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t D OCUMENT NUMBER SER - 7 20 11 S e quen c e Nu=b e r A011 , DD Form 1 4 23 PREPARED UND ER Contr a ct Da t a Requ _ -men t s, C o n tr act N A S I-13 000 DOCUMENT D ATE D e c e m be r 4 , 1 9 7_ . . . ,. PERIO D COVERED M arc h 197 ]; - No v e mb e r 1_ ; 74 , .

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S-T2 ( eSRA ) N ASI- 1 3 000

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j ' • SUMMARY ,_ ' ' : Wi nd t u nn e l t est s of t h e s ixt h s cale m o d el of t h e S i ko rs k y / NA S A / 5 Army R o t o r Syste m s Research Aircraft (RSRA) were conducted i n the U n ited _ Aircraft Research L a boratories (UARL) l a rge subs o nic wind tu nn el d u ring _ _ ; !

the perio ds o f M a y 22 through Ju n e 2 9, 19 7 k , and Augu s t 1 2 thr o ug h , _ 0ctober k , 1 9 74. The o bjectives of the s e tests were to deter m ine, i n I _o_w_ r d flig h t, aero d y na mic ,: h a racteristics and fusel a ge _ u rface pressure -._ distributions in b o th the h e licopter an d f u ll c o m p ou n d c o nflg u rations.

" " Pa rticu l a r a tt en tio n w a s giv en to wing i n board f a iri n g co n figuratio n , powere d TF- 3 4 cant an d incidence angles, TF- 3 h support fairing sh a pe, ' ,, a n d empen na ge c o nfiguration. Neither a main nor a tail rotor was tested. :.

_ A This re p ort d o c um ents test results an d is supplemented by UARL _ _ Report N-4 3 2 3 77-I, Reference I, and UARL Report N-4 3 2409-!, Reference 2. _ ; These wind t unn el tests resulted in an R S RA configuration that _ " differed fro m t h e March 19 7_ desig n pri m aa o i l y in t h e a r eas of t h e wi ng- . _ fuselage Ju n ction and the e m p e nnage. Wi n g fusel a ge s e als w e re incorporate_ ;_ a t t he J un ctio n wit h t he side of t h e fu s e la ge and a ft of t h e wi n g c en ter box. Numerous tail it e rations r e s u lted in a c o mpound empe nn age c o nfigu- " _ ration with the vertical tail e xtended to wat e rline 3 60 , a rectangul a r : , .... # planform, 9 8.1 squar e foot, l o wer horizontal tail, and a 17. 2 squ a re f o ot ,) upper h o rizontal tail. _e helicopter configuration used the sam e v e rtical . _ t a i l wi_h a 35.4 squ a re f oo t u pper ho riz on tal t a i l and no lower t a i l . : " -.

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: ; A® }_ TABLE OF C ONTENTS : Volu m e _" r ; : S_mary i _ : List of Tabl e s ix _ _ Li s t of Figur e s x . :_ ! List of Symb o l s xx_ . : .

_,_ Intr o d u cti o n I ' _.

De s cripti o n o f T e st Facility 2 .!

D e scripti o n of Wind Tunn e l M o d e l 3 . ; [ L Balanc e D a ta 6 _ Te s t Proc e dure 6 Data Pr e se rtation 7 r _ e Balance Data Precisi o n 8 _ Aer o dynamic Tare and Interference 8 :_ , R e yn o lds Number Effects 9 _ _= ,._ H e licopter Build u p and Stability-Final Co n figur a tion 9 ._'_ _ ,. C omp o und Buildup a nd St a bility-Final Con figuration 12 Compound Nacelle Configuration Selection 18 ?

H e l i c o pter and C om pound E m pennage Down w a sh a n d Si_ e wash 18 k H e lic o pter and Co m pound Emp e nn a ge Se lection 2 0 '_ .'

Fl o w Visuali z a ti o n 21 , :, S urface Pr ess ur e Data 26 ._ T es t Pr o c e dure 26 Data Pr e s en tation 2 6 ': .

Emp enn age T o t a l Pre s sure Da t a 27 ':'° Empe n nage Anem ome t e r Data 28 ._ Instrumentation 28 "_ T e st Pr o c e dur e 2 9 Data Analysis 30 ,_;.

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! \ TABLE OF CONTENTS(Cont 'd. ) C o nclusi o ns and Rec o mm e ndati o ns 3 _ _ Re f e r e nc es 35 T ab l e s 37 _ '. Figures 159 -4 V o lume II i t ! Tabulated B_s.nce Data 1 ,i i viii PA @"='_ J , _ S A 2 _ ) RFV D

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LIST OF TABLES ;_ / T a bl e % / Numb e r Titl______e :_ I Emp e nnage C o nfigur a tions 37 _ II Tail Com p on ent I de n% i fication I nde x 3 9 _ III T a il C o mponent Matrix h O -- IV Pressure Tap Loc a tion s hl , _ V Wind Tunnel Run Log - Ph ase I 45 :, _" VI Wind T u nn el Run Log - Ph a s e II 6h i_ !

! Vl l R un N umb e r / F igure Nu m ber I n dex 9 0 I _!

i ' _ VIII Static Data Accuracy 94 f, _, IX Eff e ct o f TF-S4 Nacelle Incidence Angl e 95 X Effect of TF-Bh Nac e lle Cant Angle 95 _,_ XI Helicopter Surface Pressure Data 96 _" - I Effect of Angle Attack,_= 0 d e g _-_4) XII Helicopter Surface Pressure Data 106 Effect of Angle of Yaw, eL= 0 d e g XIII Helic o pter Surface Pr es sur e D a ta • 115 Eff e ct o f Angl e o f Yaw,_= -!0 d e g XIV H e licopter Surfac e Pr e ssur e Data 124 Effect of Angle Yaw,@ t - + I0 deg X V C om p o un d Surface Pre ssur e Data, U np ow e re d 1 33 E ffect o f Angle o f Att a ck, _ = 0 d e g XV I Compound Surface Pr e s su r e Data, Un pow e red 143 E % ct of Ang l e of Yaw,_ = 0 deg XVII Co m poun d S urface Press u re Data, Powere d 152 Effect of Angle of Attack,_= 0 de g i x PAGE _. SA29 R E VO ( t

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i" LIST OF FIGURES i Figure Number Title P age r_ 1 R SRA Sixth Sc a le Wind Tunnel Model ii j , , _ Compou n d Configu r ation : . :.

j_ ?: 2 HSRA Sixth Scale Wind Tu n nel M od el iii ; _ r H e licopter C o nfiguration - Phase I _ : - _ , ,. 3 RSRA General A r r angement v 4 UARL L a rge Subsonic Wind Tunnel 159 _ 5 Model Inst a ll a tion - Tail Al o n e 16 0 6 Basic C o nfigurations 161 7 Powered Nacelle Configurations 171 ; , / ' 8 Wing C o nfigurations 183 9 Empennage Co nfigurations 187 " _ i0 S e lected C o nfiguration Drawings 2 07 _ : a-d H o rizontal Stabilizers _ e Vertical Stabllizers :_ ! , f G e n e ral ii S ign Co nventions 213 / : / a. W i nd A x is C oo rdin a te S yste m : _ b. C o ntrol S ystem Sign Co nv en tion 12 D a t a Rep ea tibility-Configur a tion F P BNp_W7 T I o B T 2 15 , a. Lift b . Dr a g c. Pitching Moment . _._ d. Yaw i ng Mam_nt _: e . Rolling Mo m e nt ' f. Si de F o rce _ 13 Dat a R e peatibility- Con figuratio,_ FPBNp_W7T 2 221 i_ a. L ift f b. Dr a g , .

c. Pitching Mo m en t d . Ya wi n g M omen t e. R ol l ing Mome n t f. S i de For c e 14 Fu sela g e A er o d yna mi c Ta r e and I n terference C orr e ctions 2 9 7 1 _ 15 T ail Al on e A e r od y n amic Tare Co r re cti o n s 228 I _i / X PAGE " : SA 2 _ REV O _ I | I ' '_-! , t i U ' Sikor s ky I:li r cra f t R P O RT N O, SER-72011 .

_ 4 TABLE OF FIGURES (Cont'd.)

Figure Number Title _ ) _ , ,L' 16 Effect of Reynolds Number 229 a. Helicopter Configurations _ b. Compound Configurations Helicopter _ _" 17 Effect of Helicopter CoL1ponent Buildup 231 / a. Lift vs @ L b. Drag vs eL , c. Pitching M o ment vs @L ; i 18 Effect of H e licopter Component Buildup 2 3_ : !

i a. Drag vs _ _ : b . Yawing Moment vs _ c. Rolling Moment vs _ d. Side Force v s _ 19 Effect of Angle of Attack on Directional Stability, 238 i Ya w ing Mo m ent 20 Effect of A_gle o f Attack on Lateral S tability 239 a. Rolling Moment t b. Side Force 21 Effect of H o rizont a l Stabilizer Incidence-Helicopter 2 41 ( a. Lift b. Drag ' c. Pitching Mom en t ; _ 22 Effect of Vertical Stabilizer Incid e nce-Helicopter 2hh _ " a. Yawing Moment _ B. R o lling Mo me nt c c. Side Force 23 Effect cf Rudder D e flectlon-Helicopter 2_7 : , a . Y aw ing M o ment , / b. Roll i ng Moment c. Side F o rce : 2_ Effect of Angle of Yaw on Helicopter Tall Alone 2 50 a. L i ft b. D ra g c . Pitching Mom e nt PAGE SA 29 REV O ; JlII I |L -- -- L U

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i TABLE OF FIGURES (Cont 'd.) i : l Figure I Number Title Page_ - _ 25 Effect o f Angle of Attack on Helicopter V ail Al o n e 253 Mome n t I a. Yawing b. Rol l ing M om ent : c. Side Force _ .. 26 Eff e ct o f Rudder Deflection-Helicopter Tail Alon e 256 a . Drag b. Yawing Moment ; c. Sid e F o rce L 27 Ru dde r C o ntr o l Power -H elic o pter T a il Alon e 2 6 0 _ 28 Eff e ct of Sp ee d Brak es on Helicopter Tail Alone, 261 / Dr a g ve r su s @L _' 29 Effect of Sp ee d Bra ke _ on Helicopter Tail Alon e , 262 _ Drag versus . : _ C omloou nd ._ 30 Nacelle Lift and Thrust versus RPM 263 a . Nacelle Thrust vs _ P M : b. Nacelle Lift and Thrust Parameters vs RPM 31 Effect of Component Buildup-Tail Off i = 0 deg 265 : a . Lift v s b. Drag v s o ( c. Pitc h i n g M ome nt vs _ 32 Effect o f Co m pound Com p o nent B uildup-Tail Off, 268 t i = 0 deg : _ W _\. a. D rag vs b. Y a wi n g Mom e nt v s ' : c. R o ll ing Mome n t vs %v d . Si d e F o rce vs 33 Effect o f Comp on e n t Buildup - Tail On, i = 0 d e g 272 w _ a. L ift vs _. b. Drag v s c. Pitchi n g Moment vs o_ xii PAGE

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Figur e N umb e r Title Page i_ 3 4 Effect o f C om poun d C o mp on e n t B uild u p -Ta il On , 275 iw = 0 d e g a . Dr a g vs b. Y a wing Moment vs _ • _ c . R o ll in g M o ment v s _ d. Side F o rce vs _ ; 35 Effect of Tail Ro tor Hub, q = 4 0 psf 279 a. Yawi n g Mom e nt . _ b. Rolling Mom e nt c. Sid e F o rce 3 6 Effect o f Wi n g I ncidence - Ba se line C om p o und 282 a. Lift b. Drag c. Pitching Mo me nt _ _ 37 Effect o f Wing Incidence-Tail Off, _F = 0 deg 285 ' _ ' a. Lift _ b. Dr a g " c . Pitching M o m e nt i 38 Effect of Wing Incidence-Tail Off, _F = 0 deg 288 a. Yawing Moment b. Rolling Mcment c. Side Force 3 9 Effect o f Wing In cidence-Tail O n, _F = 0 deg 291 a. Lift vs _ b. Drag vs c_ c. P i tching l_mment vs e_ d. Y a wipg M o ment v s _ h O E ffect o f Wing In c idence o n Directional Sta b ility 2 9 5 a. Yawing Moment vsC &) _= 5 deg b . aYawi n 6 M om e n t ( ' "/ _ VS @_ 4 1 Effect o f Wi n g In ci d e n ce-Tai l On , _F = 0 d e g 2 9 7 a. Yawing Moment vs b. Rolling Moment vs c. Side Force ¢s _ J xiii PAGE _ _ SAZ _ REV 0 ! ,

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TABLE OF FIGL R ES ( Con t'd.) 1 . F igure N umb e r T itl_.__e Page 42 Effect of Angle of Attack on Directional S_ability, 300 " = -9 de g !

• _ iw _ : 4 3 Effect o f Angl e of Att a ck o n D ir e cti on al S t a bi l ity, 3 01 "' i = 0 deg w ,j " ' 4h E ffect of Angle o f Yaw o n Directional S tabi l ity - 30 2 = 0 de E, 8 = 3 0 d e g " Ta i l On, i w F ; 45 Eff e ct of Nacelle Thr u st - Tail O ff, i = - 9 d e g 303 ; w .,_ a. L ift ." b. Drag c. Pitching Mom e nt . 4 6 Effect o f N a celle Thrust-T a il O ff, i = 0 de g 3 06 ._ w , a _ Lift b Drag c . Pitc h i n g Moment 47 E ffect o f Nacelle Thrust - Tai l Off, iw = 7.5 d e g 30 9 o_ a. L ift , b. Dr a g • c. Pitc h ing M o ment h8 E ffect o f Nacelle Thru s t-T a i l O ff, i = 15 d e g 312 w ' a . L ift _ b. Drag • c. Pitching Mome n t _9 Effect o f Nacel l e Thru s t - T a i l Off, i = 1 5 d e g, 315 _F = 30 d e e w , : . a. L ift _ . b. Drag c. Pitching M o me n t 5 0 E ffect o f N a cel l e Th ru s t-Tail O ff, i = -9 d e g 3 18 w a. Yawing Moment O. Rolling Moment c. Side Force

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Numb e r Titl e Page _ 51 Effect o f Nacelle Thrust-Tail Off, i = 0 d e g 321 w a. Yawing M o ment b. Rolling Moment i c. Side F o rce '_ -- 5 2 Effect o f Nacell e Thr us t - Tai l O ff, iw = 7.5 d e g 3 2 4 '4, _ a . Y a wi n g Mom e nt b. Rol l ing M ome nt c. Side F o rce 53 Effect o f Nacelle Thrust-Tail Off, iw 15 d e g 327 , a. Yawing Mom e nt b. Rolling Moment t c. Side F o rce 54 Effect of Nacelle Thrust-Tail Off, i = 15 d e g, 330 _F = 30 deg w a. Yawing Mo m ent b R o l l ing M ome nt c. Side F o rce 55 Effect of Nacelle Thrust-Tail On, iw = -9 deg 333 a. Lift b. Drag c. Pitching Moment 56 Eff e ct o f Nacelle Thru s t-Tail On, iw = 0 d e g 336 a. Lift b. Drag c. Pitching Moment 57 Effect of Nacelle Thrust-Tail On, iw = 7.5 deg 339 a. Lift b. Drag c. Pitching M o ment 58 Effect o f Nacelle Thrust - Tai l On , iw = 15 d e g 342 a. Lift b. Drag c. Pitching M o ment XV m PAGE SA 2g _ EV 0 I _ . ] . .

