Advancement of proprotor technology. Task 2: Wind-tunnel test results

ADVANCEMENT OF

PROPROTOR TECHNOLOGY

TASK 11- WIND-TUNNEL TEST RESULTS

(NASA CONTRACT NAS 2- 5386)

REPORT 300-099-004

N S A C - 1 1 3 6 3 ) hCVAHCEBISI 0 3 PROP601'CR N74- 11634 ZECEUOLOG1. T A S K 2 : E I I D - T U d l E L TEST 229 p HC B E S U L T S ( B ~ l i iielicopter Co.)

$ 1 3 75 CSCL 01i Unclas G3;C2 23271 Page I - 1 I1 I - CrSCRIPTIOX O F TEST HARIHARL A , PWIPWIVR AND W T R O L S 1 , D e s c r i p t i o n 2 , N a t u r a l F r e q u e n c i e s 8. =ST STANDS 1, Perfomarrce S c a d 2 , Dynasic T e s t S t a n d s

C . INSTRWEWAT ION

IV. P E R F O r n C E A. GENERAL B- T W DESCRIPTION IV- 1 D . SPINNER TARE I#TA E R F m T S AND OORRELATION 1. Hover F l i g h t Mode 1w-4 2 . H e l i c o p t e r - C o n v e r s i o n F l i g h t Mode 1f'-5 IV- 5 3. A i r p l a n e F W . i g h t Mode

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V . DYNAHIC STABILITY A . MODELING T?CHNIC:IJES i . J u s t i f i c a t i o n ot S e w i s p a n C a n t i l e v e r Xode 1 2 . ftedscea S t i f f n e s s T e s t Stand 8. TEST PROCEDURES C. MEASURED STABILITY CHARACTERISTICS 1. D e s i g n - S t i f f n e s s T e s t S t a n d 2 . O n e - F o u r t h - D e s i g n - S t i f f n e s s T e s t S t a n d TABLE OF tXSTE3TS - C o n t i n u e d Page

D. OORREUTION OF THEORY WITI: MEASUR€D v- 5

D W I C STABILITY CHARACf&KI!3'ICS VI, BIADE FLAPPIMG 1 - M e a s u r e d D e r i v a t i v e s 2 , C o r r e l a t i o n 3f Theory w i t h M e a s u r e d FLa pp i ng B, FLAPPING OONTROLLE R INVE-ST1 CATION 1. T e s t R e s u l t s 2. C o r r e l a t i o n of T h e o r y w i t h N e a s u r e d S t a b i l i t y a n d Perforaance v11-1 I BLADS AND CONTIEDL SY!7EN LOADS v11-1 B, MEASURED BLADE XMD CCSFf(lDL LOADS v11-1 - V I I -2 1, Loads i n H e l i c o p t e r Mode, a m , * - +75 Degrees 2 . C o n v e r s i o n Mode, amS~ = +60 D e g r e e s v11-2 +30 Degrees, a n d + L 3 Degrees 3. A i r p l a n e Mode, Q ~ S T = 0 Degrees v11-2 C , CORRELATION OF THEDRY WITH MEASURED V I I -3 OSCILLATORY LOADS V I I I . NOISE AND VIBRATION A . NOISE 1. V i b r a t i o n maracteristics of t h e T e s t S t a n d s ? u r i n e N o r m a l P r o p r o t o r Ope r a t i o n 2 . V i b r a t i o n During S t ~ p - S t a r t O p e r a t i o n 3. C o r r e l a t i o n of T h e o r y uf t h M e a s u r e d V i b r a t i o n 1x-1 TX. CONCLUSIONS I X - 1 A . GENERAL 1x-1 B. PERFORMANCE

TABLE OF (X)XTEXTS - Continued

P a g e C, DYNAMIC STABILITY ix-2 D, RLADE F I A P P I N C ix-2 E, B U D € AXD CO?*TEEOL S Y S T M I-OADS TX-2 F, WOISE AND VIBR4TION 1x-3 X . RclFFRENCFS X - 1 APPePDIX ?3OHeJcYATURc, POWERED T E n W T A L I S T I N G S DRAWINGS ILLtfSTRAT ION S F i g u r e Page 1-2 P r o p r o t o r D y n a m i c T e s t i n NASA-Araes 1-1 L a r g e - S c a l e T u n n e l P r o p r o t o r P e r f o r a a n c e T e s t i n M S A - A m e s L a r g e - S c a l e T u n n e l , H e l i c o p t e r Hode f r o p r o t o r P e r f o r n a n c e T e s t i n XASA-Aaes L a r g e -Scale T u n n e 1. A i r p l a n e Mode XV-3 C o n v e r t i p l a n e D y n a a i c n o d e 1 T e s t i n g B e l l H o d e l 266 Army C m p o s i t e R e s e a r c h A i r c r a f t E k l L M e 1 300 T i l t - R o t o r A i r c r a f t P r o p r o t o r C o l l e c t i v e Mode N a t u r a l F r e q u e n c i e s , N o n r o t a t i n g P r o p r o t o r C y c l i c Mode N a t u r a l F r e q u e n c i e s , N o n r o t a t i n g P r o p r o t o r C o l l e c t i v e Mode i z a t u r a l F r e q u e n c i e s , R o t a t i n g P r o p r o t o r C y c l i c Mode N a t u r a l F r e q u e n c i e s , R o t a t i n g MSTRAN S t r u c t u r a l Model of P n w e r e d T e s t S t a n d NASTRAM S t r u c t u r a L Model of Dynawic T e s t S t a n d Hode S h a p e s - Dynamic T e z t S t a n d 111-7 D y n a a i c T e s t S t a n d S t a b i l i t y B o u n d a r i e s 111-8 IV- 1 Computer R e p r e s e n t a t i o n o f t h e 2 5 - F o o t P r o p r o t o r B l a d e P r o p r o t o r B l a d e S e c t i o n D a t a , A i r f o i l S e c t i o n No. 1, BLade S t a t i o n 0.075 t o 0.45 P r o ~ r o t o r B 1 ~ d e S e c t i o n D a t a , A i r f o i l S e c t i o n No. L , B l a d e S t a t i o n 0.45 t o 0.70 ILLUSTRATIONS - C o n t i n u e d P r o p r o t o r B l a d e S e c t i o n Data, A i r f o i l S e c t i o n No. 3, B l a d e S t a t i o n 0 . 7 9 t o 0.90 P r o p r o t o r B l a d e S e c t i o n Data. A i r f o i l S e c t i o n N o - 4 . B l a d e S t a t i o n 0.90 t o 1 -00 S p i n n e r Drag T a r e , A i r p l a n e F l i g h t Mode IV-7 S p i n n e r L i f t T a r e V e r s u s Mast A n g l e o f A t t a c k , H e l i c o p t e r - C o n v e r s i o n Mode I V - 8 S p i n n e r Drag T a r e V e r s u s Mast A n g l e o f A t t a c k , H e l i c o p t e r - C o n v e r s i o n Mode IV-9 D i ~ e n s i o n a l H o v e r P o w e r R e q u i r e d IV-LO Nond i a e n s i o n a l H o v e r Poue r R e q u i r e d IV-11 N o n d i m e n s i o n a l H o v e r Power R e q u i r e d IV-12 H o v e r i n g F i g u r e o f Merit V e r s u s H e l i c o p t e r T h r u s t C o e f f i c i e n t / S o l i d i t y C o r r e l a t i o n of N o n d i a e n s i o n a l H o v e r P o w r R e q u i r e d N o n d i a e n s i o n a l H e l i c o p t e r - C o n v e r s i o n P e r f o m a n c e N o n d i a e n s i o n a l H e l i c o p t e r - C o n v e r s i o n P e r f o r n a n c e N o n d i a e n s i o n a l R e l i c o p t e r - C o n v e r s i o n P e r f o ma n c e V o n d i a e n s i o n s l He1 i c o p t e r - C o n v e r s i o n P e r f o r m n c e N o n d i r n c n s i o n a l H e l . i c o p t e r - C o n v e r s i o n P e r f o r m a n c e N o n d i a e n s i o n a l R e l i c o p t e r - C o a v e r s i o n P e r f o r m a n c e S h a f t H o r s e p o w r V e r s u s T i p P a t h P l a n e A n g l e of A t t a c k , He1 i c o p t e r - C o n ~ e r s i o n Mode Fipu re Page I V - 2 1 Shaft Horsepower Versl s Tip Fath P;ane IV-27 AngLe of Attack. ;ieli,-opter-Conversi n Node Comparison of Measured Horsepower with Calcalated Horsepow~r, Hel-icopter- Conversion Mode Propulsive Force and Ff f iciency Versus Horse pow r Propulsive Force and Ff f i c iencv Versas Horse pow! r Propulsive Force and Ff f iciency V2rszc Horse poue r Porpulsive Force and Efficiency Versus Horsepoue r IV-33 Propulsive Force an6 Efficiency i'ersus Horse pow r IV- 34 Propulsive Force and Efficiency Versus Horsepower Propulsive Force and Efficiency ircrsus Horsepower Propcis ive Force and Efficiency Versus Horsepower Propulsive Force and Efficiency Versus Horsepower Propulsive Force and E C f iciency Versus Horse powe r Propulsive 5f f iciency V ~ r s u s Power Coefficient, 129 Ynots Propulsive Yff iciency Versus Power Coefficient and Advance Ratio, 160 Knots Propulsive Efficiency Versus Power Coefficient and Advance Ratio. 185 Knots

ILLUSTR1TIdNS - C o n t i n u e d

F i g u r e P a g e IV-36 Cornparisoir o f Y e a s u ~ e d H o r s e p o w e r w i t h IV-42 C a l c u l a t e d H o r s e p o w e r , A x i a l F c i g h t Mode Root L o c i o f C a n t i l e v e r w i n g C o a p 3 r e d t o C o m p l e t e A i r c r a f t R o o t L o c i C o m p a r i s o n o f R o o t L o c i f o r S e a i s p a n C a n t i l e v e r ru'ing a n d O n e - g o u r t i l - S t i f f n e s s T e s t S t a n d Wing Beaa a n d Chord D a a p i n g V e r s u s S i m u l a t e d A i r s p e e d V i n g B..aa B e n d i n g Mode S t a b i l i t y V e r s u s A i r s p e e d T r a c e s of Decay o f Wing Beam V i b r a t i o n s f o r O n e - F o u r t h - S t i f f n e s s T e s t S t a n d Wing B e ~ m F r e q u e n c y ~ n d h a p i n g V a r i a t i o n w i t h T u n p e l V e l o c i t y f o r S t a n d a r d S t i f f n e s s T e s t S t a n d Wing C h o r d F r e q u e n c y a n d Dampiag V a r i a t i o c w i t h T u z q e l V e l o c i t y f o r S t a n d e r d S t i f f n e s s T e s t S t a n d Wing T o r s i o n F r e q u e n c y a n d Damping V a r i a t i o n u i t h T u n n e l V e l o c i t y f o r S t a n d a r d S t i f f n e s s T e s t S t a n d Wing B e a a F r e q u e n c y a n d Damping V a r i a t i o n w i t h A n g l e - o f - A t t a c k f o r S t a n d a r d S t i f f n e s s T e s t S t a n d Wing Beam F r e q u e n c y a n d Damping V a r i a t i o n w i t h P r o p r o t o r RPM f o r S t a ? d a r a S t i f f n e s s T e s t S t a n d Wing Beam F r e q u e n c y a n d D a n p i n g V a r i a t i o ~ w i t h T u n n e l V e l o c i t y f o r O n e - F o u r t h - S t i f f n e s s T e s t S t a n d Wing C h o r d F r e q u e n c y a n d Damping V a r i a t i o n w i t h T u n n e l V e l o c i t y f o r O n e - F o u r t h - S t i f f n e s s T e s t S t t a d Wing T o r s i o n Frequency a n d Damping V a r i a t i o n w i t h T u n n e l V e l o c i t y f o r O n e - F o c r t h - S t i f f n e s s T e s t S t a n d 004 v i i i

ILLUSTRATIONS - C o n t i n u e J

Page v - 2 0 Wing B e a m F r e q u e n c y a n d Damping V a r i a t i c n w i t h A n g l e - o f - A t t a c k f o r O n e - F o u r t h - S t i f f n e s s T e s t S t a n d Wing Bea% F r e q u e n c y a n d Damping V a r i a t i o n w i t h P r o p r o t o r RPM f o r One - F o u r t h - S t i f f n e s s T e s t S t a c d , 1 5 0 K n o t s W'ng Beam F r e q u e n c y a n d Damping V a r i a t i o n w i t h P r o p r o t o r RPM f o r O n e - F o u r t h - S t i f f n e s s T e s t S t a n d , 1 7 0 K n o t s Wing T o r s i o n F r e q u e n c y a n d Damping V a r i e t i o n w i t h P r o p r o t o r R P M f o r O n e - F o u r t h - S t i f f n e s s T e s t Starid P r o p r o t o r L o n g i t u d i n a l a n d L a t e r a l F l a p p i n g V e r s u s A n n l e o f A t t a c k , S i m u l a t e d 2 6 5 K n o t s , 458 RPM C o r r e l a t i o n o f F l a p p i n g T h e o r y w i t h M e a s u r e d T o t a l F l a p p i n g D e r i ~ a t i v e P l e a s u r e d R e s p o n s e t o S r e p I n p u t f o r V a r i o u s V a l u e s o f I n t e g r a l Gain a t a S i m u l a t e d 185 K n o t s Summary o f S t a b i l i t y a n d R e s p o n s e C h a r a c - teristics w i t h I n t e g r a l G a i n F e e d b a c k G u s t R e s p o n s e f o r V a r i o u s V a l u e s o f L a g g e d P c s i t i o n G a i n a t 92 K n o t s M e a s u r e d F l a p p i n g R e s p o n s e a t S i m u l a t e d 185 K n o t s a n d 2 6 5 K n o t s , f o r Combined I n t e g r a l a n d L a g g e d P o s i t i o n G a i n s ( S t e p I n p u t ) R e s p o n s e t o S t e p G u s t f o r Combined I n t e g r a l a n d Lagf,ed P o s i t i o n Gains ( 2 4 0 K n o t s , S i m u l a t e d ) R e s p o n s e t o T r i a n g u l a r P u l s e G u s t f o r Combined I n t e g r a l a n d L a g g e d P o s i t i o n G a i ,s (240 K n o t s , S i m u l a t e d ) C o r r e l a t i o n Between T h e o r y a n d M e a s u r e d L o n g i t u d i n a l F l a p p i n g F r e q u e n c y R e s p o n s e C o n v e r s i o n C o r r i d o r Based on M e a s u r e d i3lade L o a d s

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ILLUSTRATIONS - C o n t i n u e d

P a g e V I I - 5 B l a d e a n d C o i r t r o l L o a d s Waveforms i n H e l i - c o p t e r , C o n v e r s i o n , a n d A i r p l a n e Modes S p a n w i s e O s c i l l a t o r y B e n d i n g Moment D i s t r i b u t i o n i n H e l i c o p t e r Mode B l a d e S t a t i o n 5 2 . 5 Beam O s c i l L a t o r y B e n d i n g M o a e n t V e r s u s T h r u s t i n H e l i c o p t e r Mode S p i n d l e C h o r d O s c i l l a t o r y B e n d i n g Moment V e r s u s T h r u s t i n H e l i c o p t e r Mode B l a d e P i t c h - L i n k O s c i l l a t o r y L o a d V e r s u s T h r u s t i n H e l i c o p t e r Mode O s c i l l a t o r y L o a d s V e r s u s A i r s p e e d .

