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C.P. No. 1261

MINISTRY OF DEFENCE (PROCUREMENT EXECUTIVE)

AERONAUTICAL RFSEARCH `PAPERS COUNCll

CURRENT

NPL 9615 and NACA 0012

A Comparison of Aerodynamic Data

BY

N. Gregory Aerodynamics and

.

P. G. Wilby NPL

'

Division,

LONDON:

HER MAJESTY'S 1973

STATIONERY

OFFICE

PRICE 8Op NET

CP No. 126;~. Rowember, 7968 NPL 9615 AND rucA co12 - A COMPARISON AERODYNAMIC OF 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 on

, The NACA four-digit series of wing sections was first reported in 1932 , and the following formulae for the thickness distribution and leading edge radius are taken from Ref. 2. f y/c = t/c0.20 0.29690 v'x/c - 0.12600 x/c - 0.35160 (x/c)'

( + 0.28430 wc)3 R,/" =

1.1019 it/c12

- 0.10150 wc)4 >

t/c = 0.12 in the above equations, and the NACA 0012 is obtained by putting ordinates are listed in 'IYLBLR together with surface slopes and curvatures. .l 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 [email protected]? for the standard. for a NACA four-digit thickness NACA OOi2, (a figure ,rhich is reduced to O.r)M distribution with the II.% thickness/chord ratio of NPL 9615). Cne of the &ms of the design features of NPL 7615 W.SSto reduce the to that for NACA 0012, in the region rihere curvature of the upper surface, relative This helps to reduce the maximum supersonic floz develops at incidenoe. 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 These were 31 in. (0.79 m.) apart throughout the length of the xrorkxnG section. 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 lift n 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 Furthermore, the oq~r,rison hatvrcen the 5.~0 sections sho!GS not be Mach number. 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 llthough the end effects 3re much further removed tunnel on a model of NAG OOl2. from m-La-span wzth a planform aspect ratio of 3.6 compared sith 1.4 for the the tests revealed e gain in CL ImY of over 0.1 when 36 m. x 14 in. iflnd tunnel, 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 DLrect-shadow photographs revealed +&t with the band present, the other. 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 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 1 for WL 9615 at Mach numbers betneen C/I and TA%?? 4 contains 0.3 and 0.85 with incidences between -2' and the stall, 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 Curves shoving the variation at CL, CD, md %I& to 3.75 x 10s at M = 0.85. with incidence are plotted in Figs. 3, l+ and 5 forNPL 9615, and m Figs. 6, 7 and 8 for NACA 0012.

characteristics

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

5.

Some Comparisons

The effect of the profile chanss refer-ed to inSection 2 on the aerodynamic characteristics of the 'GJOscctiors 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 particular, at incidences belo>; I', NPL 9615 exhibits a in their location. pronounced drag creep preceding the more rapid rise, as can be seen at zero lift This figure serves to emphasise an&her in the drag comparison of Fig. 12. This is that the values of CD, cm and OL differ limi'ation of Fig. 11.

____________________-------

- -____________

*2X-2,0 mesh carbmund~m lm~lies grains that were sieved through a %?u%e vrlth 231 wlms to the linear Inch, but This ,m,,lles 'inIlls that passed throu&!h P 59"pTe which were retained by P ea"ee with 270 wires to the Inch. 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 values mvst be 1lferr-d from the Tables, but are shown the boundaries: in sore more de%iled ?ompar~sons vhlch follow. NPL 9615 has dightly larger zero-lift &a& than NJ&A Sl? et Mach Mxh number for rrpid numbers below dreg rise, (Fig. 12), though its crlticsl &rag rise is marginally grextw, an& beyowl t:us point Its ratr of increase of This latter featxre rriurs also at zL+ze;-o values drag is apprecjably less. 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 CL 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 (The dlstributfon for WL 9615 was number of 0.6 anI1 a CL wlue of 0.76. 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 (Fig.

15)

in CL max of between 0.08 and 0.14 despite the reduction

at Mach numbers below 0.65 ratlo.

15, 18).

in thickness/chord

An increase

of about 0.02 In the drag-rise

Mach number (Figs.

A reduction of drag in the hlgh-drag region, particularly vhere supersonx flow is present. At values of CL m the region of CL for NACA 0012, max the saving in drag can amount to as much as 3& (Fig. 13). The maxunum values numbers (Fig.14). of lift/drag ratio are increased at all Mach

Conclusions/

- 5Acknowledgement 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.

References

I

&.

1

Author(s) Eaetman N. Jecobs X. E. iYard and 9. h!. Pinkertau

I. R. Abbott and

Title.

etc. sectlons wind

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

Theory

2

A. E. van Doenhoff

of Cng sections, IdcGrerr H111, I 949.