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t : ) ' TABL E OF FIGURES ( C ont 'd. ) Figure _ / N u mber Title FaK e ; 59 Effect o f Nacelle Thru s t-Tail On, i = 15 deg, 345 F = 30 de g w a. Lift _ b. Drag c. Pitching Moment r " 60 Effect of Flap D e flection-Tail Off, i = 0 d e g 348 w a. Lift _ b. Drag c. Pitching M o ment ' 61 Effect of Flap D e flection on Lift-Drag Ratio 351 " a. L / D vs eL _ b. L / D v s C L 6 2 Effect o f Flap D e flection-Tail O n, iw = 7.5 d eg , '_ 353 : _ a. Lift I b. Drag _ c. Pitc h ing M ome nt F _ ' 63 Effect o f Wing Incidence on Longitudinal Stability , 356 ! ' - _' & F = 30 deg _ : 64 Effect o f Angle o f Attack o n D!rec+i o nal S t a bility, 59[ i = 15 : 30 a e g deg, W : 65 E ffect o f Ang l e o f Attack o n Lateral S tability, 3 58 !

iw = 15 d e g, ___F = 3 0 deg i; < a. R oll ing Mome n t _ b. Side F o rce S [ 66 Effect of Em pe nnage on Ailer o n Co ntr ol 36 0 [ L_ a. R o lli n g M om ent vs _, i = 0 d e g I _ b. Rolling Mom e nt vs _, iw 0 deg c. R o lling M o ment vs _, iw = 15 deg, _ . F = 3 0 deg d. R ol li n g Mo men t v s _, iw w 1 5 deg, _F = 3 0 de g 6 7 Effect o f Ailer on D e flection, Tail Off, i = 0 deg 364 a. Lift vs @& w b. Dr a g vs e L _ ; c. Pitching Moment vs d. Y awi n g Mom e nt vs _ r e. R o ll i ng M o ment vs _- f. Si de F o rce vs / xvi P A GE ? S A 2 9 REV 0 1 , 'i

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Figur e _ N u m be r Titl e Page I 68 Effect o f Aileron D e flection, Tail Off, i = 0 d e g, 37 0 _ _ _ _ = 5 d e g w a . Lift vs @_ b Pitching M o m ent v s _ c. Yawing M o ment vs @( _ d. R o lling M o ment vs @4 " 69 Effect o f Aileron Deflection, Tail Off, i = 0 d e g 374 W a. Lift v s _ - b Pitc h ing Mom e nt vs _ ; , _ c . Yawing M o ment vs ;• d . R ol ling M om ent v s V; _ 7 0 Effect o f Ail e ron Deflection, Tall Off, i = 0 d e g, 3 78 @_ = I0 d e g w _ , a. Lift b. Pitching M o ment , ._ _ c . Yawing Mom en t " _ _ L o) d. R o lling Mo me nt _ 71 Eff e ct o f Ail e ron D e flection, Tail Off, iw = 15 d eg 3 82 a . Lift v s c_ b. Pitching M o ment vs @( c. Ya w ing Mo m e nt vs @( ; d . R o lling M o m e nt vs @ < 72 Effect o f Ai l e ron D e flection, Tail O ff, i = 1 5 d e g 3 86 W a. Lift vs b. Pitching Moment vs _ c. Yawing Mo me nt v s V; d . Ro l ling Mom e nt v s 7 3 Effect o f Ailer o n Deflection, Tail Off, i = 15 d eg, 390 ' _ _F = 3 0 deg w a. Lift v s _ 4 b. Pitching Moment vs c. Yawing Moment vs_ d . R o lling M ome nt vs _ |!

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TABLE OF FIGURES ( Co nt'd. ) Figure N u mb e _ _ _ __r Titl __ ___ e Page 74 Eff e ct o f Ail e r o n Deflection, T ail Off, i = 15 d e g, 39 4 _F = 30 deg w a. Lift vs b. Pitching M oment vs c. Yawing M o ment vs -- " " d . Ro lli n g Momen t v s 75 M om e n t In cr emen t s Due t o A iler on De f le cti on 398 a. Ro ll i n g Mom en t v s Win g An g l e o f A tt ac k b . Yaw i n g Mom e n t vs W i n g Angle of Att ac k c . R o l l i n g Mom e n t vs Ang l e o f Y aw 7 6 Effect o f L o w er H oriz on t al Tall Inc ide nc e on 4 01 Pitc h ing M ome n t a. iw = -9 d e g, _F = 0 deg b. iw = -9 deg, _F = lO deg c. i w = 0 d e g, _F = 0 d e g ' , d. i w = 0 deg, 6F = l 0 deg e. iw = 7.5 deg: _F = 0 deg f. i w = 7.5 deg, 6F = 3 0 d e g g. i w = 15 deg, _F = 0 deg h. i w = 15 deg, _F -- 3 0 deg 77 Effect o f Uppe r Ho ri_ .o ntal T a i l In cid en c e on Pitching 4 09 Mom ent _ _ a. iw = -9 deg , _F = 0 deg b . iw = -9 deE, 6F = lO deg c. iw = 0 deg, 6F = 0 dee d. iw = 0 d e g, _F = i 0 d e g 1 e. iw = 7.5 d e g, _F = 0 deg f. iw = 7 . 5 d eg, _F = 30 deg 6. iw = 15 d eg , _ = 0 d e g t h. iw = 15 deg, 6F = 3 0 deg xviii PAGE _A _ q r_ V r t

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_" TABLE OF F IGURES ( C ont'd.)

< I Figure < Number Title Page :. 78 H o riz o ntal Tail I n terference-Effect o f Empenn a ge 417 B u i l dup, i = - 9 deg, Pitching Mo m e n t k >, W < 7 9 H or iz o ntal Ta il I n te r ference-Effect o f Empe nna ge 418 _ Buildup, i = 0 deg, Pitching M om ent i W i.. .. •- - - 8 0 H o rizonta l Tail In terference-Effect o f Empennage 419 " _, Bui l dup, i = 7. 5 deg, Pitc h ing Mo m e n t 81 H o ri zon ta l Tail Interference-Effect o f Empennage 420 _ Bui l dup, i = 15 deg, Pitching Moment W • 82 Effect o f Vertical T ail In ci d e n ce on Y awi n g 421

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Mo m e n t , i = 0 de g W ! a . T r i m Th r ust , b . M a ximum Thru s t, O ne Engine Inop er a ti v e _ 8 3 E ffect o f S p l it F la p Elev a t o r 42 3 : i , a. Li ft |_ ) b. Drag - c. Pitching Moment 84 Effect o f Rudder Deflection on Yawing Moment 426 _ 85 E ffect o f Ang l e o f Yaw o n Com p o un d Tai l Al o ne 427 a . Lift

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_ c. Pitc h ing M om ent _ ; 86 Effect o f Angl e o f Attack o n C om p o und Tail Al o n e 4 30 ; : a. Yawing M om e n t b. R o lli n g Mo m ent c. Side Force _ 87 E ffect o f R udder Deflection, Co m pou n d Tail Al o n e 4 33 a. Dr a g b. Y a wing M om ent c. Side Force ,_' J 88 R ud der Co ntrol P o wer, Co m p o und Tai l Al o ne 4 3 8 '_ 8 9 . Effect o f Speed Br a ke s o n C o mp o und T a il Alone, 4 39 Dr a g v s _ : , 90 Effect o f Sp e e d Br a kes o n Com pound T a il Alone, 44 0 _ _ Dr a g vs _ , xlx !;:

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$ " TAB LE OF FIGURES ( Co nt'd.) : Figure _ Number Tit l _____ e . Page :i 9 1 T a i l A l o ne Co nfigurations 441 _ _ _ °a . Lift #s _ ,ji b. Drag vs _ :i c. Pitching M om ent vs 0_ J _ 92 Tail Alone Configu r ations 444 i _

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d. Side F o rce v s _ _ii 9 3 Effect o f Wing In ci d ence-Landing Gear D ow n, 448 ? .

Tail Off, _F -- 30 deg _ _" a. Lift _ b. Drag _ _ : c . Pit ch ing M om e n t , 94 Effect of Wing Incidence-Land i ng Gear Down, 451 4 = 30 deg _ -- _. Tail Off, (_F !.

a. Yawing M o ment :_ b. Rolling Moment c. Side Force 95 Effect of Wing Incidence-LandingGear Down, 454 .

T a il On, _'F = 3 0 d eg _ a. Lift _, b. D rag ,_ .

c. Pitching M om ent _ 9 6 Effect o f Wing I ncidence-L a ndingGear D o w n , 457 _ : Tail On, _F = 3 0 deg "_i " a . Ya wi n g Mo m ent " b. Rolling Moment c. S i d e Force 97 E ffect o f Ang l e o f Attack o n Dir e cti o nal S tability, 46 0 iw = 7.5 deg, _F = 30 deg 98 Effect of A n gle of Yaw on Direction a l Stability, 461 iw = 7.5 deg, _ ' y = 30 deg _.

9 9 Effect of Lo wer H o rizontal S tabiliz e r o n Pitching h62 Mo m ent, Landi n g Ge a r D o w n, iw = 15 deg, _F = 30 d eg '_ I: xx ._ _ @E ._

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, ] _) Sikorsky A ircraft . ® .EP O AT . O. S -72 0 1 , -[ T ABT, e. OF FI G U R ES (Co nt 'd . ) Fig u r e L_ Number Title Page i00 Effect of Nacelle Fairing-Tail On, i = 0 deg 46 3 _ a. Lift " b. Drag c . Pitching Moment 101 Helicopter Sidewash and Do w nwash Angles 466 102 Comp o und Sidewash Angl e s 467 103 Compound Lower Horizontal Stabilizer Downwash Angles_ 468 Do w n w ash vs Angle of Attack i0_ Compound Low e r Horizontal Stabilizer Do w n w _sh Angl e s, 469 Do w nwash vs Lift Coeffici e nt _ 1 0 5 L a nding Gear Effect o n Lower H o rizontal Stabilizer 470 Down w ash Angles 1 0 6 Compound Upper H o rizontal TaI ]* D ow n was h Angle 471 i' ] 107 Wing Oil Flow Patterns - Unp o wered 47 2 _ _ - 108 Wing Oil Fl o w With Wing Fences In s talled - Unpowered h81 : 109 Wing Oil Flow Patterns - P o w e red 48 2 i _, ll0 Wing Oil Flow Without Fences - P o wered 489 lll Wing Oil Fl o w With Fences - P o w e red 491 i I I12 Wing and Nacelle Flow vs Angle o f Attack, I = 15 d e g 494 ; _ _ w ll3 Wing Fl o w v s Angle o f Attack, i = 0 d e g 49 8 W ll4 Wing and Nacelle Flow vs Angle of Attack, i = 0 deg 501 W ll5 Wing Flow vs Angle of Yaw, i = 0 deg 506 W ll6 Wing and Nacelle Flow vs A n gle of Yaw, i - 0 deg 511 W ll7 Wing and Nacelle Flow vs Angle of Attack - Powered 516 ll8 Empennage Flow-Configuration FPBNphWsTB T 519 I' 1 19 Emp e n n age F lo w -Co nfigurati o n FPBNp_W7T 3 6B T 522 I i 1 2 0 Emp e n n age Fl ow - C onf i gur a tion FPBNpsW7T 3 7B _ _ 5 2 9 121 Emp e nn a g e Fl o w - C on f i gur a tion FPBNp sW7 T 29 B T 5 3 5 1 22 f a il Flo w Envir o nment 541 xxi I _' PA GE

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Slkorsky l :llrcraft ------ -

_@ =m_ . A , m . _POR Y NO . SE_R-72011 _" TABLE OF FIGURES (C o nt'd . ) ., Figure : Number Title ' Page 123 Helicopter Surface Pressure Distribution, 543 i c_ = 0 deg, _= 0 deg ' 12h Helicopter Surface Pressure Distribution, 545 4, i _ = -i0 deg,_= 0 deg I 125 Helicopter Surface Pressure Distribubion, 547 'L- _ _ = i0 d e g,_' = 0 deg _ i 126 Helicopter Surface Pressure Distribution, 549 ._ _ _= 0 deg,_= i0 deg I 127 Helicopter Surface Pressure Distribution, 551 _ - _ o ( = i0 deg,_ = i0 d eg _:- I 128 T o tal Pre s sure R a ke Pr o be Locations 55 3 '_ J | 1 2 9 T o tal Pressure Rake Data, Run 3 h 3 , Configuration 554 _ FPBNpIW5T 9 , iw = - 9 d eg, _F = 0 d eg, Trim P o wer, _ 130 T o t a l Pr es sure R a ke Dat a , Run 3_5, C o nfiguration 558 _ I FPRNpIWsT9' iw = - 9 d e g, _F = 0 d e g, Trim P o w e r, _' _ = 0 d e g i I 1 3 1 T o ta l Pre ss ure R a k e D a ta, Run 3 _6, Configuration 56 0 FPBNpIW . = 0 deg, Trim P o wer, ,_" 5T9, i = 0 d e g, 6F . , W '_= 0 deg 1 3 2 T o ta l Pr e ssur e Rake Data, Run 3 47, Co nfiguration 564 FPBNpIWsT9, iw = 0 deg, _F = 0 . . d e g, Trim P o wer, _ = 0 dee _' 1 33 T o tal Pr e ssur e Rak e Data, Run 3 _8, C o nfiguration 566 FPBNpIWsT9, iw = 0 deg, 6F = 0 deg, Trim Power, _ =.lO deg 1 3 4 Total Pr e ssur e Rak e Data, Run 3 4 9 , Co n figuration 56 8 FPBNpIWsT 9 , iw = 15 d e g, _F = 0 deg, Tri m Power, _ @ = 0 d e g 1 35 Total Pr e ss,Are Rak e Dat a , R un 35 0, C onfig u ration 572 FP B N p IW5T 9 , iw = 15 deg, IF = 0 deg, Trim Pow e r, _= 0 d eg " xxii b \

Sikorsky Rlrcraft . tP o . T . . o. S -72011

._, j TABLE OF FIGURES ( C ont'd.)

S I Figu r e '_ N umber Title Page _ * %, " * i 136 Total Pressure Rake Data, Run 351, Configuration 574 :; i = 15 deg, = 30 deg, Windmill FPBNpIW5T9 ' _F _ * w _= 0 deg _ 137 Total Pressure Rake Data, Run 352, Config u ration 578 .

_ FPBNpIW5T9, iw = 15 deg, _F = 30 deg, Windmill, : - _= 0 deg ! .

- _ 138 Total Pressure Rake Data, Run 353, Configuration 580 * FPBNpIW5T9, i = 15 deg, _F = 30 deg, Trim Power, 139 T o tal Pres s ure Rake Data, Run 3 54, Configuration • 584 (,,, _" FPBNpIW5T9, iw -- 15 d eg, _F = 30 d eg, c g = 0 deg , . ) i k O T o tal Pressur e Rake Data, R un 3 75, C o nfiguration 586 , _ _' FPBW5T 9, i = 15 deg, _F = 0 deg, _= 0 deg * _

w _

141 Total Pressure Rake Data, Ru n 376, Config u r ation 591 FPBW5Tg, i = 15 deg, _F = 0 deg , . (= 0 deg _ lh2 Tot a l Pres s ure R a ke D a t a , Run 3 77, C o nfiguracion 595 _i_ FPBW5Tg, i = 15 deg, _F = 0 deg, _= 0 deg i k 3 T o tal Pre ss ure Rak e D a ta, Run 378, Co nfiguration 6 00 FPBW5T 9, iw = 15 deg, _F = 0 deg, 0( = 0 deg _%" " l_k Total Pressure Rake Data, hun 379, Configuration 604 J FPBWsT9, iw = 0 deg, _F = 0 deg, _- 0 deg :i_.

145 Total Pres s u r e Rake Data, R u n 3 8 0 , C on fig u ration 6 09 _- FPBWsT 9, iw = 0 deg, _F = 0 deg, _ = 0 deg 'i_, i_ 6 T o ta l Pressure Rake Da ta, R un 3 81, Con figurati on 61 3 _ FPBW5Tg, iw = 0 deg, d F = 0 deg , _-- I 0 deg _ 147 T o ta l Pres s ure Rake Data, Run 3 82, C o nfiguration 617 FP.BW5T 9, iw = - 9 deg, 6 F = 0 deg, _ = 0 d eg . , _!

lh8 T o tal Pr e ssure Rake Data, Run 3 8 3 , C onf i guration 62_ _ FPBWsTg' -w = -9 deg, _F = 0 deg , _( = 0 deg ; { L PAGE SA _ 9 R EV O ., "' - " I ' r I.

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Slkorsky . E P O . T NO . S . - 72011

. TABLE OF FIGURES (Cont'd.)

Figure Number Title 149 L o cal Total Pre s sure Ratio, Waterline ] . 84, 626 i = 0 deg, Trim P o wer i w 150 Dynamic Pressure Loss - 300 Inch Span 627 I Lower Horizontal T a il 151 Dynamic Pressure Loss - Compound and 628 ; Helicopter Upper _Horizontal Tails ? 152 Dynamic Pressur e Lo_s - V e rtical Tail 62 9 "_ 153 E ffect of Tr a vers e , i = 0 deg 630 _ W , . a. Lift '- b. D rag _!

- c. Pitching Moment _ ..

154 Empennage Flow Characteristics, Configuration 63 3 ' FPBNp8W7TII 155 Empen na ge Fl o w Ch ar a cteristics, C o nfiguration 654 FPBNp W 7TI! 1 156 Emp e nnage Flow C h aracteristics, Co n figuration 661 FPBW7TI I 157 D y n amic Press u r e Ra ti o C o ntour, Co nfiguraticn 68 0 : FPBNps_,T TI I, c_ = 0 d eg 1 58 Dy n amic Pr es sur e Rati o C ont o u r , C on figuratio n 6 81 . _'_ FP BNp8W7T I I , c_ . = - 5 deg -_" 15 9 D y n amic Pre s sure Ratio C ont o ur, Configurati o n 68 2 _ F PB NpsWT TII , ,M = - i 0 deg _:_ 16 0 A n gl e of Att a ck at t h e H o rizo n ta l Stabili_ . er 683 , , 1 6 1 In te g r a ted Horizont al S t a bi l izer" L ift Co e ffi c ient 681, .'} 1 62 Mea n D o w n was h at t h e Horiz on ta l S tabi l izer 685 | ! J f :• i < ' , XX i V _ 1 P A GE _; %

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Sikor s ky A ir c raft = v , . .. o.. , = . _ . ,=-.... . EP O . T . 0. Sin-72011 _

_ ' LIST OF SYMBOLS °' b Wing Span - ft e Wing Chord - ft _ ; _ Mean Aerodynamic Chord - ft , , i ' CD Drag C oefficient, Wind Axis CDS Drag C oefficient, Stability Axis CL Lift C oefficient, Wind Axis _ CLS Lift C oefficient, Stabil_ty A xis : C_ Rolling Moment Coefficient, Wind Axis " ; C_S Rolling M o ment Co e fficient, Stability Axi s C_ Pitching M o ment C o efficient, Wind Axis ' I C_ f S Pitching M o ment Coefficient, Stability Ax i s , _ Yawing Mom e nt Co e fficient, Wind Axi s , C_S Yawing Moment Co e fficient, Stability Axis ', Cp Pr e ssua'e C o efficient Cy Side Force Co e fficient, Wind Axi s / Cys Side Force C oe fficient, Stability Axis i.

