account of its camber, NPL generally show a nose-down bias compared with that for the symmetrical section: this also shows up at zero-lift, Fig. 17. At high Mach numbers, the pltchlng moments on both sections change rapidly.

6. Conclusions

A slightly drooped extension wlth larger radll of curvature at the leadIng edge and on the upper surface has been fItted to NACA 0012. This reduces the maximum velocities in the supersonic region at high Mach numbers and also the strength of the shock wave that terminates it.

The modification has reduced the thxkness/chord ratlo from lz toll.J:, but a zero-lift pitching-moment coefflclent at M = 0.3 of -0.008 has arisen because of the droop (Fq. 17). The followlngbeneflts have been secured by the redesign.-

An mcrease in CL max of between 0.08 and 0.14 at Mach numbers below 0.65

(Fig. 15) despite the reduction in thickness/chord ratlo.

An increase of about 0.02 In the drag-rise Mach number (Figs. 15, 18).

A reduction of drag in the hlgh-drag region, particularly vhere supersonx flow is present. At values of CL m the region of CL max for NACA 0012, the saving in drag can amount to as much as 3& (Fig. 13).

The maxunum values of lift/drag ratio are increased at all Mach numbers (Fig.14).

Conclusions/

- 5-

Acknowledgement

The work herein reported was carried out as a team effort. Particular acknowledgement must be made of the parts played by Mr. V. G. Quincey in charge of the 36 in. x 14.in. wind tunnel and blr Miss E. M. Love In preparing the smoothed aata and the Figures.

&. Author(s)

1 Eaetman N. Jecobs X. E. iYard and 9. h!. Pinkertau

2 I. R. Abbott and A. E. van Doenhoff

References I

Title. etc.

The charac+eristxcs of 78 related sectlons from tests in tie varrable-density wind tunnel. Nxx Rpt. 460, 1932.

Theory of Cng sections, IdcGrerr H111, I 949.

-6- TABI. I

NACA 0012 CRDIXXFS

X (t

z e 0 P I

0 0 63,291 cl. 0005 0.00~0 SF.5 0.0010 0.0056 70 0.0025 0.0087 59.6 0.0050 0.0122 49.8 0.0075 c.Ol49 43.5 0.6100 O.Ol7O 39 21.062 0.0225 o.cl89 35.6

0.015 3.0206 0.02 0.0236 :i=: 0.03 0.0284 :3:,

10.893 6.976

0.04 0.0323 19.5 4.9E5 0.05 0.0355 16.g TJ.p15 0.06 0.0383 14.8 0.08 0.0430 II .8 0.10 0.0469 9.6 q.663 0.12 0.0499 7.9 0.14 0.0524 6.4 0.16 0.0544 5.2 0.18 0.0560 lc.2 0.20 0.0574 3.3 0.741 0.225 0.0586 2.3 0.644 0.25 0.0594 1.4 c.568 0.275 0.0599 0.7 0.505 0.3 0.0600 0 0.~52 0.325 0.0599 -0.6 0.407 0.35 0.0595 -1 .I 7 0.368 0.375 0.0588 -1.68 0.4 o.o:Eio -2.13 0.305 0.425 0.0563 -2.55 0.45 O.0558 -2.93 0.475 0.05ut -3.29 0.5 0.0529 -3.61 q.zq 8 0.55 0.0495 -4.19 0.6 0.0456 A70 0.165 0.65 O.04l3 -5.14 0.7 0.0366 -5.56 0.138

0.75 0.0715 0.8 0.0262 2 5'; -&

0.1 3-l C.85 0.0205 ::;, 0.0080 0.0145 -7.53 -7.10 0.142

I .oo o.ool3 -8.02 0.169

0 :: Surface slope, degrees

P = Surface curvature = C/';r

Leading edge radns: 3o/o = l/p" = O.ol5P

0

90 50 45 w 35 32.5

2: 26 24 22 20 19 18 17 16 15

I 53.1 0 s3.1 0.00443 53 0.00586 18 0.00857 I o.m359 9 o.q726

2; 0.021~2

6:2 0.02589 0*03065

5*6 0.03602 5.0 0.04209 4.4 0.04905 4.03 p.05297 3.76 3.49 3.22 2.95

0.05723 0.06~83 0.06682 z:,o7227 3.07823 X0&52 LO8495 2.08858 k09wk ~96% woo98 0.10574 3.1107G 3.1 I 622 0.12239 3.12928 3.13688 1 3,14520 3.15s.03

14 ' 2.68 13.5 2.545 j3 2.41 12.5 2,275 12 2.14 II l 5 2.005 II 1.87 10.5 A.735 IO

1 I ~33

9.5 I A79 9 t L3lY a.5 I .I90 8 ~085 7.5 0.996 7 0.922 6.5 0.857 6 023 5.5 0.752 5 4.5 x0: 7 4 ok37 3.5 3 2.5 I.82 0.637 0.66 0 I