-6TABI. I NACA 0012 CRDIXXFS

X (t 0 cl. 0005 0.0010 0.0025 0.0050 0.0075 0.6100

z 0

e I

P

0 0.00~0 0.0056 0.0087

63,291 SF.5 70 59.6 49.8 43.5

0.0122

c.Ol49 O.Ol7O

0.0225 0.015 0.02

0.03 0.04 0.05 0.06 0.08 0.10 0.12 0.14 0.16

0.18

o.cl89

3.0206 0.0236 0.0284 0.0323 0.0355 0.0383 0.0430 0.0469 0.0499 0.0524 0.0544 0.0560 0.0574 0.0586 0.0594 0.0599 0.0600 0.0599 0.0595 0.0588 o.o:Eio

0.0563 O.0558 0.05ut 0.0529 0.0495 0.0456 O.04l3 0.0366 0.0715 0.0262 0.0205 0.0080 0.0145 o.ool3

39 35.6

:i=:

21.062

10.893 6.976 4.9E5 TJ.p15

:3:, 19.5 16.g 14.8 II .8

9.6

q.663

7.9 6.4

5.2

0.20 0.225

0.25

0.275 0.3 0.325 0.35 0.375

lc.2 3.3 2.3 1.4 0.7

0

-0.6 -1 .I 7

-1.68

0.741 0.644 c.568 0.505 0.~52 0.407 0.368 0.305

0.4

0.425 0.45 0.475 0.5 0.55 0.6 0.65 0.7 0.75 0.8

C.85

-2.13

-2.55 -2.93 -3.29 -3.61 -4.19 A70 -5.14 -5.56 2 5'; -& -7.10 -7.53

q.zq 8 0.165 0.138 0.1 3-l 0.142

::;, I .oo

-8.02

0.169

0

::

Surface slope,

degrees =

3o/o

P =

Surface curvature

C/';r

= l/p" = O.ol5P

Leading edge radns:

0 I

90 50 45 w 35 32.5 2: 26 24

22 20

iL c

53.1 0 -0.0136~ -0.oa55 co. oooq 0.00268 0.00649

C. 00893

s3.1

53 18

0.00443 0.00586

0.00857 o.m359 o.q726 0.021~2 0.02589 0*03065 0.03602 0.04209 0.04905 p.05297

-6.320 -6.702 -7.098

-7.525 I;-;;;.

I 9 2; 6:2 5*6 5.0 4.4 4.03

ml17 0.8588 WQ59 O-P529 0.9765 1 ,oooo

0.0247 ma93

o.ol36

0.007c; wm5 am3

o.ot-r63 0.01392 0.07 633 O.olW

o.oa4JJ

LWER SURFACE -90

-80.5

0 0.0003

19

18 17

0, fJ21c.07 0. WY+5

0.02687 0.02832

16 15

14

13.5 j3

12.5 12

'

II l 5

II 10.5

IO

9.5 9 a.5 8 7.5 7 6.5 6 5.5 5 4.5 4 3.5 3 2.5 I.82 0.66 0 -0.618 -:.68c

1

t

I

-2.133 -2.942 -3.611

I:$;;

-5L43 -5.557 -5.943

0.06~83 0.02y80 0.06682 z:,o7227 0.0313l 2.95 2.68 3.07823 CL03306 2.545 X0&52 0.03365 LO8495 O.OYt7~ 2.41 0.03528 2,275 2.08858 0.03612 2.14 k09wk 0.03698 2.005 ~96% 1.87 woo98 0*03?F36 A.735 0.10574 0.03876 0.0396? I ~33 3.1107G 0.0~~065 I A79 3.1 I 622 0.04=165 0.12239 L3lY 3.12928 I .I90 0.04271 3.13688 1 o.oh-3&1 ~085 0,@@+94 3,14520 0.996 0.04609 0.922 3.15s.03 0.04724 0.857 oA-838 023 0.04Y50 0.752 k~Wd+ O*-l97% 0.05059 x0: 7 OSQ995 0,05ak ok37 0.22358 [email protected];4 0.23703 O.C5~f7 0.05424 0,[email protected] 0.~6450 O.wtY? 0.05565 0.637 0.28333 0.0564 0.375 0.3409 1 o.c%5 9.0564 0.3642 0.4115 0.0554 0.0546 2.4351 0.0525 3.482~ 3.5292 I 0.0498 0.04-66 2.5763 3.62% 0,043" 0.038? x6iO4 3.7176 O-W+5 U'297 3.7646 3.22

3.76 3.49

0.05723

-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

2.5 2.0

~6

1

0.0008 O*Ool9 wo33 0.0056 0,008~ o.ol30 o,o? 811. . -0.Of2L5 0.023F; -co>15 0.02go -o.o3y$3 0.03535 -3.03652

-0.ul366 -0.cq600 -cwlm -0.0208o -0.02~0 I-c,o25w -0,027m -0.03010

0.03884

"15

&O

- 5.75 - 5.5 -5 - It. '5 -4 - 3.5 -3 - 2.5

-2

0.55 0.43 O&7

- 1.5 - 1 - 3.5 0

Ithen I 1 L/C I,eading

0.4l7

0.113

Hge rtiius:

0.04330 -0.03781 -0.03870 0*049~5 -0,03y86 ~.05900 0,07?32 -0.041 L3 0.09030 -O.OLyi-0 OJOO55 -g.ow6 0~1oy8 -0.04-547 0.13182 -0.04736 0.~5269 -0.04%1 -0.05065 OJ7355 o.~g&!&c,O5203, o.?g&!& -(?,05203, 0.2~532,-W5321 0.23624. I -0.05422 0.25715 I -0.055& 0.27806 -o,Oy$8 0.29839 -0.05615 -0.0,5615 o.j1990,-0.056~~ O.jl990, -0.05643 0.3.m90] -0.0565fi

with 8 +

ve.

as upper surface, ;

,Y - ve.

C

b/c

NIth centre at x/c = 0.~ 883, yf c = -q.r>137 f+7f fle is circular for LO' of arc on upper surface

P ,

= 1 fpo

= q.01883,

`mf i le joins smoothly with NACX0012 shape at x/c = 0.28333 on the unper surface crnd at x/c = 0.3409 on the lower surface

1

1

TABId3 =

0.30

a0 CL cD cm

0.35

0.40

I -1 &i-i I -0.0079 -0.008-r

-0.m83

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

5

-0.243

cD

cm

I I

cL

cD WI 02

-0,236 -r& -0.185

-2

-I 4

0

0.0096 -O.Olll 0.0096 -0.c098 -&I 34 0.0097 -0.0090 -0.083 O.OQY? -0.0085

-0,032

0.0099 -0,008l -0.1 9-l Q.QQYY -0.0080 -0,139 0.0099 -0.~080 -0,087 o.0100 -0,00&l

-0,250

-0.197 -0J43

-o.oycr

Q.0lQ-l O,crlol

w3lo2

3 t 13 2

24

0.070

03 9 0.1a 0.172

-0.0078 -0.oo75

-0.~75 -0.0075

~O,OO&r -0.035

o.Olo2 -o,o082 -0.037 0.028 O.mo3 -0.0087 0.017 0.071 o*ol OL -0.0080 awl 0.123 1 0.Q 04 -0~1080 OAT5 -o,o08o OJ?? 0.175! aa03 -0,008O -0.008J -0.0081 -0.00~ -0,008l -0,008l -0.0080 -0.0078 -0.0066 0.233

0,285 0.342

O*OlO3

0.a 0:

awl3

0.0102 o.clo2

-0.0053 -O,OO83

-c.ooK? -0.0082 -o,ao52 -0,0082 -0,00&1 -0. cml -0.0081 I 1

j+ $

5 ti

3

0.223 0.274 0.326 0.377 0.4-29

0.480

OAO7 -0.0075 0.0106 -0.0076 0.q 05 -0.0077 0.0105 -0.0078

o,ola 4OQ79

0,228 o.oi 04 0.281 O.olO4 0.334 O.VlO5 0.387 0mo5 0.0~06 0.W 0*4?3 0.546 0.599 0.652 0.706

0,812

0.0102 LolQ3

0.397

0.m

0.M05

0.m 06 0.a 08 O,[email protected]

0.0112

6

65 7

0.a c3 -0.0078 o.53l 0.6100 -0,006l 0.583 0.0098 -0.0048 0.635 omv6 -o,oogr, 0.687 O.OO97 -0.0070

o&w6 0.q 05 0.q 02 w-ml O.olOQ

0.0106

0.615 0.670 -0, @O.&f30.725 0.780

U34-

0.506 0.561

@.a133 [email protected] oJ3 09 -0.0070 0.q 05 -O,OOL5

0.q 07 -0,[email protected] Q.0112 -0.004=1

O.cllY

-Q.OQU

73

8 a

9

IO [email protected]

l-1 113

0.841

@[email protected]

0.738 0.790

-0.0078 -0.0077 -0.0064

-0, ooy

0.759 0.0102 -0,0048

0,864 O.OI? 6 -0.0~52 0.915, O.CG8 -0.0046 0.963 omh3 -0,003

LO09

-o,om?

9% 0.939

-0.0037 -0,002~

`0.0006 o.oQlo o.ocG

omtk.