H e 1 i c o p t e r Mode B l a d e S t a t i o n 52.5 O s c i l L a t o r y Beamwise B e n d i n g Moment V e r s t i s T h r u s t i n C o n v e r s i o n Mode Yoke S p i n d l e Chord O s c i l l a t o r y B e n d i n g M o a e n t , C o n v e r s i o n Mode VII-LO P i t c h - L i n k O s c i l l a t o r y Load V e r s u s T h r u s t i n C o n v e r s i o n Mode B l a d e S t a t i o n 5 2 . 5 Beam B e n d i n g Moment a n d V I I - L L P i t c h L i n k L o a d s V e r s u s P r o p r o t o r T h r u s t i n A i r p l a n e Mode B l a d e a n d P i t c h L i n k L o a d s V e r s u s A n g l e o f A t t a c k i n A f r p l a n e Mode P i t c h L i n k S t e a d y L o a d V e r s u s A i r s p e e d i n A i r p l a n e Mode C o m p a r i s o n o f Y o d e l 300 P r o p r o t o r a n d C o n v e n t i o n a l R o t o r N o i s e L e v e l s M e a s u r e d i n Ames 4'3- by 8 0 - F o o t Vind T u n n e l ~ s t i m a t e d E x t e r n a l N ~ i s e L e v e l s -4111 - 6 VIII-7 P y l o n T h r e e - P e r - R e v V i b r a t i o n L e v e l Versus P r o p r o t o r RPM f o r S t a n d a r d S t i f f n e s s T e s t St,qnd ( A i r p l a n e Mode)

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ILLUSTRATIONS - C o n t i n u e d

F i g u r e P a g e P y l o n T h r e e - P e r - R e v V i b r a t i o n L e v e l V e r s u s V I I I - 8 V I I I - 4 T u n n e l V e l o c i t y f o r S t a n d a r d S t i f f n e s s T e s t S t a n d ( A i r p l a n e Mode) P y l o n T h r e e - P e r - R e v V i b r a t i o n L e v e l V e r s u s T u n n e l V e l o c i t y f o r S t a n d a r d S t i f f n e s s T e s t S t a n d (Airplane Mode - Yaw L i n k O u t ) T h r e e - P e r - R e v Beamwise V i b r a t i o n a t P y l o n S t a t i o n 36 V e r s u s S i m u l a t e d A i r s p e e d ( A i r p l a n e Mode ) T h r e e - P c r-Rev Beamwise V i b r a t i o n a t P y l o n S t a 2 i o n 36 V e r s u s A n g l e o f A t l a c k ( A i r p l a n e Mode) Mast C a s e T h r e e - P e r - k e v V i b r a t i o r t L e v e l s V I I I -12 V e r s u s Mast At1gLe of A t t a c k , H e l i c o p t e r a n d C o n v e r s i o n Modes Time H i s t o r y of F e a t h c r - S t o p f o r Q u a r t e r V I I I - 1 3 S t i f f n e s s T e s t S t a n d , Z e r o D e g r e e s A ~ q l e - o f - A t t a c k , T u n n e l S p e e d = 92.5 K n o t s Time H i s t o r y of F e a t h e r - S t o p f o r Q u a r t e r S t i f f n e s s T e s t S t a n d , F o u r D e g r e e s A n g l e - o f - A t t a c k , T u n n e l S p e e d = 9 2 . 5 K n o t s Time H i s t o r y o f F e a t h e r - S t o p f o r Q u a r t e r S t i f f n e s s T e s t S t a n d , Z e r o D e g r e e s A n g l e - o f - A t t a c k , T i l n n e l S p e e d = 1 3 2 K n o t s D i r e c t i o n of A r t g l e s a n d F o r c e s

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I. SUMMARY An advanced-design 25-foot -diameter f lightworthy proprotor was tested i n the NASA-Ames Large-Scale Wind Tunnel under N A S A Bell Helicopter Company designed and manu- Contract NAS2-5386.

the government f o r the factured the proprotor and loaned it t o t e s t s . These t e s t s , Task I1 of the Advancement of Proprotor Technology Program, have t-erif ied and confirmed t h e theory and design solutions developed a s p a r t of the Army Composite A i r c r a f t Program. This report presents the t e s t r e s u l t s and compares them with t h e o r e t i c a l predictions. (Reference 1 reports the r e s u l t s of Task I , the design study.)

Figure 1-1 shows the proprotor i n s t a l l e d i n the tunnel i n prepa- r a t i o n f o r the July 1970 dynamic t e s t s . This t e s t i n g (which reached a simulated speed of 408 knots) showed that s t a b i l i t y increased with airspeed. The meastired damping of the major coupled modes was a s predicted from theory and from the r e s u l t s of one-fifth-scale model t e s t s .

During performance t e s t s (Figures 1-2 and 1-3). conducted i n November 1970, the r e s u l t s met o r exceeded predictions. Hover t h r u s t 15 percent g r e a t e r than the predicted maximum was measured.

I n airplane mode, propulsive e f f i c i e n c i e s (some of which exceeded 90 percent) agreed with Theory.

S t r u c t u r a l l y , the proprotor behaved much a s predicted. Blade

loads were. acceptable i n a11 f l i g h t modes, altbough they exceeded the i n i t i a l l y predicted values a t intermediate conversion angles.

After t h e t e s t s , however, computati-onal :-nput e r r o r s were found i n the predicted values. The elimination of these e r r o r s put the t h e o r e t i c a l (predicted) loads s l i g h t l y higher than those measured.

Blade loads i n the airplane f l i g h t mode were very low.

These t e s t s took the flightworthy t e s t a r t i c l e through mar?y of the conditions i t would encounter i n a c b a l f l i g h t operat ions, and found i t operationally and s t r u c t u r a l l y s a t i s f a r t o r y i n a l l of them. The proprotor is therefore ready f o r t h e next l o g i c a l s t e p i n the development of t i l t - p r o p r o t o r technology--flight t e s t i n g on a research a i r c r a f t .

M i L l t a : ~ and c i v i l p l a n n e r s a r e becoming aware of the need fsr VTOL q i r c r a f t i n a v a r i e t y o f m i s s i o n s and r o l e s . The Irl'OL concept which a p p e a r s t o o f f e r t h e g r e a t e s t proraise of f i l l i n g t h i s need is t h e t i l t - r o t o r a i r c r a f t , I t has a h;gh hover pay- load and causes olnimun n o i s e and do-awash, I t s a v e s weight and complexity by u s i n g t h e same p o e r p l a n t f o r both hover and f o r - ward f l i g h t , The r o t o r system, similar t o that o f t n e h e l i c o p t e r , g i v e s t h e p i l o t t h e sam p r e c i s e c o n t r o l i n t h e hover and at low f l i g h t speeds that a pure h e l i c o p t e r would g i v e him; y e t i n high-speed f l i g h t , t h i s same r o t o r f u n c t i o n s e f f i c i e n t l y as a p r o p e l l e r , range and endurance o f a g i v i n g t h e a i r c r a f t t h e fixed-wing a i r c r a f t . The side-by-side a r r a ~ g e m e n t of t h e p r o p r o t o r s g i v e s the L i f t system a l a r g e o v e r a l l s p a n , which i n t u r n g i v e s it e x c e p t i o n a l l y good SrOL c h a r a c t e r i s t i c s , Many o f t h e d e s i r a b l e c h a r a c t e r i s t i c s of t h e a i r c r a f t d e r i v e from t h e Low d i s c Loading o f t h e t i l t - r o t o r s y s t e a - Indeed, s t u d i e s have shown t h a t low-disc-loading GTOL a i r c r a f t . t h e tilt- m t o r i n p a r t i c u l a r , w i l l make t h e most economical mi3-range VTOL t r a n s p o r t s f o r her c i v i l o r m i l i t a r y a p p l l c a t i o n s .

( S t u d i e s by B e l l , Vestland, Lockneed, Sikorsky, Hoeing, and t h e ?4arine Coros have supported t h i s - p r e d i c t i o n . See References 2 through LO- 1 The Law disc Loading also minimizes noise and d u s t ; t h e low d m w a s h v e l o c i t i e s co~trlbilte t o t h e s a f e t y of ground personnel.

The simple conversion p r o c e s s makes t h e a i r ~ r a f t e a s y t o f l y .

Conversion a a y be s t a r t e d , s t o p p e d , or r e v e r s e d a t a n y p o i n t , w i t h power on o r o f f . The r o t o r - l i f t e d speed range o v e r l a p s t h e wing-lif t e d speed range ; t h e r e f o r e t h e conversion c o r r i d o r is wide, and a i r s p e e d and pcver need n o t be programmed w i t h con- v e r s i o n a n g l e . I n c a s e of complete power f a i l u r e , t h e a i r c r a f t c o n v e r t s t o h e l i c o p t e r f l i g h t and makes an a u t o r o t a t i o n a l d e s c e n t t o a h e l i c o p t e r - l i k e f l a r e and Low-velocity l a n d i n g .

The t i l t - r o t o r concept has been under development f o r two de- cades. E a r l y work l e d t o t h e e s t a b l i s h m e n t of t h e j o i n t Anny- A i r Force XV-3 C o n v e r t i p l a n e Proeram i n 1951 ( F i g u r e I f -1). The e v a l u a t i o n tests were completed i n 1961 and r e p o r t e d i n R e f e r - e n c e s 11 through 13. The program i n c l u d e d more than 375 hours of wind-tunnel and ground-run t i m e , and more t h a n 250 test f l i g h t s i n 125 hours of f l i g h t t i m e . The test a i r c r a f t was flown by t e n government test p i l o t s and two B e l l p i l o t s , who mads a t o t a l of more than 110 f u l l conversions. Five of t h e government t e s t p i l o t s made p0w.r-off r e c o n v e r s i o n s from c r u i s e t o h e l i c o p - t e r a u t o r o t a t i o n a f t e r simulated engine f a i l u r e . The f l i g h t e v a l u a t i o n demonstrated t h e soundness and s a f e t y of t h e conver- s i o n p r i n c i p l e and showed t h a t a p r o p r o t o r could be used e q u a l l y w e l l f o r l i f t and p r o p u l s i o n . I t a l s o d e f i n e d dynamic s t a b i l i t y ?robLevs t h a t requ! red f u r t h e r study. probleas involving dynamic s t a b i l i t y of the proprotors and coupline e f f e c t s between the

proprotors and the a i r c r a f t - A t t h a t t ~ a e these phenomena =re

l i t t l e understood and there was no theory f o r predicting them with any degree of accuracy.

Specifically, the probless were high p r o ~ r o t o r flapping duri3g airplane aaneuv --s. proprotorr/pylon s t a b i l i t y ( l a t e r f outld t o be c l o s e l y related t o propeller-nacelle w h i r l f l u t t e r ) , and a i r - c r a f t dutch r o l l and short-period Longitudinal s t a b i l i t y . I n the e a r l y 1960's B e l l Zielicopter Company i n i t i a t e d an extensive theory and aodel research program t o resoive these problems and

t o develop technology f o r the design of f a t u r e a i r c r a f t - The

program yielded a fundamental understanding of proprotorJpy1on phenomena and explained the behavior of the XV-3 i n f l i g h t and i n the wind tunnel (Reference 1 1 ) .

The ;iroprotor/pylon s t a b i l i t y behavior of t h e XV-3 i n the 1962 tunnel t e s t was f i r s t simulated with a simple sodel, A s theory was f u r t h e r refined and more e f f e c t s and degrees of freedom were added t o the equations, wind-tunnel t e s t i n g uent on t o evaluate blade f l e x i b i l i t y , a i r f raae a e r o e l a s t i c modes. and a i r p l a n e f l i g h t degrees of freedom (References 15 and 16). Some of the aodels used i n the dynaaic aodel t e s t program are shown i n Figure 11-2.

Severai d i f f e r e n t desllgn approaches f o r dynasic s t a b i l i t y evolved and were studied i n d e t a i l . These included positive p i t c h f l a p coupling (negative 6 3 ) . high wing s t i f f n e s s , swashplate/pylon coupling, a focused r o t o r , and automatic flapping control.

I n 1965 the Amy established the Conposite A i r c r a f t Program t o combine i n one a i r c r a f t the good hover c h a r a c t e r i s t i c s of the helicopter and the e f f i c i e n t high-speed c r u i s e c h a r a c t e r i s t i c s of the f ixed-wing a i r c r a f t . The XV-3 experience and subsequent theory and aodel work provided a foundation f o r design e f f o r t i n t h a t program.

The Bell Model 266 a i r c r a f t design shown i n FZgure 11-2 resulted frcm t h i s work. (This work is documented and reviewed i n References 17 through 22).

The exploratory d e f i n i t i o n phase of the Composite A i r c r a f t Pro- gram w a s completed i n 1967. The research a i r c r a f t prograrir which w a s planned t o follow would have established t h a t the l e v e l of technology was adequate f o r an a i r c r a f t system. This program was not begun, however, primarily because of a lack of R I D fund- ing and the absence of a well-defined mission requirement.

Recognizing the need f o r large-scale v e r i f i c a t i o n of t h e tech- nology that had developed from the time of the XV-3, Bell authorized as part of its I W D program the design and fabrica- t i o n of a 25-foot-diameter rctsr- -..d drive system ccrmponents s u i t a b l e f o r t e s t i n g i n the NASA-Ames Large-Scale Vind Tunnel.

This was i n 1968.

I n 1969 the XASA-Ams Research Center and t h e A m y Aeronautical Laboratory contracted with Bell (Reference 23) f c : . t u n q e l t e s t s

@ BELL H E - -nr

of the 25-foot r o t o r and f o r design s t u d i e s of a t i l t - r o t o r proof -of -concept a i r c r a f t . The all Mrdel 300 a i r c r a f t design shown i n Figure 11-4 resulted from the design study (Reference 1).

The 25-foot-diaaeter proprotor completed i t s f i r s t tunnel t e s t i n July L970 (Figure 1-11. The dynamic investigations shotled t h a t the systev was s ~ a b l e t o the aaxiaum tunnel speed of 202 knots with a fulL s t i f f n e s s wing s t r u c t u r e . Flight was simulated t o 408 knots with a one-fourth-stiffness wing s t r u c t u r e . &aping of a l l modes was good and trends -re predicted a c c u r a t e l y by theory.

Control and s t r u c t u r a l investigations showed low blade, hub, and control s t r e s s e s . Investigations with a n automatic flapping con- t r o l device uere i n agreement with theory.

The most recent tests of the 25-foot proprotor took place i n the Ames 40- by 80-foot tunnel i n November 1970 (Figures 1-2 and 1-3).

During these t e s t s , the rotor was power driven i n the helicopter mode, i n the airplane mode, and over a range of conversion angles between these modes, Test r e s u l t s c o r r e l a t e d well with predic- t i o n s as is summarized i n Rzference 24 and shown i n d e t a i l i n the following sections of t h i s report, The dynamic s t a b i l i t y of the r o t o r was excellent i n a l l modes, and blade loads were acceptable.

The r e s u l t s of the t e s t s show t h a t the 25-foot-diameter proprotor i s mady f o r t e s t i n g on a complete a i r c r a f t , 111. DESCRIPTION OF TEST :IARD'dARE A. PROPROTOR AX3 CON'I'ROLS The 25-foot thlee-bladed o r o o r o t o r used f o r t h e s e t e s t s i s scmi- r i g i d , with the hub gimbal mounted t o t h e mast t o provide blade f l a p p i n g freedom. The all-bonded b l a d e s a r e made w i t h high- s t r e n g t h , h e a t - t r e a t e d s t a i n l e s s s t e e l . Blade p i t c h mction and r e t e n t i o n a r e ~ r o v i d e d by needle bearings and wire s t r a p s .

S t a i n i e s s s t e e l Liners, bonded t o t h e t i t a n i u m yoke, prevent f r e t t i n g . fie s t i f f t i t a n i u m yoke p l a c e s a l l blade bending f r e - quencies above r o t o r speed a s i s d i s c u s s e d i n Subsection III.A.2 of t h i s r e p o r t .

The geometry of t h e blades w a s developzd with t h e h e l p of two- dimensional t e s t s ic subsonic and t r a n s o n i c wind tunnels. The blades haveanNACA 64-208 a i r f o i l a t t h e t i p and a h i g h l y cam- bered, 27-percent t h i c k s e c t i o n a t t h e r o o t . A combination of t w i s t and camber was chosen t o meet t h e aerodynamic requirements f o r both h e l i c o p t e r and a i r p l a n e f l i g h t , and t o permit t h e blade s p a r s t r u c t u r e t o have a uniform t w i s t r a t e . The i n t e g r a l blade and g r i p e l i m i n a t e t h e need f o r an aerodynamic c u f f a t t h e rooc of t h e blade, t h e r e b y saving weight and minimizing performance Losses.

An e l a s t o m e r i c hub s p r i n g is u t i l i z e d t o i n c r e a s e t h e c o n t r o l and damping moment c a p a b i l i t y of t h e o r o p r o t o r . The hub s ~ r i n g i s l o c a t e d i n t h e n o n r o t a t i n g system t o e l i m i n a t e f a t i g u e loading on t h i s component. The spri:~g i s a t t a c h e d d i r e c t l y t o t h e t o p case of t h e transmission and t o t h e hub yoke through a bearing.

c y c l i c c o n t r o l is achieved through a monocyclic ( f o r e and a f t ) swashplate below t h e proprotor. A r i s e - a n d - f a l l c o l l e c t i v e head assembly above t h e proprotor mcves t h r e e walking beams t h e r e b y providing c o l l e c t i v e c o n t r o l , Hydraulic a c t u a t o r s p o s i t i o n t h e c y c l i c and c o l l e c t i v e c o n t r o l s . The s e r v o v a l v e s of t h e 1500 p s i h y d r a u l i c a c t u a t o r s were p o s i t i o n e d by a 28-volt e l e c t r o - mechanical a c t u a t o r t o provide remote c o n t r o l f o r these t e s t s .

I n t h e event of h y d r a u l i c p r e s s u r e l o s s , t h e e l e c t r i c a c t u a t o r provided a mechanical back-up t o c a r r y t h e c o n t r o l loads u n t i l t h e t u n n e l could be shut down.

Table 111-1 provides a summary of t h e p e r t i n e n t d a t a concerning t h e proprotors. Both aerodynamic and dynamic d a t a a r e included.

B e l l drawings 300-960-002, 300-010-001, and 300-010-100 included i n t h e appendix show some of t h e c o n s t r u c t i o n d e t a i l s of t h e proprotor. Reference 1 provides a complete d e s c r i p t i o n of t h e p r o ~ r o t o r including mass and s t i f f n e s s d i s t r i b u t i o n s .

TABLE 111-1. PROPROTOR DESCRIPTIVE D A T A Number of Blades per Proprotor 3 Diameter 25.0 f t Disc Area p e r Proprotor 491 s q f t I 14 i n b a s i c blade Blade Chord 17 i n cuff r o o t a t

i

0,0875R Tapering t o 14 i n at 0.25R Blade Area ( 3 Blades) S o l i d i t y Blade A i r f o i l Section

Root ( 5 Mast) NACA 64-935 a = 0.3

Tip NACA 64-208 a = 0.3 Blade Twist (See Figure I V - 1 for ~ s t r i b u t i o n ) -45.0 deg Hub Precone Angle +2.5 deg -15.0 deg Underslinging 0 deg Mast Moment s p r i n g Rate ( p e r Rotor) 2700 in Lb/deg Flapping Design Clearance 1 2 . 0 deg Blade Flapping I n e r t i a (per Blade ) 105 Slug f t 2 Blade Lock Number 3 . 8 3 T i p Speed ( f p s ) ( r p ) H e l i c o p t e r 740 56 5 Conversion Airplane 2. Natural ~ r e a u e n c i e s p r i o r t o t h e wind-tunnel test^ t h e p r o p r o t o r n a t u r a l f r e a u e n c i e s were measured i t 1 a n o n r o t a t i n g v i b r a t i o n survey and t h e i r l o c a - t i o n s confirmed d u r i n g a w h i r l t e s t a t t h e c o n t r a c t o r f a c i l i t y .