-0.618 -:.68c -2.133 -2.942 -3.611

I:$;; -5L43 -5.557 -5.943

k~Wd+ O*-l97% OSQ995 0.22358 0.23703 0,25@94 0.~6450 0.28333 0.375 0.3409 1 0.3642 0.4115 2.4351 3.482~ 3.5292 I 2.5763 3.62% x6iO4 3.7176 3.7646

-0.0136~ -0.oa55 co. oooq

0.00268 0.00649 C. 00893 o.ot-r63 0.01392 0.07 633 O.olW o.oa4JJ 0, fJ21c.07 0. WY+5 0.02687 0.02832 0.02y80 0.0313l CL03306 0.03365 O.OYt7~ 0.03528 0.03612 0.03698 0*03?F36 0.03876 0.0396? 0.0~~065 0.04=165 0.04271 o.oh-3&1 0,@@+94 0.04609 0.04724 oA-838 0.04Y50 0.05059 0,05ak o.@2f;4 O.C5~f7 0.05424 O.wtY? 0.05565 0.0564 o.c%5 9.0564 0.0554 0.0546 0.0525 0.0498 0.04-66 0,043" 0.038? O-W+5 U'297

then as upper surface, with 8 + ve. ,Y - ve. C

L/C ; 0.113

,eading Hge rtiius: b/c = 1 fpo = q.01883, NIth centre at x/c = 0.~ 883, yf c = -q.r>137

+7f fle is circular for LO’ of arc on upper surface

I

I

1

I

f

P

, c

1

‘mf i le joins smoothly with NACX 0012 shape at x/c = 0.28333 on the unper surface

rnd at x/c = 0.3409 on the lower surface

1 TABId3 =

-6.320 -6.702 -7.098 -7.525

I;-;;; .

-90 -80.5 -71 -63.5 -51 -41.5 -32.5 -27 -20.5 -15.5 -1 I -9 - 8.5 -8 - 7.5 -7 - 6.5 -6 - 5.75 - 5.5 -5 - It. '5 -4 - 3.5 -3 - 2.5 -2 - 1.5 - 1 - 3.5

0

ml17 0.0247 0.8588 ma93 WQ59 o.ol36 O-P529 0.007c; 0.9765 wm5 1 ,oooo am3

LWER SURFACE

2.5 2.0 ~6 1 "15 &O 0.55 0.43 O&7

0.4l7

0 0.0003 0.0008 O*Ool9 wo33 0.0056 0,008~ o.ol30 o,o? 811. 0.023F; 0.02go 0.03535 0.03884 0.04330 0*049~5 ~.05900 0,07?32 0.09030 OJOO55 0~1oy8 0.13182 0.~5269 OJ7355

I

-0.ul366 -0.cq600 -cwlm -0.0208o -0.02~0 -c,o25w -0,027m -0.03010 -0.Of2L5 -co>15 -o.o3y$3 -3.03652

.

-0.03781 -0.03870 -0,03y86 -0.041 L3 -O.OLyi-0 -g.ow6 -0.04-547 -0.04736 -0.04%1 -0.05065

o.?g&!& -(?,05203, o.~g&!&c,O5203, 0.2~532,-W5321 0.2~532,-W5321 0.23624. 0.23624. -0.05422 -0.05422 0.25715 0.25715 -0.055& -0.055& 0.27806

I

I 0.27806 -o,Oy$8 -o,Oy$8 0.29839 0.29839 -0.05615 -0.0,5615 o.j1990,-0.056~~ O.jl990, -0.05643 0.3.m90] -0.0565fi 0.3.m90] -0.0565fi

iL c

a0 CL

-2 -0,236 -r& -0.185 -I -&I 34 4 -0.083

0 -0,032 3 03 9

t 0.070 13 0.1 a 2 0.172

24 0.223 3 0.274 j+ 0.326 $ 0.4-29 0.377

5 0.480 ti o.53l 6 0.583 65 0.635 7 0.687

73 0.738 8 0.790 a 0.841 9 @.@-I 9% 0.939

IO 0.986 I@ I.% l-1 1.074 113 Lll4 12 I A49

12* I .? 80 13 ~207 u$ 1.223 ?4

0.30

cD cm

0.0096 -O.Olll 0.0096 -0.c098 0.0097 -0.0090 O.OQY? -0.0085

OAO7 -0.0075 0,228 o.oi 04 -0,008O 0.0106 -0.0076 0.281 O.olO4 -0.008J 0.q 05 -0.0077 0.334 O.VlO5 -0.0081 0.0105 -0.0078 0.387 0mo5 -0.00~ o,ola 4OQ79 0.W 0.0~06 -0,008l