0.888 0.940 0.988

0.a 29 -0.0039 @.OI 61

cJ.q 0.9 79 yy

O.ol45 -0,oq 8 0,0(?06

o.ooyt

0.986

I.% 1.074

12 12*

Lll4 I A49 I .? 80

~207 1.223

0.0027

1.053 0.0172 ~096 @a187

I .I 9 I ~66 1.193

-0*00?4 0,0006 0.0~27 0.0207 0.0050 0.~232 0.0076

0.a57 o.olo2 o,ul12

I.032 1.073 1.113 1.139

I .I 9, 1.162 i

0.0224

0.0261

0,@064 0.0096

Q.Ql23

0.0068

u$

?4

13

0.0095 CL0307 o.ooy4

W273

0.0262 1.205/ 0.0298

; t

I I

TABLE 3 (Contfi>/

-9TAE!JX 3 (CGNTD)

CPI 1

-2

cL

':,

0.0106

1

'rr

-0.0080 -c.ooi32 -C.O085 -0.0087 -0.0088 -0.W87 -0.0087 -0.GO86 -0.0085 -0.0084 -0.0083 -0.0083 -0.0082 -0.0~80

;

'L

-0.264 -0.208 -0.15l -0.0?4 a.037 0.019 G.G76 0.133 0.190

(

'D

0.0100 0.0106 0.0105 o.olo4 0.a OL c.mo3 G.0103 o.olo3 o.Olo3 0.0102 G.olO3 0.0105 o.Olo7 o.otc8 0.0110 G.oil3 0.0117 0.0-l 21 0.0126

1

'IT

-0.0086 -0.Ori88 -0.GG80 -0. GO90 -0.C~90 -0.0090

'L

-0.271

)

'D

0.0112 0.0108

1

%

-0.0097

-I+

-1 -3 0 + 1 1% 2 2% L

-0.257 -0.092

-0.202 -0.147

0.0104 0.013 0.0103

-0.213 -0.155 -0.006 -0.038

0.021 0.080 0.139 0.198 0.258 0.317 0.376 0.436 0.496 0.557 0.619 0.682 0.744 0.805 0.865 0.924 0.982 q.038 1.068 1.078

-0.~96 O.Ol05 -0.0095 0.0105 -0.0094 0.0105 -0.OC94 -0.00?4 -0.0093

-0.00% -0.0089

-0.07 O.OlO3 0.018 o.Olo3 0.073 o,clr)2 0.129 0.0102

0.184 0.2$cl o.olo2 0.0102

-0.0c90 -0.0088 -0.0087 -0.0386 -C.OG05

-c.O083 -G.CO&' -0.0~78 -0.0075 -0.0071 -0.33& -G.O054 -0.024

0.0104 0.0103 0.0:03 O.CHO: O.Olo3 o.Olo3 c.0105 @.a07 o.mo9 0.cl11 0.0115 0.01'9 0.0124 G.Ol30 0.0140 0.C163 O.Cml 0.0241

4 4; 5

0.296 0.3% 0.407 0.463

0.520

o.olo3 c.Ol05 0.0106 0.0108

0.0110

0.247 0.304 0.361 0.419 ~ 0.476

-0.0057

-o.coeg -0.0083 -0.0080 ~ -0.0076 -o.c071 -0.0064 -0.0054 -G.OGw) -0.OC21 0.0003 0.00% G.WO

-0.0078

-0.oo.77 -0.0075 -0.OC62

2" h-:

7 84 9 Y$

IO 104 11 116 12 12$ 13 134 14

0.576 0.632 c.68V 0.745 0.801 0.857

0.912

O.Oll3

0.0116 0.0120 0.0122 0.0123 0.0123 o.Ol25 o.ol33 0.0147 [email protected] o.Ol94 0.0225

0.534

0.592 0.650 0.708 0.767

-c.o047

-0.00:9 -0.0037 -0.0026 -0.0007 0.0022

0.0070

0.965 I.013 1.055

1.089 1.116 I.132 I .I 32

O.Ol43 c.m57

TIBLF

3

(ContaV

- 10 -

TABLE 3 (COIITD) NPL 9612

TABIX 3 (Contd'l/

TABIE

-11 3

-

(CONCL)

NPL 96.15,

M I I

I

0.75

T

cm

-0.0208 -0.q 99 -0.0171 -&or33 5J -0.375 -0.317 -0,241 -0.145 -0.a 9 0.082

0.163 0.228

0.80

cD

x0364 0.0-l 54 0.0140 c.olw 0.M4-8 c

0.85

a0 I

cL

cD

m

cD -0.129

-0.100

C In -0.07%

-2 -0.392 *Q -0.298 -I -0.207 -3 -0.117 0

-0.0035 -0.0070 -0.ol24 -o.q60 -0.0194 -0.0337 -0.0366 -0.0354 -0.0306

-0.0692

-0.083

-0.071

-0.0553 TO*0334 -O,Oll4 +o,olo2 0.0316

8 I

I& 2

-0.035 0.042

0.118

0.195 0.273 0.356 0.448 0,494 0.5l7 0.520 0.521

o.or13 0.0!11 0.0110 o.Ol14 o.q 28 o.a69 0.0241

[email protected]

-0.01A2

-0.0111 -0.oll3 -o.c-lj 9 -0.ol36

0.0165 -0.0234 0.01 yo -0.0281

0.02%

-0.063 -0.058

-0.055

0.274 0.297 0.308

0*0,417

2+ 3 32 4 4; 5 6: ? 1.