~ i g u r e s 111-1 and 111-2 compare t h e measured n o n r o t a t i n g f r e - q u e n c i e s w i t h t h e c a l c u l a t e d f r e q u e n c i e s . I n BHC terminology t h e c o l l e c t i v e modes a r e t h e symmetric modes of t h e p r o p r o t o r , i , e , , p o l a r symmetric about t h e mast. T ~ L c y c l i c modes a r e t h e asymmetric modes. For t h e v i b r a t i o n s u r v e y , t h e p r o p r o t o r was mounted on a t e s t f i x t u r e which was e f f e c t i v e l y r i g i d i n t h e d i r e c t i o n normal t o t h e plane o f t h e r o t o r , b u t which had a n i n p l a n e mode c l o s e t o t h e f i r s t c y c l i c out-of-plane mode. Coup- l i n g o f t h e two modes was n o t i n c l u d e d i n t h e c a l c u l a t i o i ~ but i s e v i d e n t i n t h e measured f r e q u e n c i e s .

The f i r s t t o r s i o n a l mode of t h e blade was measured a t 80 cps a s shown i n F i g c r e 111-1. The v i b r a t i o n s u r v e y t e s t s t a n d provided a n extremely s t i f f t o r s i o n a l r e s t r a i n t , hence t h i s i s e f f e c t i v e l y t h e t o r s i o n a l n a t u r a l frequency of a c a n t i l e v e r e d b l a d e . The c a l c u l a t e d f i r s t t o r s i o n a l n a t u r a l frequency f o r t h e c a n t i l e v e r e d b l a d e is 87 cps. With t h e c o n t r o l system f l e x i b i l i t y i n c l u d e d i n t h e c a l c u l a t i o n s , a mode which i s r i g i d body f e a t h e r i n g on t h e c o n t r o l system i s i n t r o d u c e d a t 36 cps. Coupling w i t h t h i s mode f o r c e s t h e t o r s i o n mode up t o 122 cps.

Wher, t h e p r o p r o t o r was w h i r l t e s t e d on a h o r i z o n t a l t e s t s t a n d , t h e blade n a t u r a l f r e q u e n c i e s were f u r t h e r e s t a b l i s h e d by r e s o n a n t c r o s s i n g s and harmonic e x c i t a t i o n of t h e swashplate. Figures 111-3 and 111-4 compare t h e s e f r e q u e n c i e s ' i d e n t i f i c a t i o n s w i t h f a n p l o t s based on e x t r a p o l a t i o n of t h e f r e q u e n c i e s measured d u r - i n g t h e n o n r o t a t i n g v i b r a t i o n survey. Note t h a t s e v e r a l modes a r e i n d i c a t e d t o be i n resonance, nameiy t h e t h i r d c o l l e c t i v e mode and t h e second and t h i r d c y c l i c modes. These resonances were c l o s e l y monitored d u r i n g t h e wind-tunnel t e s t s , but were never a problem.

TEST STANDS 1. Performance Stand a . D e s c r i p t i o n Performance t e s t i n g of t h e Model 300 p r o p r o t o r was accomplished on t h e N A S A p r o p e l l e r t e s t r i g a s shown on F i g u r e s 1-2 and 1-3.

The power module of t h i s r i g c o n s i s t s of two 1500 hp e l e c t r i c d r i v i n g a n R-2800 e n g i n e motors mounted i n tandem on a frame, r e d u c t i o n gearbox. The power module was mounted a t appr oximately t h e c e n t e r of t h e t u n n e l t e s t s e c t i o n on s t r u t s a t t a c h e d t o t h e t u n n e l balance frame. Angle of a t t a c k could be changed on t h e module by means of a remotely-actuated t a i l s t r u t . The angle range a v a i l a b l e was from 0 degrees (mast h o r i z o n t a l ) t o 85 degrees (mast 3 e a r l y v e r t i c a l ) . Bell Drawing 309-018-013, i n - cluded i n t5e appendix, shows t h e p r o p r o t o r mounted on the t e s t r i g .

To adapt tl-e Model 3 0 0 r o t o r t o the t e s t r i g gearbox, a new ge.2rbox f r o n t cover was v a i e . A mast support case c o n t a i n i q g a a l l t h r u s t bearing and a r a d i a l r o l l e r t z d r i n g was a t t a c h e d t o t h e s p e c i a l f r o ~ t cover. The bearings i n the mast case were separated by 13.5 inches t o c a r r y Eoments generated t y the r o t o r .

A c o l l e c t i v e p i t c h i n p u t s l e e v e was i n s t a l l e d between t h e bear- i n g s . This sleeve contained a r o ~ a t i n g - t o - r o t a t l n g bearing s e t , and a t t a c h e d t o a c o l l e c t i v e tube i n s i d e t h e mast. The c o l l e c - t i v e c o n t r o l c y l i n d e r , i n s t a l l e d below t h e gearbox, a c t u a t e d t h e s l e e v e through a l e v e r . The upper end cf t h e w i ~ d - t u n n e l mast c a s e was s i m i l a r t o t h e Model 300 t r a n s m i s s i o n mast c a s e , and had attachment p o i n t s f o r t h e Mode, 300 swashplate and hub rnosent s p r i n g . An instrumentation s l i p r i n g was i n s t a l l e d i n s i d e the mast c a s e , above t h e upper b e a r i n g .

The mast used during t h e wind-tunnel t e s t s was i d e n t i c a l t o t h a t of t h e Model 300 above t h e mast c a s e . I n s i d e the mast c a s e , t h e wind-tbnnrl mast was s l o t t e d t o permit o p e r a t i o n of t h e c o l L e c t i v e s l e e v e . The mast wall t h i c k n e s s i n t h e a r e a of t h e s l o t s was i;-.cLeased over t h a t of t h e Model 300 i n o r d e r t o maintain t h e mast s t r e n g t h . The lower end of t h e mast had a d r i v e s p l i n e t o which a modified R-2800 p l a n e t a r y g e a r c a r r i e r a t t a c h e d . The c a r r i e r con- t a i n e d s t a n d a r d R-2800 p l a r e t a r y g e a r s .

The Model 300 proprotor and r o t a t i n g c o n t r c l s tittached t o t h e wind-tunnel p a s t and mast case without m o d i f i c a t i o n . The ModeL 30U s p i n n e r f a i r e d t h e forward p o r t i o n of t h e r i g ; n o n s t r u c t u r a l f a i r i n g s , a t t a c h e d t o t h e t u n n e l f l o o r , f a i r e d t h e remainder of t h e r i g .

b . Natural Frequencies A v i b r a t i o n a n a l y s i s of t h e p r o p e l l e r t e s t s t a n d , w i t h t h e 25- foot p r o p r o t o r i n s t a l l e d , was made p r i o r t o t h e powered t e s t .

The t e s t stand s t r u c t u r e was modeled on t h e NASTRAN s t r u c t u r a l a n a l y s i s (Reference 25) and t h e n a t u r a l f r e q u e n c i e s through e i g h t per rev c a l c u l a t e d . Figure 111-5 shows t h e NASTRAN s t r u c t u r a l mode 1.

A v i b r a t i o n survey of t h e p r o p e l l e r t e s t s t a n d was conducted with t h e stand i n s t a l l e d i n t h e t e s t s e c t i o n . For t h e survey t h e proprotor blades = r e removed and replaced by e q u i v a l e n t weights.

A Lazan e c c e n t r i c mass v i b r a t o r was i n s t a l l e d i n a blade g r i p t o provide e x c i t a t i o n . Surveys were made a t n a c e l l e a n g l e s of 0 and 60 d e g r e e s . The i n f l u e n c e of n a c e l l e angle was very s m a l l .

The measured s t a n d n a t u r a l f r e q u e n c i e s a r e compared t o t h e calcu- l a t e d f r e q u e n c i e s i n Table 111-11. While t h e frequency c o r r e l s - t i o n i s reasonable, n e g l e c t i n g t h e f l e x i b i l i t y of t h e s t a n d torquemeter i n t h e math model caused a considerable e r r o r i n t h e TABLE 111- 11. PROPELLER TE3 STAND NATURAL FREQ'JENCIES abIAST = 0 DEGREES Mode No. Measured Calculated i d e n t i f i c a t i o n (cps) (cps) 1 1 . 8 1.66 Strut first l a t e r a l bending 2 3.8 3.52 Nacelle yaw 3 3 . 5 2.72 S t r u t f o r e and a f t bending 4 5.2 1C.47 Cross tube v e r t i c a l bendir-g 5 16.g2 22.3 Mast l a t e r a l ending 6 22.12 26.7 Past v e r t i c a l bending

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' ~ o r q ~ e meter s o f t n e s s a t i n t e r f a c e between stand & mast case n o t reprr5ented i n b ! 'TRAN model f r e q u e n c i e s o f the vast vertical a n d Lateral b e n d i n g m o d e s , r h e L o c a t i o n o f t h e s e modes is v e r y i m p o r t a n t s i n c e t h e y are l o c a t e d it. t h e t w o - p e r - r e v a n d t h r e e - p e r - r e v excf tatiori f r e q u e n c y range.

D u r i n g t h i s t e s t , t w o - p e r - r e v L o a d s a t 535 rpln p r o v e d a l i n i t a - t i o n a n d p r e v e n t e d e x t e n s i v e o p e r a t i o n a t that r p . However, t h e r e wa* P.O p r 2 b l e a ir. o p e r a t i n g a t rotor s p e e d s h i g h e r or lower t h a n 535 rpw.

2 , 3 y n a a ~ c T e s t S t a n d s a . D e s c r i p t i o n The W e 1 300 p r r r a t o r was t e s t e d f o r d y n a a i c s t a b i t i t y on a s t r u c t u r e v h i c h s i m u l a t e d the H o d e l 300 s e a i s p a n v i n e and n a c e l i c .

B e l l Drawing 309-0;8-941, i n c l u d e d i n the a p p e n d i x . s h o w s t h e test assesbly. This w i n g was 9 o u n t e d v e r t i c a i l y i n t h e w i n d t u n n e l , a n d a t t a c h e d d i r e c t l y 5s t h e t u n n e l b a l a n c e . T h e w i n g a n g l e of a t t a c k couLd be v a r i e d r e s o t e l y ; alsc, t h e n a c e l l e c o u l d be p i t c h e d r e l a t i v e t o t h e w i n g f r o a O t o 20 d e g r e e s m a n u a l l y .

E x t e n s i v e u s e was aade of the resote a n g l e o f a t t a c k v a r t a t i o n c a p a b i l i t y d u r i n g the test, h o w e v e r , 211 d y n a m i c test - s t a n d runs -re made w i t h the nacelle i n the a i r p l a n e a o C e .

The basic w i n g s t r u c t u r e c o n s i s t e d of a beam o f r e c t a n g u l a r c r o s s s e c t i o n , a n d equal i n b e n d i n g a n d t o r s i o n a l s t i f f n e s s t o t h e actual Hodel 300 w i n g d e s i g n . A s e c o n d bear, having o n e - f o u r t h t h e s t i f f n e s s e s of t5e f L i g h t w i n g , was a l s o p r o v i d e d . The w i n g " a i r f o i l " vas z s i q p l e aerodyriasic fairing corilplete w i t h 1igk.t n o n s t r u c t u r a l c m p o n e n t s . T h e f a i r i n g s would a d a p t t o e i t h e r wing beaa- ALL c o a p o n e n t s of the beam a n d f a i r i n g s were a l u m i n u m .

The test s t a n d p a r a m e t e r s - - w i n g c h o r d , s p a n , -*ei@t, s t i f f n e s s , a n d s w e e p - - w e r e h e l d as c l o s e t o t h o s e of t h e Mode! 300 a s p o s s i - b l e . The a i r f o i l s e c t i o n . h o w e v e r , w a s s i a ~ l i f i e d t o r e d u c e m a n u f a c t u r i n g cost. Its t h i c k n e s s was 13-5 p e r c e n t as c o m p a r e d t o t h e Model 309 w i n g ' s 23 p e r c e n t .

The b a s i c s t r u c t u r e o f t h e n a c e l l e was a steel s ~ l d v e n t ~-t:ich -.

a t t a c h e d t o the w i n g t i p b y lreans of a c o n v e r s i o n s p i n d l e . lne b e n d i n g s t i f f n e s s of t h e s p i n d l e was t h e same as t h e Model 309 c o n v e r s i o n s p i n d l e . The c o n v e r s i o n a c t u a t o r was r e p l a c e d b y a L i n k . A yaw L i n k w a s a l s o p r o v i d e d t o s i m u l a t e t h e p y l o n - t o - w i n g a t t a c h a e n t when t h e p y l o n is f u l l y c o n v e r t e d so t h a t a d o w n s t o p is e n g a g e d . The s a a t s c p p o r t case u s e d on the p o v e r test r i g a t t a c h e d t o t h e f r o n t of t b e u e l d r a e n t . Two a l u m i n m b u l k h e a d s , z t t a c h e d t o t h e u e l d m e c t , s u p p o r t e d a fiberglass f a i r i c g o v e r t h e s t r u c t u r e . The f a i r i n g c o o t o u r was made t o t h e Model 300 n a c e l l e l i ~ e s .

The c o l l e c t i ~ ~ c c o n t r o l roechanis* w a s the same as was u s e d o n t h e power rig. The c y c l i c h y d r a u l i c c y l i n d e r was a l s o u s e d , b u t was c o n t r o l l e d b y a small h y d - ~ l i c . c t u a t o r ( S C A S U n i t ) e q u i p p e d with 3? e i e c t r i c a l l y c o n t . o l ! e c s e r v o v a l v e . The SCAS c y l i n d e r was u s e d for t h e f l a p p i n g c c ? t .oiler i n p u t and as a c y c l i c c o n t r o l s y s t e T s h a k e r .

In addition t o the c y c l i c shaker, an aerodynamic shaker was a l s o provided. T h i s shaker consisted of a z ~ d L 1 a i r f o i l mounted on a s h a f t , attached near rhe a f t end of t h e nacelle, The shaker was o s c i l l a t e d by another S U S hydraulic cylinder, b, Scaling The s c a l i n g f a c t o r s of the two design stands a r e given i n Table IIT-JII f o r reference, Note t h a t data from the design s t i f f n e s s stand may tz used 2irc:tly. Data from the one-fourth design s t i f f n e s s t e s t s t a ~ d must be muliLplied by the appropriate f a c t o r from rable 111-iII t o obtain the f u l l - s c a l e equivalellt value, c. Test Stand Natural Frequencies The t e s t stand n a t u r a l frequencies were calculated p r i o r t o the dynamic s t a b i l i t y t e s t using a f i n i t e element s t r u c t u r a l model of the stands. Figure 111-6 shows t n e layout of t h e s t r u c t u r a l model and t h e element properties and node point masses a r e given i n Tables 111-IV and 1 1 1 - V , A vibration survey of the design-stiffness t e s t stand was c m - ducted f i r s t a t the c o n t r a c t o r ' s f a c i l i t y and again when the starid was i n s t a l l e d i n the t e s t section. A v i b r a t i o n survey of t h e one-fourth-design-stiffness stand was a l s a canducted i n the t e s t section, 3uring these surveys the proprotor blades were removed and r e - placed b ; * equivalent weights. A Lazan shaker was used t o e x c i t e the system natural frequencies. Hand e x c i t a t i o n of the fundamen- t a l reoder was used t o obtain modal damping r a t i o s , m r i n g t h e design s t i f f a ~ e s s stand v i b r a t i o n survey the tunnel balance n a t u r a l frequencies were i d e n t i f i e d . S t r a i n gaged beam transducers mounted a t the corners of the balance frame were used t o determine the balance mode shapes, The liacural frequencies for t h e Model 300, the calculated stand natural frequencies, and the measured frequencies a r e tabulated i n Table 1 1 1 - V I , l'he measured damping i s a l s o indicated. Also l i s t e d i n Table 1 1 1 - V I a r e the frequencies of t h e balance modes.

The mode shapes f o r the four lowest modes of the design-stiffness t e s t stand a r e shown i n Figure 111-7. These modes a r e the same f o r t h e one-fourth-stifffiess stand with the exception of f o u r t h modes, which is a pylon yawing mode. The balance modes were e s s e n t i a l l y r i g i d body and uncoupled, d. Test 3tand Dynamic S t a b i l i t y Boundaries Figure 111-8 shows the calculated dynamic s t a b i l i t y boundaries f o r the t e s t stand a s a function of proprotor rpm. These a r e based on the t e s t stand measured n a t u r a l frequencies. Ztto TABLE 111-111. DYNAMIC TEST STAND SCALE FACTORS' Design s t i f f n e s s Z One-Fourth S t i f f n e s s 3 Parameter T e s t Stand Test Stand I - Length Mass Time Velocity Acceleration Frequency Force ' ~ u l t i ~ l ~ model data by s c a l e f a c t o r to obtain equivalent f u l l - s c a l e value

2 ~ c a l i n g - Mach N o . 1 : l

Froude No. I : l Lock N o . 1 : l

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Froude No. 0 . 2 5 : l Lock No. 1 : l TABLE 111-IV, DYNAMIC TEST STAND MASS PROPERTIES Mass and I n e r t i a * Node Mz 18 x 0 . 4 3 0 . 6 0 . 2 0 - 6 0 . 2 0 . 6 0 . 2 0 . 6 0 . 5 1 1 0 . 6 0 . 0 1 0 . 3 1 0 . 0 1 0 . 0 1 1 . 0 8 7 772.0 2 . 1 3 800.0

1.036 --

-- 112 -96

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A l l units i n l b - i n - s e c system.