0.a c3 -0.0078 0*4?3 o&w6 -0,008l 0.6100 -0,006l 0.546 0.q 05 -0.0080 0.0098 -0.0048 0.599 0.q 02 -0.0078 omv6 -o,oogr, 0.652 w-ml -0.0066 O.OO97 -0.0070 0.706 O.olOQ -0, @O.&f3

omtk. 0.0068 W273 0.0095 CL0307 o.ooy4

~O,OO&r -0.0078 -0.oo75 -0.~75 -0.0075

-0.0078 -0.0077 -0.0064 -0, ooy -0.0037

-0,002~ ‘0.0006

o.oQlo 0.0027 o.ocG

0.35

5 cD

-0.243 0.0099 -0.1 9-l Q.QQYY -0,139 0.0099 -0,087 o.0100

-0.035 o.Olo2 0.028 O.mo3 0.071 o*ol OL 0.123 1 0.Q 04 0.175! aa03

0.759 0.0102 0,812 0.0106 0,864 O.OI? 6 0.915, O.CG8 0.963 omh3

LO09 0.a57 1.053 0.0172 ~096 @a187 I .I 9 0.0207 I ~66 0.~232

1.193 0.0262 1.205/ 0.0298

c m

-0,008l -0.0080 -0.~080 -0,00&l

-o,o082 -0.0087 -0.0080 -0~1080 -o,o08o

-0,0048 -o,om? -0.0~52 -0.0046 -0,003

-0*00?4 0,0006 0.0~27 0.0050 0.0076

o.olo2

0.40 I

cL I

cD

-0,250 WI 02 -0.197 Q.0lQ-l -0J43 O,crlol -o.oycr w3lo2

-0.037 O*OlO3 0.017 0.a 0: awl awl3 OAT5 0.0102 OJ?? o.clo2

0.233 0.0102 0,285 LolQ3 0.342 0.M05 0.397 0.m 06 0.m 0.a 08

0.506 O,Oll@ 0.561 0.0112 0.615 @.a133 0.670 oJ3 09 0.725 0.q 05

0.780 0.q 07 U34- Q.0112 0.888 O.cllY 0.940 0.a 29 0.988 O.ol45

I.032 @.OI 61 1.073 cJ.q 79

1.113 0.9 yy

1.139 0.0224 I .I 9, 0.0261

1.162 i o,ul12 ; I t I

I -1 &i-i -0.0079 I -0.008-r -0.m83

-O*OO& -O,OO& -0, 0084 -0.0053 -O,OO83

-c.ooK? I

-0.0082 1 -o,ao52 -0,0082 -0,00&1

-0. cml -0.0081 -@.o08j -0.0070 -O,OOL5

-0,30@3 -0.004=1 -Q.OQU -0.0039 -0,oq 8

0,0(?06 o.ooyt 0,@064 0.0096 Q.Ql23

TABLE 3 (Contfi>/

-9- TAE!JX 3 (CGNTD)

1 c CPI L ':, 1 'rr ; 'L ( 'D 1 'IT 'L ) 'D 1 %

-2 -0.257 -I+ -0.202 -1 -0.147

-3 -0.092

0 -0.07 + 0.018

1 0.073 1% 0.129 2 0.184

2% 0.2$cl L 0.296

0.3% 4 0.407 4; 0.463

5 0.520

2” 0.576 0.632 h-: c.68V 7 0.745

0.801 0.857

84 0.912 9 0.965 Y$ I.013

IO 1.055 104 1.089 11 1.116 116 I.132 12 I .I 32

12$ 13 134 14

0.0106 -0.0080 -0.264 0.0100 -0.0086 0.0104 -c.ooi32 -0.208 0.0106 -0.Ori88 0.013 -C.O085 -0.15l 0.0105 -0.GG80 0.0103 -0.0087 -0.0?4 o.olo4 -0. GO90

O.OlO3 -0.0088 a.037 0.a OL -0.C~90 o.Olo3 -0.W87 0.019 c.mo3 -0.0090 o,clr)2 -0.0087 G.G76 G.0103 -0.0c90 0.0102 -0.GO86 0.133 o.olo3 -0.0088 o.olo2 -0.0085 0.190 o.Olo3 -0.0087

0.0102 -0.0084 o.olo3 -0.0083 c.Ol05 -0.0083 0.0106 -0.0082 0.0108 -0.0~80

0.247 0.304 0.361 0.419

~ 0.476

0.0102 -0.0386 G.olO3 -C.OG05 0.0105 -c.O083 o.Olo7 -G.CO&’ o.otc8 -0.0~78

0.0110 -0.0078 O.Oll3 -0.oo.77 0.0116 -0.0075 0.0120 -0.OC62 0.0122 -c.o047

0.534 0.592 0.650 0.708 0.767

0.0110 -0.0075 G.oil3 -0.0071 0.0117 -0.33& 0.0-l 21 -G.O054 0.0126 -0.024

0.0123 -0.00:9 0.0123 -0.0037 o.Ol25 -0.0026 o.ol33 -0.0007 0.0147 0.0022

O.@l69 o.Ol94 0.0225

-0.271 0.0112 -0.213 0.0108 -0.155 O.Ol05 -0.006 0.0105

-0.038 0.0105 0.021 0.0104 0.080 0.0103 0.139 0.0:03 0.198 O.CHO:

0.258 O.Olo3 0.317 o.Olo3 0.376 c.0105 0.436 @.a07 0.496 o.mo9

0.557 0.cl11 0.619 0.0115 0.682 0.01'9 0.744 0.0124 0.805 G.Ol30

0.865 0.924 0.982 q.038 1.068

1.078

0.0140 0.C163 O.Cml 0.0241

-0.0097 -0.~96 -0.0095 -0.0094

-0.OC94 -0.00?4 -0.0093 -0.00% -0.0089

-0.0057 -o.coeg -0.0083 -0.0080

~ -0.0076

-o.c071 -0.0064 -0.0054 -G.OGw) -0.OC21

0.0003 0.00% 0.0070 G.WO O.Ol43

c.m57

TIBLF 3 (ContaV

- 10 -

TABLE 3 (COIITD)

NPL 9612

TABIX 3 (Contd'l/

M I I I

a0 I cL

-2 -0.392 *Q -0.298 -I -0.207

-3 -0.117

0 -0.035 8 0.042

I 0.118 I& 0.195 2 0.273

2+ 0.356 3 0.448 32 0,494 4 0.5l7 4; 0.520

5 0.521 6 :

? 1.

IO I@ 11 ?I& 12

128 ?3 13& 14

0.75

cD

-0.0208 -0.375 x0364 -0.0035 -0.129 -0.07% -0.q 99 -0.317 0.0-l 54 -0.0070 -0.100 -0.0692 -0.0171 -0,241 0.0140 -0.ol24 -0.083 -0.0553 -&or33 -0.145 c.olw -o.q60 -0.071 TO*0334

o.or13 -@.a-18 -0.a 9 0.M4-8 -0.0194 0.0!11 -0.01A2 0.082 0.0165 -0.0234 0.0110 -0.0111 0.163 0.01 yo -0.0281 o.Ol14 -0.oll3 0.228 0.02% -0.0337 o.q 28 -o.c-lj 9 0.274 0*0,417 -0.0366

o.a69 0.0241

-0.ol36 -0.0216 -0,027o -0.0293 -0.0302

0.297 -0.0354 0.308 -0.0306

c m

-0.0296

T

-11 - TABIE 3 (CONCL)

NPL 96.15,

5 J

0.80

cD c m

-0.063 -O,Oll4 -0.058 +o,olo2 -0.055 0.0316

0.85

cD C In

0 0.0 3.0103 o.oOfl4 0.0 o.Olo3 O.OGO5 0.0 O.OlO3 0.0006 6 0.050 O.OlO3 o*rmo4 0.05q 0.6103 0.0005 0.053 o.c-l0_7 0. cc06

-l 0.1 ol 0.0lo3 0.0004 0.303 O.QO3 0.0005 0.106 O.OlO3 0.0007 1% 0.152 0.0103 0.0005 O-158 o.cno3 0.0006 0.160 o.c-lo3 0.0008 2 0,201r. o.olob 0.0007 0,208 o.cno4 O.OOO7 0.213 O.o;% o,oooy

$ 0.255 o.oro4 a0009 0.260 o.qc& fl.oqo 0.266 o.cno4 o.ool2 3 0.306 0.~~06 o.oti-2 0.32 0.~06 o.Ool3 0.320 0.0-I 06 O.OOl4 3i 0.358 0.0 07 o.ml5 0.366 0.0Io7 o.ool6 0.375 0.0108 o.w7 4 0,409 0.0-l 09 0.c019 OAY o.qog 0.0020 0.430 0.0110 O.OO21 4* 0.460 0.0110 0.0025 0.472 O*Qll o.oc126 0.486 o.mj2 0.0028

" i:g ok56

O.oll~ 0.0111 0.ul11 0.0033 0.0041 0.0048 0.523 0.565 0.616 O.Oll4 o.Ol14 O.O-ll3 o.ooy, o.oou 0.0050 0.630 0.538 0.585 xTl;z 0:0l17 O.OO35 0.0045 0.0056 6& 0.0142 0.0056 0.667 0.0114 0.0059 0.684 o.a17 0.0066 7 0.709 0.0114 o.m63 0.722 0.01:6 0.0067 0.740 O.OllS 0.0076