-0.0216

-0,027o -0.0293 -0.0302 -0.0296

IO [email protected] 11 ?I& 12 128 ?3 13& 14

0 6 -l 1%

2

0.0 0.050 0.1 ol 0.152

0,201r.

3.0103 O.OlO3 0.0lo3 0.0103 o.olob o.oro4 0.~~06 0.0-l 09 0.0110

o.oOfl4 o*rmo4 0.0004 0.0005 0.0007 a0009 o.oti-2 0.c019 0.0025

0.0 0.05q 0.303 O-158

0,208

o.Olo3 0.6103 O.QO3 o.cno3 o.cno4 o.qc& 0.~06 0.0Io7 o.qog O*Qll

O.OGO5 0.0005 0.0005 0.0006 O.OOO7 fl.oqo o.Ool3 o.ool6 0.0020 o.oc126

0.0 0.053 0.106 0.160

0.213

O.OlO3 o.c-l0_7 O.OlO3 o.c-lo3 O.o;% o.cno4 0.0-I 06 0.0108 0.0110 o.mj2

0.0006 0. cc06 0.0007 0.0008 o,oooy o.ool2 O.OOl4 o.w7

O.OO21 0.0028

$ 3 3i 4 4*

0.255 0.306 0.358 0,409 0.460

0.260

0.32

0.0 07 o.ml5

0.366

OAY 0.472

0.266 0.320 0.375 0.430 0.486

" 6& 7 7s

8

i:g ok56 0.709 0.765 0.819 0.874

0.ul11 0.0111 O.oll~ 0.0142 0.0114

0.0122

0.0041 0.0033 0.0048 0.0056 o.m63 0.0070 0.0077 0.0085

0.616 0.565 0.523 0.667 0.722 0.777 0.834 0.858

O.O-ll3 o.Ol14 O.Oll4 0.0114 0.01:6 0.0124 0.0135 0.0150 o.ol8-l 0.0165 0.0200

0.0050 o.oou o.ooy, 0.0059 0.0067 0.0075 0.0384

0.0092

0.585 0.538 0.630 0.684 0.740 0.796 0.852 0.904 0.998 0.955

0:0l17 xTl;z o.a17 O.OllS 0.q 26 o.0138 o.a53 o.or70 o.ol9l

I

0.0056 0.0045 O.OO35 0.0066 0.0076 0.0087 0.0098

8$

c.0133 O.O? 47 0.0162 o.ol77 , o.olv3

0.0212

0.6109

0.0138 O.O-I23

k

10 11 II& 12

? 2$

0.979 0.928

I.020 1.090 I.?20 1.144

0.003 0.0099

o.olo7

o.oz16

0.940 0.988

I.024 4 ~56 :[email protected] 1.106

I .ooo

o.a:z 13.0099 '0.0126

o.ol4E; o.cn69

IO& 1.056

:.g

,:080

0.02 3

0.a 59

O.ol85 O.02O.L

0.0233 0.0259 0.02Y4

0.0125 0.0~35 o.m49

13

13s

A4

!rABLE

4 (co-ltal/

-13

-

TABLE 4 (CONTD) NACA 0012

XI a0 -2 -I3 -I -3 0 4 :& 2 2; 3 3i 4 4;

$

0.45

0.50

0.55 k I m

cL

cD

%

cL

'D

cL

'D

Cm

0.0 0.054 0.164 0.108 0.218 C.273 0.33c 0.387 O.&j 0.500

;*g;

0.0-l O? 0.0102 0.0102 o.cloj o.oror, O.olO7 0.0,09 0.0111 0.0114 0.0116 0.0118 0.0120 0.0122 o.qz+ 0.0133 0.0146 0.0l66 o.c%92

0.0007 0.0007 o.colo 0.0009 0.0012 o.m4 0.0017 0.00l9 0.002L O.OC~30 0.0050 0.0039 0.0063 0.0077 0.0092 o.olnf; o.ol20 o.ol39

O.ol62

0.0 0.056 0.167 0.112 0.225 0.283 0.342 0.399 0.458 0.518 0.631 0.578

n.684

O.CIOI , 0.0101 o.Olo2 o.olol 0.0102

0.0008 0.0009 0.00l3 O.Ooll 0.0016

0.0 0.058 0.116 0.174

o.EJ+

o.oloo 0.0100 O.MO2 0.0101 0.0!03 o.ao5 0.0107

0.0110

o.ooog O.Coll o.oOl7 0.0014 0.0021 0.0026 0.0030 ' 0.?035 0.0045

o.oE4

0.01 or, ! 0.0019 0.0107 0.0022 o.ol1o 0.0026 0.0l:3 o.ncfio 0.0116 0.0037

0.0124 0.0120 o.or!y 0.0045

I 0.295 0.356 0.416 0.478

o.y+o

0.0ll3 0.0!17 0.0l30 d.Ol23 0.0!39

o.ol54 0.0174

0.662 0.602 0.723 0.733 0.8W

0.891

o.no90 0.0068 0.6114 0.0141

[email protected]

