Document
---------------------------------------------------------- ~ ------ , , J NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ORIGINALLY ISSUED February 1946 as Restricted Bulletin L5K02 AERODYNPJ{[C CHARACTERISTICS OF FOUR NAeA AJRFOn.
S'ECTIONS DESIGNED FOR HELICOPI'ER ROTOR BLADES By Louis S. Stivers, Jr. and Fred J. Rice, Jr.
l' Langley Memorial Aeronautical Labor atory Langley Field, Va • • Trl:RmcAl U8RARY CAtRETT CORP. AJRrSEARCfI MFB O/V 9851-99S1 Sepulveda 8 Vd.· • IDs Angeles , Cillt. 9Ua WASHINGTON NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were pre- viously held Wlder a security status but are now Wlclassified. Some of these reports were not tech- nically edited. All have been reproduced without change in order to expedite general distribution.
L-29 NACA RB NC' . L5K02 RESTRICTED NA:l'IONAL ADVISORY COMMITTEE FOR AERONAUTICS RESTRICTE.D BULLETDJ ASRODYlIJ" lMIC CEA.-qACTSRISTICS OF ?Ot.JR N AC A AIRFOIL S3CTIO~;S DESIGN~D FOR H~LICOPTE2 ROTO~ BLADES By LOUlS S . Stivers , Jr ~ and Fred J . Rice, Jr .
SUMMARY Four NACA airfoil secti-:ms , the NACA 7 - H- 12 , 8 - H- 12 , ~ - H - 12 , and lO - H- 12 , suitable for use as rotor - blade sections for helicopters and other r tary - wing aircraft have been derived and tested . These airfoil secti o ns have comparatively low drags in the ra..'1ge of low and moderate llfts and small pitching moments that are nearly c o nstant up to maximUil1 lift . The undesirable adverse changes in aerodyn~~ic characteristics at higher lifts and the undue sensitivity t ,J rou g hness, which were found for the alrfoil sections reported in NACA CB No . 3113, are minimized for the airf o il sections presented . A com - paris on f calculated profile - drag losses for a rot or successivel) incorporating the NAC 3 - E- 13 . 5 (reported
in NAC 4 CB No . 3113) and the 8 - H- 12 airfoll sections
showed that the FACA 8 - H- 12 had smaller profile - drag losses in nearly every operating condlti on presented .
From aero-dynamic considerations , the NACA 8 - H- 12 and 9 - H- 12 airfoil sections appeared more promlslng f or use as r ., tor - blade sections than any other airfoils thus far tested at the NACA laboratories .
INTRODUSTION The desirable aerodynami characteristics of airfoil sections suitable for use as rotor - blade secti o ns are : (1) nearly zero pitching moments, (2) low drags thr c ugh - out the range of low and moderate lifts, and (3) moderate drags at high lifts . ·ith these characteristics in mind several rot r r - blade sections were derived with special emphasis on obtainLlg high lift - drag ratios . These air - foils, data for which are presented in reference 1, had max .:mwn llft - drag ratios nearly twice as large as those of the NACA 230 - series airfoils at che same Reynolds number .
RSSTRICTED 2 NACA RB No . L5K02 They showed, however , some undesirable characteristics : namely , sensitivity to rou ghne ss and abrupt adverse changes in drag, lift - curve slope, and p itching mo me nt in the vicinity of the high - lift end of the range of l ow dra g s .
The 9u~pos e of the p resent work is to extend the pr e v ious investigation and to derive additional airfoil sections desi g ned to rr:in:imize the undesirable character - istics of the previously te s ted airfoils . The tests of t hes e add itio~al airfoils were ~ad e in the Langle y two - dimens iona l low - turbulence turmel (LT'r ) .
By the use o f the p rocedure g iven in refer e n c e 2 , p rofile - dra g lo s~e s have b ee n calculated for a ty p ical helicopter rotor o pe r 8 ting in v ar io u s flight conditions and succe3sively incorporating the airfoil sections develo ped in the pr esent invest i ga tion . These calc~ l at ions oe rmit the evaluRtion of the re18tive profile - dra g losses associated with th e use of the v a rious a i r - foil sections .