-- I n d i c a t e s c o o r d i n a t e was n o t r e t a i n e d .

TABLE 111-V, DYNAMIC TEST STAND STIFFNESS PROPERTIES Segment (in) (15-in2) ~ 1 0 ' ~ (lb-in2) x10" (lb-in2) ~ 1 0 ' ~ *Multiply El ~ n d GJ x 0.25 for one-fourth design s t i f f n e s s stand .

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boundaries a r e shown f o r the one-fourthaesign-stiffness stand, one when the scaled one-fourth-design-stiffness hub r e s t r a i n t i s used, and one when the design-stiffness r e s t r a i n t i s employed.

Prior t o t h e t e s t i t was planned t o switch t o a one-fourth-design- s t i f f n e s s hub r e s t r a i n t when the wing spars were switched, How- ever, while the t e s t was i n progress a decision was made t o r e - t a i n the design-stiffness hub r e s t r a i n t for the one-fourth design s t i f f n e s s stand t e s t s , i n order t o conserve occupancy time. Cal- culations indicated the s t a b i l i t y c h a r a c t e r i s t i c s were not s i g - n i f i c a n t l y affected by the hub r e s t r a i n t and t h a t the time required t o change t o the one-fourth-design-stiffness r e s t r a i n t would not be j u s t i f i e d . The difference i n one-fourth s t i f f n e s s t e s t stand dynamic c h a r a c t e r i s t i c s w i t h hub r e s t r a i n t i s dis- cussed f u r t h e r i n Section V.

INSTRUMENTATION conventional instrumentation was used t o measure loads, def l e c - t i o n , vibrations, and pressures during both t e s t s , The t r a n s - ducers included s t r a i n gages, potentiometers, accelerometers and pressure sensors. Proprotor r o t a t i n g system instrumentation channels u t i l i z e d a 52-ring s l i p r i n g t o provide 2 e x c i t a t i o n power channels and 24 data chaanels. Table 1 1 1 - V I I is a summary of the d a t a channels f o r the dynamic and tne powerec! t e s t s .

There were 27 channels available during the dynamic t e s t and 41 channels available during the powered t e s t . The channel £re- quency range is shown f o r reference. system accuracy is e s t i - mated t o be within -+ 3 percent (based on channel f u l l - s c a l e w i t h e r r o r s being :he square root of the sum of the square of the individual e r r o r s ).

B t a were recorded using d i r e c t -write oscillographs, and during the dynamic t e s t s magnetic tape records were a l s o made. Two ld-channel oscillographs were used during the dynamic t e s t s and four of the same type oscillographs were used during the powered t e s t s .

@ BELL HEUCOPTER

TABLE 1 1 1 - V I I - INSTRUMENTATION S U M M A R Y r Chan ne 1 Dynsaic S t a b i l i t y T e s t Frequency Item Transducer Range ( c p s ) Mast p a r a l l e l bending S t r a i n gage 0-6CI ( i n red b l a d e f l a p p i n g s e n s e ) Yoke beam bending S t r a i n gage 0-60 Red p i t c h L i n k a x i a l Load S t r a i n gage 0-135 Xed blade s p i n d l e chora bending S t r a i n gage 0 -69 Blade bean S t a 52.5 ( 3 5 per- S t r a i n gage 0-135 c e n t R ) bending Blade beam Sta 7 5 ( 50 p e r c e n t R) S t r a i n gage 0-135 bending Blade chord Eta 5 2 . 5 ( 3 5 p e r - S t r a i n gage 0-135 c e n t R ) bending Red blade f l a p p i n g R o t a t i n g pot 0-135 Fore c?d a f t f l a p p i n g S t a t i o n a r y pot 0-190 L a t e r a l f l a p p i n g S t a t i o n a r y pot 0-190 Wing beam inboard bending S t r a i n gage 0-160 Wing chord inboard bending S t r a i n gage 0-60 Wing t o r s i o n inboard bending S t r a i n gage 0-60 Wing chord outboard bending S t r a i n gage 0 -60 C y c l i c tube a x i a l load S t r a i n gage 0-60 C o l l e c t i v e tube a x i a l load S t r a i n gage 3-60 Conversion l i n k a x i a l l o a d S t r a i n gage 3 -60 Yaw l i n k a x i a l load S t r a i n gage 0-60 Shaker beam S t r a i n gage 0-60 Pylon i n t e r n a l s t a t i c p r e s s u r e P r e s s u r e 0 -60 I n l e t t o t a l head p r e s s u r e P r e s s u r e 0-60 Hydraulic p r e s s u r e P r e s s u r e 0 -60 0 -60 Fore and a f t a c c e l e r a t i o n A c c e l e r o m ~ t e r Pylon S t a 0 beam a c c e l e r a t i o n Accelerometer 0-60 Pylon S t a 0 yaw a c c e l e r a t i o n Accelerometer 0-60 Pylon S t a 36.0 beam a c c e l e r a t i o n Accelerometer 0-60 Pylon S t a 36.0 yaw a c c e l e r a t i o n Accelerometer 0-60

@ B E U H E U C O C O M P A N Y

TABLE 111- V I I . CONTINUED Powered T e s t Channel Frequency I tem Transducer Range ( c p s )

-

Mast p a r a l l e l bending S t r a i n gage 0-69 ( i n red blade f l a p p i n g s e n s e ) Mast p e r p e n d i c u l a r bending S t r a i n gage 0-69 (normal t o red blade f l a p p i n g s e n s e ) Red p i t c h l i n k a x i a l load S t r a i n gage 0-60 White p i t c h l i n k a x i a l load S t r a i n gage 0-60 Green p i t c h L i n k a x i a l load S t r a i n gage 0-60 Red blade s p i n d l e beam bending S t r a i n gage 0-63 Red blade s p i n d l e chord bending S t r a i n gage 0-60 GXite blade s p i n d l e beaa bending S t r a i n gage 0-60 White blade s p i n d l e chord bending S t r a i n gage 0-60 Blade beam S t a 22.8 ( 1 5 per- S t r a i n gage 0-60 c e n t Rj bending Blade beam S t a 7 5 ( 50 percent R) S t r a i n gage 0-60 bending Blade beam S t a 112.5 ( 7 5 per- S t r a i n gage 0-60 c e n t Ri bending Blade chord S t a 52.5 (35 per- S t r a i n gage 0-60 c e n t R) bending Blade chord S t a 75 ( 5 0 p e r - Str-iin gage 0-60 c e n t R! bending Blade chord S t a L i 2 . 5 ( 7 5 per- S t r a i n gage 0 -611 cenL R) bending Blade t o r s i o n S t a 112.5 (75 per- S t r a i n gage 0-59 c e n t R) bending R 2 J blade f l a p p i n g Rotating pot 0-60 Rt\d blade f e a t h e r i n g S t r a i n gage 0-60 Fore and a f t f l ? p p i n g S t a t i o n a r y pot 0-60 L a t e r a l f l a p p i n g S t a t i o n a r g pot 0-60 Xed b l a d e t r a i l i n g edge s t r e s s S t r a i n gage C y c l i c tube . ~ x i a l load S t r c i n gage 9-60 C o l l e c t i v e tube a x i a l load S t r a i n gage 0-60 C o l l e c t i v e s l e e v e p e r p e n d i c u l a r S t r a i n gsge 0-60

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TABLE 1 1 1 - V I I . CONCLUDED Cha nne 1 Frequency I tern Transducer Range ( c p s ) C o l l e c t i v e s l e e v e p a r a l l e l S t r a i n gage 0-60 bending Collective p o s i t i o n Linear pot 0-60

I Cyclic p o s i t i o n Linear pot 0- 60

Spinner upper support arm S t r a i n gage 0-60

I bending

Spinner lower support arm S t r a i n gage 0 -60

; btnding

i L i v e r load S t r a i n gage 0-60

Mast torque S t r a i n gage 0-60 Test stand i n t e r n a l s t a t i c P r e s s u r e 0-60 pressure

I Hydraulic p r e s s u r e Pressure 0-60

1 Tunnel a i r s p e e d P r e s s u r e 0-60

1 Mast case a x i a l a c c e l e r a t i o n Accelerometer 0-60

I Mast case Seam a c c e l c r a t ion Accelerometer 0-60

%st case yaw a c c e l e r a t i o n Accelerometer 0-60 Accelerometer 0-135 T e s t stand a f t v e r t i c a l a c c e l e r a t i o n Test s t a n d a f t l a t e r a l Accelerometer 0-135 a c c e l e r a t i o n T e s t stand forward v e r t i c a l . Accele roae t e r 0-60 a c c e l e r a t i o n Test stand forward l a t e r a l Accelerometer 0-60 accelerometer

0 B E U HEUCOPTER COU-NV

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BLADE T I P COL7,ECTIVE PITCH - DEGREES

F i g u r e 111- 1. P r o p r o t o r C o l i e c t i v e >!~de l i a t u r ~ ~ l Frequencies, Nonrotating.

BELL HELICOPTER COMPANY

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BLADE T I P COLLECTIVE PITCH - SZGPZES

Figure 111- 2 . P r o p r o t o r C y c l i c Mode N a t u r a l F r e q u e n c i e s , N o n r o t a t i n g .

ROTOR RPM Figure III- 3 . P r o p r o t o v C o l l e c t i v e Mode Natural F r e q u e n c , ? ~ , Rotating.

@ BELL HELICOPTER corrprur

nO'2OR RPM F i g u r e 111-4. Proprotor C y c l i c Mode Natural Frequencies, R o t a t i n g .

300-099-004 111-19 (a) TCP VIEW (b) SIDE VIEW (c) FROST VIEW Figure 111- 5 . NASIRAN Structural Model of Powerad T e s t Stand.

COO RI) INAT E SYSTEM

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ELEMENTS 26 ZND 27 ARE PIX ENDED.

Figure 111- 6 . NASTRAN Structiiral Model of Dynamic Test Stand.

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STABLE 5 . ... UNSTABLE

. - . - ..

m .

..

..

. .

..

. . .

.. . ..

S T I F F N E S S DYNAMIC 2R = 30G FT/SEC

--------

200db WITH DESIGN S T I F F N E S S HUB RESTRAINT 1/4 DESIGN

I

STIFFNESS HUB

I RESTRAINT

100.-

I

( MAXIMUM WIND SPEED CAP-ILITY O F 40 x 80 FOOT TUNNEL

I

0 6 0 100 200 300 400 500

WIN3 TUNNEL VELOCITY - KNOTS

.gure 111-8. Dynamic Test Stand Stability Boundaries.

@ BELL HEUCOPTER c o v - w r

IV. PERFORMAXCT The BHC 25-foot-diameter proprotor r-ras t e s t e d i n the NASA-Aqes 40- by 80-foot wind tunnel t o e v a l u a t e proprotor performance and loads i n a l l t h r e e modes of f L i g h t : hover. h e l i c o p t e r - conversion, and z f r p l a n e . The t e s t r e s u l t s c o r r e l a t e d well with p r e d i c t ions.

The t e s t s showed t h a t t h i s proprotor maintains good propulsive e f f i c i e n c y i n a i r p l a n e f l i g h t node with no compromise of i t s performance i n t h e hover o r i n t h e h e l i c o p t e r f l i g h t .node. The propulsive e f f i c i e n c i e s c a l c u l a t e d f o r the c r u i s e c o n d i t i o n s a r e approximately 75 percent. These p r e d i c t i o n s were v e r i f i e d by t h e t e s t d a t a . E f f i c i e n c i e s i n e x c e s s of 90 percent were obtained f o r t h e higher t h r u s t requirements. The h i g h e s t measured r o t o r 9ean l i f t c o e f f i c i e n t i n hover ( a s i n d i c a t e d by the value of CT~:''~) was 18 percent h i g h e r than t h e p r e d i c t e d maximum. Furthermore, t h e helicopter-conve r s ion f Light d a t a were --Lightly better than p r e d i c t e d .

P r e d i c t i o n s of perfornance were c a l c u l a t e d p r i o r t o t h e t e s t s .

using B H C ' s standard p r e d i c t i o n techniques Where t e s t r e s u l t s and p r e d i c t e d perfomance d i s a g r e e . t h e t e s t perfornance i s u s u a l l y b e t t e r . For example, the power required f o r a given L i f t w a s not a s high d u r i n g the t e F t s a s %las p r e d i c t e d .

Data a r e presented here by f l i g h t node. Where a ~ p l i c a b l e , i - e . , f o r t h e proper o p e r a t i n g t i p speeds, d a t a a r e shown f o r t h e t y p i c a l o p e r a t i n g requirewents of a LG,3~0-pound -gross weight v e h i c l e . Because e n g i n e e r s with d i v e r s e backgrounds w i l l be using them, t h e d a t a have been nondiaensionalized i n accordance with the two d i f f e r e n t conventions of a i r p l a n e and h e l i c o p t e r a n a l y s i s . The appendix d e f i n e s the n o a e a c l a t u r e .

B . TEST DESCRIPTION The wind tunnel had equipment f o r measuring p e r t i n e n t p a r a a e t e r s during p o e r e d t e s t s . B i C supplied some a d d i t i o n a l i n s t ruaen- t a t i o n . The six-component balance provided d a t a which were converted t o r o t o r t h r u s t , H-force, and torque. There was a l s o a load c e l l on t h e t e s t s t a n d f o r q e a s u r i n g t o r q u e . The r o t o r mast ( s h a f t ) w a s a l s o strain-gaged t o give yet a n o t h e r means of measuring torque. Generally, t h i s r e p o r t uses torque and power f i g u r e s obtained by means of the strain gages because t h e y seemed t o be more a c c u r a t e and were. furthermore, c o n s e r v a t i v e i n t h a t t h e y showed t h e r o t o r t o be consuaing gore power than t h e o t h e r instriimentation showed. I n a d d i t i o n t o t h e s e measure- g e n t s , t h e proprotor rpm, c o l l e c t i v e p i t c h , c y c l i c p i t c h , and f l a p p i n g were measured. For a .nore d e t a i l e d d e s c r i p t i o n of t h e proprotor i n s t r u m e a t a t i o n and force/angle r e l a t i o n s h i p s , see Subsec'iion 111. C and t h e appendix r e s p e c t i v e l y .

The a a j o r t e s t variables were t i p speed. tunnel s w e d , and s h a f t angle. Table I V - I shows the ranges t e s t e d f o r each of the t h r e e f l i g h t modes. For the majority of t e s t conditions, the collec- t i v e pitch was varied while the other variables were held approx- imately constant ( c y c l i c was adjusted t o hold fore and a f t flapping cons tart t a t zero 1.

TABLE 13-1. RANGE O F VARIABLES FOR POWERED TEST I Tunnel Mast

. Tip

Speed Angle Speed ( knots ) ( f p s ) (deg)

Hover 0 - 20 600 - 740

9 - 7 5

80 - 140

He1 icopter- 700 - 740 15 - 7 5

Conversion

Airplane 120 - 185 400 - 740 9

I i Complete predictions of performance f o r t h e t e s t conditions were computed and supplied to XASA p r i o r t o tunnel entry.

These predictions used B e l l Helicopter Company program F35.

This program employs blade-element-momentum theory with non- uniform inflow f o r hover and a x i a l f l i g h t and uniform in£ low f o r the o t h e r f l i g h t conditions. i t uses two-ditnensional a i r f o i l d a t a i n conjmcti on with the blade geometric c h a r a c t e r i s t i c s .

Reference 26 provides f u r t h e r discussion of the theory. Blade geomztry i s defined i n Figure I V - 1 .

A s shown, the 5lade has been divided i n t o four major segments, the maximum number of d2-visions allowed by the program f o r the input of geometric and a i r f o i l s e c t i o n s properties. For calcu- l a t f o n purposzs, each of the major segments is subdivided in such a way t h a t each blade is represented by 22 elements. The program l i m i t a t i o n of four segments resulted i n an i n a b i l i t y t o match exactly the inboard t w i s t and chord.

The a i r f o i l s e c t i o n data f o r each of the four segments a r e shown i n Figures IV-2 through IV-5. Also shown a r e modifications made to these datn i n an attempt t o obtain b e t t e r correli.tion with t h e naver d a t a , as discussed i n Subsection I V . E.L. The unmodified section d a t a a r e used throughout t h i s report except a s noted f o r hover. The a i r f o i l d a t a f o r the f i r s t s e c t i o n were obtained from two -dimensional t e s t s conducted i n the General Dynamics -Convair low-speed wind tunnel. The d a t a f o r the outboard t h r e e s e c t i o n s were obtained from t e s t s conducted i n the United A i r c r a f t Research Laboratories high-speed wind tunnel.

I n addition t o the geometric properties of the r o t o r and the a i r f o i l s e c t i o n d a t a , other inputs include the number of blades, t i p speed, speed of sound, freestseam v e l o c i t y , a i r density r a t i o , mast angle, and c o l l e c t i v e p i t c h o r t h r u s t .