CG Aircraft C enter of Gravity i iHT Horizontal T a il I ncidence - d e g ' : iN N a celle I ncidence _gle - d e g _' iVT V e rtica_ Tall Incidence - deg iw Wing Incidence - d e g M _ ll r _) / _ , Pitching M o m e nt Slop e - Ft3 / d e g P N Local Static Pres s ure - r f J XXV _' AGE ' ¢ S A 29 R EV D

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i LI ST OF SYMBOLS ( C on t 'd .) : P Tunn e l S tatic Pres s ur e - psf

t o

, } q ½_ V 2, Dynamic Pr ess ur e - p s f q o Wi nd Tun n el Dyn a m i c Pressure - p s f i qT Tail D ynami c Pre ssu r e - p s f f | . ; V Free S tream V e locity, fps ! _ Angle o f Attack - deg i _a Ail e r o n Defl e ction - deg "i " _e Elevator Defl e ction - deg

J

Flap Defl e ction - d e g _ , !

_r R ud der D e fl e ction - de g 6SB Speed Rrake Deflection - deg D o wnwa s h Angle - deg _ Density - Slug s / Ft 3 _" Sidew ash Angle - deg _ Na c elle C an t Angle - deg ', i _ Angle o f Y aw - deg "i Configuration Nomenclature B Main Rotor Hub BT Tail Rotor Hub ;, F Fuselage F 2 Fus e lage with Landing Gear Fairing Re m oved L Landing Gear . x x vi PAGE t Sa Z 9 REV D I

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i : LIST OF SY MB OLS _ N Unpowered Nac e lles , Np Powered Nac e lles _ , ,_ Npl Pow e red Nac e lles an d Aft Pyl o n Fairings _.

and Forward Pyl o n Fairings _ _ Np2 Pow e red Nacelles and Aft Pylon Fairings (Plate Only) . }: _ L Np3 Npl and Nos e Plug < _" Nph Npl and Nos e and Tail Plug :- . Np5 Pow e r e d Nacelles with Full Support Fairing : , . : _ , Np6 Np5 With Vented Fairi n g (i / 4 inch h o les) _.

:_ Np7 Np6 With Leadin_ Edg e O ff and Trailing Edg e Truncated !

Np8 Np5 With Trailing Edg e Truncated

I

.}_ NRI Larg e Ring Nac e lles , NR2 Small Ring Nacelles : P Main Rotor Pylon _,. TXX S ee Tables I, II, and III, and Figures 9 and I0 °_ for Tail Identification Nomenclature W Wing WI Wing an d End Plates on Flap s _ ' 2 Wing a nd I 0 inch l o ng fences l o cated 12 and 17 Inches Outboa_.d of Wing Root W3 Wing and i 0 inch l o ng f e nces located k and 12 inches _' Outboard o f Wing Root , '_ Wh W 3 and W1 W5 W 3 and r o ot plate fairings e xt en ding fro m wing :_ l e a d ing e dge to l e a d ing e dge of flaps ,. , _ * _ xxvii _ PA G E , f SA 29 REV D ,° i _ .

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,_ Sikorsky I::lircraft _ - ' -_ -"' _ "_'' " REP ort, , o. sm-720n , _ , __

_,... ': LI S T 0_'SYMBOLS ( Co nt 'd . ) / , ! " W6 W5 and W1 W7 W i ng at Phas e II l o cation wi th f e nces a n d -_ root fairings ; i o W8 W7 Without fences g "T : 2 f / J , / F xxviii : PAGE

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• Sikorsky .EP O .T .0. S -72 01 l

INTRODUCTION i Th e s ixth scale model o f the Sik o rsky / NASA_rmy RSRA was t e st e d in the ! 1 8-foot se ction o f t h e United Ai r cr a ft Re se arc h L a b o r a tories (UARL) L a rg e Subsonic Wind Tu_ u _el for the purpos e of obt a ining basic d a t a for the R S RA program in the area. - , of perform a nce, st a bility, an d body surface loads.

These d a t a are required to substanti a te and u pd a t e curr e nt a n a lytical estimates. This report is the fin a l report d oc u menting the dat a a n d test proc e dur e s of Phases I and II of the RSRA wind tunnel te s ting.

The model was m o unt e d in the tunnel on the struts arr a nge d in t a n d em.

Basic testing was limited to forward flight with angles of yaw from -20 to +20 degrees a nd a ngles of a tt a ck from -20 to +25 degrees. Tunnel test speeds were varied up t o 172 knots (q = 96 psf). Interference d a t a were derived from the tenth scal_ Utility T a ctic a l Tr a nsp o rt Aircr a ft System (UT T AS) wind tunnel testing, R e f e rence 3, an d Sikorsky S-67 dat a , Reference _. Test data was monitored through a high spee d static d a t a a cquisition system (STADAS), linked to a PDP-6 c o mputer. F n is system provided i m medi a te records of a ngle of a tt a ck, angle of y a w, six component force and mom e nt da ta, and st a tic a nd total presssure inform a tion. The t e st parameters were stored on magnetic t a pe for off-line processing.

Th e w ind tunn e l mod el wa s c o nstructed o f al u mi n um s tr u ctural me mb e rs with aluminum, fibergla s s, and wood skins. Included in the te s t progra m were tip driven fans t o simulate airflow thr o ugh the R S RA's TF- 3 4 thr u st engines.

This report i n clu d es t a bul a t e d forc e an d mom en t data , flow vis u l a ti o n photogr a phs, t a bulated surf a ce pressure d a t a for the b a sic helicopter a nd compou n d configurations, and limited discussions of the results of the test.

" In a d dition to the authors the following personnel were major contributors to the wind tunnel test program: R. Bl a uch - Test Oper a tions B. Gol d iez - Test Operations and Report Prep a r a tion D. Clark a nd R. Moffitt - Hot Wire Anemometer Operation J. Rorke, R. Monteleone, N. Heslin - Test S upervision F , Moore, R. 8 atte_hwaite, J. Hassel - NA S A / Army Representation l PAGE SA 29 RE V O i

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@ , i DESCRIPTI ON OF TEST FACILITY I Larg e Subsonic Wind Tunn el t The Large Subsonic Wind Tunnel is a single-return, closed-throat _ faci l ity wit h interchangeable 18 a n d 8 - foot octagonal t e st s e ctions. Th e tunnel is adaptable to testing models and components of airplanes and he lic o pters, ful l-s cale and mo de l missile s , propellers and h e licopter :. r o t o rs, p o w e rplant install a tions e ngine inlets and e xhaust nozzles, and air induction s ystems. Maximum tunn e l v e l o cities ar e approximately 2 00 _"" mph f o r the 18 foot t e st section. Tuz m el stagnation pr e ss t tre e quals ; _.- atmo s p h eric pre ss ur e , an d t he s t a gnation temperature o f t h e airstream can o o : be h e ld c on stant i n t h e 6 0 t o 1 5 0 F ra n ge by m e ans of air e xchang e r valves. Vacu u m and 40, 100, and h00-psig air supplies are available to use in inl e t and nozzle t e sting. El e ctric p o wer may b e supplied t o the test model by motor generator sets which develop a maximum of 750 hp at fr e quencies up t o 4 00 Hz. Bala n ce, support, and control mechanisms permit a wide range of test installations. Model installation, access to the model, and visual observation of the test from the control room are : facilitated by the d es ign o f the t e st s e cti on s. Th e in s talled model i s shown in Figur e s 1 and 2 and the te s t facility is shown in Figure 4.

F o r this test the tunnel was configured to p e rmit a tandem tw o s trut mo unti n g system con s isting o f a main f o rward strut and an aft pitch strut. _ Both struts were surrounded by self-aligning airfoil shaped fairings (see Figure 1). Separation between strut centers was 70 inches in Phase I and 57 inches in Phase II. The angle of attack range was -20 to +25 degrees and the angle of yaw range was -20 to + 20 degrees.

Tail alone testing was made possible by attaching a structural forebody to the aircraft empemlage components. Strut arrangement is shown in Figure 5. Strut separation was 38.875 inches and permitted an angle of attack range of -30 to + 3 0 degrees with an angle of yaw range from -30 to +30 degrees.

Computer Facilities A high speed static data acquisition system (STADAS), located in the Large Subsonic Wind Tunnel control room, recorded six compon e nt force and moment balance data and static and total pressure data. This data system is linked to a PDP-6 computer located in th e UARL computation laboratory and provides immediate on-line data monitoring capability.

In addition, data were recorded on magnetic tap e' to provide a permanent test record, and for off-line computer processing on the UNIVAC lll0 computer.

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Sikorsky A ircraft . o sin-72 0 11

i- J @ r x DE SCRIPTION OF WIND TUNNEL MODEL % .

: '_ / A e rodynamic M odel Th e sixth scale w ind t u nnel model was designed by Siko r sky Aircr a f t a nd f a b r icated f r c_ alu m inum, fibergl a ss, steel, a nd wood m a te r ials.

Design restrictions for the m odel we r e dict a ted by the a ir lo a ds _ : exp e ct e d at wind tunnel sp e eds o f 175 kn o ts in f o rward flight and by the safety factors required by the test facility. The mo d el w e ight is "_" 749 pounds (configuration FPBW_TB_ Physical dimensions of the RSRA ._" aircraft are shown in Figure 3. _ Basic m o del c o nfigurations ar e shown _ in Figu r e s 6 th ro ug h i0 .

,j L _ The c o ckpit se c tion was cons tructed of mo l d ed fiberglass c ont o ure d t o ,, f o rm the o uter f o rward s u rface o f the mod e l. The interior was h o llow : _. t o permit installation o f static pr e ssure taps and the pow e red nacelle _ a ir s u pply pl e num chamb e r a nd control syst e m. Th e c o ckpit s e ction J was bolted to the forward bulkhead of the cabin s e ction. Th e cabin, transition, and tailcone s e ctions contained the main aluminum structural _ . m e mbers of the body including nac e lle and wing attachment p o ints. Thr ee _ s e ts of al_nin u m cabin skins w e re fabricated f o r the various wing o n / off , _ _ and nacelle o n / off configurations. T he transition and tail_one skins " w e re fiberglass. Th e main rotor pyl o n, which includes the T58-GE-5 engines was also fiberglass.

Two sets of TF-34 engine nacelles were used for this test. The first was of solid wood with inlet and exhaust fairings to simulate airflow ., ar o und the nacelle. The s e c o nd se t o f nac, lie s w e re tip driv e n 8 inch t fan units supplied b y NA S A and manufactured by T e c h D e v e l op ment, Inc., Dayton, Ohio. These were d riven by 400 psig a ir, br o ught into the mod e l through the forward strut. These nacelles ar e 4.8 scale or 25% ov e rsize for the sixth scale model. Powered nacelle configurations are sh o wn in _" ._ur e 7.

The RSRA model wing incorporates variable incidence, and includes a slotted flap and ail e rons. T he inci d ence and d e flection a ngle r a nge s a re shown below. All surfaces were manufactured fr_ n solid alumin u m and wood. Wing f e nces and r o ot s e als are shown in Figure 8.

S A 29 REV D '_I / A r- ! .

; Sikors k y l::llrcraft , EP o, r .o . S - 72011 : .

4- -- _: , The baselin e e mp e n n ag e co ns i s ted of a v e rtical tail with a mo vabl e :_ " ru d der, a va r iab l e i n ci den ce h ori zon tal tail wit h mo vab l e elev ators, a nd . _ .

_ - a sp e e d bra ke . Co n tr o l ra n g e s ar e sh o w n b e l o w. E _pe nn age co n figurations :::.

' are s ho wn i n Figures 9 a n d i0, with dim en si on al d ata pr e se n te d i n Table I. _.

j : Tables II a nd III provi de cr o ss r e f e r en c e i nde c e s to aid i n i den tifying • th e c omp o n e nts for each tail n u m ber.

Airfoil I n ci den c e Control D e flection , Section deg deg i- ,/ Wi n g 6 3 2_1 5 - 9 t o +1 5 0 t o + 40, Fl aps _',? -2 0 t o +20, A ilerons _ -' !

_- V e rtical Tail 00 15 0 t o k .5 -25 to +25 _. Bas e line Horizontal Tail 63 1 .A 21 2 -9 t o +9 -25 to +25 ; - Spe ed Brak e s - - 0 to +55 : _ A dd itional c om po n e n t s bu ilt f o r t he R SRA mode l , s ho wn in Figure 6 ; and defined in the llst' o f s_mb o ls are: i _, ) - A mai n r o t o r he ad a nd a t a i l rot o r h ea d • R otor H ead s (B & BT :'- w e r e d e-_ -- _e_ _ s imulate r o t o r hea d wa k es. Ro tor d o wnw a sh was not _ si m ul a te d duri n g ;his t e st. _ _t_ La nding G e ar and Oleo Struts (L) - Th e ex t ende d ma Jn l an ding g ea r i pr o xi m ity t o th e pitc h strut attach m ent point of t h e mo d el .

i The fo llo wing table pr e s en t s basic mo de l dime ns i on s, s upp l ementing an d its str u ctural me m b e rs. Th e tail gear was n o t test e d because o f its I I t he d a ta s h ow n in Figure 3 .

• , L ENGTH F u selage 12 0 . 3 in. - _' Overal l ( n o s e t o aft poi n t on s tabilizer) i k l .2 in.

W IDTH C a b in S e ction 1 3 . 33 in.

Overall (H o riz o ntal stabilizer spa n ) 5 0 . 0 0 in.

H E I G HT , : C a b i n S e ction 1 5 .1 in.

Overall (V e rtical Sta b ilizer t o wh e e l s) 3 8.8 i n .

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" Slkomky Rlrcraft U REP O RT .O. S - 720 l l

' _ !_ ' _ ® : MODEL Rk_ 4 0LV I NG CENTER : - _ Fuselage S t ation (A /C dimensi o n) 309 in.

_ W a ter l ine (A / C d imen sion ) 22 3 in. • l Buttline (A / C d ime ns ion) 0 in.

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_, BALAN C E RESOL VI NG CENT E R _ F us elage St a ti o n (A / C dl m en s i o n) 2 65.5 i n. _ W a ter l ine (A / C dimens ion ) 2 0 8.0 i n .

• - , - -' PIVOT CENTER _, .

Fu s elage S tation (A / C d ime ns i on ) 2 65.5 i n .

?

_ . W ater l ine (A / C d imensio n ) 178.0 in. ' _ ' " REFEREN CE LENG TH S , A R EAS , AND V O L UM E S h Mo d e l Vo l u me ( us e d for t u nne l _ . bl o ckage co rre c ti o n ) m :; Fuse la ge 7.16 cu. ft.

Wing i.i0 c u . ft.

: Nacelles 1 . 8 0 cu. ft. .._ " i Empennage . 62 c u. ft. : _" Mai n R o tor Pyl o n .8 9 cu. ft. _ • C r o ss S ectional Area 1.8 5 sq. ft.

_ , . 3 Reference S pa n 7.h72 ft. ;: Reference Ch o r d 1 . 400 ft. _ _ The tail al o ne c on figuration uti l ized t h e UTTAS m o unti n g s ystem _.

_ - with a s trut separation of 3 8.875 inc h es. This i ns ta ll ation ha s a piv o t :: center as foll o ws: i_ .

_ F u selage S t a ti o n (A / C dime ns i on ) h52.8 i n . _ ' Waterline (A / C dime ns i o n) 1 82. 0 in. ._ _' M od el Vo lu m e ,_.

: Empen na ge a n d F o rebody ( T3 ) . 9 5 c u . ft. . :_ ' ? . _: ' F o rebody (Th) .5 3 cu. ft.

_ The wi n d tunnel m o del was equippe d wit h 16 3 pre ss ure tap s t o ii m e a sure pre ss ure distributions o n t h e rig h t side o f t h e c o c k pit, ca bi n , ._ _ tr a n s ition a nd t a ilcone s ections, a nd t h e ma i n rotor pylon. T h e t a ps i c o nsist o f s t a inless s teel tubing b o nde d flush to t h e surf a ce a nd connected by flexible pl a stic tubing (Ge o n) t o f o ur hS-tap S c an iv a lve 'i units mounted i n t h e f o rward part of t h e c a bi n . L_c h S c an iv alv e unit _ _° c o nv e rted t h e pr e s s ur e t o a n e l e ctrical s ignal w h i c h w as r e c o rded by STADAS. Only nine lead lines from th e model were r e quired for pressure : re ad ings, a nd s ince t h ese li n es were enc l o s ed wit h i n t h e strut f a irings, they did not affect model aerodynamic forces. Pr e ssure data were therefore , a cquired d uring st a bility t es t runs without t h e nee d for s ep arate pr e s s ur e test runs. Components having pressure taps were removed from the model by _ unc o upli n g t h e S c a niv a lve c o n n ectors a t t h e Sc a ni va lve o r by removi n g the j : : tubi n g on ea ch c o mponent. The loc a tions o f t h e fuselage pre s sur e t a p s are _ .