6' 6+ 7 74 8 86

Q

0:652 0.706 0.761 0.8-16 0.871 0.924 0.973 1.004 I.013

0.736 0.790 0.848 0.899 0.943 0.767 0.935 0.845

0.0128 o.ol32 0.0139 0.0151 0.017L c.0216 0.0300

O.bO7j 0.0098 O.Ml9 0.0142 0.0167 0.0197 0.0232 0.0111 4.9 3

0.938 0.900

0.844

0.0205 0.0360 0.03E4

0.0200 c.Oz25 0.0!?4 -0.0022

'9; IO [email protected] 11 11; 12 12; 13 1% 14

0.0l90 0.0228

TAECE

4 (Co&d)/

- 14 -

TABLE (CONTD) 4 NACA 0012

t

M

u" -2 -13 -1 -$ 0 6 1 I$ 2 2; 3 34 :;

0.60

cL

T

cD

-r

c a

0.65

T

cL

0.70

-I-

-r

1

C m

cL

cD

cD

3.0 0.060 0.120 0.182 0.245 0.309 0.373 0.439 0.504 0.569 0.634 0.700 x2 Ok82 0.876 0.836

0.0100 0.0010 0.q 00 0+0ol3 0.0101 O.ml7

0.0102 0.0103 O.cnO6 0.a 08 0.0110 0*0114 O.oll9 0.ol27 0.0150 0.0-l 8.4 O.C230

0.0293

0.0022 0.0028 o.ooyc 0*004l 0.0052 0.0064 ,3.00& 0.Llo4 O.Ol3l 0.ol58 0.q 90 aoa7

3.0 3.063 3.128 O-193 0.260 3.328 0.397 3.470 ;:z% 3.691 3.752 3.796 3.805 3.78O

0.0100 0.orc-l 0.q 02 0.M03 o.wo4 o.oro7 0.0111 0.0126 0.0123 o.mn 0.0209 0.0275 0.0356 0.0475 0.0679

0.0011

o.ool6 0.0021 0.0027 0.0036 0.0046 o.oO58 0.0072 0.0085 0.0099 0.0116 0.0135 o.or46 0.0102 O.mlL

1.0 1.068 3.138 l.213 I.288 1.366 ).&I.2 I.517 I.593 I.648 1.672 I.667 1.657

o.oiol 0.0102 0.0lo3 0.0lo4 0.0106 0.0114 0.0l35 0.0175 0.0245

0.0012 0.0019 0.0025 0.0034 0.0044 0. ows 0.0060 0.0075

C.3068

0.0w o.coo4 ~0.0046 .o.ooy_s

0.03Yl

,O.0l78 ,O.0048

IO IO+ 11 11s 12 129 13

-r

TABuE 4

(coda)/

-

45 4 (CONCL)

TABI

NACA OQ2

M

-

0.75

0.80

0.85

C m I I cL

IP

cL

?n

+

C D

1

Cm

o.ool 0.08cl

c.~60

0.0015

0 l 0024. 0.00~

CL005

0. 00 0201

0.265

0.2& 0.328 0.408

0,@2

0.0c22 4.0004 -c. 0037 -0.0078 -0.p 2' -0.0166

0.275

o.ol3 0.a 76 o.wl8 0.0200 -O.OO~& 0,020 0.02q -0.m 2c 0.020 0.0370 -0.0163 -0.0208

4.

ocog +0.014

0.040

0.500 0.500

t

.

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 ( p

of upper surface

curvature

distributions chord)

non-dimensionalised

on actual

30657 FIG. 3

I.4

I

I NPL

, 9615

,

I

I

I

I

I' o-

O-8cL O-6, 0.75 0.775 _

0'2

025

0

(O-85

-0.2

of -*1, d

-2

/

6

b

I 6Q$y8 I ,,`a to I I2 I 14 I I6

-0-4

Variation

of CLwith

o( and ,M for NPL

9615

30657 FIG.4

0~0400 NPL 9615

-

0.0

O-0300

0.

CD

0*0200

1

5 &

O-7

.6

O-5

0 -1

I

I

I/ / .

/ /

o-0 100

0

1 I

2

a u"

IO

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

I

1

I

I

I

I

1

1

NACA

0012

,O-

,6-

02 5

-4 2

I 0 (Scale for

I 2

I 4 M=O*3 I

0

, 6 qo

I 8

I IO

I I2

1 I4

J

16

Variation

of CL with

o( and M for NACA

0012

30657

FIG.