Tip l o s s i s normally calculated by equating t o zero the l i f t of the blade segment outboard of t h e spanwise s t a t i o n determined by the r e l a t i o n

&

B = l - - b The program a l s o allows the user t o e n t e r any a r b i t r a r y constant t i p l o s s f a c t o r . The program then u t i l i z e s t h i s f a c t o r a s

explained above; i . e . , no l i f t is produced on the blade outboard

of the s t a t i o n s e t by the t i p l o s s f a c t o r .

D . SPINNER T m DATA

Figures IV-6 through IV-8 show :he L i f t and drag t a r e s used during the t e s t . Approximate t a r e d a t a were obtained i n i t i a l l y with the blade spindles protruding from the spinner. After the t e s t was completed, these spindles were removed and d a t a obtained f o r both a s m o ~ t h spinner and one with the blade holes open. I n general, curves a r e f a i r e d through a l l of the d a t a since n e i t h e r case i s t r u l y representative of the a c t u a l t e s t conditions. Figure IV-6 shows the e f f e c t of the spinner base pressure on the drag t a r e when the proprotor is i n the a x i a l f l i g h t node. Most of the spinner drag was due t o skin f r i c t i o n r a t h e r than base pressure since d u r i n g the t e s t the measured ..W/q value never exceeded 9.1. These low base pressures were probably a r e s u l t of the one- t o one-and-one-half-inch gap between the spinner and t e s t stand shroud. Also shown i n t h i s figure a r e the tare data previously measured on a o n e - f i f t h - s c a l e proprotor model.

Fjgures IV-7 and IV-8 show the l i f t and drag t a r e d a t a used i n During the spindle-off reducing the helicopter-conversion data.

t a r e t e s t s , the l i f t balance d a t a become questionable. I n order t o obtain the proper l i f t t a r e f o r t h i s condition, the incremental values determined from the one-fifth-scale t e s t s were applied to the f u l l - s c a l e d a t a .

E. RESULTS AND CORRELATION Figures IV-S through IV-36 show the cal-culated d a t a and powered t e s t r e s u l t s . The d a t a a r e separated i n t o the hover, helicopter- conversion, and a x i a l o r a i r p l a n e f l i g h t modes f o r c l a r i t y of preserrtation. These d a t a a r e shotm i n both dimensional and nondimensional form. Nondimensional d a t a have been shown i n h ~ + h a i r p l a n e a ~ d h e l i c o p t e r nomenclature where applicable. A l l c o e f f i c i e n t s a r e i n a i r p l a n e t e n s u n l e s s noted. D e f i n i t i o n s f o r a l l c o e f f i c i e n t s and parameters a r e found i n the appendix. Also i n the appe.ldix i s a cornputer L i s t i n g of a l l dimensional test d a t a .

ALL c o r r e l a t i o n was aade i n t e r a s of f o r c e s and horsepower because t h e y a r e of p r i a a r y i n t e r e s t . C o r r e l a t i o n with r e s p e c t t o q a s t angle has been shown by s t a t i n g the value of a a s t angle where a p p l i c a b l e ( h e l i c o p t e r - c o n v e r s i o n aode ) . As t h e d a t a i n d i c a t e , good c o r r e l a t i o n was o b t a i ~ e d when t h e f l a p p i n g angle was Less than 2 0 . 5 degree. The c o r r e l a t i o n between the c a l c u l a t e d and aeasured d a t a becomes Less a c c u r a t e a s the s h a f t a n g l e i s i n - c r e a s e d , e s p e c i a l l y beyond 60 d e g r e e s . I t should be noted, how- e v e r , t h a t the disagreenent i s one of conservatistn, i . e . , more lift and/or propulsive f o r c e =as produced f o r a given horsepower, Two c o n t r i b u t i n g f a c t o r s a r e ( 1 ) t h e d i f f i c u l t y of determining t h e a b s o l u t e a a s t angle under Loaded and dynamic c o n d i t i o n s , and ( 2 ) the e f f e c t s of t h e proximity of t h e t u n n e l c e i l i n g f o r t h e h i g h e r a a s t a n g l e s .

No tunnel a n g u l a r i t y c o r r e c t i o n s have been made t o t h e t e s t d a t a .

Such c o r r e c t i o n s might have some s i g n i f i c a n c e f o r t h e c o n d i t i o n s where t h e proprotor was a t Large mast a r g l e p o s i t i o n s ( 6 0 - 7 5 degrees) and the blade t i p s approached t h e tunnel r o o f . To put i n p e r s p e c t i v e a l l t h e v a r i a b l e s . i n c l u d i n g the t i p - p a t h - p l a n e angle (aTPp) and t h e t i p c o l l e c t i v e angle ( ~ T ~ P ) , maps of c a l c u - l a t e d d a t a a r e shown in Figures IV-20 and I V - 2 1 f o r t h e h e l i c o p t e r - conveision f l i g h t modes a t 80- and 140-knot tunnel speeds.

1. Hover F l i g h t Mode h i g u r e s IV-9 and I V - 1 2 show t h e r e s u l t s of t h e hover t e s t s . Data e r e taken f o r proprotor t i p speeds LI 600 and 7 4 0 f e e t per second and f a r a a s t angle s e t t i n g s of 9, 3 0 , 60, and 7 5 degrees. Overall c o r r e l a t i o n between the t e s t and predicted data is good. For t h e lower mast angle s e t t i n g s , t h e induced flow i n t h e tunnel i s r e s p o n s i b l e f o r che i n c r e a s e i n power r e l a t i v e t o t h e c a l c u l a t e d d a t a . The tunnel f a n s were operated i n reverse t o minimize flat-7 through t h e t e s t s e c t i o n , but were not e n t i r e l y e f f e c t i v e . The d a t a c o r t h e h i g h e r mast qng-Le s e t t i n g s may be s l i g h t l y o p t i m i s t i c f r o % a power s t a n d p o i n t due t o the r e c i r c u l a t i o n i n t h e t e s t s e c t i o n , although t e s t s e c t i o n top doors were p a r t i a l l y opened t o minimize t h a t e f f e c t . The r o t o r plane was more t h ~ n one r o t o r diameter above t h e tunnel f l o o r , ~ t h i c h would o r d i n a r i l y axclude ground e f f e c t : however, s i 9 c e the t u n n e l i s rounded on t h e s i d e s , o t h e r r e c i r c u l a t i o n e f f e c t s may have been p r e s e n t .

The p r e d i c t i o n i f hover p e r f o m a a c e and ( t o a l e s s e r e x t e n t ) p o t e n t i a l aaoeuve r i n g performa;~,.e of a new r o t o r depends upon t h e accuracy of p r e d i c t i o n s of che maximum t h r u s t of t h e r o t o r .

P r e d i c t i o n s of +ne maximum t h r u s t of t h i s r o t o r have been more c o n s e r v a t i v e a s t h e y shouid have been, with t h e predicted rnaximun values f a l l i n g well below t h a t measured. On the assumption t h a t

@ BELL HEUCQ-R cow-rur

the p r e d i c t i o n methods were p r i m a r i l y r e s p o n s i b l e f o r t h e non- exact c o r r e l a t i o n , not t h e t e s t procedures o r d a t a r e d u c t i o n .

t h e p r e d i c t i o n methods were examined.

O f s p e c i a l :'t,:+rest was t h e de2ree t o which some of the i n p u t v a r i a b l e s s f r v e t e d the p r e d i c t i o n s . This was i n v e s t i g a t e d by modifying the i n p u t s a s follows: - Tip Loss f a c t o r = 0.97 - T i p l o s s f a c t o r = 1 - 0 0

- T i p l o s s f a c t o r = 0.97 and modified blade element

c h a r a c t e r i s t i c s This was a d e p a r t u r e from t h e use of t h e expression given i n 1V.C f o r c a l c u l a t i n g the t i p l o s s f a c t o r . The 0.97 f a c t o r has been used i n t h e p a s t t o improve t h e c o r r e l a t i o n between t h e predicted and t h e a c t u a l performance of h e l i c o p t e r t a i l r o t o r s .

Like t h e t a i l r o t o r , the proprotor has a comparativelj- high d i s c l o a d i n g , and t h e r e f o r e t h e t i p Loss f a c t o r which is a p p l i c a b l e f o r one should be c o r r e c t f o r t h e o t h e r .

The 0.97 t i p l o s s f a c t o r d i d improve t h e c o r r e l a t i o n , but not s u f f i c i e n t l y . Tho 1 .r)O fisctor was a l s o inadequate, and f u r t h e r - more was r a t h e r u n r e a l i s t i c . T h e r e f o r e , t h e t i p l o s s f a c t o r was r e t s r n e d t o 0.97 a t ~ d the blade element d a t a modified s l i g h t l y a s shown i n Figures IV-2 through IV-5. This combined adjustment put the c a l c u l a t e d performance i n good agreement with t h e meas- ured perf ornance .

2 . Helicopter-Conversion F l i g h t Mode Figures IV-14 through IV-19 show (nondimensionally) t h e v a r i a t i o n of horsepower a s a function of proprotor l i f t and propulsive f o r c e f o r mast a n g l e s of 1 5 , 30, 6G, and 75 d e g r e e s , a t u n n e l speed range of 80 t o 14G k n o t s , and t i p speeds of 700 and 740 f e e t per second.

Generally, c o r r e l a t i o n i s b e s t a t t h e lower mast a n g l e s . with some d e v i a t i o n s o c c u r r i n g with i n c r e a s i n g mast a n g l e . The d e v i a t i o n s i n a b ~ l i t y t o a t the higher mast a n g l e s may be a t t r i b u t e d t o an determine t h e t i p - p a t h - p l a n e angle with s u f f i c i e n t accuracy and t o t h e f a c t t h a t a c o n s t a n t mast angle does not i n s u r e a c o n s t a n t t i p - p a t h - p l a n e angle a s shown by t h e following r e l a t i o n : An a r b i t r a r y l i m i t of i 0 . 5 degrees has been placed on t h e value of f l a p p i n g a n g l e , and a l l d a t a o u t s i a e t h i s l i m i t have been flagged. A t t h e higher mast a n g l e s t h i s angle range lirnit could show the s e n s i t i v i t y of the c t h e r v a r i a b l e s t o t h e t i p - p a t h - p l a n e and t i p - c o l l e c t i v e a n g l e s .

@ BELL HELICOPTER c--Nr

Figure IV-22 shows the c o r r e l a t i o n between measured and c a l c u - l a t e d horsepower f o r a l l of the h e l i c o p t e r - c o n v e r s i o n d a t a . The data were c a l c u l a t e d f o r the measured l i f t and propulsive f o r c e f o r t h i s c o r r e l a t i o n . The f i g u r e shows a conservative t r e n d f o r the c a l c u l a t 2 d d a t a .

3. Airplane F l i g h t Mode Figures IV-23 thrsugh.IV-36 c o n t a i n a i r p l a n e mode ( a x i a l ) d a t a .

Figures IV-23 through IV-32 show t h e propulsive force and e f f i - ciency d a t a a s a f u n c t i o n of horsepower. Data a r e presented f o r tunnel speeds of 120, 1960, and 185 knots and t i p speeds ranging from 400 t o 740 feed per second. C o r r e l a t i o n f o r t h i s d a t a i s good and where a d e v i a t i o n e x i s t s , t h e c a l c u l a t e d d a t a a r e u s u a l l y c o n s e r v a t i v e .

Figures IV-33 through IV-35 show t h e above-mentioned d a t a i n t h e nondimensional form of propulsive e f f i c i e n c y a s a f u n c t i o n of power c o e f f i c i e n t .

Figure IV-36 show a c o r r e l a t i o r : between c a l c u l a t e d and measured norsepower f o r a l l t e s t c o n d i t i o n s . The c a l c u l a t e d horsepower d a t a were e v a l u a t e d f o r t h e measured v a l ~ l e s of propulsive f o r c e .

This f i g u r e shows t h e t r e n d of the c a l c u l a t e d d a t a t o be more conservative a t t h e h i g h e r values of horsepower.

0 BELL HELICOPTER ConapANr

I 4

8 - 4 P I -

BLADE STATION - INCHES

/ Figure I V - 1 . Computer R e p r e s e n t a t i o n o f t h e 25 - F o o t P r o p r o t o r Blade.

I

I

,MACH NUMBER = - 7 I I I I I I I I STAWARD DATA I I , -

-- - - MOD IF1 ED DATA

I I

I I I I I

I

I II

.6 ,r. / I

I

I

a 4 -

I L

I

a 3 i

I

j I-

m LIFT COEFFICIENT Figure IV-2. Proprotor Blade Section Data, A i r f o i l S e c t i o n No, 1, Blade s t a t i o n 0 . 0 7 5 t o 0.45.

---- MODIFIED DATA

LIFT COEFFICIENT F i g u r e I V - 3 . P r o p r o t o r B l a d e S e c t i o n D a t a , A i r f o i l S e c t i o n No. 2 , B l a d e Station 0 . 4 5 t o 0.70.

LIFT COEFFICIENT Figure I V - 4 . P r o p r o t o r BLade S e c t i o n Data, A i r f o l l S e c t i o n No. 3 , Blade S t a t i o n 0 . 7 0 t o 0.90.

----

HODIFIED DATA

- -

, 4 -6 - 8 1 . 0 1 . 2 LIFT COEFFICIENT Figure fV-5 Proprotor B l a d e Section Data, Airfoil Sectibn No. 4 , Bladc Station 0 . 9 0 t o 1-00.

BASE PRESSURE C!IEFFICIENT (A P/q 1 Figure IV-6. Spinner Drag Tare, A i r p l a n e Flight Mode.

0 20 30 60 80 100

MAST ANGZE OF ATTACK - DEGREES

Figure IV-7. Spinner Lift Tare Versus Mast Angle of Attack, Helicopter-Conversion Mode. .

1)RAC;/DYNAMIC PRESSlIRE - SQUARE FEET

0 DEGREES, RUN 1 = 30 DEGREES, RUN 2 , 7 , 8 0 2000 4000 60CO 8000 10000 120

THRUST/DENS ITY RATIO - POL'NDS

Figure I V - 9 . aimensions1 iiover Power Required.

DEGREES, RUN 1 = 30 DEGREES, RUN 2.7.8 THRUST COEFFICIENT (CT) Figure I V - L O . Nondimensional Hover Power Required.

@ BELL HELICOPTER E~U-rur

.-

I

I

I

I I

-

-

- CAEUIATED

I I zwT = 0 DECREES, RUN 1

-

lHATT = 30 DEGREES, i(UN 2,7,8 zMAST = 75 DEGREES, RUN 14,15

- 1 ' I

I

! I I 1

, I I I

I

a

I

I

i l i l l l D

w TYPICAL OPERATING

@ REQUI REMENTS

i V

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I 0 o /

v

/

f

v

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04 06 .08 .lo • 1 2 ,14 16 .18

HELICOPTER THRUST COEFFICIENT/SOLIDITY (cnH/O ) F i g u r e IV-11 .

Nondimens i o n a l H o v e r Power R e q u i r e d .

@ BELL HEUCOPTER c - - r

h

I

I

- CAUULATED

zMAST = O DEGREES, RUN 1 O = 30 DEGREES, RUN 2,7,8 'MAST = 75 DEGREES, RUN 1 4 , 1 5 A n M A s ~ 1.0- I .

.

0 -04 -38 . 12 ,16 , 20

HEUCOPTER TFAUST COEFFICIENT/SOLIDITY (cTdo)

Figure IV-12. Hovezing Figure of Merit Versus Helicoptzr Thrust Coefficient/'Solidity.

7 5 DEGREES, RUN 14, 15

- C U U LATE D DATA-

TIP LOSS = "STANDARD"

---- TIP LOSS = 0 - 9 7

---

TIP LOSS = 1-00

------

Irc Irc W ,012 P1 -04 .06 -05 .10 12 14 -16 .18 HELICOPTER THRUST COEFFICIENT/SOLIDITY ( c T ~ / ~ ) Correlation of Nondimensional Hover Figure IV-13.

Power Required -

CAUULATED MAST ANCLE = 30' F ' N O M = .241 (RUN 9) QR = 7i39 fps LIFT COEFFICIENT (CL) Figure IV-14. Nondimensional Ilelicopter-Conversion Perf onnanca .

LIFT COEFFICIENT (CL) Figure I V - 1 5 . Nondimensional Helicopter-Conversion Performance.

MAST ANGLE = 1 5 ' 0 & I N O M = ,319 (RUN 28) QR = 740 FPS

FLAGGED S Y M B O ~ - lalS1 > . 5 '

LIFT COEFFICIENI' (CL) Figure IV-16. Norldimensional H e l i c o p t e r - C c n - - 2 r s i nn Performance.

FLAGGED S Y M B O U - 1 alSI > - 5 '

LIFT COEFPICI ENT (C L) Figure IV-17. Nondimensional Helicopter-Conversion Performance.

--

C A E U LATED p 'NOM = ,228 (RUN 12) MAST ANGLE = 6 0 0 (3 I3 = -319 (RUN 13) O P = 740 FPS LIFT COEFFICIENT ( CL) Figure IV-18. Nondimensional Helicopter-Conversion Performance.

LIFT COEFPICIEST <CL?

Figure IV-19. Nondimensional Helicopter-Conversion Perf ormacce.

TIP PAlH P U V E ANGLE OF A T i . 4 K - DEGREES

Figure IV-20.

C a l c g l a t e d S h a f t Horsepower Versus T i p Path Plane Angle of Attack, Helicopter- Conversion Mode, 80 Knots.

@ SELL HELICOPTER oovrm

PI CAL OPERATZ N

TIP PATH PLANE AVGLE OF ATTACK - DEGREES

F i g u r e IV-21. C a i c u i a t e d Shaft Horsepower Versus T i p P a t h ?lane Angle of Attack, Helicopter- Conversion Mode, 140 Knot s, FLAGGE3 SYMBOIS DENGrE ZR = 700 f p s AND PROPULSIVE FORCE Figure IV-22. Comparison of Measured Horsepower w i t h Calculated 3orsepower, Heliccpter- Conversion ?-?ode.