_,'_ give n in T a ble I V. 5 _.."

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Sikors ky I:: U r craft . e o . t . o. sm-72Oll

BALANCE DATA ANALYSIS _ T est Procedure ; i j Th e RSRA wind tunnel test procedure was arranged t o provide data in i f o ur main ar e as o f interest. Th e s e areas consisted o f m o del buildup, wing i p e rformance, emp e nnage s tudies, and m o del c o mponent interference e ffects. : Mod e l c o mponents, including main and tall rotor heads, wing, nacelles, main r o tor pylon, and landing g e ar, w e r e added t o the m o del individually ; _; provide incremental e ff e cts. Data w e re taken f o r a maxim u m angle o f _ . _ _tack range o f -2 0 t o + 2 5 degre e s, and a maxim u m angle o f yaw range o f " _ " - + 20 d e gr ees . Th e tunn e l wa s op e rated at a dynamic pressure of 55 p s f _- (about 13 0 kn o t s ) e xc e pt f o r R e ynolds number and nacelle thrust studies.

f " The tunn e l balance data w e r e r e corded o n magnetic tape using the ! S TADAS data r e duction system, and immediately r e duced using the UARL PDP-6 c o mputer and print e d on-llne. The amount of output printed on-lin e : wa s c o ntrolled t o o btain o nly those values n e c e ssary for data checks.

( ' The s e data w e r e m o nit o red during e ach run and questionable points _run .

_ Final data w e re c o mpiled fr o m the tapes o n the UNIVAC lll0 c o mputer. The _ ' data presented in this r e port ar e in f u ll scal e param e tric form, i. e ., ; _ , normalized by the dynamic pr e ssur e , q, and in c o efficient form, non- !

dim e nsionalized by q and wing area and chord or span .

¢ In addition t o f o rce and m o m e nt data, fl o w visualization studies . . w ere made on the a erodynamic model using tuft s a nd oil. Tuft s were used to determin e surface flow characteristics on th e forward sections, main r o tor pylon, e ngines, wing, and e mp e nnage. Ph o t o graphs w e r e taken o v e r • a range o f f o rward flight attitudes. A tuft rake was us e d to investigate ; . thr e e - dim ens i o n a l flo w a r o und t h e e mpenn a _ p at a l ow tunnel s p e e d. 0il _ fl o w s tudies w e r e c o n du cted by di s s o lving lamp black in SAE 30 o il and pl_cing the s o lution o n critical surface ar ea s o f the model. N o rmal : tunn e l q wa s t h e n mai n tai n ed f o r a ppr o xi ma tely 5 m inutes with the mo del ' : in a fixed flight attitude. Aft e r shutd o wn the m o d e l surface c o uld b e : _ ph o tographed to record fl o w patterns.

_" Fl o w c o nditions at the e mp e nnage were m e asur e d using two types o f vel o city measurement apparatus. A total pre s sure rake was mounted on the m o d e l t o s urv e y the fl o w dynamic pr e ssure. A tri-axial hct wire probe was m o unted b e hind the model t o measure fl o w dynamic pressure and local flow angles. Thi s e quipment and resulting data are discuss e d in a later s ecti o n o f this r e p o rt

REPRODUCIBIL I TY OF TH_

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Slkorsky Rlrcraft _ - - @ - - _ - _m_ - - --- * _ . EP O.T . 0. SER -7 20n

d, • Dat a Presentation Aerodynamic force and moment data have been reduced from encoder _ values to parametric or coefficient values in the wind axis syst_n, Figure ii, and coefficient values in the stability axis system. The steps in the data _ reduction process are listed below: _ i. Convert encoder forces to forces and moments in units of i pounds or foot-pounds.

2. C orrect f o rces an d mom ents f o r start z e r o s and static -_ moment variations.

3. Transfer mo ment s to the m od el r e so lving center (Puselage station 309 , waterline 223, and buttline 0).

_ h. Corr e ct forces and moments for aer o dynamic tare and inter- _ ferenc _.

5. Compute forces and moments in parametric f o rm by dividing by dynamic pr e ssure, q, and correct for model sc a le using i_ a factor of 36 for forces and 216 for m o ments.

:_ Data is presented in parametric form throughout this vol u me and J in Vo l ume II.

Force and moment dat a were also computed in wind a xis a nd stability i axis c o efficient f o rm using the f o ll o wing equati on s: i WIND AXIS : CL = L / qS

cD = D / _S

c_- _s_

c_ = _? / qSb ,,

c x -

c r - Y / as

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} _ " STABILITY AXIS ' _ : _ CLS = CL '_ ! ' f CDS = C D c os % U - C x s in_ /i: , % :• % S _ C ; W c°s%_- (b / _) C z sin W _iI _, C_S = C _ f : _ 'L[ .! C_s = Cz c o s _ + (S / b) C_wsln _ u ',':':ii!

_,_ I "! J l C y S = C y c o s_ + C D sin_ "_,_ _I Table Vll has be e n pr e pared t o as s ist in the loc a tion o f data runs _% • ._ among the plotted da t a . Only t h ese run s w h ose da t a are pr es ented in plotted : I_ form are li s ted.

"- Balance Dat a Precision ' 1 ; • During the c o urse of the t e st, data runs w e re r e p e ated t o e st a blish th e s tatic start and e nd z e ros, and data points a_ z e r o pitch and yaw i the confidence l e vel o f the balance data. In addition t, r e peated runs, _..,_ i angles pr o vide r e peatability i n f o rmati o n. _ , UARL es tablish e d the static data accuracy for a 95% c o nfidence l e vel f o r Phase s I & II t es ts in R e f e r e nc e stl & 2. This information is r e produced in , _ T a b l e VI II , w h ich al so incl ude s t h e a ccur a cy r an ges for test s r e p o rted in j _ R e f e re n ces 3 a n d 5. Variations d ue t o t h e f lo w o f c o mpre ss ed air thr o ugh _ , !_'_ ' t h e cr o ssover s y s t em fr o m air su pp l y t o t h e bala n ce influenced the accuracy o f p o wered nacelle balance data, esp e cially the pitching m o ment component , Whil e tunn e l and nac e lle run-ups can r e duce th e pitching m o m e nt shifts, the shift s do n o t affect data sl o pes and the displacement is small r e lative to the range o f pitching m o m ent me asur ed . _ / Figur e s 1 2 and 13 show the data r e peatability f o r all six force and m om e nt param e ters for c o nfigurations FPBNp_W7TI0 _ and FPBNpsW7T 2, i ' respectively.

Aer od_mamic Tare and Interference T he forc e an d mo me nt da t a we r e co r r e cte d for d e rived tar e a nd int e rference e ff e cts, s i n c e m o de l c ons truction di d n ot provide f o r m o de l inv e rsion t o g e n e rate t a re a nd interference d ata . To de fine tar e a nd I int e rf e r e nce (T &I ) corr e ctions i n t he he licopter m o d e a co mp a ris o n w as i ma de b e tw ee n t he uncor re cte d RSRA d a t a a nd t h e 1 / 12 sc a l e 8 -67 win d tunn e l i t e st r e sult s (R e f e r e nce _). 'For lift, sid e f o rc e , ro ll_ng mome nt a n d 1 y a wing m o m ent t he dat a w e r e s imilar a nd ther e fore th e tar e an d int er f e r e nce

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cont r i b uti o n to th e d a t a s h ou ld b e r e l a tiv e ly insignific a nt, Th e r e was a :_ m e asurabl e change in pitching mom e nt, wh e re th e application o f T&l's w o uld . produce a more st a ble sl o p e , and ther e fore pitching mom e nt T&l's were "_ cons e rvatively not applied. Drag tare and interference corrections are J significant, and w e re the o nly cnes applied to the h e licopter data. Th e • i' drag T&I was d e t e rmined by forcing the drag o f the Phas e I fus e lage alone i c o nfiguration at zer o angle o f attack and y a w equal to the e stimat e d fuselage drag (3.10 sq. ft.). Th e sh a pe of the dr a g T&I curv e as a function o f angle of a tt a ck and yaw w a s then d e v e loped from the 1 / 12 scale S-70 pilot tunnel t e st results (R e ference 3). Fuselage ae r o dynamic tare " _._ . ,_ and interference corrections are sh o wn iz Figure i_.

ii_ F o r th e c o mpound c on figuration a s urv e y was mad e o f the tar e and :; interference c o ntribution f o r s e v e ral c o mpound c o nfigurations pr e viously _ ' t e sted by Sikorsky Aircr a ft. Th e results o f this surv e y indicated the _ T&I c o ntribution is similar t o that discussed f o r the h e licopter c o nfigur a tion.

A tar e c o rrection c o rr e sp on ding t o the f o rce and m o ments o f the f o r e body and aft tailcone (C o nfiguration T4_ was applied t o the tail alone :_ data. Th e s e tar e corrections are pr e s e nted i n Figure 15.

Reynolds Number Effects The e ff e ct of tun ne l sp e e d was investigated e arly in the t e st t o provide info r mation th a t led t o t h e s e lection o f a d ynamic pr e ssure o f 55 psf for n o rmal running. The trend o f dr a g wit h tunnel s p e ed is s ho wn in Figure 16a in te_ms of Reynolds number per foot, where a Reynolds number of 1.28 x lO- corresponds to a dynamic pressure of 55 psf. The Reynolds number values presented are not corrected for the Large Subsonic Wi n d Tunnel turbulence factor which is approximately 1 .l_. The Reynolds numbers presented herein should be multiplied by the factor prior to any c o mpariso n s with corrected dat a . Figure 1 6b s h ows t h e effect of R eyn o lds , _ numb e r o n lift and drag f o r c omp ou n d co n figurations. In clude d is th e S effect o f leading edge r o ughn e ss o n the wing a n d tails (grit size i_ 0 ). _ Helicopter Buildup and St a bility - Final C onfiguration H e licopter performance and stability parameters for the final Phase II c on figuration, an d t h e c o mpon e nt bui ld up to t h i s c on fi g uration, t ar e pr e s en te d in Figures 1 7 thr o ug h 29 . The RS RA h el ico p ter mee ts or i exceed s the pitc h ing mo ment c r iteria e stab l i sh ed f o r t h is test o f - _ 0 cu. ft / deg with t h e 35._ squar e f oo t upper h o rizontal tail.

' , Figures 17 a@d 18 show the effect o f indivi d ual c o mponents during I _ a helicopter b u i l dup. The a dlition o f t he main rotor h e ad and _y lo n do not have a significa n t impact on lo n gitudinal stability, hut does reduce lateral stability f o r angles of yaw l e ss than + 8 degrees, as shown in i Figure 18b.

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" " (_ , _ ". _ ' The par"slte drag of sever a l RSRA components can be evaluated from the wiz: d tunnel results. To establish incremental drag levels the measured drag at zero angles of attack and yaw were listed and averaged. . _ Analysis of incremental drag is then possible by subtracting the averaged : drag f o r two c o nfigurations, o ne with and the second wit h out the component. .. , _ Since the drag tare and interferencevalue was derived by adjusting : the tested fuselage (FT2 - Phase I) drag to an estimated value of 3.10 s q uar e f ee t, the test cannot b e used to confirm this valu e . Modifications i made to the fuselage and main strut for Phase II testing increased the .

ati o ns FT2, FPBT2 - Phase I, a n d FPBT2 - Phase I I ). The sourc e o f t h is differ e nc e has no t bee n de termine d . _ ' f u selage drag by 2. 6 1 square fe e t t o 5 .71 s quare feet (based o n c o nfigur- : Th e main r o tor pylo n co n tribut e s 2 . 2 3 s quar e feet e l drag, co m par e 4 , _ ,_" to the estl m ated value of 1. 5 0 square fe e t. The tested main and tail _ r o tor head s , do n o t fully represent the actu a l RSRA c on fig u ration, The main __ : _ r o tor head tested and predi c ted d_-ags were 3.73 and 8. 9 3 square feet, respectively. The tail rot o r head tested and pre d icted drags were 1.64 and 1.76 s q uar e f ee t. i_,._ ' i_oe drag of empen n age components was derived from tail alone D / q ' , : da ta multiplied by the dynamic pressure ratio at the tail surface in the following mann e r: _ _ , Vertical Tail ([T65BT -(T6 3 Bt- T6_BT)]- BT_c_ / q 'i

- .i '

W h ere qT / q = 0. 68 for % he h e li co pter ( _ ' igure 1 5 2) " a nd = 0.86 for the compou n d(trim power) _: This calculation res u lts in vertical tail parasite drag_ of 0.86 and 1.09 i s quare feet for t h e hel ic op ter a n d co m po und c on fig u rati o ns, respectively, compare d to a pr ed ict ed val ue of 0.9 6 s q uar e f e et f o r th e u n e x tended i: vertical tail, or i.i0 s q uare fe e t f o r the e x tended vertical tail. _ } : The d r a g of _ a c h h ori z o n t a l tail i s c a lcul a t ed as a bove : ' H o rizontal Tail Al o ne _ i Component Tail ....... D r aK _T " q , Para s ite Dra_ 17.2 m q . ft. Upp e r 0.69 0 . 9 7( C ompound) 0 .6 7 3 5 . 5 _q. ft. Upper 0. 72 0. 9 7(H e licopter) 0.70 98 . 1 sq. ft. L ower 1. 3 0 O .81(Co m po u nd) 1. 0 _ i ' 10 PAGE ,. . ' , , b

S ikorsky I::llrcra t "--- '- o . P o . T N o , s .R-72on

:T C X M _,_g l A_I O N ; The predicted d r ag of th e origin a l " 10 square foot horiz o nt a l t a il w a s 1.02 square feet. It shou l d be noted, however, th a t D _ . eof the ta i l surfaces tested were exactly as defined for the RSRA; . ,efinal tail t" _ surfaces should have a lower parasite drag. _ The e ffectiveness of the s peed brakes is a function of the dynamic i < pressure in proximity to the tailcone. Det a iled measurements in this , ' region were not taken, although a measure of the helicopter qm / q can be I i determined from sp e ed brake data from helicopter a nd tail alo fi e c onfig- : , urations. Thes e dat a ar e presented b e low.

I • :. . .. a ,T a il Alone ,Hel_,c p pter cJT / cl 35 7 . 7 1 7. 3h . 9 5

_. D / ,, T _ 15 2. ,' 2.2 3 1.00 '

" I _ e_'_,l,u_,, 55 14 . 45 1 3. 3 9 9 3 "_ . ' _ O :, 0 2 0 _ FO 6 o J i • C ompound dynamic pre s sur e lo s s is expected to be h ig h er t ha n th a t f o r = the h e lic o pter.

Figures 1 9 and 2 0 pre s ent th e effect of a ngl e o f a ttack on i a _ e ral _ and direction a l charact e ristics. Figure 1 9 shows the eff e ct o_ " _bo main , l rotor pylon on the vertical tail due to positive body attitude. , _t ¢ ¢ = -10 degr e es t h e _a w ing m om e n t slo pe is c o n s t a nt a t -86 c u ft / deg. A c e <. -- 0 _" d e gree s th e slope is re d- ced to - 2 h cu ft / deg, for -84 _ P < 8 degre e s, and at e _ = +I0 de@-e e s the slope is neutr a l t_ slightly un s table.

Figure s 2 1 and 22 s h ow the effect s o f h o riz o ntal a nd vertic al tail i n cidence, w hich appear nearly symet_Ical. Th e se d a ta w e re use d in t he calculation of d own w ash a nd si d e w ash angl e s (see Figure i 0 ).

Figure 23 presents the effects of rudder deflection. The rudder effectiveness, b _ / _ p , is 37 cu ft / d eg at _( = 0 degr e es. Isolate d : ".

tail dat a (Tail A Lo ne) is pr e sented in Figures 2h-2 9 , a nd i n c lu d e s th e .!.

impact of th e sp e ed brake ,s on drag (Figures 2 8 and 29 ). Compari s on s wit h co mpound tail a lo ne d ata may be found in Figures 9 1 and 9 2.

ii '.i PAG _ S A 2 g REV D ,, ,, \.

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Sikorsky Rlr c raft . EP o . T N o. sin-72011

_Chm_ s oun d Buildup and Stability - Final Configu ration { 4 The e ffects o f the w ing a nd n a c e lle wak es o n the em penn a ge f o r c ed : z, f t_o-tail s o lution f o r t h e c om pound to meet a longitudinal stability criteria u se d for this t e st Of -6 cu ft / deg, for the full range o f wing " incidences. The p e rformance and stability parameters for the final Phase II compound configuration are presented in Figures 30 through 5_.

f ; Critical t o t h e e valu a tion o f c o mpound p e rf o rmance a nd stability is the proper assessment of nacelle forces. Provisions were not m a de to pe r forz a n is o lated calibr a tion o f e ach nac e lle, but the t . hrust was m e a s ured w ith the n a cel les i n st a lled o n t h e mo d el, b o t h s tatically and at a dynamic _ press r _e of 55 psi. Th e resulting d a ta i s s h o wn in Figure 30. On Figure _ 30a th e m e asur e d thrust at q = 55 psi has b een adjusted by the drag of the c o nfigur a tion without n a c e l l es, FPBW-T - B , whic h is 2 6 4 squar e fe e t, o r i L j proximately J_ 0 _3pounds. Th e nacelle t_ust presented is th e refo r e the - r : n e t thrust o f the nacelle / nacelle fairing system. In this form, the thrust ; at q = 55 p s i is on a clos e ly comparable base t o the static thrust. Actual _ i m e a s ured thrust is s h o wn in Figure 30b in t e rms of the thrust parameter. ; The thrust produced i s s function o f wing lift and nac e lle fairing _ - !i i c_. ours. Wing / nacelle s e paration is v e ry important, aud there are _i_ indications that small differences in the internal contours of the nacelle _ fairing (N _) may have caused flow disturbances to produce nac e lle lift at _ r _ a z e ro nace_leP angl_o of attack. Variation of total lift, for the nacelles " at an attack of zero, is s hown in Figure 3 0 b for the range of fan spe e ds _ .