I 7s 0.765 0.0122 0.0070 0.777 0.0124 0.0075 0.796 0.q 26 0.0087 8 0.819 c.0133 0.0077 0.834 0.0135 0.0384 0.852 o.0138 0.0098 8$ 0.874 O.O? 47 0.0085 0.858 0.0150 0.0092 0.904 o.a53 0.6109

k 0.928 0.979 0.0162 o.ol77 0.003 0.0099 0.940 0.988 0.0165 o.ol8-l 13.0099 o.a:z 0.955 0.998 o.ol9l o.or70 O.O-I23 0.0138

10 I.020 , o.olv3 o.olo7 I.024 0.0200 '0.0126 0.02 3 0.a 59 IO& 1.056 0.0212 o.oz16 4 ~56 o.ol4E;

:.g O.ol85

11 1.090 0.0233 0.0125 :.@87 o.cn69 ,:080 O.02O.L II& I.?20 0.0259 0.0~35 1.106 12 1.144 0.02Y4 o.m49 I .ooo

? 2$ 13 13s A4

!rABLE 4 (co-ltal/

-13 -

TABLE 4 (CONTD)

NACA 0012

XI 0.45 0.50 0.55

a0 cL cD % cL 'D m k cL 'D m C

-2 I -I3 -I

-3

0 0.0 0.0-l O? 0.0007 0.0 O.CIOI 0.0008 0.0 o.oloo 4

o.ooog 0.054 0.0102 0.0007 0.056

:& 0.108 0.164 0.0102 0.0102 0.0009 o.colo 0.112 0.167

, 0.0101 0.0009 0.058 0.0100 O.Coll

o.olol o.Olo2 O.Ooll 0.00l3 0.174 0.116 0.0101 O.MO2 0.0014 o.oOl7 2 0.218 o.cloj 0.0012 0.225 0.0102

!

0.0016

I

o.EJ+ 0.0!03 0.0021

2; C.273 o.oror, o.m4 0.283 0.01 or, 0.0019 0.295 o.ao5 0.0026 3 0.33c O.olO7 0.0017 0.342 0.0107 0.0022 0.356 0.0107 0.0030 3i 0.387 0.0,09 0.00l9 0.399 o.ol1o 0.0026 0.416 0.0110 ' 0.?035 4 O.&j 0.0111 0.002L 0.458 0.0l:3 o.ncfio 0.478 0.0ll3 0.0045 4; 0.500 0.0114 O.OC~30 0.518 0.0116 0.0037 o.y+o 0.0!17 o.oE4

$

6'

;*g;

0:652

0.0118 0.0116 0.0050 0.0039 0.578 0.631 0.0124 0.0120 0.0045 o.or!y 0.602 0.662 d.Ol23 0.0l30 o.no90 0.0068

0.0120 0.0063 n.684 0.0128 O.bO7j 0.723 0.0!39 0.6114 6+ 0.706 0.0122 0.0077 0.736 o.ol32 0.0098 0.733 o.ol54 0.0141 7 0.761 o.qz+ 0.0092 0.790 0.0139 O.Ml9 0.8W 0.0174 o.q@

74 0.8-16 0.0133 o.olnf; 0.848 0.0151 0.0142 0.891 0.0205 0.0200 8 0.871 0.0146 o.ol20 0.899 0.017L 0.0167 0.938 0.0360 c.Oz25 86 0.924 0.0l66 o.ol39 0.943 c.0216 0.0197 0.900 0.03E4 0.0!?4 Q '9;

0.973 o.c%92 O.ol62 0.767 0.0300 0.0232 0.844 -0.0022 1.004 0.0l90 0.935 0.0111

IO I.013 0.0228 0.845 I@

4.9 3

11 11; 12

12;

13 1% 14

TAECE 4 (Co&d)/

- 14 -

TABLE 4 (CONTD)

NACA 0012

t -r T 1 M 0.60 0.65 0.70

u” cL cD T c a

-2 -13 -1

-$

0 3.0 0.0100 0.0010 6 0.060 0.q 00 0+0ol3

1 0.120 0.0101 I$

O.ml7 0.182 0.0102 0.0022

2 0.245 0.0103 0.0028

2; 3 34

:;

0.309 O.cnO6 0.373 0.a 08 0.439 0.0110 0.504 0*0114 0.569 O.oll9

o.ooyc 0*004l 0.0052 0.0064 3.00&

0.634 0.700

x2 Ok82

0.ol27 0.0150 0.0-l 8.4 O.C230 0.0293

0.Llo4 3.691 0.0209 0.0116 O.Ol3l 3.752 0.0275 0.0135 0.ol58 3.796 0.0356 o.or46 0.q 90 3.805 0.0475 0.0102 aoa7 3.78O 0.0679 O.mlL