SYMB O LS s e ction angle of attack c chord se c tio n dra g coeffici e nt section lift coeffici e nt design s ec tion lift coef fi cient c1,i r.1axipmIr!. lift - drag r at io ( C1,/Cd)max c mo:r:ent c0 3ffi ci en t about ae rod ynam ic cent o r m H·a . c.
moment co effi ci e nt about th e quarter - cho~d ,o int horsepower hp fr ee - stream tot a l press ure 1 0 'T cri tica l Ma ch nu. mbe r r ·· cr NACA RB No . L5K02 p local stat ic pressure on airfoil surface free-stream dynamic p ressure Reynolds number R
p\
HO - !
p ressure c oefficient S qo
/ ~
airfoil thickness ratio t/c x distan ce alon g chord froIT. l eading edge distan ce perpendicular to chord y
( cos a'\ )
tip-speed ratio \V OR The followin g symbols are used only in tables VI , VII, and VIII : forward speed V rotor disk lo~ding, pounds per square foot
W/s
f paras ite - drag area , square feet ",J R speed of axial flow through rotor disk (positive u pward ) rotor angular velocity, radians per second R rotor blade radius a angle of attack of rotor disk a solidity ; ratio of total blade area to swep t - disk area pitch angle of blede element , degrees difference between hub and tip pitch ang l e s , degrees ( pos itive when tiD angle is greater ) NACA RB No . L 5K02 D E~IVAT I ON AND TESTS In the derivation of the four airfoil sections tes t ed in this investigation several basic thIckness distribu - t i ons having a naximum thickness of 0 . 12c ",'ere empl C'l yed .
The NACA 7 - H- 12 airfoil sectio:c1 has an NACA 0012 thickness distr.Lbution , the NACA 8 - ri - 12 and 9 - H- 12 airfnil sectio n s have thickness distribut i o ns tnut have their m i nimlli~ pressure at 0 . 3c at zer l") l ift , and the l'~ACA lO - H- 12 air - foil section has a thickness distribution that has its minimlli'TI pressure at 0 . 5c at zero lift . The mean l i nes of these four airfo i l sections were designed so that s~al l pitching moments and extensive favorable p r essure grad i ents alone; the 10wer surf aces ' \,wre prC'lduced . In the des i gna - tion of these airfoils the fir:;t number is a serial num - ber , the H indic&tes that the aIrfoil has been designed for use on helicopters and other rotatlng - wi ng aircraft, and the last two digits designate the thickness in percen t of the chord . Ordinates for these airfoil sect i ons a r e given in tables I to IV .
The m ~ dels , const r ucted ~ f nahogany lam in ated i n t he chordwJ . se dlrecti o n , had .q chord t"l f 24 inches and a span of 35~ inches . In preparation for the tests the su r faces of the models were sanded in a eLordwise di r ection with rTo . 400 carborundum paper in order to obtain aerodynam i - ally smooth surfaces . Each model was tested in the Langley two - dimensional low - turbulence tunnel . This wind tunnel has a closed throat with a rectangular test sec -
tion 3 feet wide and 7} feet hIgh and is des i gned to test
models c l") mpletely spanning the wldth of the tunnel in two - dImensional fll")'.'; . The low - turbulence 18vel amounts to rmly a few huY!oredths of 1 ' Je rcent and is a hieved by the large c t"l ntraction ratio (approx . 20 t l") 1) and by the introduction rf a nwnber of fine - wire smal l - mesh turbulence - redu Lng 8creens in the w1dest part ('If the entrance cone . The maximum speed of this VJ ind tunnel is approximately 155 miles per hour .
The 11ft and pitching mOllJents were obtained fr ~ m balance readings ; the drags were o btained from measure - ments of pressures in the w&ke . The pressure - distributi on measurements were obtaine by the use of a static - pressure tube placed &t cnnvenient positi ~ ns a10ng the airfril surface . The roughness , applied all")ng the span to the N .'\'C 1,. RB No . L5 ·K02 leading edge of the models , consisted of a l - inch - wi de strip of carborundQm - covered cellulose tape . This rough - ness was suffic len t to cause transi tion virtual l y at the lea ding edze , ~ift, drag, and pitching - moment data were obt a ine d at Rcynolcs numbers of 1 . 8 an d 2 . 6 x 10 for each of t he mode l s in the smooth and rough cond i tions .
Pressur e - d i st ribution measurements we re made for each airfoil at an ang l e of attack co!'respona.ing approximate l y t o the design lirt c oe rricient .
All pitching moments were obtained in the tunnel abou t the quarter - cl'lOrd position and were tr E.ns ferred to the aerodynamic center . Only the pitching mom ents abou t the aerodyna mic center are presen te 'd . "The l i. ft J dr8.S , and pitching - momep t data ha ve been corrected ror tunnel - wall interference by f a c to!'s that inc lude correc tions due to the shape, size, and the effe ct Up t"l D th e v e loci t y measur ed by fixed static - pressure ori fi ces in the tunnel walls of the air ;'cil model mounted in the tunr.:el . F'or the airfoils of the pre sent r eport t he correctIons reduce to th e follo 's in g form , in wh ich the primed quan ti ties ref er to the v al u As measured in t he tu nnel : airfo i l N ACA 7 - U- 12 , 8 - E- 1 2 , and N ACA 10 - H- 12 Bi rfoil sections '] - :rr - 12 section c L ' 8c cL
= 0 · 977 =
c"" 0 . 9 7 "" ' a a 1. 015 ao ' 1 . O15ao '
=
=
o o c c
=
o. 992clTIC/ 4. ' = O. 993 /4 '
mc
cmc /4
m c/ L~ - - - - Cd O· 99 2 c d ' Cd O. 9 93c d ' R~SU1!:'S AND D IS CUSSION The results of t~e te sts of the NACA 7 - H- 1 2 , 8 - H- 12 , 9 - B- 12 , a~d l O- H- 12 airfoil sectlons are pre sent e d in
fi gure s 1 to 4, res pec ti ve l y . ~acb figure is divided
into t~o par ts ~ the first part p res ent s t he lift, drag , and pitchlng - ~ome nt data ~nd the s econd part presents the pressure distribution obta ined at approximately the design li f t co eff icient .