"_ t e sted. _ Th e effects o f the compound component buildup, beyond that of the ,_ h e licopter configuraticn, ar e shown in Figures 31 and 32 for tail off . i ; configurations, a n_ 33 and 3 4 w_t h the t a il o n. Ea c h powered n ac el l e run ; _ . consisted of 8 "windmill" data point at _ = • = • 0 d e gr ee s, normally taken ' _ i_ prior t o the data f o r e ither "trim thrust" or "maximum thrust," or with one e ngine in o perative ( O EI). I ndivid u al w indmill p o ints, when available, ._ _ ar e shown with solid symbols. The "trim thrust" condition was d e fined ! : a s the fan spe e d n e c ess ary to balance the total drag of the model and support struts at @(--_--0 d e grees. "Windmill" RPM was the f a n speed _ ; .

_. resulting fr o m the f o rce o f tunnel air. "Maxim '-, m thrust", 23 ,000 RPM on the s e fans, r e sulted in a scaled thrust e qual to 7 0 % of the actual TF-3h % _ maxi m um thru s t. OEI c o rre s p o nd s t o maxim u m thrust o n the right fan, w ith _ " th e l e ft fan windmilling. Unless o th e r wise noted on th e figures, the nacelle • _ i n cidence i s-3.5 d e grees an d the c a nt . ang l e is z e r o . 'ALl c on t r ol ' "_.

s u rface d e fl e ctions ar e z e ro unl es s not e d. ' !_ i_ Th e drag of s e v e ral nacelle and nacelle fa i ring co nf i gurations wa s _ d e t e rmined. Th e basic nac e lle configuration, N p _ at iN = -3.5 d e gre e s, i _ had v e ry good drag characteristics. Either a n _ficr e ase or d e crease in _.

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nac e ll e incid e nc e incr e a se t he z e r o an gl e o f a ttack dr a g by ab ou t 0.3 _ squar e f e et / deg ree o f n a c e ll e incidenc e . Th e o nl y f use lag e / n a c e lle f a iring t ha t had le s s dr a g w a s t he sm a ll a i r foil- sha p ed fa i r i n g a roun d t he , p o w e re d n a celle a i r s upply pip e ( N_). Th e v e nt e d f a i r ing N_ h ad t h e sa m e incre a se d dr a g _ v dr a g a s Nps. Trunc a ti o n o f t h e tr a iling e d ge ( P l us v en ti n g_ by a n a n a d d iti o nal 3 . 0 sq u a r e f ee t. Th e break d own of n a c e lle, n a c e lle ' f a iring, and w indmill dr a g i s n ot p os sible u sing tunn e l d a t a , b u t it c a n b e es tim a ted an alytic a lly . Usi n g c oe fficients an d value s fr om th e e stim a ted : dr a g, t h e f o ll owi ng n a c e ll e d ra g br eakdo wn is a ss u m e d f o r c o nfig u r a tio n Np5 a n d Np .

N ace ll e P a r a site D ra _ .... _ Item - RS R A De s ign Np5 Np R e m a r ks :_ , Isolate_ TF- 3 h, sq . ft. 2 . 00 _.1 3 3 . 13 S cal e .j.

c o rr e ction Support F a iring, s q. ft. 0 .6k 0.9 6 0 .1 3 Size and s ha pe _ ____ _ _ corre c ti o ns , Int e rference, s q . ft. 1.7___8 1.7_88 1.33 i : , _ Sub-T o tal, sq . ft. k .h2 5 . 8 7 4.59 ' _ Windmilling Drag, so.. ft. n o t 3 . 96 5 . 96 a pplicable T o tal h .42 ii .8 3 10.55 H_ l ' 7 J : i; PA 0 "_ ' _ ' : 't S ikor s ky REP ORTNO . SE R- 72 0 11 ,, Figure 35 _ e s e nts limit e d tail r otor hu h affect data. The tail " _ rotor huh is th e only _o_ p onant of the B S RA wind tunnel moael that is, c by design, not symmetric. Since th e data for configuration s w_thout the _ tail rotor ar e not symm e tric, the model component construction _ust not ._ be _netric. Examination ofthe wing sh o wsthat the left wing flap is _ _ slightly warped t_ward, resu B ting in del e ted stall on the left wing. • : _ , • Early in th e t e st it was d e t e rmined that the required lift could • '_ : . n o t be achieved in t h e la n ding c on figuration (15 degr e e wing incidence). _ Fl o w visualization identified the source of the r e duc e d lift, and wing _ ' : f en ces w e re a dded t o t he wing w l t h a resultant i n cre a s e in lift t o t h e _ ( req u ired valu e . (S ee Fl o w Visualization s e cti o n for further discussion. ) i_ : Ad d itional f en ce studies w e re p e rf o rm e d at the begi n ning o f Ph,_e II : _ t e sting, s_nd ag a i n s h owe d t h e fe n ce s were nece ss ary o n th e mo de l t o o btain _i _ - the r e quired lift. At a 15 degree wing incidence, th e f e nces produc e a 15% increase in maxim u m lift without th e f laps defl ec ted, and a 24% : : : i n cr e as e in maximum lift with th e flaps d e fl e cted t o 3 0 degrees. _.

Figure 36 pres e nts th e e ffects o f wing incidence and angle o f _ _ . _ attack o n li f t, drag, and pitching m o m en t f o r th e baseline configuration, _ _ _ with th e o riginal ii0 sq u are f oo t horiz o n tal tail and e xt e nded spa n and _ c ho rd v e rtical tail. The l on gitudinal instabilities, with trim thrust _ .

}. o n t h e p o wered nac e lles, are clearly sh o wn in Fig-are 5 6 c . _ : Figures 3 7 e nd 3 8 present t h e tail o ff data f o r the c o _ po und , _ configuration at trir_ thrust, while Figur e s 39 - 41 sh o w similar information for the co_ p ound with the final Phase II tail (T_0) installed.

The I_RA has l o ngitudinal stability with t h is t a il for -15 t 5 +17 d egrees fus e lage angle o f attack, except for the -9 d e gr ee incid e nc e wing which _ is neutrally stable from +i0 to +15 degrees f u selage angle of attack and unstable at _igher attitudes. Fig u re 40a shows direetiona_ stability f o r -2 0 _ +_ 0 at z e r o angle o f attack. Figure 40b shows dir e ctional % stability f o r 0° _5 ° f o r a n gles of attack fr o m -20° to +ii °, wh e re the ; _ yawing m o ment i ncrement becomes p o sitive at a wing incidence of zero ' _ de gre e s . Oth e r t es t e d wing incidences remain s table t o b_y o nd +17 degrees. _ Figur e s 42-44 pr e sent ad d itional stability trends.

Nacelle t b_ -ust level impacts on all f o rce and m o ment param e ters.

Figures 45 - 49 _ressnt th_s e e ffects versus angle of attac k for wing incidences o f -_-, 0 -, 7.5-, and 15 - , and f o r the 15 degree wing with B 0 _ ' d e gre e s o f flap deflection, all with the tail o ff. Figures 5 0 - 54 pr e se n t , - .

similar d ata versus angle of ys_. Tail on data comparabl e to Figures 45-49 are sh o wn in Fig u res 55-59.

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The effect o f nac e lle inci d ence and cant angle wa S i nve s tigated l d uri n g t h e t es t. A s c an h e seen in T a b le I X a t a ze r o w ing in c i d ence i a s m all Sot a _i l izing e ffe c t i s realiz e d _y changing engine Incid en ce t o " : . ! 0 -7 o r 0 ^ . H o wever, t h i s e ffect i s _ e ver se a at h igh _rlng incidence "" (! W = 1 5 ). An znc r e as e to po s itive engi ne z n c z dence C 5 ) res_Llts in a de - s t a bilizing e ffect f o r both wi n g i n cidence s . T h e baseli n e n a celle _ : , inc z de_c e o f - 3.5 y ze lds m ini mu m d r ag wltb a drag penalty o f t_ t o I. 7 ft- by either increasing or decreasing nacelle inciaen c e from this poi n t.

_ ' " C anting t h e e ng i ne tailplpe outboard (_w = "5 ° ) produced a -: - r e lativel y larg e de-stabilizing effect a l o ng wI_ h a reduction in p_ T , asite _ drag (s e e Table X). I nboard cant of t h e en gine tailpipe (_ = + 5 ) did ; _. sh ow a significant improv_ ne nt in stability but als o pr o duc e _ a significant d r a g p e n a lty.

;: _a s ed o n t h ese results an engin e cant a n gle o f 0° and incidence _- o f -3 .5- were selected as o ptim u m. , l " The nacelles te s ted on t h is m o del w e re oversized by 25%. T h e fans _ that w e r e tre ed l a ter in the La n g l e y Res e arch C e nter t e sting of this < model permit proper scaling of the n acelles. These were not avMlable i ,-.

_ ) to Sikors k y Aircraft f o r the test doc um ented h e r e in. TO get an indication _ t o f the scale e ff e ct, ring nacelles (s ee Figures 9 t - w ) w e re fabricated ¢ : an d tested. R e sults o f t h e s e runs i n dic a te th a t a pr o perly sc a led , n acelle will have a stabilizing pitching m om e nt i n cre m ent of from -22 t o " - 30 cu ft / deg. " - - The RSRA's s l o tted flaps w o rked well o n t he m o del, as evid e nce d _: i" by F igures 60-65 . Additio n al fla p c o m pari so ns can be f o und i n Figure s 3 6- : 4h, 45, 54, and 5 9 . Figure 60 shows t h e effect of fl a p d e flecti on f o r :- " a z e r o wing in c i d e n c e . Figure 6 1 pr ese nts the sa m e infor m ation in ._ t e rms o f lift-_ r a g r a ti o , d e rived with correcti o ns f o r pow e red n acelle r ;: _ ;" lift and n e t pr o p_ G slve f o rce. ..

_ ' Ail e ro n co n trol p o wer was evaluated f o r wing in c ide n c es of z e r o _ and 15 de grees, and a t an incide n ce o f 15 d e gr e e s with 3 0 d e grees of flaps. < T hes e data a re pre s ented i n Figures 66 thr o ug h 75. i !

O nly t h e rig h t ail e ron wa s d ef le cted d uring the P h as e II te s t, and the d ata o n t h e ai l ero n figures are f o r o nly right ail e ro n d e flections. : . : TO ohtai ' n c o mpl e t e a _ r or B f t T o ll _ n g _ u me nt, a_ the rol l i n g mo me nt ca u s ed : By a d e fl e ction of t h e r_g h t ai le r o n t o t h e r o lling m o m e nt for a d e fl e ction - _ o f t h e opp o site sign t_mes a g e ari n g ratio (If different tha n I : I). F o r _ i " - ' F_OF THE ,_ ' REPKODtJCIBILI :, _ 15 O RIGNAL PAGE IS POO_ : , PA g E %_ .... SA 29 REV D I i . .

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example: _ : Total _ / q = (_ / q @ _ = i 0 O) + (_ / q @_ = -lO°x gearing ratio) ' Figure 66 shows the effect o f the empennage on ailer o n c o ntr o l . While " : : there are shifts in th e rolling moment when the tail is added to the -_ i , model, the change in rolling moment with re s p e ct to aileron defl e ction _ increments (_ l q l _) remains basical_y unchanged. Sinc e , !_ s ho uld not be a ffected by the e_annag e , t h e aileron trend runs we r e _ L_.! m ad e wit h t he t a il o ff. T he aileron rol l ing mo me nt increments are _ i uniform exc e pt near positive and negativ e wing stall, where the dlsy_etry _!

_ o -i in wing c o nstructi o n can ca u se a non- m liform r o lling mo m ent. ' t_ : T he aileron dat a h av e been s u mmariz e d f o r all condition s te s ted '_!

:" in Figure 75 in terms of the change in roll i ng moment from the _= 0 i_ ; [_ degre e c as e to the deflect e d ailer o n cases. The absci s s a in Figure 75 a . !_ : ' fuselage angle of attack into one var i able. No control reversals exist is wing angl e of a ttack which effect i vely combines th e wing incid e nce and _ _ for the r a ng e of angle of att a ck t e sted. Figure 75 " o show s th e eff e ct of ._. .

right ail e ron d e fle c ti o n on ys_ing moment. Using this data th e ail e ron _ , , deflection a ngles that r esult in adv e rs e or prov e rse ys_ at each a ngle of _ : ' attack can be determined. The sketch b e low shows t h e aileron defl e cti o n i_ ' ,_ angles a nd angles of attack w he re adver s e and provers e y a w exist, and sh o ws _ _'_ , .

those areas where a il e r o n gearing c a n reverse the sign o f _he y_cing moment. _ ,_ Figur e 75 c show s t h e effect of angle of yaw on the rolling moment in c rement . :_ . _ t_ _" PRo _ / i;RSE YAW t • \ A L ,,'WAI $ _ ' J:_ ' : DF . FI.GCTION |O _ _&_ . . _ I _AD VER S E YA W "¢ OF RI_;HT : " - t 1 0 - tO O IO ZO ,_, W ING AN G LE OF AT T A CK ) _ _, DEG

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' The eff ec t s o f tail ' incidenc e we r e determined for all tail surfaces to permit the calculation of integrated sidewash and downwash angles, ' and to s how any und e sirable tail characteristics. So that o n e horizontal _ r tail did not bias th e tru e c h aract e ristics o f t h e o th e r, th e l o w e r • horizontal tails and upper h orizonta l tails were run separately .

¢ :' Figure 76 show s the characteristics of the lower horizontal tail, - T61, for ranges of wing incidence and flap d e flection. These curv e s w e r e _ - - , .. - us e a to d e rive Figure s I0 3 and 104, discuss e d later. Similar information for the compound t_per h o rizontal tail is . presented in Figure 77, and .", r e sults in Figure 1 0 6. ._- L L Figures 78-81 c o mpare the pitching moment c o ntributions of th e '_-" lower and upper h o rizontal tails. T he das he d l l ne o n ea c h of these .. figures r e presents the summation o f the pitching moment increment b e tween T41 and T_ and the pitching m o ment of T - , and s h o uld approximate the data ..

"_ zo r the co2mplete tail T6 n. Good agree m e61t i s se en in Figures 78-80, and the differences between _he _ - rived and t e sted lines appears to be a =_ function of experimental accuracy, not inter-tail interference. Figure 81 does not d em o nstrate this agre e ment, but a comparison wit h other data _ _. indicates that there is pr o bably a data s hift in either the T60 or T61 _ data, or pos s ibly both sets o f dat%at this wing incidence. '_ Figure 82 shows the e ffect of vertical tail incidence on yawing _ ; .

moment for trim and OEI thrust l e vels . The sidewash tr e nds resulting from th e se data are presented later. _ & ' Only the original ii0 square fo o t low e r h o rizontal tail was equipped • with m o veable el e vators . To obtain an appr o ximation of e l e vat o r effects o n ; the family of smaller surfaces, a split flap e levator _as fabricated and . - _.

installed on tail T . (see Figures 9C and 9D). The resulting data are _ , ' esented in Figure _0 83 . '_ Rudder e ffectiveness in the c o mp o und c o nfiguration (Figure 84) i s , ,_ j - comparable to that o f the helicopter discussed e arlier. Rudder data f o r ._ t he isolat e d t a ll is pr ese nted in Figur e 87. _ Figures 85-92 present tail alone data f o r the comp o und tail. Th e _ " eff e cts o f y_ angle f o r a range of angles o f attack are shown in Figur e 85, _ _ with the e ff e cts of angle o f attack on yaw angle shown, in Figure 86. Rudder '_ ' : performance is presented in Figures 87 and 88, and spe e d brake drag is _ shown in Figures 89-90. The p e rf o rmance o f all "solution" tails i s '_: summariz e d in Figures 91-92. %' PAGE : L ' k SA 2 9 REV D i _: • ! " / "+ U

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_ _ A limit e d am o unt o f compound data with the landing g e ar ext e nded I w as acquir e d during the test to repres e nt th e landing configurati o n .

, , _ A sampling o f t h is data is s hc _n in Figur e s 9 2 - 99 . Th e landing _ gear is shown to impact on th e lift, drag, and pitching moment, but some of #) t he tr e nds, espe c ially drag, i n di c at e significant tar e m d int e rf e r e nc e + _ _ e ff e cts du e t o th e g e ar's pr o ximity toth e main strut. Th ,- r e f o r e , all I landing gear data sh o uld b e us e d with caution.