0.876 0.836

0.03Yl O.0l78 O.0048

IO IO+ 11 11s 12

129 13

,

, ,

-r

-I- -r

cL cD cL cD C m

3.0 0.0100 0.0011 1.0 o.oiol 0.0012 3.063 0.orc-l o.ool6 1.068 0.0102 0.0019 3.128 0.q 02 0.0021 3.138 0.0lo3 0.0025 O-193 0.M03 0.0027 l.213 0.0lo4 0.0034 0.260 o.wo4 0.0036 I.288 0.0106 0.0044

3.328 0.397 3.470

;:z%

o.oro7 0.0046 1.366 0.0111 o.oO58 ).&I.2 0.0126 0.0072 I.517 0.0123 0.0085 I.593 o.mn 0.0099 I.648

0.0114 0.0l35 0.0175 0.0245

0. ows 0.0060 0.0075 C.3068

0.0w

1.672 o.coo4 I.667 ~0.0046 1.657 .o.ooy_s

TABuE 4 (coda)/

M

IP - -

cL

o.ool 0.08cl c.~60 0.2& 0.328

0.0015 CL005 0 l 0024. 0. 00 0.00~ 0201 0.0c22 0.265

4.0004 0.275

0.408 -c. 0037 0,@2 -0.0078 0.500 -0.p 2' 0.500 -0.0166

0.75

?n

- 45 -

TABI 4 (CONCL)

NACA OQ2

0.80

C D +

.

0.a 76 0.0200 0.02q 0.0370

t

C I m I cL

o.wl8 o.ol3 -O.OO~& 0,020 -0.m 2c 0.020 -0.0163 -0.0208

0.85 1 C m

4. ocog +0.014

0.040

EM

Blend point

9615*OOlZ 1

NPL 9615

Position of 0012

BI end point

9615~0012

Comparison of profiles ahead of position of maximum thickness

30657

FIG 2

Comparison of upper surface curvature distributions

( p non-dimensionalised on actual chord)

30657 FIG. 3

I.4 I I , , I I I I

NPL 9615

I' o-

O-8-

cL

O-6- , 0.75

0.775 _

0’2

0

-0.2 /

025

(O-85

of -*1, d 6 b

-0-4 I I I I I -2 6Q$y8 ,,‘a to I2 14 I6

Variation of CLwith o( and ,M for NPL 9615

30657 FIG.4

0~0400

O-0300

CD

0*0200

o-0 100

0

NPL 9615

0.0 0.

1 2

O-7

5 1 &

I / / .

a

.6

u”

/

/

I IO

O-5 0 -1

-

I I

Ii

Variation of C, with o( and M for NPL 9615

30657 FIG 5

-2 0 2 4 a 14 16

Varlatlon of C, c/4 with o( and M for NPL 9615

30657 FIG.6

,O-

,6-

I 1 I I I I 1 1

NACA 0012

02

5

-4 I I I 0 , I I I 1 2 0 2 4 6 qo 8 IO I2 I4

(Scale for M=O*3 I

Variation of CL with o( and M for NACA 0012

J 16

30657

FIG. 7

o-04

0 *030(

CD

0*020(

JACA )OI2

l-7

-I-

Variation of C, with o( and M for NACA 0012

30657 FIG.8

o-02

o-o

0

-0-O

-0-o:

t-

2-

2- 0 2 4 I 6 8 I( ) a0

0 .d

7- 0-35

/

/

0

, ( I.3

Variation of C mc14

with ac and M for N ACA 0012

0.2 -

30 657 FIG. 9a

0(=--O 1

0.6

0.5

o-4

0.3

/ I I I I I 0 o-2 O-4

x/c Oe6 0.8 I.0

Upper surface

Lower surface

NPL 9615 Pressure distributions, K--Z’

0.2

0.4

P

%

0.t

0.0

I4

30657 FIG.9b

o(=-IO

Upper surface

a=-1 0

o-4

P

Ho 0.6

IA M=0*85 -

o-4 X/C

O-6

Lower surface

NPL 9615 Pressure distri buttons, a=-lo

30657 FIG .9c

O-2 c

a=oo

O-8 O-5

I 0.4

3 -

1 I I I I I

0 o-2 0 4 O-6 0.8 I.0 x/c

Upper surface

0’4

P

Ho

O-6

2 0.3

I .o I I I I I I

0 o-2 O-4 O-6 0.8 I-0 x/c

Lower surf ace

NPL 9615 Pressure distributions, a=Oo

30657 FIG.9d

0.2

O-0

02

0.4

P

Ho

06

a= to

I I I I

0.2 0 -4 x/c

O-6

Upper surface

o-0 I-0

o-2 o-4 x/c O-6 0.8

Lower surface

I -0

NPL 9615 Pressure distributions OC= lo

30 657 FIG. 9 e

a-2 0

. \ bw..- M=O-825 1

O-8

I I 0 o-2 0 -4 a,c O!b 018

Upper surface

I

O-b

O-8 o-5

o-4 / 0.3 /

I.01 0

I o-2

I I o-4

xlc 0.6

Lower surface

I O-8 I -0

,

NPL 9615 Pressure distributions, a=2’