NACA HB No . L5K02 All the pi telling moment3 abont the aerody-namlc center for tne c.irfoils 1")[ this report are essentially constant up t il maximu.m lift and show no breaks in the curves as were shown for some of the airfoils of ref9rence 1 . The data show ttwt the 1-;ACA 7 - H- 12 and 5 - 3 - 12 airfoil sectlons (figs . l(a ) 2~d 2(a )) have pitctin0 moments that are nearly zero throughout an extensive lift range . The NACel. 9 - H- 12 and 10 - B- 12 airfoil sections (figs . 3(a) and 4(a) ) have s~nall negatl.ve pi tcning maments u~ to 'c:Ce stEtll . The addi tion of roughness on the leadl.rf?: edg.e of each of the airfoil sec tions ilas very 11 t tle effec t on the r.Jagni tude of the pi tching momen.ts except in tbe region of maxinlU .. '11 lift .
A comparison of the maximu.:'11 lifts of the airfoils of thls 1::1Vssti2,&tlon wi t:'"l tl10se of the NACA 0012 and 2301~ airfoil sections (1 . 3~ at a Reynol~s nw~ber of 2 . 5 x 10 for the N AC-\ 0012 rirfoil sec tiOll and 1 . 6 at a Reynolds number of 3 . 0 x 10 for the I~ AC A 23012 8 irfoil sec tion) shows th-i.t the maximum lifts for tY,e a,rfoils of the present report are sliehtly lowe~ than fo r the NACA 0012 airfoil sec tion ancl materially lO'Ner than fo r the j\J 1-\C A 23012 airfoil sec tion dt the s arne P.eyno l d s number .
T~e lift curves for the airfoil sfcti0ns investigated herein, hmvever , BY'e more rounded near the peak , eS0eciaLly for the NACA S - H- l2 and 9 - H- 12 air'foils where higr~ lifts are maint8.:ned over a consic:ierablo, rb.'.1ge of angles both in the smooth and rough conditions. In t~e rough cundi - tion the maxl.mum lii'ts of each 0_ the four airfoils reduce to essentially the sWne value (1 . 13) . Data for other air - foil sectlons at' approximately the same Reynolds number indicate a slmllar value of maxJJnum llf't for most airfoi l sectlons wltr~ leE..cing - odoe roughness . There is no measur - able change in lift dl:e to rougr-.J.1ess at the small angles of attack .
The rninlmum drags of t;'e NACA 8 - H- 12 , 9 - H- 12 , &nd 10 - 5 - 12 airfoll sections co~re3pond very closely to the minimum drags of air'.':'oi1 scctions for \I:hich th'3 thick:-less distributions lave their minirnu..""1 pressure at 0 . 5c at zero lift . The aforementioned airfoils of the present report also show a definl te ra'1!,c of lifts for low drags .
For the NACA 7 - H- 12 airfoil sect!on (having an NACA 0012 thickness dis tribution) t~1e minimwl1 dreg is sonewhat ~igher then for the other three airfoils but is less tha
thc."C ex .. _ec te d of the r AC A 23012 or 0012 airfo il sec tions
at tr~ same Reynolds number.
NACA RB ~o . L5KJ2 The drags for each of the airfoi ls of the present rsport increase rapidly at high lifts, but this increase
is much snallcr than that of the NACA 3-H - 13 . 5 airfoil
section reported in reference 1 . The addition ef rough - ness on the leading edge results in an increase in drag , for the four airfoils , similar to that found for othe r airfoil sectlons.
A S1..Ul1Inary of the im por tlmt characteristics of the NACA 7 - B- 12 , 0 - n - 12 , 9 - H- 12 , and lO - H- 12 a i rfoi l sections is 61 ven in t able V in vvhich the aerodynamic char ac ter;( ist ics are presented for a Reynolds nmnber of 2 . 6 X 10 0 .
Data for the NACA 3 - H- 13 . 5 airfoil section, as obtained from table II and figure 8 of refereLce 1 , have been included for comparison .