Co_ound Nacelle Fairing Configuration Selection Five basic pow e red nac e ll e falrings wer e studie d during this test ;_' "" (Np,N , N. , _ , N_) . O f th e s e o nly f o ur ar e practical for u s e with th e T_-_4-2P_ n ginF_. _ o nfiguration Np, w h ich shows th e best stability, is _ " n ,) t a practical s o lution.

Th e nac e ll e fairing c o nfiguration did not significantly aff e ct aircraft lift, but do e s impact on drag and pitching m o m e nt. This information > is pr e sent e d in Figure i 00 for th e compl e t e compound configuration with th e o riginal Ii 0 squar e foot tail . Venting the fairing d oe s n ot affect th e -7 pitching mom e nt slop e or drag. Th e minimum fairing (_7) improv e d th e + pitching moment sl o p e by -15 cu ft / d e g., but e . tth e exp _ns e of 4.6 s q . f e et + : + of drag . Th e fairing with th e trun c at e d trailing e dg e (N ) did n ot affect _ t he pitching mome n t, but did incr e as e drag by 1.6 s q . f e e_8 Of th e _+ fairings test e d th e existing fairing provides th e b e st o v e rall solution . + . !

H elicopter and Com p .ound Empennage Downw as h and Sidewash Downwss h and s idewas h angl e s at th e e m p e nnag e may be calculat e d using balanc e data or "dir e ctly using v e l oc ity o r angle data. 0 nly limited v e locity data is availabl e for th e s e calculations, and th e r e f o re the • balanc e data has b ee n us e d exclusively in th e d e rivation o f th e s e angles in this section.

K Figure i01 shows the changes in sia e _ash angle versus angl e of ys_, and downwash angl e v e rsus angl e of attack for the h e li c opt e r configuration.

The sld e wash curv e should b e symmetric and pass through th e origin+ . of the ' _ curv e . Th e shift in th e curv e cann o t b e e xplained by r e ference t o tail + rotor hu b e ffects _nd ther e for e it is r e comm e nd e d that the curv e b e translated up to the origin. The sid e wash characteristics of the co_ound are similar (FigRre 102) and it is recommended that this curve als o b e translat e d thr o ugh th e o rigin. Figur e 1 0 2 sh o ws that th e on e e ngin e i no perativ e ( O EI) c o nditi o n doe s n o t h av e a s ignificant imp a ct o n si dewash angles.

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Figure 103 sho w s the d ow n w ash angl e s f o r the c o mpou n d c o nfig ur ation for a ngles of wing incidence and flap deflection. The wing do w n w ash dominates the downwash trends , and results in a i - ag / _ of . 45 for angles of attack in the range of + - 5 degrees, improving for angles outside of this range. The d_nwash information found in Figur e 103 was replotted in Figure 104 versus lift coefficient. Th e use of lift co e fficient as the abscissa tends to collaps e the data into on e line, but the body and nacelle contributions to downwash prohibit complete generalization.

_ " Install a ti o n of the landing gear reduces the d o wnwash angle s a s s h own in Figure 105.

Figure 106 clearly shows that the compound upp e r horizontal tail is not effective at high positive attitudes , _nd is de-stabilizing b_ond _ = 12° f o r maz_ r c o mbinations o f wing and flap a ngles. T h e upper horizontal tail is e ffective in the angle of attack range of ± 5 degrees (with I- _ G / _ ranging in general from . 55 to . 70), and can therefore _ provide the margin of stability necessary in that range. :' I r i * I PAGE . ' _ ! 1 e , _ : , _ " U ""

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' H .: " _ He licopter and-Oo_ound Empe nn age Selection '- A t o t a l o f 6 5 emp e nnage c onfig u r a tion s w ere ev a l ua te_ du ring t h e _:- te s t program. I n a dditi o n to s pecial configurations for tail alone testing : and tail fl o w e nvir on ment surveys, many c o mbinati o ns o f vertical tails, _ I lower h o riz o ntal tails, an d upper horizontal tails w ere assembled and _ _, tested. Repr e sentative configurations a re s h o wn in Figures 9 and i0 . ,_ _" To e valuate t h e relative merit of each tail surface, an effective- ,!_ ; n ess coefficient, appr o ximating the stabilizer lift c o efficient, w as - ' __ : defined as a function o f the change in pitching m oment between tail on , , and tail o ff configurations, the distance from the CG t , o t h e tail, and the , ._i_i " i tail ar ea . , : ,_ _.

: "| ". Thi s effectiveness c o efficient was t h en e v a l u ated for g e neral s l o pe = ; characteristics irat h e angle o f attack range o f -+i 0 degree s , where tail __ "_ performance at a wing incidence of 0 degrees, is critical. For the latter _ _ case the change in @_ from -lO degrees to + l0 degrees was used for a ....

relative comparison. _ ; , To isolate individual tail surface effects, tail configurations .... ; _ that c o nsi s ted o f o nly low e r h o riz o ntal tails or of upp e r horizontal _ _ tails were used where p o ssible. Wh e n necessary combined tail configur- I_'!_ , _ t _ , / , _J ati o ns were used, but thi s sh ould have little effect o n individual tail , ._, performance as s hown in Figure 79 . _i .

/ The ratings prese n ted b e low illustr a te the p o rtions o f the airstream _ _'- at the empennage t h at provide air at m o re opti m t_n dynamic pre s sure s and _!_' ' _ , downwash angles. While airfoil c o nstruction techniques varied fr o m tail :ii i :, , t o tail, variations due t o airf o il shape see m ed only s ignifica_,_ for the _ upp e r horiz o ntal tails. '; The f o llowing table shows the effectiveness rating of each l ow er " : _ h o riz o ntal tail in the co_ o und c o nfiguration: ,_ ,_ , 20 ,-, P AO E _ ;

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_ Lo w er Hori zontal ; Taii Designat ion Effect iven ess Ratin_ / (See Table I) (From -i0 ° _o +i0 °) (% of Ta l l VIII (T61) _ I .441 to .502 74 " II (.639 )* 99 -_ III .621 to .645 98 t - IV .647 lO1 V (.526)* 82 ; , • _, VI •639 99 ' : _. V I I .624 97 .,.

: - . VII I (T61) .644 i00 :_ ,_ IX •407 6 3 , , L _ , X •571 89 : ;, XI (. 582)* 90 _ .. , XIl .468 7 3 ' ; i _ I ] XIII 643 i 00 ¢" xrv . h82 75 _ . : * Based o n data with both a lower and upper h o riz o nt_l tail. i_

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" C om parison o f ratings f o r t h ese tails results in consistent trends . f o r the effect of span, c ho rd vari a ti o ns, and s.nhedral a n gle. The above t_ " table s h ows t h e selected compoun d lower tail is among t h e best o f the tail configurations t-_ted. Span increases from 250 to 300 inches im D roves the effectiveness by 26%. Examination o f Figure 154 shows that the airflow becomes more uniform for span segments at buttlines in excess of llO inches.

Use of the good flow _n the area of the tail tip can be made by increasing t_p chord, although the effectiveness of the increase is only slightly (less than 2%) improved (i.e., the lift increase is nearly proportional to the stabilizer area increase). Root chord reductions remove ineffective portions of the stabilizer. The change from tail III to IV resulted in a 3% improvement in effectiveness. In general chord extensions over the full tail span degrade effectiveness by .5 to .7 p ercent for each percent inc"ease in chord.

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: ® C" ; L_wer horizontal tail anhedral improves effectiveness by I0 - 11% : for a i0 anhedral angle. Anhedral has two beneficial effects on the RSRA.

: First_ it increases the difference between the wing and stabilizer axes _ _' beyond the existing angle of seveu degrees (resulting from wing dihedral) ; l _ so t ha t smaller s egments of the tail are:. in the adve rs e flow field of the wing at aw one angle of attack. Sec6n_ anhedral keeps the tip i segments of the stabilizer farther from the engine efflux in the flight attitudes w h ere red u ced e ffe c tiveness d u e to wing downwash and engine : efflux on the inboard tail exist. A_suming that the effects of anhedral ._ _ an_ span are linear, a five degree anhedral angle would save 5% of the :'_ "_ tail a rea or 5 squ a re feet, if chord were reduced, or 3.5% (3.5 square ; feet) if span were reduced. '_ :_ Increased size, in general, seems to improve the effectiveness of . - upper h o rizontal tails, but c o nstruction techniques ms_ be the d o minant item infl, _ncing effectiveness. The small upper tall (T4_) is simply a _ " i / 4 inch _ uminum plate with rounded leading edges. This _"airfoil" has il i poor drag characteristics (see Figure 91b) which enhance its effectiveness _ by increasing the pitching moment. Stall occurs on this "airfoil" at L -+9 degrees angle of attac_, but because of the high downwssh angle for ' the compound configuration, stall is experienced only for negative angles. 4 _: t !

; The geometrically similar 40, 80, and 120 square foot upper tails _ have aimilar characteristics, with effectiveness increasing about 9% between _ i the 40 and 120 _quaUre foot designs The upper tails with areas between 30 and 36 sqaure feet were fabricated by cutting back the trailing edge I . and reducing span of the 40 square foot tail. This cut-back increased the I trailing edge angle and increased the thickness ratio which might be ' responsible for the 30% reduction in effectiveness experienced between the 30 and hO square foot upper horizontal tails.

Increased span and end plating improved the effectiveness of the vertical tasl significantly. The effectiveness of each configuration was compared in the angle of yaw range of + 5 degrees, w here tail performance _ is reduced by large main rotor pylon interference effects, and in the I range of +- i0 degrees, representing a normal operating range for the tail.

! The 40 inch extension improved the effectiveness by 33% in the -+5 i_ degree r a ng e and 11% in t h e -+ I 0 degree range. T h e large i m pr o vement in 1 the smaller ys_ range indicates that a major portion of the original tail is in disturbed air around _ = 0e, but gets into "cleaner" air : for angles of yaw beyond 5 degrees. This effect was apparent in all of the L l _ 22 PAGE } ' ;A 2 9 RF . V0

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0 modifie d vertical tails . A further 10 inch extension improves t h e _ effectivenessby 8 and 7 perc e nt for th e a bove r ang e s. Since the eff e ctive- r _ ness is only slight l y improved by the i0 inch extension at larger yaw .

i angles, the vertical tail m_st be in relatively clean air above waterline S6Q.

The large chord rudder improved effectivenessby 7 percent (for the ' + 5 degree yaw range). Vertical end plates mounted at the tip s o f the lower horizontal tail i_proved effectivenessby 16 percent in the range of _ ± 5 degree s ,but detr a cted from t h e effect i veness outside of t h is range, f_ The ventral fin extension _- 3 snot effective, and even reduced stability _ for angles of yaw less than 5 degrees, probably due in part to model strut : interference.

The horizontal tails contribute to directional stability in i_o _ ways First, the drag on the horizontals is stabilizing, with the • extensions of the lower horizontal tails contributing about 8 percent to vertical stabilizer effectiveness. The upper tails not only provide I_ this increment to stability, but also provide end plating for the vertical I_ tail. These result in an overall improvement in vertical tail effective- ! • ness of 16 and i0 percent for the -+5 and . ± I0 degrees yaw rang e s, j , respective l y, if the area of the upper tail is not included in the v e rtical _ , t tail area. Unless the upper horizontal tail is needed for longitudinal stability, end plating would not be an effective means of achieving _ directional stability. Including the upper horizontal tail area as part of the vertical area r e duce s the net effectiveness of the vertical tail.

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' k , , ! FLOW VISUALIZATION i Four m e thods of fl o w visualization were used throughout this te_t 1 to get a pictoral view of f low on and around th e model. Th e se m e thods were: r _ i. Oil 2. T U f t S _', 3. Tuft Rake ..i .

4. Smok e 2 _ "" Oil flow, using a mixtur e o f S A E 30 o il and la m p black, was used i_ ih to show the e ff e cts o f nacelles and f e nces on wing flow, as shown in _ Fig1_res 107- iii.

Figur e 107 shows the flow separation patterns on the wing surface nacelle are apparent, with the formation of vortices on the aft wing surface.

Closing the gap between the fuselage and wing improves the flow, but a for runs with and without the s o lid nac e lles. Th e adverse e ffects of th e 1 vortex still is present when the nacelles are installed. Landing gear and " nacelle fairing sp6ilers have little e ffect on the wing upp e r surface flow J Figure 108 sho w s the flow straightening e ffects of wing f e nces. _ Figure 1rig. p r e s e nts similar fl o w s t u dies wit h t h e pow e r e d nacelle s _.

, instal l ed in place o f the solid nac e lles. Th ese phot o graph s s how th a t the nac e lle fairing has an impact on the wing surface flow. The air supp_v pipe fairi n g configuration (N_) has little impact o n the flow patterns, but the simulation of the full fairing resulted in flow patterns similar t o t h ose s ee n in Figure 1 0 7. I nstallation o f f e nces helps t o straighten the _ " w flow. Figures i 0 9i and J show the impact of fences located at th e flap -' mid-span and 5 inches further outb o ar&. Moving th_ o utb o ard fenc e to a , ,!.

location 7.5 inches inboard of the Aap mid-span fence further impr_ ; es Figure ii 0 shows t h e wing fl o w pattern with o ut fences for the Phase II wing location with nacelle fairing N_ Fi@ure Ill shows the flow i_prove- :,' m e nts due to the f e nces for similar_6_ditlons.

t e ! Figures 112 - 118 show the nacelle _d wing flow patterns using

tufts. Figure 112 illustrates the flow for wing W, at an incidence of _ 15 degrees. Inboard sections of the wing can be seen to stall prior to the tip sections. Figure 113 shows flow patterns for a zero degree wing incidence with the nacelles off. The effect of the wing flow on the fuselage . , is _pparent, and shows that an object such as _h e nacelles must have an impact if placed near the upper surface of the wing. The wing stalls more uniformal_v with the nacelles off. Figure 114 shows the effect of the nac e lles at a z e ro d egree wing incidence.

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Figures ll_ and 116 show the effects of an angl e of yaw variat.-'on with the nacelle off and on. No u nusual effects are evident.

Figure ll7 shows the wing flow wit? the powered nacelles installed.

Figures ll8 - 121 illustrate the flow _onditions on the tall surfaces. _ , Figur e ]_18shows that the lower horizontal tail stalls about (_ = 17.5 degrees, while Figure I19 shows that the upper horizontal tail does r -'_ stall at _ , .. - high positive angles, which i s explained by examination of _: " high down- wash angle at th(.ts_l at those angles (see Figure 106). F' . oares 120 and 121 show simila r r e sults. ; The tu f t rake, Figure 122, shows the flow directicns in the vicinity of the empennage. Areas of high downwash can be seen and the nacelle fairing vort e x flow, quantified in the anemometer t e sting., appears o n th e = : _"ake as bl u rred tufts (inboard in tuft ro w s i0 - 13). _ The use of smoke confirmed the flow patters described above. Tunnel c o nditions did n o t permit c le a r ph o tographs of smoke flow, but b a sic < observations are listed below: i . Go o d flow around w ing except w here nac e lle interference 1 _ I caused inboard flow separation.

2. Flow around the nacelle and nacelle fairing is good.

Flo_ up w ssh d ue t o n acelle incid e nce (iN = -3.5 degre es ) w as vis ible. , !.

3. The d / st u rban c e from the m a in rotor head is large and e x tends about t w o feet a tove the m o d el.

4. No fl o w separator v o rti c es co ul d be distingui sh e d. _ .j 5. Upper h orizontal tail is in a high downwash field a.*an g les ._ of attack greater than 0 degre es , b u t the flo w i s o-_nerwise _ undisturbed. $ L i f i " C mW PAGE f

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i, SL_FACE PRESSURE DATA ': Test Pro c edure ' ' The BSF L _ m odel was instrumented with 163 pressu r e taps on the right !

side of the aircraft. The press_re taps o n the fu s elage and main r oto r ' pylon were ar r anged in a matrix form which ].ends i t self to crossplo tt ing and in t erp o lation. J "/ . . _ Pr e ssure data were acquired f o r a limited number o f representative i configurations. The installation of the model and % / .eplacement of , " electrical leads for the pressure tr a nsducers m a de it possible to obtain : : p ressu r e da t a a t th e s a m e t i me a s m odel f o rces were measured. The use : o f Scanivalves incorporating a pressure transducer allowed o n-line data " pro c essi n g as well as off-line outpu t . Each static pressure transducer : , _ , had a range of t 7.5 psi. : i Data Prese ntation -_ t_ -" " Pr ess ur e data - _ _: i s present e d in terms of pres s ure co e fficient where : _:_ _ .

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I'qL_ , The pr e cision o f the pres s ure data was evaluated f,_r c o nf:.gnarati o ns _} with o uz nacelles at dynamic pressures o f 80 and 55 p,_f. For a 95% _ ' : c o nfidence level, the C_ precision is _+ ,0 2 5 at a dynamic pressure of 4 _.

80 psf based o n the ana_ysls o f 82 pressures at o c = _ = 0 degrees _ ' on th e c o ckpit s e ction f o r config_ration FPBTB_. Similarly, the Cp pr e cision is -+ .035 based o n the analy s is o f i_ ' 8pressures at a q = 55 psf f o r configurations FPBT, Ft _ BT o and FPBWT 2 . Thes e t o lerances co.ar e fav o rably with those of th e YU_-60A tests, Reference B. ',i" Tables ":Tthrough ](IVpresent surface pressure data for the helicopter configuration F P BTB T for the following a , _rcraft attitudes: Table XI _ = 0° _ eL = +_ 20° Table XTI _= 0° j _ = + 20°

TableXrV +_ °

Table XIII _= -i0°_ _= +- 20° , 1 !