30657

FIG.9 f

o-2

O-4

P

Ko

0.t

0-e

a=3’

0.2 o-4 x/c

O-6

Upper surface

0.2 I

oc=3O

0.6

@4 -

r 0.3 -

/ I I I I

0.2 0.4 x /c

O-6 0.8 I, 0

Lower surface

N PL 9615 Pressure distri butions,a-3’

30 657

0.2

o-4

11. Ho

0.6

O-8

I-C

a=4'

Upper surface ---- Lower surface

,- ye-- __L--- ---------- . -yjy--- ) __----- _____-- -------- .

I I I I I 0 0.2 0.4

XIC O-6 O-8 I’0

0.2

a.4

P

Tib

0.6

O-6

I .c

: I I ____ -__-------

)----- - --

‘. I I

0 0.2 0.4 x/c 016

9 NPL 961 5 Pressure distributions, a=4'

h NPL 9615 Pressure distributions, oc=5’

30657 FIG.9 i a j

a=6’

-----_

___----

0.2

0.8

I-0

i NPL 9615 Pressure distributions,a=6’

a-7’

-_----

\

0 o-2 o-4 x/c

O-6 O-8 I-O

i NPL 9615 Pressure distributions,a=7’

30657 FIG.9k l 1

a=BO

0 o-2 o-4 x/c

O-6 O-8 I-O

k NPL 961 S Pressure distributions,aa’

3

,6

,8 -

0 o-2 0’4 x,c O-6 O-8 7.0

a-9'

1 NPL 9615 Pressure distributions, a=9’

30657 FIG. 9m

o-2,

o-4

P

Ho O-6

a o-2 04 X/C

O-6 O-8

m NPL 9615 Pressure distributions, cx-IO0

30657 FIG.9 n l o

oc-II”

0.4

P

Ho

O-6

I-O 0 0.2 0.4 x,c O-6

n N PL 961 5 Pressure dlstributlon oc-ilo

0 NPL 9615 pressure distributions,oc=l20

30657

FIG 9p sq.

a=i3'

o-2 04 x/c 06 08

P NPL 96 15 Pressure distributions. a= 130

02------ -- - a= 136’

04

08

I 0

O-4 x/c 0.6

9 NPL 9615 Pressure distributions, a=i31/2’

30657 F-a&b

o-

P RI

0 *6 -

C

I --- - ,

\

( b60 / y OJiO - --_

7 n-40 - .-

I I.0 I I I I J 0 0.2 0.4 x/c O-5 0.0 I.0

0 NACA 0012 Pressure distributions , Q = 0’

0.2 a = I0

1 Moth No.

- Upper surface

-- - Lower surface

a=oo

- 0.30

I.0 I I I I 0 0.2 0.4 x/c 0.6 0.8 I.0

b NACA 0012 Pressure distributions, a=l"

30657 FIG.10 c 6 d

act0

I*01 I I I I 1 0 o-2 0.4 x/c 096 0.0 I-O

C NACA 0012 Pressure distrlbtiions , a- 2’

0.2,

I Cl=3 0

I-OF I I I I 0 O-2 0*4 x/c O-6 04

d NACA 0012 Pressure distributions , a = 3’

Mach No. 0.7r5 0.75 o-70

0.60 0.50 0.40 0.30

0

30657 FIG IO e 4 f

o-4

1 Ho

0.6

0-I

I4

0 Ho

0-t

04

Lower surface

dlstributionr at

fl omitted

M-0.675 ,0.725

i- I I I I 0.2 0.4 x/c O-6 0.0

e NACA 0012 Pressure distributions, a= 4O

a ~5’

A

,/- --_

I I I I o-2 0.4 x/c O-6 0.0

Mach No. 0.75 0.725

E-o.6

0.60

75

0.50 0.40 0. JO

0

Mach No. 0.70 p 675

O-65 0.60 050 D-40

f NACA 0012 Pressure distributions, a= 5’

30657 FIG. IO g L h

9 NACA 0012 pressure dhtributians. a = 6’

a=7O

J I 1 I 3 o-2 0.4 x/c O-6 O-0 I,

Mach No. 0.70

O-65 O-60 o-50 0.40 0.30

1

Mach No. 0 -65

O-60 o-55 D* 50 3.40 I.30

h NACA 0012 Pressure distributions, a= 7’

30657 FIG.10 i L j

0.4

P K

0.6

0 0.2 0.4 x/c O-6 0.8 I.0 i NACA 0012 Pressure distrlbutions,a = 8’

a=9’ I

Mach No. 0.60

2 o-40 o-30

O-8

No.