AlthougJ:-~ t.te floVJ conditions over aE airfoil section mounted rigidly in tl'!e V'i ind tur.nel are different from those over a section of a rotor blade in operation, the sectlon characteristics measured in t~3 ~ind tunnel , particularly for l0V! 8...ll - moderate ai.'l8::Les of attack , a re , not expected to be very different from those exhibited by the rotor - bl&de sectlon . Becrulse the greatest part of the profile - drag losses occurs while the blLdes a re operatin~ in the region of low to moderate ang::Les of attack , less accuracy is required for ca~c1.-1ations at tne hi6her angles of attack . It is thdrefore concluded that relatlve merits of rotor - blade sections may be evaluated from aL""· oil section data . The reI c:tive merit of a particular airfoil section depends 1~r3ely on the operating cond'tions and the deslsn of the rotor . In reference 2 rotor ci1aracteri3tics and flight condition s tIl. '::1.t were bel leved t y"pic al were as sumed , and we ightlng factors were obtained for each condition to permit the rotor - blade :~rof'il'3 - dra,-, loss to be calculated . Table VI preser..ts t' cse E.ts3umed rotor cha.C'8cterl.stics and flight condi tions fer the sbo.'11ple helicopter . Ey the use of the welghtl~g fsctors the profile - drag losses were calculated for a rotor tl:at success::'vely incorpor::cted the NACA 7-H - 12 , 8 - H- 12 , 9 - H- 12, and lO - H- 12 alri"oil sections in the smooth and rough ·conditions . For COmp&rlSOn , calculations were als . rrade of tne rotor - blade profile - drag l o sses of a rotor incorporating an NACA 2)012 BJr.foil sec tion in the smooth condition , Dra~ d&ta for each airf oi l were incom - plete at r:i~h angles of attack: and. a iilethod .:;iven in reference ~ 1:a3 used to extend these data . The rotnr - blade profile - dra~ losses were calcul~ted for several N AC i~ 813 No . L5K02 fligh t conditions &nd the results gl v en in t able VI I show the effect of. loading lD hoverir.B; and in forward fligh t a~Q the effect of tip - speed ratio .
A co~parison of the values given in table VI fo r the smooth airfoil :"ndicates thst the NACA 8 - F -1 2 and 9 - 11 - 12 airfoil sections hav2, in ger-eral , the le as t profile - drag losses for the flight conditions presented .
For a particular d~sk loadin or tLp - s~8ed rati ~ a choice of one of the ot~nr airfoil sections miz~t be ir..dicated .
At very h:"gh pitch settings ~n hoverlng flight (condi - tion 3 ) and the hig~ tip - speed ratio (~ = 0 . 3 , condi - tion 6 ) the NACA 23012 airfoil section has the leas t profile - drag losses . In these conditions , sections of the rotor are operating at high ant.,les of attack where t:le NACA 23012 ai'''foi1 sectlon has low",r drags than the all"foils presented herein , wllich accoll...'1ts for t he l ower profile - drag losses . For the dirfolls of this repor t in the rough condi tion , the vb.hl.es of prof lle - drag loss differ very little . In eith~r the smooth or rough sur - face condition , the airfoil sectl ~ ns having , i n general , the least pl'o';:'ile - arag losses for the ol.erating condi - tIons presented herein are the NAc:J 8 - 1I -1 2 and 9 - H- 12 sections . Preference , however , would probably be given the lIT ACA 8 - H- 12 airfoil sec tion bee Cl')Se i t has a sm& ll er Y pi tchin2:>mO ,ltmt coeffie ient BCOut tile aer·odynamic center (0 . 005 ) tf.tan tl:e NACA 9 - H- 12 o.irfoil section ( - 0 . 012 ).
In order t C' provide a comparison of the calculations of rotor - blade profl1e - dra~ loss glven in this report with simila.r c alculbtions for the li:.OSt promising &lrf011 sect::"o~ of reference 1 , det" for the }TA CA 8 - H- 12 and 3 - H- 13 . 5 &1rf011 sections in the smooth condition are pre sented in t able VI II . 1hs v 81 ut? s for the N AC A 3- H-13. 5 airfoil section were o':Jtainpd from table I of reference 1.
The ~TAC.!~ 3 - H- 13 . 5 airfoil sectlon h&d the larger value of
:naXlITlUITl lift - crE~8 rCltio (see table V ), but the profile - drag losses fer the NACA 8 - H- 12 airfoil secti on are less in every condItion presented except fo~ the disk loadings 0.:.' ~ . 33 and 2 . 5 in hoveri:1g fli[}lt (conditions 2 and 4).
The rraznitude of the lift - drag ratio in itself therefore is not a reliable indication as to the re l ative merit of airfoil sections intended for u~e in rotor blades . The NACA 8 - H- 12 airfoil s'3ction show's c.. la:c'ge r eduction in profile - dra2: loss as cOT?lpared witn that of the NAC A 3-H- 13.5 airfoil see"GJ.on at t.he l::.i:shest piteL setting in hovering flight (condition .3) . L;;,rge red:J.ctions are als o shown NACA RB Nn . L5K02 at the high disk l oading in c r uislng flight (c o ndit i ~ n 8)
and at high speed (condit i on 6). These reductions were
made possible by the l ower drags at the high angles of attack .