POUNDS

IJROPIJU 1 VE FORC E/I)E~NSI TY RATIO -

HORSEPOWEWDENSI TY RATIO Figure IV-24, Propulsive F o r c e and Efficiency Versus Horsepawer, A i r p l a n e Mode.

F i g u r e IV-25. P r o p u l s i v e F o r c e a n d Efficiency V e r s u s H o r s e p o w e r , A i r p l a n e Mode.

P R O P L L S I V E FORCE/DENSITY R A T I O - POUND2

I-' I-' h, h, W W F Se rQ 0? 0 E CO h, m 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0

EFFICIENCY Cq) - PERCENT

PROPULSIVE VI m V 00 W E 0 0 0 0 0 0 PROPULSIVE POI(CE/DENS IJY RATION-- P O U N E r r o o C 00 rG 0\ 0 & Oa h, m 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 tic- ; .

* G L I - 90.

-

I H z 1 W I U / e : W &70 I n C

V I i

I

/ p PF; 0 '

G60 NON- STANDARD z Lil H r TYPICAL CPERATING U H

/ REQUIREMENTS NOT

R APPLICABLE.

&SO > / I

?i

S4OL I / pG C U U IATE D & / RUN 23

I J = ]..I5

30' MAT = 0.71 160 KNOTS QR = 740 FPS

C

. - L 400 600 800 1000 1200 1400 16 ( HORSE?OWER/;)ENSITY RATIO Figure IV-28. Propuls.ve Force and Efficiency Versus Horsepower, Airplane Mode.

PROPULSIVE E F F I C I E N C Y (q) - PERCENT

Ci, E Vc 0 V Oc 0 0 0 0 0 0

P R O P U I S I V E FORCE/DENSITY R A T I O - POUNDS

PROPULSIVE FoRCE/DENSITY RAT I 0 - POUNDS

F i g u r e I V - 3 2 . P r o p u l s i v ~ F o r c e and E f f i e i e n c y V e r s u s Horsen er, Airplane Mode.

@ SELL H E -

POWd COEFFICIENT ("p) F i g u - e IV-33 P r o p u l s i v e Tfficirncv Versus Power C o e f f i c i e n t , 120 K n o t s , Airplane Mode.

IV-39

Figure IV-34. P r o p u l s i v e E f f i c i e n c y Versus Pow 1

Figure IV-35. Proptllsive Efficiency Versus Power C -06 0 . 8

+

- I

-02 -04 -06 -08 -10 COEFFICIENT (Cp) Power C o e f f i c i e n t and Advance Ratio, 185 Knots, A4rplane Mode, :V-41 - .

-

LINE 3F ZERO DESCREPANCY-7

zoo

.r nnn 0 200 4 OC C 30 660 L~~~ 1 2 ~ ~ 14nn CAEULATED IIORSEPWER/DENSI TY RATIO Figure IV-36. Comparison of Measured Horsepower w i t h C a i c u l . a t e d Horsepower, A i r ~ l a n e Mode.

@ BELL HE"a3FrER ar.rur

V. DYNAMIC STABILITY A . MODELING TECHNIQUES

1. - J u s t i f i c a t i o n of Semispan Canxilever Kodel -

Semispan c a n t i l e v e r models are often used i n the study of fixed- wing a e r o e l a s t i c problems. One reason f o r t h i s i s t h a t the a e r o e l a s t i c c h a r a c t e r i s t i c s of t h e c a n t i l e v e r nodel a r e gene r a l l y conservative ( f o r example, f l u t t e r car( be triggered a t lower speeds on the c a n t i l e v e r model than on the f r e e - f r e e wing).

Theory indicates t h a t t h i s i s a l s o t r u e f o r propl-otor/pylon d p a n i c s t a b i l i t y (sometimes c a l l e d & i r l f l u t t e r ) . Figure P-1 compares the calculated va:iation of t h e c h a r a c t e r i s t i c roots of a semispan c a n t i l e v e r with airspeed v i t h those of t h e complete a i r c r a f t . The %odes a r e i d e n t i f i e d a s b i n g s)armetric o r asymmetric f o r c l a r i t y . The math model used f o r theory i n - cludes the fundamental beam, c h ~ r d , and t o r s i o n aodes of the wing, the proprotor flapping and f i r s t inplarge bending nodzs.

For the cosplete a i r c r a f t , the s i x rigid-bodv degrees of freedom and associated a i r f rase zerodynamics a r e included, together w i t h the drive system ( including engine and interconnect s h a f t i n g ) .

The proprotor and airframe aerodynamics have been corrected f o r compressibility- The most notesorthv dffference between root l o c i of the c a n t i l e v e r wing and those of the complete a i r c r a f t i s the lower frequency and damping of the wing bean, chord, and t o r s i o n nodes of the c a n t i - l e v e r wing. This indicates t h a t the f l u t t e r c h a r a c t e r i s t i c s of the c a n t i l e v e r wing w i l l be conservative ( a s References 20 and 22 explain i n some d e t a i l ) .

2. Reduced S t i f f n e s s Test Stand For a portion of the dynamic t e s t s , a t e s t s t ~ n d was used having a s t i f f n e s s of one-fourth t h a t of the normal design s t i f f n e s s .

The use of t h i s stand and o p ~ r a t i o n a t one half the standard pronrotor rpm preserves the frcqzeacp reletionships between the blade flapping end wing modes and the proprotor inflow angles but not the blade. e l a s t i c ;node frequency relationships and com- p r e s s i b i l i t y e f f e 2 t s on the prcprotor ael-odynamics. However, calcutations indicate t h a t the l a t t e r e f f e c t s a r e not s i g n i f i c a n t for prop!-ctors w e r e the lowest blade e l a s t i c mode i s above one p e r rev and f o r airspeeds up t o 400 knots. Figure V-2 compares the root variation with airspeed of the one-fourth-design-stiff- ness t e s t stand with t h a t f o r the cantile-7er wing of Figure V - 1 .

The c h a r a c t e r i s t i c s of the wing fundamental beam and chord mode roots aro, nearly i d e n t i c a l , while those of tb.2 wing torsion mode a r e more heavily d a ~ p e d . The root c h a r a c t e r i s t i c s of the blade flapping and inplane e l a s t i c modes a r e d i f f e r e n t , but the prin- c i p a l moaes of i n t e r e s t are those of the wing beam and chord since they a r e the vast l i g h t l y damped.

The damping of the wing fundamental beam and chord modes i s compared f u r t h e r i n Figure V-3. Note t h a t the damping of modes on the one-foilrth-stiffness stand is only s l i g h t l y higher than t h a t of the c a n t i l e v e r wing a t airspeeds above 300 knots.

This is due t o compressibility e f f e c t s not being f u l l y represented i n the one-fourth-stiffness stand. (Also shown i n Figure V - 3 i s t h e small difference i n damping when the standard s t i f f n e s s hub r e s t r a i n t is used instead of the scaled one-quarter s t i f f n e s s hub r e s t r a i n t , as was the case i n the dynamic s t a b i l i t y t e s t . ) Clearly, the dynamic s t a b i l i t y data from the one-f ourth-design- s t i f f n e s s t e s t stand, along with that of the design-stiffness stand, can be used t o f o r e c a s t the dynamic s t a b i l i t y character- i s t i c s of the complete a i r c r a f t .

Figure V-4 is a summary of the measured damping of the wing bean m d e f o r both t e s t stands up t o a simulated airspeed of 408 knots. Note t h a t the data f o r the one-fourth-design-stiffness stand a r e not a t constant rpm; the c o l l e c t i v e p i t c h encountered its mechanical L i m i t at a simulated 310 knots, hence rpm increased with increasing airspeed (constant inflow angle). The apparent f a l l - o f f i n damping above 320 knots is due t o the increasing rpm.

For reference, the prcdicted damping i s shown f o r constant rpm and f o r the t e s t rpm range. Also shown i n Figure V-4 are d a t a from tests of a one-fifth-scale semispan a e r o e l a s t i c model. Thtse data c o n f i m the high l e v e l of proprotor pylon s t a b i l i t y predicted f o r the Model 300 a i r c r a f t .

TEST PROCEDURES Most of t h e dynzmic s t a b i l i t y data were taken with the wind tunnel balance frame Locked out. This was done because it ha?

become apparent during the vibration survey of the design- s t i f f n e s s t e s t stand i n s t a l l a t i o n t h a t there was r e l a t i v e l y strong co.~pling between t h e stand modes and the balance modes.

The frequencies of the fundamental beamwise, chordwise, and torsion modes were s l i g h t l y lower with the balance f r e e . Also, t h e damping of t h e wing chord mode was 12wer with the balance f r e e . For the one-fourth-design-stiffness stand the l a t e r a l t r a n s l a t i o n frequency of t h e balance (1.8 cps) was predicted t o have strong coupling with the beamwise bending frequency of the stan? (1.4 cps) and t o influence the dynamic s t a b i l i t y character- i s t i c s .

The frequency and damping of the t e s t stand modes were determined a t each data point by means of the two-square-foot ae20dynamic vane iocated on the nacelle f a i r i n g (mountecl v e r t i c a l l y i n the tunnel sense). The f requeocy of o s c i l l a t i o n of the vane i n p i t c h was adjusted a t each point so as t o force the selected mode a t resonance. ( A l i s s a j o u s figure was generated on a dual-axis cscilLoscope t o locate resonant frequencies .) When the response turned o f f and the was s u f f i c i e n t l y high the vane e x c i t a t i o n was r e s u l t i r t g fecay of the t e s t stand motion anaLyzed t o determine the mode frequency and damping. Figure V-5 shows a t y p i c a l decay h i s t o r y and i l l u s t r a t e s the method used t o c a l c u l a t e f req uency ar,d damping.

This method proved highly s a t i s f a c t o r y f o r t h e stand fundamental t?am and t o r s i o n modes, but did not permit ready determination cf the stand chordwise bending node c h a r a c t e r i s t i c s . I n the case of the chordwise mode the e x c i t a t i o n was low compared t o the beamwise and torsion modes. However, it was a l s o evident from the absence of any t r a n s i e n t response i n t h e chordwise bending mode t h a t i t s damping was r e l a t i v e l y high. (With the proprotor blades removed and replaced by equivalent weights, the chord mode damping was low and frequency and damping were measured.)

The sysLca natural frequencies were also excited by o s c i l l a t i n g the swashplate. This method was s a t i s f a c t o r y , but had t h e same limitetfons a s the aerodynamic vane i n s o f a r as the cling chord mode was concerned.

A frequency sweep i n the frequency range of the proprotor f l a p - ping and blade inplane modes conf i m e d t h a t these modes were heavily damped. Frequency sweeps up t o 45 cycles per second showed l i t t l e response i n any modes o t h e r than the wing funda- mental beam, chord, and t o r s i o n aodes.

C. MEASURED STABILITY CHARACTERISTICS 1. Design-Stiffness Test Stand The contribution of the wing and pylon a i r l o a d s t o t h e t e s t stand s t a b i l i t y c h a r a c t e r i s t i c s w a s de t e mined by rzmoving the prop- r o t o r blades ~ n d replacing them w i t h equivalent lumped weights.

Frequency and damping were determined f o r wind speeds u p t o tne tunnel maximus speed (204 knots) and f o r angles of a t t a c k up t o 18 degrees a t 100 kncts an2 8 ciegzees st '185 knots.

With the proprotor blades i n s t a l l e d , w i t l d speed and angle of a t t a c k sweeps were made, a s wzll as proprotor rpg swceps up t o the overspeed rpn. Two pylon configurations were tested: (1) a conf i g ~ r a t i o n s i r a ~ ' a t i n g the pylon -to -wit?g attachment when the gylon is f u l l y converted s o t h a t a downs+op s engaged ( t h e

yaw 1 i n k i n configctration) , and ( 2 ) a configuration simulating

a confition where the ~ v l o n i s ??most I u f l v converted. but the domsto;, is not yet er.i&ed ( t h e yaw l i n k but c ~ c f i ~ u r a t i o n ) .

The dynamic s t a b i l i t y was found t o be e s s e n t g a l l y the same fjr bo' h conf iruracions.

Figures V-6 through V - 8 s u m r i z e the meastired v a r i a t i ~ n i I t e s t stand beamwise, chordwise, ,nd torsion c nod^ damping with a i r - The static frequency and da..ipZng a r e sho~vn f o r esch mode.

speed.

The icfluc 7ce of w i n g h a s t angfr of attack on tke dyrramic s t a - a i l i t y was measured a t two z 'rspeeds. The beam mode frequency and damping variation with angle of a t t a c k a t the higher a i r - speed, 185 knots, i s shown i n Figure V-9. The v a r i a t i o n i s negligible. Witlg chord and torsion mode e x c i t a t i o c s were not made. The influence of proprotor rpm on beam lnode s t a b i l i t y is shown i n Figure V - L O .

2. One-Fourth-Design-Stiffness Test Stand Tne contribution of the wing-pylon aerodynamics t o the system frequency and damping was established by removing the blades and replacing them with equivalent weights. The wind speed was w r i e C up t o the tunnel maximum speed and mast angle of a t t a c k sweeps made a t 92.5 knots (proprotor inflow equivalent t o the 185 knot data from the design s t i f f n e s s t e s t ) and 132.5 knots (simulated 265 knots).

The airspeed and mast angle of a t t a c k sweeps were repeated w i t h the proprotor blades i n s t a l l e d . I n a d d i t i o n , proprotor rpm s w e p s up t o a simulated 1200 r p m were made a t 150 and 170 knats, simulating 300 and 340 knots respectively. The pylon configura- t i o n t e s t e d simulated the pylon f u l l y converted and on a down- stop.

The v a r i a t i o n i n frequency and damping of the stand beam, chord, and t o r s i o n modes with wind speed is shown i n Figures V - 1 1 through V-13. Data from BHC t e s t s of a one-fifth-scale dynamic model a r e shown, a s -ell a s the predicted frequency and dampi-ng.

Note t h a t the one-fourth-design-stiffness data a r e not f o r a constant rpm; the upper l i m i t on the c o l l e c t i v e p i t c h was reached a t a wind speed of 150 knots and the rpm therefore increased a s wind speed increased (time considerations precluded reindexing the blade pitch hcrns t o achieve the c o l l e c t i v e re- quired f o r operation a t 229 rpm up t o the maximum tunnel speed).

The downward trend i n s t a b i l i t y a t blind speeds above 150 knots is associated w i t h the increasing rpm. A t a constant 229 rprn the s t a b i l i t y would increase monotonically -7ith wind speed as shown by the predicted damping a t constant rpm. The limited data on the chord mode frequency and damping a r e due t o the i n a b i l i t y t o e x c i t e the chord mode with the aerodynamic vane.

Figure V-14 shows tne v a r i a t i o n i n freauency and damping of the wing beam mode with angle of a t t a c k a t a simulated airspeed of 265 knots. The angle of a t t a c k data taken a t 92.5 knots were i n good agreement with the 185-knot design-stiffness stand data.

The measured v a r i a t i c n i n frequency and damping of the wing beam mode with proprotor rpm a t 150 knots and 170 knots i s shcwn i n Figures V-15 and V-16, respectively. The damping trend wit' r p ~ f o r the 170-knot i 3 4 0 knots simulated) condition indicsted hat n e u t r a l s t a b i l i t y would occur a t a somewhat higher rpm thav t h e maximum t e s t e d . The maximum t e s t e d was 420 rpm and simulated 840 rpm, o r 180 percent of the normal operating rpm. Damping

of t h e w i n g torsion mode showed a s i a i l a r reaction t o rpm, as

shown i n F ~ g u r e V - 1 7 .

@ BELL HEUCO-R -rrm

D. CORRELATION OF THEORY WITH MEASURED DYEU'AMIC STABILITY CHARACTERISTICS For each t e s t c o n d i t i o n the frequency and damping of t h e t e s t s t a n d modes were p r e d i c t e d by means of two independent a n a l y s e s .

These predicted c h a r a c t e r i s t i c s a r e shown i n Figures V-6 through V-17. The l i n e a r a n a l y s i s , BHC P r e p r o t o r S t a b i l i t y A n a l y s i s , DYN4 (computer program DRAL06). i s based upon p e r t u r b a t i o n t h e o r y . The nonlinear a n a l y s i s , BHC P r o p r o t o r A e r o e l a s t i c Analysis , DYNS (computer program ARAP08), i s based on n o n l i n e a r open-form theory. A b r i e f d e s c r i p t i o n of each a n a l y s i s is given i n t h e paragraphs below. Complete d e t e i l , i n c l u d i n g t h e equa- t i n n s of ~ o t i c n , i s contained i n References 27 and 2 8 .

- Proprotor S t a b i l i t y A n a l y s i s , D Y N 4 Program DYN4 i s a 1 i n e a r , tGlenty-one-degree-of -f reedom proprotor s t a b i l t t y a n a l y s i s . It can determine t h e proprotor/pylon, blade motion, a n d f l i g h t mode s t a b i l i t y c h a r a c t e r i s t i c s of a t i l t - r o t o r v e h i c l e . A t i p - p a t h - plane r e p r e s e n t a t i o n is used f o r the p r o p r o t o r , and l i n e a r aerodynamic functiorrs a r e assumed. D e t a i l s such

a s p i t c h - a x i s preconing, underslinging , p i t c h - f l a p

coupling, and f l a p p i n g r e s t r a i n t a r e i n c l u d e d . The f i r s t inplane blade mode i s represented . Control system f I e x - i b i l i t y may a l s o be simulated. Five coupled wing/pylon e l a s t i c modes a r e represented: wing \earn, chord, and t o r s i o n , and pyloc p i t c h and yaw. 3ix r i g i d body degrees of freedom a r e included t o a l l o w s i ~ u l a t i o n of f r e e - f r e e bcdy c o n d i t i o i ? . ~ and t h e a i r c r a f t s h o r t period f l i g h t m r Ces .