7

o-04

JACA

)OI2

-I-

0 *030(

CD

l-7

0*020(

Variation

of C, with o( and M for NACA 0012

30657 FIG.8

0 .d 0 o-02 7- 0-35 / / o-o , (I.3

0

-0-O t-

-0-o: 2-

2-

0

2

4

I

6 a0

8

I( )

Variation

of

C mc14

with

ac and M

for

N ACA

0012

30 657 FIG. 9a 0.2 0(=--O

1

0.6 0.5 o-4 0.3 / 0 I o-2 I O-4 Upper I x/c Oe6 surface I 0.8 I I.0

Lower

surface

NPL

9615

Pressure

distributions,

K--Z'

30657 FIG.9b 0.2

o(=-IO

0.4 P %

0.t

0.0

I4

Upper surface

a=-1

0

IA o-4 P

-

M=0*85

Ho

0.6

o-4

Lower

X/C

O-6

surface

NPL

9615

Pressure

distri

buttons,

a=-lo

30657 FIG .9c O-2 c

a=oo

O-5 O-8 I 0.4 3 1 0 I o-2 I 04 Upper I O-6 I 0.8 I I.0

x/c surface

0'4 P Ho O-6

2

0.3

I .o I 0

I

o-2

I O-4

Lower

I

x/c O-6

I

0.8

I

I-0

surf ace

NPL

9615

Pressure distributions,

a=Oo

30657 FIG.9d 0.2

a= to

O-0

I

I

I

I

0.2

0 -4 Upper

x/c

O-6

o-0

I-0

surface

02

0.4 P Ho 06

o-2

o-4 Lower

x/c

O-6

0.8

I -0

surface

NPL

9615

Pressure

distributions

OC= lo

30 657 FIG. 9 e a-2 0

.

\

bw..-

M=O-825

1

O-8

,

I 0

I

o-2

0 -4 a,c Upper

O!b

018

surface

I

O-b O-8

o-5 / / o-4 0.3

I 0 I I I

I.01

o-2

o-4 Lower

xlc

0.6

O-8

I -0

surface a=2'

NPL

9615

Pressure distributions,

30657

FIG.9

o-2

f

a=3'

O-4 P Ko 0.t

0-e

0.2

o-4

x/c

O-6

Upper

0.2

I

surface

oc=3O

0.6

@4 r / I 0.2

0.3

I O-6 I 0.8

I 0.4

x /c

I, 0

Lower N PL 9615

surface distri butions,a-3'

Pressure

30 657 0.2 a=4' Upper surface Lower surface

o-4 11. Ho 0.6

----

O-8

I-C 0

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

I 0.2 I 0.4

XIC

-yjy--I O-8 I I'0

O-6 distributions,

9

0.2

NPL 961 5 Pressure

a=4'

a.4 P Tib 0.6

O-6

: I I

`. ____ -__-------

)-----

-

--

I .c

0

I

I

0.2

0.4

x/c

016

h

NPL 9615

Pressure

distributions,

oc=5'

30657 FIG.9

i aj

a=6'

-----_ ___----

i 0.2

NPL

9615

Pressure

distributions,a=6'

a-7'

0.8

\

-_----

I-0

0

o-2 NPL

o-4 9615

x/c Pressure

O-6

O-8

I-O

i

distributions,a=7'

30657 FIG.9k

l

1

a=BO

0 k

o-2 NPL 961 S

o-4

x/c

O-6

O-8

I-O

Pressure

distributions,aa'

a-9'

3 ,6

,8 -

0

o-2

0'4

x,c

O-6

O-8

7.0 a=9'

1

NPL 9615

Pressure

distributions,

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 n 0.2 0.4 x,c O-6 dlstributlon oc-ilo N PL 961 5 Pressure

0

NPL 9615

pressure

distributions,oc=l20

30657 FIG 9p sq. a=i3'

o-2

P NPL 96 15

04

Pressure

x/c

06

08 a= 130

distributions.

02------ -04

a= 136'

-

08

I 0

O-4

x/c

0.6

9

NPL

9615

Pressure

distributions,

a=i31/2'

30657

F-a&b

--

Upper Lower

surface surface

a=oo

oP RI 0 *6 1Moth No.

I ,

C

/

---

-

\ (b60 y --_ OJiO

7

n-40 - .I.0

II.0

I 0

0

I

0.4 0012 x/c Pressure

I

O-5 distributions

I

0.0 , Q = 0'

J

0.2

NACA

0.2

a = I0

I.0 0

b

0.2

NACA

I

0.4

I

x/c

0.6 distributions,

I

I

0.8

a=l"

- 0.30

I.0

0012 Pressure

30657 FIG.10

c 6 d

act0

I*01 0

C

o-2 NACA

I

0.4 x/c 0012 Pressure

I

I

096 distrlbtiions

I

0.0 , a2'

1

I-O

0.2,

I

Cl=3

0

Mach 0.7r5 0.75 o-70 0.60 0.50 0.40 0.30

No.

I-OF 0

d

I

O-2 NACA

I

0*4 x/c 0012 Pressure

I

O-6 distributions

I

04 , a = 3' 0

30657 FIG IO e 4 f

Lower o-4 1 Ho 0.6 fl

surface

dlstributionr at M-0.675 ,0.725 omitted

Mach 0.75 0.725 E-o.6 0.60 0.50

i-

No.