The weight i ng curves of reference 2 p r ovide not nly a means for the calcu l ati ~ n of the rot or - blade profile - drag loss but als o a direct indlca~ion as tn the relative imp ('I rtance of regi l'"l ns of the drag of a r ~ tor airfoil section . For the assumed conditions these weighting curves indicate that for h ~ vering wlth a rotor - blade
pitch angle h f approximately 6° to 100 and f or low fo r -
ward speeds , the regi ~ n of sectl nn drag coeffi ients c o rresp nding to cL ~ 0 . 2 t n 0 . 6 has the greatest effect upon the magnitude of the rotor profile - drag l ~ ss . For high disk loadings in hovering and low f ~ rward speeds and for high forward speeds at norm a l or high disk l oadings , the same re ion of drags still has the greatest effect , but the drags at high lifts are als ~ prom i nent in affactmg the magnitude of the pr n file - drag l r ss .
CONCLUDIKG RSMA.RKS The NACA 7 - H ~ 12 , 8- H- 12 , 9 - H- 12 , and 10- E- 12 Rirfoi l secti ns , derived for use as rot or - blade sectinns f r tary - wing air raft , have been tested in the Langley tw r- dimensinnal I nv - turbulence tunnel . These airf ~ il sections had omparatively l e ' dregs in the range ~ f 1 w and moderate lifts and nearly cnnstant pitching moments up t n maximum lift and the aer o dynamic characterist i cs were not unduly sensitive to rougrmess . The NACA 2 - H- 12 , 9 - H- 12 , and 10 - H- 12 airfoil sections had a definite range nf lifts for 1 0 ,- , drags and had !:! inimum drags correspond.L.'1.g clos e ly t " the minimum r.rags of 8.irf o il sections fl'"lr which tbe thickness distributi lt IlS have their minimlL.11 pressure at O. 5c at zer n lift . r'r m a omparisnn of the calculated profile - drab l l'"l sses of & typical heliccpter r tor successively incorporatin g the airfoils of this report, the NACA 8- H- 12 and 9 - H- 12 air-foil sections hA.d , in general , the least l ~ sses in the op~ rating onditi e ns presented . The NACA 8 - H- 12 airfnil sActlon , having the "maIler pitching moments , would ~r0bab ly receive prefer - ence as a rotor - blade sec Ion . corr.parec with the
NACA 3- H- 13 . 5 airfeil section report8d in ~ITACA CB
IJo . 3I13 , the NACA 8 - H- 12 airfoil sho1Ncd smaller profile - drag l ~ sses in nearly every operating conditi on presented .
10 N AC A RB -No , L5K02 Frnm aer ", dynarnic consideratinns the NACA 8 - 8: - 12 and 9 - H- 12 airfoil sections Rppeared nore promising for use as rotor - blade sections than any other airfoils thus far tested at the NACA laboratories .
Langley Memorial Aeronautical Laboratory Nati0na1 .r\.dvisory CI")ITuni ttee for I eronautics Langley Field, Va , R~?E;{ENC~S 1. Tetervin , (eal : Tests in th3 NACA Two - Dimensinnal Low - Turbulence Tunnel of Airfoil Sectinns ~ esigne i to Have Small Pitching I\~oments and High Lift - Dl'ag qatios . N AC,~ C:3 ~~o . 3 I13, 194- 3 .
2 , GU8tafS t"l l1 , F . B .: :2ffect on Helicopter Ferformance of M0difi ations in Profile - Drag Characteristics of R, ") tr)r - Blade Airfoil .sections. ACA ACR _ro . L!j.E05 , 1944 .