Figures 123 - 127 ar e repre s ents+ i re ss_mples o f plotted pr e ssure data at buttline zer o _nd waterline 190. Fig_ ' e 123 i , , eludes the r e sults o f : a pot e ntial flow cal c ulation u sing Comput e r program Y179 (R e f e ren e 6) ' , ¼ _' Tabl e s XV thr o ugh XVII pr e sent s urface pres s ure data for c o mp o und co_.figurations, with and with o ut the powered nacelles. :L _ I

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EMPENNAGE TOTAL PRESSURE DATA _ , The flow at th e emp e nnage (fuselage station 664) was s urvey e d : during Phase I of the t e st to d e t e rmin e the dynamic pressure envir o n- ment f o r the v e rtical tail, and for a range o f three horiz o ntal tail ] ocations. The total pressure rake designed to surv e y the em p ennag e was attached to the model in place o f th e tail surfaces as shown in Figure 9i - 9k. Th e pre s sure tubes of th e rs_kew e re .063 h ypo connected to a manometer board in the contr o l room by pl a stic tu l ing. Tube l o cations, in the full scale coordinate system, are sh o wn in Figur e 128. Polaroid photographs of the manometer board at each test point were taken t o o btain a p e rman e nt r e c o rd of the data.

Empennage total pressure data was acquired for a range of wing incidences f o r model c o nfigurations F P BW_T^ and FPBN _. W_ T_ . Var ia bles i**cluded angle of attack, angle of yaw, _l_p deflect[_n > a_gle, and nacelle thrust (the latter two at iW = 15 degrees on_y) . M a nometer board photographs, showing tunn e l t o tal a_d static pre s sur e, and l oc al e mpennage total pressures are shown in Figures 129 through 148. This data can be converted to a dynamic pressure ratio using the following relatio n ship:

% / = i +

ocal PS - H _ : w here % qT / _ is the local dynamic pressure ratio local

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H is tBe tunnel total pr e ssure !

PS is th e tunnel static pressure [i

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HT is th e measured t o tal pr e ssure f o r e ach _ ': positl_e tap Th e r e sult o f thls calculation is presented, for exampl e , in !

Figure 149 for configuration FPBN p lW_T Q, iw - 0, Trim Thrust, 0 e

•

for a range of angles of attack. -Dy_afaic 9ressure losses near h tailcone can be seen in thls figure. Also evident is the impact of the Jet efflux a_ _, = -8 d e grees.

Th e spa n v lse int e grati o n o f the dynamic pressure ratio at each tail surfac e locati o n I s used in the simulation o f helic o pter flight using "tail alone" and "tail off" balance data. The total pr e ssur e rati o , / q, has been calculated f o r all RSRA tails. The results are pr e s e nt e d Figures 150 to 152 f o r e ach tail surface. The data is pr e s e nt e d as measured; no adjustments have been made to lin e arize or sh J ft data fairlngs.

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• ENP _ _N NAGE _,%_ ._C_ . _TER DATA ": The flo w velocity magnitude and angularity w e r e measured in th e , :_ : vicinity of the R SRA horizontal tails w ith a triaxial hot wire probe g _" anemometer system. The probe .location was remotely controlled with a - _ traversing mechanism which permitted the probe _ , b e positioned w ithin _ a 3 0 " x 30" vert! al plane l o cated at fuselage station 664 (at sm angle. _ o f attack o f 0 degrees). F o r each data c o ndition e valuat e d, th e axial v e locity, the vertic c l vel o city, and the sidewash v e l o city w e re m e asured within the traver s ing grid. All velociW measurements were performed '_ wit ho ut the h o rizontal t r i s ( C onfigur a tions a n d ). : } FPBNp5WTTII FPBW7TII ' , Analysis o f t h e resulting data indicated th e presence of a :_ d es tabilizing do w nwash region ov e r a substantial portion of the lo we r _ _ trailing from th e Junction of the nacelle pylon and the nacelle b o dy. i' '. ; h o rizontal tail. The int e rf e ring do w nwash is created by a strong vort e x _ T h e str e ngth o f the fnt e rferring vort e x did not vary appr e ciably as th e .nacelle fairing w as m o dified. .i_ ; Flow survey s . wit h out t h e nacelle and nacelle fairing indicated _ , ; onl y min o r flo w distortion at the lower tail location due t , "wing llft d_ - _ " ._ induced d o wnwash. _'w _ ,_ : Instrumentation _:_ The triaxial hot w ire probe, which was used to measl_e the flow _ vel o cities, c o ntained three orthogonal wire elements . F_ch element had 2_ .

A I :EMO._TE_ INSTUMENTATION , _ .% X-Y-Y ' Plot'te_ "_ Anemometers i Lineari zers _ Digital Voltmeters 'Triaxial Anemometer Computer. _ ( , , _,_-- q REPR O DU CIB IL rrY OF TH E _.i'" ORIGINAL PAGE IS PO O B _::.

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a s ensitive wir e l en gth o f 1. 25 _llfmeters a n d a dia m eter o f five m icrons .

/ • The flow se n sitive wlre le n gth was li_ted by gold plati n g which c o ver e d the e n d of the wire s u pp o rt pro n gs and a small se gment of each wire end. The _ plating reduces t h e i n terference _,ong t h e t hree wires, and betwee n the s u pport pro n gs and w-lre s . Thls i n terfere n ce is so m eti m es experience d "_ in a skewed flow f_eld.

The pr o be w_res wer e heated _y . three c o nstant temperature anemo- : me t e rs . Th e v o ltage s requir e d t o _i n taln the c on stant temperature of the : wire s , a measure of th e velocSty n ormal to e a c h wlre, wer e passe d t hroug h , sig n al li n earizer9 t o a s peci a lized a n alogue c o m puter , Thi s comput e r / circuit transferred the ii n earlz e d v o ltages for the three wires fr o m th e ,.- probe coor d inat e system t o the wi n d tun n el coordi n a t e system. Thes e final ' ; _ velocity voltages were th e n us e d t o d rive t h e velocity ordinate o n X-Y " _ _' plotters , T h e conversi on from th e pr o b e c o oFdinate system t o the tu n ne l :, coordi nat e s y st em was required sin ce th e probe wir e t riad form ed a 54.7_ ' cone with respect to the tunnel axls system. I J The tr a verse m ec h anism w as drive n by two ½ h orsep o wer e l e c t ric '" mot o rs which vere mounted on th e trav e rse _r_me an d operated remotely 1 fro m the win d tun n el c o ntr o l room. Probe positi o n was mon it o red by i_r o _ ,_ pote n tlo m eters geare d to t h e drive m ec h anism. The output v ol tage s ig n als i from thes e potentiometers w e re wir e _ _ir e ctl y to t h e pl o tt e rs r e c o rdi n g t h e velocities and _sed to drive t h e position axis.

Test Procedure I _ The i n strumentati o n set-up pe r mitt e d co n tinuo us data acquisiti o n i as the probe we s traversed parallel to the horizontal tail plane. For _ _'

each mo del data condltlon(bod y &ttitude, and nacelle configuration), !

horizontal data sweeps were obtained at vario u s vertical probe :'heights in order to define the flow con_tions which would be e n cou n tered at different horizontal tail mounting positions. Since the mechanism was positioned du ri n g this test to braverse to the right and up from fuselage station 209 at zero angle of attack, velocity data below the f u selage could only be obtained at negative bod_ attitudes.

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T h e o n- l in e v elo city d a t a was continuously " m o nit o r e d f o r I significant trending inform a tion and data quality during a cquisition.

As a r e sult of thi s insp e ction, it b e cam e appar e nt ths_ th e calibration of ; th e h o t wir e el_n e nts w e r e sl o wly _rlfting w_th time. Sil_ce the sensitivity it of hot wir e prob e s is s u sc e ptable to drift from oil cohtsmination, and b e caus e the mod e l apparatus showed e vidence of surface oil, it w a s concluded • _ that th e wir e calibration shift was probably du e to o i l film build-up. Th e probl e m was corr e cted by periodically trav e rsing the probe to a position o utside the model flow _nt_rf e rence area and 1-eadJusting the wire overheat ratio's until th e orlglnal calibrations w e r e matched. Th e original cali- _ brations w e r e p e rformed by th e probe supply v e ndor. As a r e sult of th e '_ ov e rheat ratio adjustments, slight velocity scaling e r rors were intro- duced in t he data . F o r mo st run conditions, t h e e rror do e s not exc ee d : -+ 5% of q. The flow angularity data was not effected by the calibration s hift.

.-" Additi o nal e rr o r ms M b e introduced by t he traver s e system a nd its s u pp o rt struct u re. Figur e 15 3 s h o ws the e ff e cts o f the traverse on lift, _ _ drag, a n d pitc h ing mom e nt. T h e e ff e ct o n drag i s significant. Th e i_ _ anemometer data its e lf show e d a decrease in flow v e locity bey o nd buttline 1 3 0, o r within 8 inch e s o f the traverse supports.

The v e l o city data from s e lected c o nditions w e r e reduc e d to yi e ld ' plots of loc a l downwas h ( _ ), sid e w a sh (•), a nd dyn m mic pr e ssur e ratio (qT / q) vers u s full scale buttline. Thes e data ar e pr ese nted in Figures 154 thr o ugh 15 6 • The fus e lage stations corresponding t o th e data varies slightly with body attitud e since the prob e was trans- lated in th e hori z ontal / vertical tunnel axis plane. T he prob e , however, _ was locat e d at the fus e lage station of quarter chord of the bas e line low e r h o rizontal tail at z e ro d e grees of bod y attitude. .,, t o the large qua n tity recorded and the fact that the major tr e nd s w e re _ n o t substantially altered as nacelle fairing c o nfig u rations w e r e changed. _ A large porti o n o f t h e o n -li n e data o bt a i n ed was n ot r e duced du e The specific ru n con d itions selected for a na lysis were cho s en to illustrate the major characteristics of the wing-fuseSage- n acelle inter- fere n ce in the vicinitM of the horizontal tail. The data, reduce d to q_ / q, 6 , a n d 6 " , are pre s ented i n Figures 154 - 156. I n the s e figures, s _ d ewash is p ositive to the right, a n d the n o r m al flow, labeled downwash- is positive up. Thi s i s i n dic a ted on Figure 154a.

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'_ Thi s dat a consists of Run 527 (F P B _8W7TII a t _ = 0°, -5°, -i0 ° ), R un 5 28 (FPB_ W TTII at _ = - 5 ° ) an d Run 5 35 ( ?P BW7T II at _ = 0°, - 5 ° -i0°). Run s 527 and 5 35 c o ntras t the flow enviro n ment obtained with and wit h out n a celles w h ile runs 527 a nd 528 c o ntr as t t h e ma x im u m * va r iation in the n a celle interfere n c e e x perienced bet w een the tested c o nfi gurati o ns.

The flow interference measured for the no nacalle configuration, :4 Figure 156, indicates on l y minor distortion in flow argularity in the vicinity of the horizontal tail. Furthermore, the downwash which is prese n t is fairly constant throughout the buttline variation of the lower bas e line stabilizer span a n d assum e s the characteristic " . ' , _ _ . / 6 _ . a s sociated with wing induced d o wnwash. There is no indication , J , of discr e te vortex flow patterns emlnating from the ui0strea m b od y .

structure. With this flow enviro n ment, on]7 m inor pitch stability ' washout from the wing induced flow would be e xp e cted.

In contra s t, the data obtained with the _ nacelle fairing configuration : pr e s e nted in Figure 154, displ a ys c_assic v o rtex inter- < f e rence ov e r the inner 50% o f the lower horizontal tail p o sition. Th e -i, position of the interfering vort e x, approximately buttline 60, indicates a nd nacelle.

" _ " " L_ upstream formation at the Junction of the nacelle fairing ; The fact that the vo_tici_ is b e ing sh e d from the outer end o f th e nacelle fairing rather than the root Junction of the pylon a n d f u selage, - c a n a lso b e c o nfirmed fro m the sens e of r o tation o f the trailing v o rtex filament. For negative bod y attitudes, and zero d e grees of wi n g incide n ce, the vcrtex induced velocities are negative on the outbo a rd side of the vortex core and positive on the inboard wid e in d icati n g a clockwise rotation which is consista n t with outboard shedding from the n a c elle pylo n . The data ex sm lned shows that the waterline location of the _per tail configurations ( 3 60) is above the nacelle fairing interfere n ce region.

^ I nspectio D of the data obtained with _ . 8 for aagles of attack of 0° - 5 9 , and -i0U rev e al s that the strength of the i n terfering vortex increases significantly as th e body attitude b e comes m o re nose d o wn.

Because of the vortex location and rotation sense, a large _ / de_ is pr,sent over the inner 50% of the baseline horizontal stabilizer which would seriously degrade the stabilizer effectiveness.

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a t -5 ° Da t a p r esen te d i n F i gure 15 5 for t h e Np n a celle p y l on co n fi gura ti o n ,: bod_ r a tt i tude is simil a r to th a t obtaA n ed wit h t h e -N _R pylo n : _ , and illu s tr a tes the consi s t an cy of t he interfer en ce , despite%u-bst a nti a l v a ri a tion in the det a il e d pylo n geometry .

Flow _t a from all of t h e surveyed nacelle pylons sh _ed large _i i d y n amic pr ess ur e i n creases as t he pr o b e en t e r e d t h e n acelle flow r e gion.

P e a k d_ n ssmlcpressures up to 2.2 times t h e fre e stream q were measure d at t h e - _ c en ter of t h e na c e l le e ff l ux. _igures 1 5 7 - 15 9 ar e co n t o ur map s o f q_ / q f o r angles o f a tt a c k o f 0°, -5 ° , an d-I0 °. A s discuss e d i n t he following _ paragraphs , th e q variatio n p a rti a lly n eg a ted t_ e flow downwas h e ff e ct s _ an d a ugm e nt e d t h e pitc_ stability w h en t he n acelle flow i_pi_g e d . o n t l., e _'_ .

low e r horizo n tal st a bilizer. ¢ Figures 160 through 162 w e r e construct e d from th e flow data obtained w_t h t h e _A na c e lle f a iri n g to evaluat e t he d e gr ee of s tabilizer ,_:_ lift was h out du e to _ he v ortex i n duced flow g e ner a ted by t h e pylo n .

Figure 160 s h ows t he tru e spanwlse angl e of a tt a c k distribution of t he _ stabilizer a t bod_ a ttltudes of 0° , -5 U, and -i0 ° . Inspection of F igure _" 160 i n dic a tes t ha t two _ s tinct flow r e gimes e r e experienc e d by t he i i nne r and o uter portions of t h e t a il. Ou t, _b o ard , flow i n t e rference d ue to J wing d ownwas h (_pwas h a t o _ = -5 and-I0 _) is pres en t. Inbo a r d , .h owever, i_ sever e upflow occur s , furt h er reversi n g t h e true a ng le o i a tt a c k gr a di e nt _ from t ha t of t h e g e o me tric st a bilizer an gle for a ngles of a tt a c k of -5 an d , , ; 1 _ -I0 de gr ee s . F or cl a rity, t h e pl an form and position of t h e bas e li n e sta_Ii_z e r _ n oted o n t h e plot. It ca n b e s ee n that the tap e re d pla n form j_ of this s tabIi_zer _tk _ n bo_rd _r e a weig h ting serves to exaggerate the adv e rs e pitch st a bility e ffect of t h e na celle f a iring interf e re n ce. __ F igure 161 illustrates t h e r e lative i n t e gr a te d lift of the _ s tabilizer as a functio n of body att i tu d e , Figure 162 shows the relatio n - sh ip b e twe en t h e mea n s tabilizer d ownwas h ( _ ) and _ he stabilizer " a r_le of a tt a c k for n eg a tive body a ttitudes b e tw e en O - and -I0-, calcul a t ed " fro m velocity i n tegrations over the baseli n e l ow e r tail planform. Data is pre s e n t ed wit h an d without consideratio n of the q e ffe c ts fro m the n a c ell e flow. The me a n u_wash flow angles were c a lculate d wit h a n area weighte d i n tegratio n of the flow angularity over the stabilizer s pa n . For the co n sta n t q case, the axi a l flow was assume d e q ual to the free stream velocity; whereas the d ata with dy n amic pressure v a riations was c a lculate d with q w e ighting _ ad dition t o t h e st a bilizer _re a w e ighting. While both tr end li n es i n d i cate increasi n g _wash betwee n 0_ an d -8 _, the effect of the a dd itio n al n a celle flow dy n ami_ pres s ure i s sufficie n t to eliminate the net stabilizer lift washout at -I0_. The figure I n&icates th a t a pproximately 7 3 % of the t a il effectivene s s (I - _ /d _ = .2 7 ) is washe d out by wi n g n a celle i n terf e re n ce between 0° and -5 °. I nclu d e d o n Figure 1 6 2 is the d ownwas h co m pute d from Phue I bal an ce data, a nd shows excelle n t : c o rrelation with _ he ane mometer da ta.