75

0-I

0.40 0. JO

I I I I

I4

e

0.2 NACA

0.4 0012

x/c Pressure

O-6

0.0 a= 4O

0

distributions,

a ~5'

0

Ho

0-t

Mach 0.70 p 675 No. O-65

04

A --_ ,/-

0.60

050 D-40

I

I

I

I

o-2

f

0.4

x/c

O-6 distributions,

0.0 a= 5'

NACA

0012 Pressure

30657 FIG. IO g L h

Mach 0.70 O-65 O-60 o-50 0.40 0.30 1

9 NACA

No.

0012 pressure

dhtributians.

a = 6'

a=7O

Mach No. 0 -65 O-60 o-55 D* 50 3.40 I.30

J 3 I 1

o-2 h

I

0.4 x/c

O-6 distributions,

O-0 a= 7'

I,

NACA 0012 Pressure

30657

FIG.10 i L j

0.4

P K

0.6

Mach 0.60 2 o-40 o-30 0

i

No.

0.2 NACA

0.4 x/c O-6 0.8 0012 Pressure distrlbutions,a

I.0 = 8'

a=9'

I

No.

O-8

__-----------0

i

--mm_-_ I Oe6 distribution, I 0.8

a=

0.30 I.0 9'

0.2 NACA

I Ob4 x/c 0012 Pressure

30657

FIG.10 k rt

a= IO0

Mach 0*50

I _._L--

No.

o-45 0.40 0.30 0.4 x/c

0012 Pressure

0 k

O-2

NACA

O-6

distrlbutlons , a ~10'

0

0.2

a = II"

1

.

Mach

No.

0140 833

0

o-2 L

N ACA

0.4 x/c

0012 Pressure distributions , a = I lo

I*0

3065 7. FIG.10 m

ot=l2O

Mach No.

_----_-------------

o-35 0.30 ,

0

I

0.2

0'4 x/c

0.6

0.8

m

NACA 0012 Pressure

distributions,

US 12'

30657 FIG.11

l-4 --A-

NPL NACA

9615 0012

. I.2 -

0.8 -

Approximate moment

0.6 -

boundarles

Region

of

09 Comparison of lImitin boundarles

30657

FIG. I2

0.020(

CL =o

0.01 6 0 CD

O*OI40

WI20 NPL 9615

0-0100

NACA 0012

O-008 0 ( 0.4 0.5 M

0.6

0.7

0.8

0.9

Comparison

of

variation

of Moth

zero

lift

drag

coefficient

with

number

-

30 657 FIG. 14

0

M = 0.6 CL = 0.76 0*2 P Ho 0.4 ,--I i NACA NPL 0012 9615 --

6O I I

p-pi-

O-0184

6-3O

0.0137

r\

\

I I I I

0.6

0.2

O-4 x/c

0.6

0.8

I.0

Comparison

of

pressure

distributions

at

CL= 0.76

M =0*6

30

80

657

FIG.1 5

9615

60

50 max 40

SO

20

IO .

0 o-3

0.4

0.5

M

I 5

moximum number lift

0.7

Comparlron

of

variation Mach

at

/ droa

ratlo

with

30657 FI G. 16 0.03 Q-

0.02 GO Cl4

O*Ol -

NACA

0012 70

W0

NPL

9615

-0103

-0.04

Comparison

of the

variation

ot Mach

the quarter-chord number

pitching

moment

with

30657 FIG. 17

NACA 1 Y 0.4 A r 0.5 (a = 0) 0.6 M

I0012

1A-j 0.7

0 a

-0-005 cm =I4 xa3 -0-o I 0

0.9 I -r

9

t

-0*015

-0*020

-0.02 5

COmpatiSOn

of

variation

of

the with

zeroMach

lilt

quarter-chord number

pitching

moment

coefficient

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

Wilby, P. G.

BBCCP No.1261 November, 1968 Gregory, N. and Wllby, P. 0. NPL 9615 ANDNM.2 0012 A COMPARISON AERODYNAMIC OF 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.

NPL 9615 ANDNACA0012 A COMPARISON AEZ.EODYNAl4IC OF WlyL Ordinates, surface slopes and curvatures are listed for the two aemfoils together with a detailed tabulation of lift, drag and pitching momentdata 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.

ARCCP ~0.1261 November, 1968 Gregory, N. and Wilby, P. 0. NPL 9615 ANDNACA0012 A CONPARISON AEBODYNAMIC OF CA'iX Ordinates, surface slopes end curvatures are listed for the two aerofoils together with a detailed tabulataon of lift, drag and pitching momentdata obtained at Hach numbers between 0.3 and 0.85 in the NPL The aerodynamic 36 in. x 14 in. traneonic tunnel. characteristics end ell the pressure distribution are plotted, with some comparisons.

C.P. No. 1261

HER

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C.P. No. 1261

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