NACA HB N'o. L5K02 TABLE 1.- ORDINATES FOR TABLE 11.- ORDINATES FOR NACA 7-H-12 AIRFOIL SECTION NACA 8- H-12 AIRFOIL SECTION [ Sta tions and ordinates given in ~tations and ordinates given in percent of airfoil chor~ percent of airfoil chord] Upper Surfa ce Lower Surface Upper Surfa ce Lower Surface St ation Ordinate St ation Ordinate Stati on Ordinate Station Ordinate 0 0 0 0 0 a 0 . 627 1.8 -1. 232 7 1.229 -.81 9' .14 Z . ~3 A 1. 802 3.19 -1. 576 1. 520 1. 2
~ : ~~ -· 946
. ~5 ~ .2 96 5·01 -1. 952 . 04 2 . 006 1. 6 96 -1 .128 ~.704 .853 6.127 1. 80 -2.173 3.020 2·941 -L41~ '~ f 10·5 9
4 4.312
6.97~ -2.3~ -1·73 ~.5Z6
t· 24
~:~6§ 8.13 15·392 O .0 6 -2' 14 -1.920 g.3B 8.816 20 . 221
t -2. 52
~ .77 9 '4 10·573 - 2.0
~ . 27 .26~ 9 . 93 25 · 068 -2·73 9 -2. 2 2 9.J.4~ 97 15·503
4 Z· 62
2O ·t
~0.059 -2.~ 96 . 605
29.9 t ~. 07
20.~3 - 2.~51
A·
·737 0 . l -2. 5~ 9 .243 25. b - 2 . 17 .7 ~ o 31 ~~:l84 ~.712 -2.85 29·9 9 30.031 t • 9. ,33 - 2 .4 55, 0.315 .326 ~9 . 685 -2.7~ 8 174 34.826 5 9. 32 -2.4 90
4 •
4.7 - 2 . 5 0 0.292 ZO . 2~ §.03 49
-2. t
§ :~~ 0.1 -2.1 93 2O
4 45.360
3.0 4 ~:Z~ -2.47
90 . 0
·t O
1. to ~~.954 50 .3 -1.4~0 - 2 .436 l' 66 ~.610 95·003 • 65 9 . 997 -. 9 9 A .613 -2· 377 ~5.3~ .~5 100.000 0 100.000 0·3 -2 .290
4' 6
t4J~~ 65 .311 .850 -2.178 L.B. radius: 1.58 3.838 -2.034 70.2~ Slope or radius through L.E.I tpo 0.443 2.838 7 • 16 -1.860 Z5.
0.11 ~.882 1.8~6 -1. ~M 85.060 1.0 -1.
.9~ 90.0 16 .343 89·9 -1. 051 94 . 995 -.11 9 95·005 -. 629 100.000 0 100.000 0 L.B. radius: 1.325 Slope or radius through L.E.: 0.344 TABLE III. - ORDINATES FOR TABLE IV.- ORDINAlES FOR NACA 9- H-12 AIRFOIL SECTION NACA 10-H-12 AIRFOIL SE CTION [ Stati ons and ordinates given in [ Stati ons and ordinates given 1n percent of airfoil chord] percent of airfoil chord] Upper Surface Lower Surface Upper Su rt'ace Lower Surface S tation Ordinate St ation Ordinate Stati on Ordinate Station Ordinate 0 0 0 0 0 0 0 .117 1.238 . 883 -.7 88 . 234 1.062 ·766 -.~06 1·537 .3 ~ 1.17~ -. 907 .451 -. 17 .7 2 . 037 1.3 ~ 1.~9 1. 73 -1. 073 .911 1.7 1. 9 -.966 1.933 003 06 2 3. 7 -1.3 2.108 2 · 5 -1.158 4.
2' ZZ2 A .427 ~ 'A 69 4·551 -1.414 -1. 5~ 5· 9 ~'l20 • 57 ~ :U~ 7.028 2 • 95 -1. 579 ~ . ~O -1.~ 7',7 10.630 -1.
~.521 1O.p
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1 "r -1. 9 7
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9. 4~ ~.~ 6
a • -2.0l 15
5 - 2 .12 ~ 9. *5 4 .
0. 277
t - 2 . 0
9 . ~5 9.8 0 9 9 -2.1 ez·~23 0 . 03~
4 • ge
45.353 8. 59 • 47 - 2 . 039 1O 5 ·1 9 .7 89 44.89 - 2. Z 50.390 8.107 0 10 ~ . 610 99 -1. $ 9 .~0 ~ . 690 - 2 .0 ~
5 3 . 607 5 '4
Z·2 8. 98 -2.0
i • g3 -1. Z i5 . 76
g -1. 1 .2p
0 .3 ~ ~4 · 631 ',24 0.,18
~~. 82 -2'0§t l · p6 65·3 . 676 ,.2 1 • 22 -1. ~65 65· 97 .503 -2.0 70 . 26 .22~ -1. 45 70.~3 -2.031 ~.405 tP4 64·~67 2OO 3.19 7 • 00 Z5. -1. 500 .1 Z5 . 1 5 -1.9 Z 0.134 2.212 -1.3 26 A 0 .236 Z~·866 2·90 -1.7 5 85.07 1.314 · 927 -1.114
~ : ~~ 1. 739 -1. 511
g 85 .1EA 90.02 .550 89.974 -. 844 90.0 2 ·723 89 . 95 -1. Itt 95 . 000 . 0bB 95·000 -·502 -. 002 94.999 95·001 -·7 100.000 100.000 0 100.000 0 100.000 0 L.E. radius; 1.3 L.E. radius; 1.000 or radius through L.E.: Slope 0.378 Slope of radius through L.E.I 0·301 IlAnONAl ADVISORY COMMITTEE FOR AERONAUTICS NACA RB No. L5K02 ~ABLE v.~ AIRFOIL SECTION CHARACTERISTICS
[R = 2.6 x 10 ]
I NACA airfoil
I
I
section 3-H-13·5 10-H-12 (refer- 7-H-12 8-H-12 9-H-12 Section ence 1) characteristics 106 163 (CL/Cd) 152 135 149 max
0.003 i -0.012 -0.022
------ 0.005
em
a.c.