Inputs t o DYN4 a r e lumped parameters d e s c r i b i n g t h e dimensions, i n e r t i a , s t i f f n e s s , and kinematics of t h e a i r c r a f t being simulated. Standard a i r c r a f t s t a b i l i t y d e r i v a t i v e s a r e used t o s t u d y t h e i n f l u e n c e of tkLe proprotors snd t h e wing/pylon dynamics on t h e s t a b i l i t y of the f l i g h t modes. Outputs a r e system eigenvalues and e i g e n v e c t o r s . Root l o c i can be p l o t t e d a u t o m a t i c a l l y .

- Proprotor A e r o l a s t i c A n a l y s i s , DYN5

Program DYNS is a n o n l i n e a r , open-form proprotor a e r o - e l a s t i c a n a l y s i s t h a t uses t h e same b a s i c mathematical model a s DYN4. This program c a l c u l a t e s proprotor l o a d s ,

v i o r a r i o t ~ , and s t a b i l i t y i n h e l i c o p t e r , conversion, and

h i g h - s ~ e e d modes. A s p e c i a l v e r s i o n of t h e program was developed f o r t h e A i r Force under Contract F33615-69-C- 1339, "Vibration i n V/STOL A i r c r a f t ,'' Reference 28.

The dynamic e q u a t i o n s of motion were derived usi-ng t h e Lagrangran method. Provisions f o r Large f l a p p i a g and f e a t h e r i n g motion a r e included i n DYN5. S1r2ll allgle assumptions a r e made on the wing-pylon and blade e l a s t i c degrees of freedom. The aerodynamic functions used i n DYNS are the same a s those used i n the Bell Rotor Perform- ance Analysis, F35. C l , C d , and Cm a r e inptit i n t a b u l a r form f o r a 180-degree range of angle of 3ttack and f o r Mach numbers d p t o 0.9. Tables f o r d i f f e r e n t p r o f i l e s may be input t o account f o r differences i n the blade section from root t o t i p , thereby properly accounting f o r blade s t a l l and compressibility e f f e c t s .

DYNS i s programmed f o r solution by d i g i t a l computer. A predictor-corrector i n t e g r a t i o n technique i s used i n t h e solution of the simultaneous equations of motion; integra- tion i n t e r v a l m y be varied a s a function of azimuth.

Input t o the program c o n s i s t s of lumped parameters ai~d the coupled normal modes of the wing, pylon, and proprotor.

The output consists of a time h i s t o r y of the wing and pylon motions, and t h e blade flapping and e l a s t i c deflec- t i o n . Generally, i n i t i a l conditions a r e input t o minimize the time required f o r convergence t o steady s t a t e t r i m .

For s t a b i l i t y investigations, the t r a n s i e n t response t o external inputs o r t o i n i t i a l conditions can b2 calcu- lated.

Figures V - 1 1 through V-16 show frequency and damping data from BRC t e s t s of a one-fifth-scale a e r o e l a s t i c model. These data j u s t i f y the hypothesis t h a t proprotor dynamic behavior can be predicted by means of smaller scale models or a minimum-sized research a i ~ c r a f t .

I n general the c o r r e l a t i o n is e x c e l l e n t . The nonlinear analysis is i n b e t t e r agreement with the a c t u a l damping than i s the l i n e a r a n a l y s i s . but both analyses a r e conservative. W; i l e few meslsured data on the wing chord mode were obtained, t h e excellent c o r r e l a t i o n with regard t o the wing beam mode indicates t h a t the s t a b i l i t y of the chord mode can be predicted. Further- more, the one-fifth-scale model chord mode c h a r a c t e r i s t i c s a r e i n good agreement with theory.

BLADE AiIVANCING BLADE ADVANCING f I N P W E MODE I

d

A HI= FREQUENCY FLAPPING WING T0RSI';IN DENOTES AIRSPEED RPH = 458 BLADE REGRESS18 BLADE REGRESSING I N P I M E m D E I ItJPLANE MODE BEAM

LOW E'REQUENCY FLAPPING f

LOW FREQUENCY FLAPPING 3 L.

- I ,

b - -

-25 -20 -15 -10 -5 0 5

DAMPING, Con - k

DAMPING, C u , - RAD/SEC

S-RiC F (a) CANTILEVER WING (DYNAMIC TEST STAND) Root Loci of Cantilever Wins C w Figure V-1.

v-7 - 1 4 0 t '-40 POWER TURBINE NG

*

I€ BLADE ADVANCING

IHPLANE MODE /'

,Iz0 I LADE ADVANCING # tJ ' A P1 cn 'RE I I HIGH FREQUENCY K A P P I N G PP HIC;ii FREQUENCY FLAPPING C C i WING TORSION WING TORSION s o ;

W I N G BEAM C /

P

a ~r WING l CHORD m R INTERCONNECT I g L M E REGRESSING INPLANE MODE 0 2 0 INPLANE MODF < FLAPPING ? R e o w FLWPING d - 0 -4 MODE

v v 4 t . - I

15 - 10 -5 0 5 -25 -20 -15 -10 -5 0 5

DAMPING, Cwn - RAD/SEC

RAFEIING, Cw,, - RAD/SEC

SYEPIET'UC FREE-FREE I'YMNL'TRIC FREE-FREE (b) COMPLETE NRCRAPP .tilever Wing Compared to Complete Aircraft Root Loci.

v - 7

FOLDOUT F F Z A ? n r

..-

2-

r3. 0 .

-

BLADE ADVANCING

INPLANE mDE i

w w 6 ' \ J Z FLAPPING > U REGRESSING z v- - c.

I M P L R N E MODE C:

109 300 -

tr

-

a 200 400 KNCTS

-.- -

AIRSPEED N n -1.0 4 C M T - m R WING 0 1/4 STIFR4ESS T€ST Z STAEJC I BLADE REGRESSING I H P U N E W3E CHORD, WING BEAM F r g ~ r e V-2. Comparison of Root Loci for Set.lspan Canti;equ-er Wing and One-Fourth-Stiffness Test Stand, 100 200 30 0 4 0 0

SIMVLi:'i'ED AIRSPEED - IXOTS

Figure V-3, Wing Beam and Chcrd Damping Versus Simulated Airspeed.

100 200 30 0 400

SIMULATED TRUE AIRSPEEC - K:;O'?S

Figure V-4. Xinq Bern Bending Mode S t a S i l i t y l7ersus Airspeed.

TER AT HAST CASE REAHWISE BEMliING #O&E?,VT AT STA 38 DABPING RATIO EXTRACTED FROM DECAY L'SING ',. .

AZ1lrlL.H BLIPS r f f f ) ~ ~ ~ i t + t t ? . w i t -

I

Figure V-5. Traces of Decay of Wing B e a m V i b r a t i o n s for One-Fourth-Stif f n e s s T e s t Stand,

-3 - 1 1

. ..

0 ROTOR OX, YAW LIXK IS

IEOTOR O N . '1IAW LIXK OCT 8 ROTOR OFF m9RY 1RCTC9 0 1 2 ' .

Y A W LIYK IT) NONLINEAR ANALYSIS n

TUNNEL VELOCITY - KXOTS

Figure Ti-6. Wing Beam Frequency and Damping Variation with Tunnel Velocity for Standard Stiffness Test Stand.

THEORY (ROTOR OX. YAW LINK IS)

TUNNEL VELOCITY - KNOTS

Wing Chord Frequency an< Damping Variation with Figure V-7.

Tunnel V e l o c i t y for Stancard S t i f f n e s s T e s t Stand.

4 0 ROTOR OR. YAW LINK I?;

0 ROTOR OX. YAW LIXK OLT

A ROTOR OFF, YAX LIXK IN

NONLINEAR, ROTOR ON. YAh LISK I N

---

LINEAR, ROTOR ON, YAW LINK IN

-

I I 1 0

A

A L 2 , L b

TUNNEL VELOCITY - KNOTS

Figure V-8. Wing Torsion Frequer.cy and Damping Variation with Tunnel Velocity for Standard S t i f f n e s s Test Stand.

MAST A&;LE OF ATTACK - DEGREES

Wing Beam Frequency and D a m p i n g V a r i a t i o n Figure V-9.

w i t h An,le-of-Attack for Standard S t i f f n e s s Test Stand.

DAMPING - PERCENT CRITICAL

FREQUENCY - CYCLES/SECOND

C, RPX = 229 EXCEPT WHERE NOTED YAW LINK I N 4 ROTOR OFF L/S-SCALE AEROELASTIC MODEL THECRY BHC I R 6 D TEST NONLINEAR ANALYSIS 2 4 6 RPM-I -

TUNNEL VELOCITY - KNOTS

F i g u r e V-11. Wing Beam F r e q u e n c y and Damping Variation with T u n n e l Velocity for One-Fourth-Stiffness Test Stand.

YAW LINK IN F i g u r e V-12. I.!Fr?g Chnrd F r ~ q l ~ e n c y a n d &??ping V a r i a t i o n with Tunnel Velocity for One-Fourth-Stiffness Test Stand.

TUNNEL VELOCITY - KNOTS

Wing T o r s i o n Frequency and Damping V a r i a t i o n F i g u r e V-13.

with Tunnel V e l o c i t y f o r One-FGurth - S t i f f n e s s Teat Stand.

0 BELL HELICOPTER COMRNr

A ROTOR OFF

0 1/5 SCALE AZROET,USTIC MODEL

M A S T ANGLE OF ATTACK - DEGREES

F i g u r e V - 1 4 . Wing Beam Frequency and Damping Variation w i t h Angle-of-Attack for One-Fourth-Stiffness T e s t Stand.

DAMPING - PERCENT CRITICAL FREQUENCY - CYCtES4/!3ECOND

DAMPING - PERCENT CRITICAL

Figuxe V-17. Wing Torsion Frequency and Damping Variation w i t ! !

Proprotor RPM for One-Fourth-Stiffness Test Stand.

a l a d e f l a p p i n : i n a i r p l a n e mode. n a n i f e s t e d as s t e a d v t i l t i n g o f t h e p r o p r o t g r d i s c . r-QS ?rea+-ured f o r e i g t i t c o ~ b i n a t i o n s cf r p an6 w i n d - t u n n e l s w e d d c r i n p t h e d y a a i c s t z b i l i t y t e s t , One c m b f c a t i o n was t . p e a t e d d f ~ r i n g t5e powered test- t o o b t a i ? a n i n d i c a t i o n o f t h e i n f l u e n c e of ~ i n g - p r c - o r o t o r z e r o d ; ~ a ~ i c iqter- f e r e n c e on b l a d e f l a p p i n g .

1. Measured D e r i v a t i v e s F i g u r e V I - L shows a n e x a ~ p l e of t h e ~ e a s u r t d b l a d e f l a p p i n g v e r s u s aast a a g l e o f atta?K. T b t n e a s c r e d l o n g i t u d i n a l f l a p p i n g ( a l ) . i s n e a r l y L i n e a r i n n a t u r e : a s t r a i g h t l i n e a p p r o x i ~ a t i o n was a a d e t o a r r i v e a t t h e d e r i v a t i v e . d a l 'daT. The lateral

f l a p p i n g I b i ) , is n e a r l y l i n e a r fro?: aT = O to n, = -6 . b u t

has a s h i f t b e t w e e n a, = 0 a ~ d a , = - 2 - T h i s s h i f t appears iri o t h e r f l a p p i n g v e r s u s aast a n g l e o f a t t a c k d a t a a n d is b e l i e v e d t o be c a u s e d b y a n i n s t r u - ; . e n t a t i o n i r r e g u l a r i t y , The n e a s u r e d Lateral f l a p p i n g d e r i v a t i v e , d b ~ 4 a , w a s est i ~ a t e d u s i n g t h e s l o p e b e t w e e n aa = 0 arrd a?, = +6 .

The f l a p p i n g , e r i v a t i v e s f o r t h e o t h e r r v - w i n d s p e d co-nbina- t i o n s are t a b u l a t e d i n T a b l e I T I - I . C o a p a r i s o n o f t h e 185-knot.

5 9 0 - r p n d a t a fro? t h e d y n a a i c s t a b i l i t y t e s t w i t h t h o s e f r o m the powered t e s t i n d i c a t e s t h a t w i n g - p r o p r o t o r a e r o d v a n i c i n t e r - f e r e n c e h a s n o i n f icence on t h e l o n g i t u d i n a l f l a p p i n 9 r e s p o n s e b u t n e a r l y d o u b l e r t h e lateral f L a p p i n g r e s p o n s e - Houever . s i n c e t h e Lateral f l a p p i n g is r e l a t i v e l y s a a l l . t k e c v e r a l l i n f l u e n c e on t h e f l a p p i n g r e s p o n s e t o a n g l e o f a t t a c k a p p e a r s t o be z e g l i - g i b l e .

T q r r e l a t i o ~ f T h e o r y w i t h ? & a s u r e d F l a p p i n g Two a n a l y t i c a l n e t h o d s a r e compared w i t h t h e o e a s u r e d f l a p p i 5 5 d e r i v a t i v e s s h o r n i n F i g u r e 1'1-1 a n d T a b l e V T - 1 - The l i n e a r t h e o r y is b a s e d o n t h e l i c e a r . sna! l p e r t u r b a t i o n . p r o p r c t o r d y n a a i c s t a b t l i t y a n a l y s i s : t h e n o n l i n e a r t h e o r y i s basec on the L i n e a r o p e n - f o r n p r o p r o t o r aeroelastic a n a l y s i s . The b a s e s , a s s u . n p t i c n s , a n d l i m i t a t i o n s o f t h e s e a n a l y s e ? , o i s c u s s e d e a r l i e r i n S e c t i o n V , a l s o appLy t o t h e f l a p p i , ~ g d e r r v a t i v e t h e o r y .

Both t h e o r i e s a c c u r a t e l y p r e d i c t t h e a e a s u r e d tong; t u d i n a l d e r i v a t i v e s . The n o n l i n e a r t h e o r y i s i n b e t t e r a g r c e r e n t w i t h t h e lateral d e r i v a t i v e s t h a n i s t h e l i n e a r t h e o r y . T h i s i s d u e t o t h e s e c o n d a r y f a c t o r s s u c h as p r e c u n e , b l a d e f ! . c x i b i L i t y , a n d w i n g - p r o p r o t o r a e r o d y n a a i c i n t e r f e r e n c e i n c l u d e d < a t h e non- l i n e a r t h e o r y .

I n g e n e r a l . , t n e c o r r e l a t i o n between t h e o r y and veclsured f l a p o i r l g is e x c e l l e n t . F i g u r e V I - 2 s u m a r i z e s t h e c o r r e l . a t i o n with r e p a r d t o t h e t o t a l f l a p p i n g d e r i v a t i v e itP/aaa.

The autoaatic flapping cc;ltrc.Lle r was tested o?ly or! the oce- fourth-desig?-stiffness stand. A s noted e a r l i e r i n Section V .

the one-fourrh-design-at i f fness n v b r e s t r a i n t was aot used during rhc one-fourth- lesign-st i f f n ~ s s stand t e s t . Since the flapping c o n r r o l l e r w3-- confipsred t o be used with the cne-fourth- design-stif fness hub r e s t r a i n t . i t s performance was s o a e h a t reduced, Furthernore, the swashr>late phasing was not opti-nun f o r the aooocyclic c o n t r o l l e r . Youever. the t e s t data a r e ade- quate t o v e r i f y c o r r e l a t i o n betweera theory and neasured control- l e r c h a r a c t e r i s t i c s and t o deaonstrate the f e a s i b i l i t y of a f lapping control i e r .

1. Test Results

-

To e v a l ~ a t e c o ~ t r o l l e r s t a b i l i t y and performance. the following procedure was employed: - The controLLer was txrned off except f o r the hydraulic a c t u a t o r .

- The c o n t r o l k r gains were selected.

- A s t e p was put i n t o the a c t u a t o r . simulating t o sore extent a s t e p v e r t i c a l g u s t . t r a n s i e n t l y e x c i t i n g both the w i n g arid the proprotor.

- The c o n t r o l l e r was a c t i v a t e d , r e s u l t i n g i n a reduction of f lapping - - The s t e p was removed. T h i s i n i t i a t e d a t r a n s i e n t response which could be cornpared t o the response with the control- l e r inactive. obtained i n Step 3 above.

- A triangular pulse was input t o the actuator t o measure the t r a n s i e n t response.

- When desired, the frequency response was measured by inputing a sinasoidal signal t o the a c t u a t o r .

Two types of feedback were tested: ( i ) i n t e g r a l feedback, where the r a t e of cyclic input from the c o n t r o l l e r was proportional t o the flapping, and (2) Lagged position feedback. Lateral b l feedback was a l s o investigated but because of the l i a i t e d bl feedback d a t a , the r e s u l t s were not conclusive.