__------------ --mm_-_ 0.30 I I I

0 0.2 Ob4 x/c Oe6 0.8 I.0

i NACA 0012 Pressure distribution, a= 9’

30657 FIG.10 k rt

I _.- _L--

a= IO0

0 O-2 0.4 x/c O-6 k NACA 0012 Pressure distrlbutlons , a ~10’

0.2 a = II”

Mach No. 0*50

o-45 0.40 0.30

0

1

0 o-2 0.4 x/c I*0

.

Mach No. 0140 833

L N ACA 0012 Pressure distributions , a = I lo

3065 7. FIG.10 m

ot=l2O

-----------

I 0 0.2 0’4 x/c 0.6 0.8

_--- _-----

Mach No.

o-35 0.30

,

m NACA 0012 Pressure distributions, US 12’

30657 FIG.11

l-4

. I.2 -

--A- NPL 9615

NACA 0012

0.8 - Approximate

moment

0.6 - boundarles

Region of

09

Comparison of lImitin boundarles

0.020(

0.01 6 0

CD

O*OI40

WI20

0-0100

O-008 0 (

30657 FIG. I2

CL =o

NPL 9615

NACA 0012

0.4 0.5 M 0.6 0.7 0.8 0.9

Comparison of variation of zero lift drag coefficient with

Moth number

-

30 657 FIG. 14

0

0*2

P Ho

0.4

0.6

M = 0.6

CL = 0.76

NACA 0012 - - --

NPL 9615

,--- I \ i r\

p-pi- 6O I I O-0184

6-3O 0.0137

I I I I

0.2 O-4 x/c 0.6 0.8 I.0

Comparison of pressure distributions at CL= 0.76

M =0*6

30 657 FIG.1 5

9615

80

60

50

max

40

SO

20

IO .

0 o-3 0.4 0.5 M

I 5 0.7

Comparlron of variation at moximum lift / droa ratlo with

Mach number

30657 FI G. 16

0.03 Q-

0.02 -

GO Cl4

O*Ol -

-0103

-0.04

NACA 0012

70

-

W0

NPL 9615

Comparison of the variation ot the quarter-chord pitching moment with

Mach number

30657 FIG. 17

0 a

-0-005 cm =I4

a -0-o I 0

x- 3

t

-0*015

-0*020

-0.02 5

I NACA 0012

1 A (a = 0)

Y r 1A-j

0.4 0.5 0.6 M 0.7 0.9 I -r 9

COmpatiSOn of variation of the zero- lilt quarter-chord pitching

moment coefficient with Mach number

ABC CP No.1261 November, 196% Gregory, N. and Wilby, P. G.

NPL 9615 AND NACA 0012 - A COMPARISON OF AEZ.EODYNAl4IC WlyL

Ordinates, surface slopes and curvatures are listed for the two aemfoils together with a detailed tabulation of lift, drag and pitching moment data obtained at Mach numbers between 0.3 and 0.85 in the NPL 36 in. x 14 in. transonic tunnel. The aerodynamic characteristics and all the pressure distribution are plotted, with some comparisons.

BBC CP No.1261 November, 1968 Gregory, N. and Wllby, P. 0.

NPL 9615 AND NM.2 0012 - A COMPARISON OF AERODYNAMIC WE4

Ordinates, surface slopes and curvatures are listed for the two aerofoils together with a detailed tabulation of lift, drag and pitching moment data obtained at Mach numbers between 0.3 end 0.85 m the NPL 36 in. x 14 in. transonic tunnel. The aerodynamic characteristics and all the pressure distribution are plotted, with some comparisons.

ARC CP ~0.1261 November, 1968 Gregory, N. and Wilby, P. 0.

NPL 9615 AND NACA 0012 - A CONPARISON OF AEBODYNAMIC CA'iX

Ordinates, surface slopes end curvatures are listed for the two aerofoils together with a detailed tabulataon of lift, drag and pitching moment data obtained at Hach numbers between 0.3 and 0.85 in the NPL 36 in. x 14 in. traneonic tunnel. The aerodynamic characteristics end ell the pressure distribution are plotted, with some comparisons.

C.P. No. 1261

HER MAJESTY’S STATIONERY OFFICE

Go”er”me”r *wkrhops

49 High Holborn. London WCIV 6HB I3a Castle Street. Edrnburgb EH2 3AR

109 St Maw street. Cardiff CFI *,w BraZemore Street. Manchester M60 SAS

50 Farfax Street. Bristol BSI 3DE 258 Broad Street, Bum,,,gham B, 2HE 80 Cluchester Street. Belfast BT, 4JY

C.P. No. 1261 ISEN 0 I I 4am9 3