0.0050 I 0.0046 o . 00~.6 0.0043
0.0055 cdmin
I
0.25 0·30 0·38 0·39 Low-drag range to to to to ------ 0.88 0.76 0·91 0·93 1.20 smooth 1.26 1.26 1.34 1·30 cz,max
Rough 1.10 1.13 1.12 1.14 --------
at 0.601 0.619 0.56 0.569 0.569 Mcr CLi
I
I
0.60 (approx. ) 0.60 0.46 0.42 0·57 CLi 0.1208 t/c at 0.25c 0.119 0.117 0.117 0.108
o-
0.261 0.250
xic 0.250
0.278 10 . 267 a.c. position I 0.021 0.021 y/c 0.020 10.025
1 --------
i
i
NATIO N AL ADVIS ORY CO M MITT EE FO R AERO NA 'G TIC S N A CA SB ?, o , L5K02 TABLE VI, - FLIG HT CO:IDITI ONS AND ASSUMED CHARACTERI STICS
OF THE SA.MPLE HELICOPTE R OF REFEREHCE 2
- Roto r diam ., 40 ft; tip speed, 400 f'ps ; ] [ g ro ss weight for 'N /S of 2,5, 31L ~ 0 Ibs
I -c o -n- d- i - ~~-o--n-' --~ - -~N- /-s-r---a--r! -e- -:- ,-8--- - ' A. f
r-- I ,- -- I----- - --+--~
i 1 ° 1. 55 ° ',0 I ~ I 7 - - - - - - - I 15
I 2 II 3.33 I 13
l ~ I i 5 . 1,2 I I I i 19
~ ,t I 2 . 5 I ,i 119 °'
j
- 0.0385
5 1° .22.5
I I,
f I 3 2 c: I ' I I 11
O - . 0695
• . J I i II
- . 0319
' '. : 1 1. 9 \ I II 7
- . 0~_69 I
3.1 \v I I,' 11
I . 2 2 . 5 0. 10 , 7 -.0 350
I IV
1 0
.3 2·5 . 07 1-8 ,alO · 5
-. 0680 ,1
, I
I
11 -. 0435 i bO
·3 2·5 . 07 i - 81 a8. 5
I
, '-----' '------'--- aMeasured at 0 .75 R .
bRotor alone .
NATIOHAL ADVIS ORY COM M ITTEE FOR AERONAUTICS t-' z :t:> CI ~ o f\) +=- f; ~ t;j Z Ll V1
I
I
) ) ght of flight of of 0 fli i ing at r r loading ward t i p- speed l oading
Effect ~ Effect r
fo S (hove
-----1
(: jl
' I.lc
I I
~T} . 1 ·6 .7
·5 .7 VARIOU L
24 21.7~ffect 25
42 23 25 29·2 SORY FOR
5 9
or: 9.
' . .
VI 1.
1.7 ---~-I AERONAUTICS -
~~ 41. 7 52
37 ·9 L~ /'5 AD -----
LOSS i
!
~-~
FOR ~;--
h c)~35 ~
. ·3 , ' · 6 ·3 . 0
·5 , 7.
DRAG 13 14
17· 8 14 1
40 23 -
I i
HELICOPTER
-t---- NATIONAL
=- --- LE ~ I . 2 .2 1 · 8 · 5 .4!
_.J..~ a~- 66. 1 33 39 · 59 COlvil.ViITTEE 35 39 .3 ' r 13 PROF --
d 138 I I i !
I
SAMPLE - ,
4 9
~ ·9 . .0 .J .8
7.
7· THE
r-_ ~
57 3 28 BL.ADE of~le _
! ! 19 ·h
~ r OF p - 11
R- 61
3~ 4120 7
O : . . .
1.
7.7 :JS ;; 57.
1+1.4 ROT 112
1 135 14 ! 13 1
I
1.U'" OF'
8 7
41~~~1 5
. .5 . ·3 ---r
. . 6 1. 2 6 .
18 16 17 21. 2 28 CONDITI O i I 1
I
Rotor - bl ade
:;l14 ~'1
. '
' .0 GHT I
30 '33.71
31.9 34 . 8 37 58 j37.4,2 ! ,73·4 I COMPARISON FL .- ·9 .0 .8 .0 .
.6 7. 2 25 22 24 22 2L~ 42 ·3 30 VII 1 I' , I :
-r-
14~1l6
T1 .
. 2 .:.I
TABLE 2.5 2·5 J 8 2 0 . 2
.)