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PA @_ I S A 2g n _V i _ ' ,_, _-_ _,_,_ . _ T _v . _ ; _,_w ....... _ * ._-_ , _ _,'_'_ _'_ ' ,_' z _ _ , m._ °'m_"_ m _ u , m _ oaw_, , ,m,, REPORT NO. SEB-7 2 011 _ Sikorsky I:llr c r a ft _, The d o w nws a h calcul ated fza m t h e an emu met er da t a d oes n ot i n clud e '_ , t h e effects of i n duc ed flow ge n er a ted by t h e tail. A n a a_g ona l d ownws s h comp o ne n t will be caus e d b y the tail_ tlp vortex a n d t h e n o n -u n iform : lift di s tribution o n the te / l. The m a gnitude of t, _ ese effects h e s n ot be en

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, i CONCLIISIONS AND RECOMMENDATIO NS ] Conclus ior_

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- - t i. T h e RSRA empennag e must h e modified t o provide th e d e sired stability l e v e ls. Th e directional stability is best i_proved by increasing • vertical tail span, since this places the added area in a more o ptimum fl o w e nvir o nment. Longitudinal stability is im p rov e d significantly by _ increases in lower h o r i zontal tall span. Horizontal tail a z Lhe dral , im p rov e s l o ngitudinal stability. T he inboard s e cti o ns o f t he tail J _ .. " t en d to d e grad e s tability and th e r e f o re ar e a i n this regi o n shoul d : _ _ b e minimi z_d.

2 . An upper horizontal tall i_pr o ves both direction and longitudinal : stability a n d , wit h out an anh e dral l o w e r h oriz on tal, is necessary f o r stability b e tveen angles of attac k o f + 5 d e gr ee s.

3 . Th e pow e r e d nac e ll e s caus e a d e -stabillzing pitching mom en t. Pr o p e rly : scal e d n ac e lles will r e duce th ls e ffect.

4. Th e d at a pr esen t ed is suffic ie nt t o furt he r analyze tail des ign if ne c essar y .

L _ 5 . T he ba s e l ine n a celle fairing (Nps) h ad t he best o vera l l c h ar a ct e ri s tics -' " T o f t hose tested.

6. Mod e l d_sy m metry has ca u sed some discrepanci e s in the test dat a , e specially t h e rolling mom en t param e ter f o r angles nea r wing stall.

_ Recommendat ions _ , _ , I. Th e t e st da t a frnm the La ngl e y Res e arch Ce nter t es ting o f t h is m ode l _ s ho ul d b e c o mpar e d with t he data pres e nt e d he r e in. Th e pitching _ mom e n t trend with nacelle slze s h o uld b e car e fully analyzed.

2. The flow behind th e nac e ll e support fairing impacts on tail p e rformanc e ' and sh o uld b e furth e r studiad in o rd e r t o minimize th e fairing's _' , impact o n the aircraft.

B . Define mod e l disymm e try and e _aluat e th e r e s,Alting e ff e ct s on balanc e d a ta.

_. Eval u ate d ata tr en ds with soli d , pow e r ed , a n d ring nac e l l es a n d d e t e rmi n e f o r mos t repre s entative c o nfiguration for u se in fut u r e testing.

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_ REFERENCES I i. Gree s on, J.O., "Wind Tunn e l Tests of a Sikors_ 1 / 6 Scale RSRA ' _ , _, Helic o pter Model in the UARL Large Subs o nic Wi n d Tunnel - P h ase I"_ _- , _ UARL Report Nh32377-i, O c t o ber 2 1, 1974.

2. Griffin, T. , "Wind Tunnel Tests of a Sik o rs_ 1 / 6 Scale RSRA Heli co pter Model in t h e UARL L arge S u bs on ic Wi n d Tun n el- P h ase I I " ; _ UARL Report Nh 3 2409-1, November 22, 1974.

3. Werner, J.V. end Flemming, R.J., "YUH-60A Quarter Scale Wind Tunnel Test Report," SER - 70531, M_ l, 197 3.

_ " 4. Gifford, J.A.M., "One Twelve Scale Wind Tunnel T e sts o n the AH-3 % : (S-67) Demonstrator A ircraft," SER-67000, June, 1970. , ' 5. Simm o nds, A.W., "Wind Tunnel Test s of a S ik o rs_ CH-53A H e lic o pter _ Model," UARL R_ort A430095-I, May 15_ 1963.

6. Logan, A.H., "A Method of Pr e dicting the Surface Velocities and , Pressures on a Helicopter Fuselage (Progr a m YI79B) - Its D e velopment

- _ " sm-5o867, June 7 , 197_ . '_

_ and Correlation,

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b Run 77 - WN, Iw = 10 , o C = 5 Figur e i 0 7. Wing Oil Flow Patterns - Unp o wered.

h72 PAGE REPF,ODUCIBILITY OF TI4 B " ' ORIGINAL PAGE IS POOR d Run 78 - W , Iw :lO , _ : 5 \ Figure 107 C o nt ] nued P AGE

\

\ f Ibm 8() - WIN , iw .= 15,_ -0 e Figur e i07 - C ontlnumi _ 0 PAG Ir . l

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!1 _ Run 81 - W l, Iw = 1,5, O ( . , = 0 ; ) Figure 107 - Continued ._ P A G - "_ R EP RODUC I BILI TY O F TH E " ' aRTGT , <AI , PAGE IS PO OR d j Run _ ( _ - W l, I w : 15 , _ : 30, =_" = 2.5 Figure i07 - Continued

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1 Ihill _ ) 2 " WIN, I w :- 15 , _= 0 } . Fl'gurei07 - Colltlnue,1 }l "¢" " [ P A GE r_ Run 97 - WIN, 1w =1 5 , G,C=O

U REPORT N O .S HI { - 72 0 11

Sikorsky A ircra f t _v.._ . _N ,,, o . _ ....o o.. _ . . . _

A , R

"L R un 1 ( )_) , W IN +Spo i ler , I w = 15 , o ( _ = 0

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r Run I () _ ), W I N -+ Spo i k ' r , Iw : I_ , o C = 0

Figure ] 0 '( - C,_ t , Ell t_ 0 0 PAGE !

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Run281 , W s N , Iw= lS , 6 -- 30 ) Figure 108. Wing O 1 1 Flow With Wing Fences Installed - Unpowered.

hSl

Sikorsky l::lircraft-,- -_ , o_i: , _ --.-- R:, o , , No . SER - 72011 i

a R u n 1 2 4 , WIN p , W i n dmil l, I w = 15, oC= 0 l

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V _ d ' b R u n 1 24 . W INp . Wi nd ll t i l l, I w - 1 5. oC - _ (1 Sikorsky A ircraft ...... _ _o_ , .o . , . , .. . . . U _ .. _ ..,_ Rep o. t NO S l ;r -72 0 11 c R u n 12 , 5 , W lNp , I ' rim , iw = 1`5 , o C = I}

t

d . Run 12_ , , \ylxP , I ' rml , I_, 13 , o c I ) '! 5 e P, un 126 , WlNp2 , Windmill , Iw = 15 , ¢: _ = 0 :...

k '. } ,. 0

f R un 127, WlNp2, Trim , Iw = 15 , o ( 2 - 0

Figure 109 - Continu e d ,_

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PAGE h Run 148, WNpl +Splitter, Trim, Iw = 15 , _(_ = 0 } Figure 109 - Contlmmd [ t185 PAGE

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j Run L 49, W2Npl, "rrun, Iw = 1,5, _ = 0 Figur e 109 _ C o ntinued O h8 6 ) PAGE l Run L 52 , W3NpI , T r i m , I w = 15 , ¢ _ = 2 5, _ =0 Figure 109 - Coni. i n ucd P AGE / ' | ' _,

L

. _ Sikorsky A ircra f t = - == ---- o_ - = -- - - = _¢p o. r NO. S E R -72 0 11 , ,m Run 155, W 3 Npl , Trim, Iw = 15, _ = 0 =, { . , .

Figure i09 - Concluded P AGE t ' / I ' .

J • . , - . = 0° ,6 = 15° b Ru}i ) _ ' ( !_,_p5_,_, _rir , 1 Power, iw , ,_ Figure Ii0. Wii_g Oil Flow ' , ;ithout F_tlees _ Powered.

P AG E i \ .< = 2.5° C . Run h73, NpsW8; Trim Power, iw = 15° , d. Run 475, Np5Ws, Trim Power, iw = 15° , F = 300 _ 2.50 Figurt II0 - Concluded.

t Rt_pR(_DUCI:B T L I TY OF THE _, ( < L _ i . . ,A . _ PAGE IS POOR . , • ' t .

b. Run 4_4, *' F5 ' 7 ' ' .*.>' i : " r" , _ , ', ' tr , _ = , o w ' _ = 15° "A Figure i[I . Wing {,i] !. ' l_ : w With ,"'enc ,, :: - r ' owercd.

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Sikorsky A ircraft ....... o. ..... o, . q c ............ REPORT 'N O . SHR-720! l

f _- " Trim ]'owr, " i = 0° = < = ]5°, iN = O° e, Run 493 , -P5_7 , ' w *'P5}_ 7, o _,o, 50 f. Run 495, ' '" "' I = 1 "_ = ] J l ; = li Figure Iii - Cox=e.lu,ici P A GE

t _t ' _,'_"_ m ODgC_q l TY OF THE

d I . ) c Run _ ) 4 - WIN , Iw = 1 5, o d" - -h d - Run _4 - WI N , I _ , 1.:_ , _ -4 F i g ure 112 - C e mt I n'4_ <I L

r

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e Rt m 9 4 - WI N , Iw = 1 5 , c _ - 2 I " f Run _ ) 4 - WIN , ! w = l h , _ () Fi _ u _ 'e 1. 12-Con t tn u ed 0 PA GE / ,

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AN e Cm A P t CC , _ . _ _ T I r _

. _ S ikorsky A ircraft ......_ ' _"° A_ REpo_,.NO ._;Er - ' / ;. ' ;il

g Ru n 9 4 - W l X , Iw !_ , c : C - -2 "

ws, : o , _ = o

b Run 416 , Iw • i Fi g ure 113. Win g Flow Vs An_'l e < _fAttack, i e • = 0 DeE . ,_ PAGE U REPORT NO. SER - 72011 AIf _ CR A FT C O R PO m A ' _ K)N

. . _ , . ) Sikorsky A ircraft °'_'_ ' _° ' ''° g_

/ %, -) r - ,, : f R un 416 , W5 , Iw () , c_ = 15 i F l gure 113 ,.Coneluded P AG E REPRODU C IBILITY OF THE _ : _ 01R I OINAL PAGE IS POOR ; fi : fj_ • • a R u n 413 , W5N, Iw =0 , _ = -5 b Run 413, WsN, Iw =0, a C =5 Figure lib. Wing and l;ace]le Fl_w Vs. Ang]e of Attack, iw = O D e g.

5O] PAGE ]I • = : _ I ___ JL ; i_ I -- --_ _- Z " 7" Z? =-== - _ f Run 4 1 -1 , WsN , i w - ( ) , ¢_ _- 0 I "lg_l_'," 1 1 1 _ - Co _l + , , lnu c ,l PAGE 1 it_ .

: L _ S E R - 7201 1 Sikors .ky I::llrcraft ' - - _ ' - "_: "_ '- " - , E ,o, , ,o . I . _- , ?

i z : g R un 414 , W5N, lw =0, c _ = 5 f , a g _ i h Ru n 414, W5 N , Iw = 0, _' = 7. 5 Fig u r e llh - Con t inu e d "_) PA G - "_ REPRO D UCIBI L I TY O F T HE ORIGINAL P A G E IS POO R * _ - l i, ' U

Sikorsky A ircraft - - = - _ - , °-=_, =-- , , , - .Ep o .v N O .SE R - 7 2 0 11

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i R u n 414 , W5N , Iw = 0 , _ = 10 ., j Ru n 4'14, WsN , Iw = 0 , o C = 1 2 .5

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Figure 114 - Concl u ded PAGE r . i El_ ..... i

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• a Run 41 5 , W 5, Iw =0, _ o = 15 • I K i1, 7 , .

d Run 415, W 5 , i w - t ) , / _ 0 2 . 5 • _ gigurc [1 5 - ,Jont, inu,',l PAGE

f Run4 l S, W 5, Iw = O , _ = - 2.5

Figu re 11 5 - Con ti n u ed 0 PA GE

197601 3 0 3 1

Sikorsky Rlrcraft - -----, - _ -.-- -. - ,_P o nT ,o. SE R- 72011 [ /

, e [ Ruo 4 15. W 5 , i w _ O. f , _ - 15

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Fi gure1 1 5 - Co n clu d ed .i k b Ru n 4 12 , W 3N , I w - -0 , _ ++ [ 0 + t'i_;,£:", + ii+. '_t++ , _ + '. +z,4 ),_,+' , ' L, : Y + "+' ": . ":° ' '- , , :" ":' ", . ',.', I = " ; , +,..,', P A_ ( )Dt:t'll_ll.l't'_." O P 'f L IP, 1_. t+:lP I_( • , ,+>b . t +'. \I. ILk_;t_] I._ P I)OP,.

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1976013031 - 54

, % :

f i _ L '" 'm

Sikorsky I:lircra f t =_= , o ., U

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• " _

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, _ , r i , _ ? t " 0 t , r _4 c _ '% 6a 1 1 :' g Ru n 41 2 W_ N, Iw = 0 , _ = - , 5 i • i Run 412 , W 5 N' Iw = O, _ ' = -15

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/ Figure 116 - Concluded

51__5

PAGE

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Sikorsky A ircraft = v . o.o.u.,.. o .,. o . . . . _ . o . .. = . .E_OR, NOSER-72011

e

• , _D Sikorsky I: : lircra f t U. ° o . EP ORT N O . SE R -7 2 011

a R un 303, WS N p4 , Iw =0 , _ = 0

. \

d Run 303, W 5 Np 4 ' lw = O, e C ' = 15

: 3

Figure 118 - Continued PAGE I Run.303 , W , 5 Np4 , !w : 0 , ¢ , C = 20 PAG E "" _ _ _ ' _ , _ ............ t ,ilmlr _ 1 II ..... _ Jl I l l I _ I -_,_ , 7 _ I I II III liJ 2 a R u n 6 5 1 , < = 0 ° , q = O ps f, Fan R PM =111 5 0 " ,9 I \ b R un 65 1 , " < = -!0 " , ( I - - 55 p sf, l 'an RI*M = 1 11 5 0 F isure 1 1 9. Empenz_ a g e Flow - Conf l gu r at[on FPB_IpbW, { T36BT.

52 2 t "

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: _ - : {

¢ Ru n 651 , '<= -,_ , Cl = 5_ p -_f , Fa n RPM = i1 1 50

J R un B , _l , - < - (), * q = 55p_f , Fa n RPM = i11,_O Figure '!9 - Cor:tt_ued 5 2 "_ P A GE

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a ,l_.p f Ru n ( _ 1, < = 7.,5 _ q_ 55ps f, Fa n R I 'M = 1 11 5 0 Fi g ure I19 - C ont l nr e d . _ 52_

R E FR U D U C_ I L r rY O F TH E !

O RI GNAL P AGE I S POO R m.. # __ .... , ,i o Sikorsky l ::llrcraft . _ . -.. o . _.- ° =° - . .= _ . ,_ .o, T . o .._l . : =_ - 7 : _ l l . _L.,_ P &GE

1976013031

"_ l:i_,ur e 11 9 - Continu e d PAGE f , , , . ,

Sikorsky l::liroraft = v. = U R e pO.T N O . SE R -7 201l

O B _ ITI[O AI RC RAFT A_ CON _ A T K)N _ _1 ( ' 2 /

m Ru n 651 , "_ = 25 ; q= 55p , q f , F an R PM = 11150 i

Figure 119 - Concluded 0 t PAOE "i_

a Run 650, < = - 13 "

Figure 120 . _anponn , _ge Flow - Conf[guralloi_ FPBq ,W T B . : ' t'> 7 37 T e PAGE o - ....... , " _ , , j ....... , \ . _tli f Ibm 05() , _ : 2. : 3 ° " _ Fi g ure 1 2 0 - C on t iuued ?

P A GE i t

Sikors A ircraft o ,v , ._, o. o ., . .o. . o . - . =_ .o . . . _

kY U REPORT N O. S ER- 7 2 01 1

. A_ L . _

I Figure 120 - Continued 0 PA O E jr _ b i Run 650, .,r= 15° j Run 65 ( I , _ =!7 . ,_° 533 .

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t ..... ...... ............. 1t! t t i it:i

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. i!iti,:: ::. :: :+::_,+ ,:..: . . +_:: _:,: : :,:::.!::::.::::!_ii_i!i_tj+i' i :: _i!_i:._ii_+!+i!_il i: : !_:::+_ : iii :: : _!! : : : :_I._ t : ::, t J : I J : :l: :t::::t:: : :lt = +7-.

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.... 11,+ ......................... +, 't ......... _ ....... + ..........................

, ++ + [ + + ,, + + ,, , , , , + ,, , + ,, ........ , ,I , _ ,, , +....... + , , ......... _....

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i;! _ +_!I_ _ii_ _ ............ ................................ '!+++

+i_++: ,+ ++++ ,'i++q+ +"+ _ .++x._.

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