= = _ I)
v "= f.L
__ /S
1.'_ ';'i i
I
~
----------
551 331 42 . .21 ·3 table 1. 1.9 '3 5 . 0 2·5 3 · 1
Condi t i ons = = =
(see ~
-----t----·i /S
----
,11/8 iV\i
!
~I 7I 8
6L 5
- --
: 1
l2J ! 151 i II I : L
TABLE VIII . - C8J<PARIS01 '~ OF ROTOR - BLADE PROFILE - DRAG LOSS OF THE NACA 3 - H -13.5 AND 8 - H- 12 AIRFOI L SECTIONS IN THE S_ OOTH CONDIT I ON F'OR VARIOUS FLIGH T CONDITIONS OF THE SAMPLE HELICOPTER
- - r
I Condi tions
(see table VI) , - - - - ~L
~tWI
1. 551
2 !
I
3 ·33 !
I i
I
! 3!
5 · 42 _i
I.L = 0
=
I'/S
. 2 I I
I : I
·3
I I
I
i 7 I,vls =
1.9 I !-l =
I !
I
I 5: 2·5 I I
I 8!
1·-+ -~+-
. C) I
(J =
0.071 !-l = ! ' I I I . 2 26 . 1 . 10 I
I
-
8 = QO I
I -
61 I.L 51 -1- . 5
0 . 1 I
I
I
8° I
42.4
I 10 I
. 31
1----1 -
=-~ , 10 I
f -
42 .4
= !-l 3
, 0 . 1
I I 11 i .3 , 35·9
I
I
NATIONAL ADVISORY COI'1MI1'TEE FOR AERONAUTICS 'XJ 1-" ~ P> CJQ t:O z o r- (Jl :;:>;: o 1'J z ~ o ~ ::0 l
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Fi g . lb NACA RB No. L5K02
2.4 ~--+---~--~---r---+---+--~---;---;--~
2.0 r---;----+--~----T_--_r--_+----~--+_--_r--~
/
"<- "" I r-- 0 Upper surface L~ LV r 0 Lower surface
f
1.6 S 1.2
lL-
.8 7
NATIONAL ADVISORY COMMITTEE FOA AERONAUTICS
I
o
o .2
.4 .6 .8 1.0
x/c {b) Pressure distribution for design lift coefficient,
c = 0.42; R = 2.6 x 10
L i Figure 1.- Concluded.
.
Pl "".l ..... (\) ~ » ~ tIl z o L' (Jl o (\) OQ z :> (") "
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Fi g . 2b NACA RB No . L5K02
2.4
h 0t--.
/
2.0
(~
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v-;=
I Lower surface
EJ
\
1.6
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f
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.4
NATIONAL ADVISORY COMMITTEE FOA AERONAUTICS
o
o .2 .6 .8 1.0
.4
x/c (b) Pressure distribution for d~sign lift coefficient,
c = 0 .57; R = 2.6 x 10 •
L i Figure 2.- Concluded.
~ ..... CJ.J Pl z :x:- l) :x:- AJ to z o l' CJ1 ::>:: o (\) oq d; chor 24-inch section, data alrrol1 9-H-12 pitch1ng~ament NACA and the or drag, lirt, Section characteristics (a) Aerodynamic 336.
;.- test LTT Figure Fig. 3b NACA RE No. L5K02
h
~ ~ 2.0
V
~)
/
o Upper surface
(
8 Lower surface
~ ~
1.6 I ~
\
r~
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(
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NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS I o o .2 .6 .8 1.0
.4
x/c (b) pressu~e dls~ribu~lon for dgsign lift coefficient, c - 0.60, R - 2.6 x 10 • Figure 3.- Concluded.
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1 l
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",><!7ff<" k: I'T T , ., 6 COMMITTEE
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,'" :--- "',1 I~I·-'b f 1Itf T" 1: 'I" _. .• position airfoil :::~ 2 I
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"f X/C " •
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I40 >c ll.~ 0.261 ;' •• ~ =
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l 1 I Lei I 1·1*I;li;J;·Tt+T"Vl+~l+lri IJ . LJEi:kI_·Ld=hL.L.,J. pitchlng_OII1ent ..
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·l ··I '1\1: :' .:. characteristics I ~ I : :". (a)
I I i\n "i0:'" '-1=1-
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Fig. 4b NACA RB No. L5K02 2 u O~--~--~ ----+---~---4-- -- +---~--~~--+---~
/'~ ' 1\ ~ 0 Upper surface
/' I '< r El Lower surface
.
1.6~~- . ~--~ --~--4---+-~j\+( ~ ---'---'---'--4
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.~
,Sl.(
. , .4~~---+--~--+---~~---r--~--+-~ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
o
o .2 .6 .8 1.0
.4
x/c (b) Pressure distribution for d~slgn lift coefficient, 6 •
c = 0.46; R = 2.6 x 10
L i Fi g ure 4.- Concluded.