Figure VI-3 shows the normalized flapping and wing beam tending response t o a s t e p input f o r various values 9 f integraL gain a t a simulated 185 knots, Note t h a t f o r a l l gains the a1 flapping 's completely eliminated. but the magnitude of the t r a n s i e n t reduction and the t l ~ e t o wash out the flapping a r e functions of the gain. Also evident is the apparent reduction i n d a ~ p i n g o f t h e b l a d e f l z p p i n g a o d e as t h c g a i n i s i n c r e a s e d . E x t r a p ( . - L a t i o n o f t h e d ~ ~ p i n g w i t h g a i ? r e l a t 2 o n s h i p s i n d i c a t e s n e u t r n l s t a b i l i t y a t a g a i n o f a p p r o x i m a t e l y ? l t o 12 degree..- p e r s e c o n d p e r d e g r e e .

A l s o e v i d e n t i n F i g u r c V I - 3 i s a n ir?crease i n t h e w i n g b e a a w i s e ~ o t i o n w i t h i n c r e a s i n g g a i n . I n t h e s i a p l e c o n t r o l l e r t e s t e d ( w h i c h c o u l d be d u p l i c a t e d u s i n g a h v d r o n e c h a n i c a l s v s t e a i n a n o p e r a t i o n a l a i r c r a f t ) , t h e r e were n o c o T p e n s a t i o n n e t w o r k s t o ~ i t i g z t e i n p u t s a t t h e test s t a n d n a t u r a l f r e q u e n c i e s . A t t h e w i n g b e a s f r e q u e n c y t h e i n p u t was p h a s e d s o ay t o i n c r e a s e f l a p p i n g a n d herice t h e w f n g r e s p o c s e . A n o t c h f i l t e r c o u l d be e a p l o y e d i n a n o p e r a t i o n a l c o n t r o l l e r t o e l i a i n a t e t h e i n c r e a s e d r e s p o n s e o r n e t w o r k s a d d e d w h i c h would e v e n r e d u c e t h e u n a u p e ~ t e d r e s p o n s e .

Data similar t o those s h o w n in Figure V I - 3 were also takenat a s i a u l a t e d 2 6 5 k n o t s . The s y s t e m became u n s t a b l e a t a p i n o f a p p r o x i a a t e l y f i v e .

F i g u r e V I - 4 s u m m a r i z e s t h e s y s t e m s t a b i l i t y b o u n d a r i e s . t h e r e d u c t i o n i n t r a n s i e n t f l a p p i n g . a n d t h e i n c r e a s e i n w i n g beam r e s p o n s e as a f u n c t i o n o f g a i n . T h e recommended g a i n l e v e l shown i n F i g u r e V I - 4 ( a ) is b a s e d o n t h e g e n e r a l l y a c c e p t e d g a i n m a r g i n o f 6 - 8 . N o t e t h a t t h e g a i n w i l l h a v e t o be v a r i e d w i t h a i r s p e e d o r t h e v a l u e a t t h e l i r , , i t d i v e s p e e d e m p l o y e d .

N o t o w o r t h y is t h e e v i d e n c e t h a t f o r a r e l a t i v e l y l o w - g a i ~ I n t e g r a l c o n t r o l l e r . a L a r g e r e d u c t i o n i n t r a n s i e n t f l a p p i n g c a n be a c h i e v e d . F o r e x a m p l e . f o r a g a i n o f ~ r n i t y , t h e t r a n s i e n t f l a p - p i n g is r e d u c e d b y a p p r o x i m a t e l y 30 p e r c e n t . T h i s g a i n c o n t r o l l e r h a s a n e g l i g i b l e e f f e c t o n s t a b i l i t y a n d c a u s e s o n l y a s a a l l i n - crease i n wing-beam b e n d i n g r e s p o n s e .

The m e a s u r e d s y s t e m r e s p o n s e t o Lagged p o s i t i o n g a i n i s shown i n F i g u r e VI-5. L a g g e d p o s i t i o n g a i n w a s somewhat T o r e e f f e c t i v e t h a n i n t e g r a l g a i n i n r e d u c i n g t r a n s i e n t f l a p p i n g , b u t p r o d u c e d more w i n g r e s p o n s e . A r e d u c t i o n i n s t a b i l i t y w i t h g a i n w a s a l s o e v i d e n t .

C o g b i n a t i 9 n s o f i n t e g r a l g a i n a n d l a g g e d p o s i t i o n g a i n w e r e t e s t e d t o d e t e r m i n e a n o p t i m u 3 c o n t r o l l e r c o c f i g u r a t i o n . F i g u r e VI-6 s h a w s t h e n e a s u r e d f l a p p i n g r e s p o n s e a t s i m u l a t e d a i r s p e e d s of 1 8 5 k n o t s a n d 265 k n o t s f o r t w o c o m b i n a t i o n s o f g a i n s . I n t h i s c a s e a s i g n i f i c a n t r e d u c t i o n i n f l a p p i n g is e v i d e n t , b u t a t t h e e x p e n s e o f a r e d u c t i o n i n s y s t e m d a m p i n g .

The r e s p o n s e o f t h e o p t i a u ~ c o n f i g u r a t i o n t o s t e p a n d t r i a n g u l a r s h a p e d p u l s e s i s shown i n F i g u r e s \'I-7 a n d -8. r e s p e c t i v e l y . A s i g n i f i c a n t r e d u : t i o n i n t r a n s i e n t f l a p p i n g is a c h i e v e d w i t h n e g l i g i b l e r e d u c t j . 0 ~ i n s y s t e q d a m p i n g . The i n c r e a s e i n w i n g beam r e s p o n s e i s s:>all w i t h t h i s c o n f i g u r a t i o n .

C o r r e l a t i o n o f T h e o r y w i t h M e a s u r e d S t a b i l i t ; ~ a n d P e r f o r n a n c e T h e BHC f l a p p i n g c o n t r o l l e r t h e o r y is b a s e d o n t h e l i n e a r , small p e r t u r b a t i o n d y n a m i c s t a b i l i t y a n a l y s i s d e s c r i b e d i n S e c t i o n V.

T h e method i n v o l v e s g e n e r a t i n g t r a n s f e r f u n c t i o n s for- L o n g i t u - d i n a l a n d l a t e r a l f l a p p i n g as a f u n c t i o n o f t h e s:.?ashplatc i n p u t .

a n d c o u p l i n g them w i t h e q u a t i o n s r e p r e s e n t i n g t h e c o n t r o l l 2 r .

T h e s y s t e a s t a b i l i t y a n d f r e q u e n c y r e s p o n s e is t h e n c a l c u l a t e d u s i n g t h e c o u p l e d e q u a t i o n s .

C o r r e l a t i o n b e t w e e n t h e o r y a n d m e a s u r e d c h a r a c t e r i s t i c s i s g o o d .

T h i s is b e s t i n d i c a t e d b y c o m p a r i s o n b e t w e e n t h e m e a s u r e d a n d p r e d i c t e d L o n g i t u d i n a l f l a p p i n g f r e q u e n c y r e s p o n s e . shown i n F i g c r e V I - 9 . Note t h a t rhe g e n e r a l s h a p e of t h e a a p l i t u d e a n d p h a s e c o n p a r e s v e r y w e l l w i t h t h e p r e d i c t e d r e s p o n s e . However, t h e L o w - f r e q u e n c y r e s p o n s e is h i g h e r t h a n p r e d i c t e d . ( T h e p r e d i c t e d w i n g r e s o n a n t f r e q c e n c y i s h i g h e r t h a n t e s t e d b e c a u s e t h e r o o t f l e x i b i l i t y i n t r o d u c e d b y t h e w i n d - r u n n e l mount w a s n o t known when t h e p r e d i c t i o n s were made.)

A c o m p a r i s o n o f t h e p r e d i c t e d a n d m e a s u r e d s t a b i l i t y b o u n d a r i e s w i t h i n t e g r a l g a i n is shown i n F i g u r e V I - 4 . The r e l a t i v e l y small d i f f e r e n c e is p r o b a b l y d u e t o d i f f e t e n c e s b e t w e e n t h e t e s t s t a n d c o n f i g u r a t i o n a n a l y z e d a n d t h e o n e t e s t e d : n a m e l y t h e b e a a w i s e f r e q u e n c y a n d t h e h u b r e s t r a i n t .

The m e a s u r e d s t a b i l i t y b o u n d a r y f o r p o s i t i o n g a i n is somewhat h i g h e r t h a n t h e p r e d i c t e d v a l u e . A t a s i m u l a t e d 1 8 5 k n o t s . t h e n e u t r a l s t a b i l i t y g a i n w i t h a n 9 . 1 8 - s e c o n d l a g is 3 . 9 c o m p a r e d t o a p r e d i c t e d 2 . 2 . T h i s e r r o r i s n o t r e a d i l y e x p l a i n e d by t h e d a t a a n d was p o s s i b l y d u e t o a n e r r o r i n t h e s y s t e m c a l i b r a t i o n .

o r t h e t h e l a g may h a v e b e e n s l i g h t l y l o n g e r t h a n 0 . 1 8 a n d e - g - , g d i n may h a v e b e e n s l i g h t l y Lower t h a n it was t h o u g h t t o b e .

ANGLE OF ATTACK - DEGREES

Proprotor Longitudinal and Lateral Flapping Figure VI-1.

versus Angle of Attack, Simulated 265 Knots, 458 rpm.

as/aa, - THEORY

Correlation of Flapping Theory with Figure VI-2.

Measured Total Fla2ping Derivativa.

- - - - - - -. - . . - - - . . . - - . - - - - - . -- - - - --- I-~- SIMUL.ATED GUST I N P U T

-, ----- -.--- A

ll

0 L 2 3 4 5 I

ELAPSbD T I M E - SECONDS

F L A P P I N G RESPONSE (a1) OSCILLOGRAPH RECORDS O F NORMALIZED WING BEAM BENDING RESPONSE M e a s u r e d R e s p o n s e to Step Input for V a r i o u s V a l u e s Figure V I - 3 .

of Integral Gain a t a S i m u l a t e d 1 8 5 K n o t s .

TEST P O I N T S STABILITY \ RECOMMENDED MAXIMUM GAIN

SIMULATED AIRSPEED - KNOTS

(a) S T A B I L I T Y LFAIRED LINE RESPONSE 4 5 I 1 f INTEGRAL G A I N 185 KNOTS -0 (b) RESPONSE F i g u r e V I - 4 . Summary of S t a b i l i t y and Response C h a r a c t e r i s t i c s with I n t e g r a l Gain Feedback.

a BELL HELICOPTER cou-nr

SIMULATED GUST INPUT

I----

LAG TIME CONSTANT = 0.18 SEC LAG POSITION GAIN = K al N3RMALI ZED FLAPPINS K a l = 1 . 0 e RESPONSE 1 2 3 4

ELAPSED TIME - SECONDS

NORMALIZED WING REAM BENDING RESPONSE Figure VI-5. Gust Response for Various Values of Lagged Position Gain at 9 2 Knots.

1-09 CONTROLLER OFF LqG TIME CONSTANT = 0.18 SZC 1 PcCITION GAIN = NORMALIZZD INTEGRAL GAIN = KIzl Kal = = 0.84 al FLAPPING CONTROLLER ON (185 KNOTS SIMULATED) 0 . 0 1 2 3

ELAPSED TIME - SECONDS

CONTROLLER OFF l.o\ INTEGRAL GAIN = I : , = 2.1 L NORMALIZi'D a1 FLAPPING CONTROLJJPR ON (1 85 KNOTS SIMULATED) COhTliOLLER CN (265 KNOTS SIMULATED; 0 - F i g u r e V I - 6 . M e a s u r e d F l a p p i n g R e s p o n s e a t S i m u l a t e d 185 K n o t s a n d 265 Kqots, for Combined Integral a n d L a g g e d P c s i t i o n & i n s ( S t e p I n p u t ) .

s w u m

i

GUST INPUT i ONTROLLER OFF

1 . 0 -

LAGGED POSITION GAIN = = 0 . 5 JAG TIME CONSTANT = 0.18 = 1 . 1 NORHALIXED FLAPPING RESPONSE I 1 0 4 .

0 1 2 3

ELAPSED TIHE - SECONDS

KOHTROLLER OFF NORI(ALIZEI3 WING BEllll BENDING RESPONSE I n t e g r a l F i g u r e V I - 7 . R e s p o n s e t o S t e p G u s t for Combined and Lagged T o l i t ion Gains (240 Knots, S i m u l a t e d ) .

300-099-004 VI-12 TRI-ANGLE INPUT SIMULATED GUST INPUT i 2 3

ELAPSED TIME - SECONDS

CONTROLLER OFF

r

IAp%ED P o S I T I O N GAIN = Kal FLAPPING RESPONSE LAG TIHE CCINSTAXT = 0-18 1-RU GAIN = K : , = 1 . 1

-

COrUTROUER ON ONTROUER ON CONTROLLER OFF NORmLXZED WING BEAM BENDING RESPONSE F i g u r e VI-8. Response t o T r i c n g u l a r P u l s e Gust for Combined I n t e g r a l and Lagged P c s i t i o n G a i n s (240 K n o t s , S i m u l a t e d ) .

@ eELL HEUCOPTER vv

V I I . BLADE A Y D COlTWL Sk'5TVI LOADS Blade 2nd c o n t r o l l o a d s were r e c o r d e d t h r o u g h o u t t h e p c r f ork7ance tests, a n d d u r i q g the a i r p l a n e a e d e p o r t i o n s o f t h e G y n a n i c tests. S e v e r a l additions; r u n s were ~ a d e f o r t h e p r i a a r y p u r p o s e o f c t t a i n i n p l o a d s d a t a .

D u r i n g t h e po-red tests. t h e a a j o r i t y of t h e d a t a were o b t a i n e d w i t h t h e p r o p r o t o r L o n g i t u d i . ~ a ? f l a p p i n g ' a l ) t r i ~ m e d t o z e r o b y t h e a p p l i c a t i o n of c y c l i c p i t c h t o s i a p i i f y t h e t e s t p r o c e d u r e .

E x t r a p o l a t i n g t h e d a t a t o s i ~ u l a t e c o n v e r s i o n a t a c o n s t a n t f u s e l a g e a c g l e - o f -at t a c k r e s u l t s i n a w i d e . i n s e n s i t i v e c o n v e r - s i o n c o r r i d o r - F i g u r e V I I - L show? t h a t c o n v e r s i o n c o r r i d o r b a s e d on t h e a e a s u r e d l o a d s .

P r i o r t o t h e test. c a l c ~ l a t i o n s o f p r e d i c t e d bLade f o a d s e r r o n e - o u s l y i n c o r p o r a t e d a s w a s h p L a t c a r r a n g e a e n t such that f o r c y c l i c i n p u t s t h e a a x i a u a b l a d e a n g l e v a r i a t i o n o c c u r r e d a t a n a z i ~ u t h of 90 d e g r e e s - The a c t u a l a r r a n g e v e n t of the ~ o n o c y c l i c s w a s h - p l a t e was s u c h t h a t the aaxiaza b l z d e a n g l e v a r i a t i o n o c c n r r e d a t a n a z r a u t h o f 75 d e g r e e s . The Loads ~ e a s u r e d d u r i n g t h e test =re i n r e a s o n a b l e a g r e e n e n t w i t h t-he p r e d i c t e d l o a t ' s f o r h e l f - c o p t e r a n d a i r p l a n e ~ o d e s . b u t were h i g h e r tlza? t h e predicted ',zz3s fnr c o n v e r s i o n n o d e . When t h e p r e d i c t i o n c a l c u l a t i n n s were c o n d u c t e d t o r e f l e c t t h e a c t u a : s w a s h p l a t e a r r p n p e a c n t . p r e d i c t e d l o a d s f o r t h e c o n v e r s i o n a o d z c o r r e l a t e d w i t h t h e q e a s u r e d f o a d s , B - MEASURED BLADE ASD COSTROL LOADS F i g u r e V I I - 2 shows t5e w a v e f o n s o f t h e b l a d e a n d c o n t r o L Loads f o r a r a n g e o f Tast a n g l e s c o r r e s p o n d i n g t o h e l i c o p t e r . c o n v e r - s i o n , a n d a i r p l a n e a o d e s . S e v e r a l c h z r a c t e r i s t i c s are e v i d e n t : ( 1 ) O n e - p e r - r e v l o a d s a r e d o n i n a n t . a n o r n a l c h a r a c t e r i s t i c o f t h e s e m i r i g i d r o t o r . a n d peak a t a n a z i n u t h o f 279 d e g r e e s b e c a u s e o f t h e d r a g I m d o f t h e r e t r e a t i n g b l a d e i n h e l i c o p t e r mode a n d t h e g r a v i t y - i n d u c e d Load as t h e p y l o n i s c o w e r t e d .

( 2 ) H i g n e r h a r m o n i c l o a d s a r e low. c o n f i r m i n g tihat t k , e Mode? 3C)Q p r o p r o t o r is f r e e o f r e s o r l a n c e p r o b l e a s . ( 3 ) The p i t c h - l i n k trace is f r e e o f s t a l l - f l u t t e r c h a r a c t e r i s t i c s . c o n f i r a i n g f r c e - dorn £ram s t a l l - f l u t t e r p r o b l e a s . The b l a d e Loads s a o o t h o u t p r o g r e s s i v e l y as the p y l o n is c o n v e r t e d f r o a h e l i c o p t e r t o a i r - p l a n e mode, r e f l e c t i n g t h e r e d u c t i o n i n skewed f l o w . I n a i r p l a n e n o d e . t h e o n l y o s c i l l a t o r y f o r c e i s frorr! g r a v i t y a t o n e p e r r e v .

C o n s e q u e n t l y , t h e a i r p l a n e mode o s c i i l a t o r y loads are e x t r e n e l y low.

Loads i n ~ e l i c o p t e r Mode, a NAsT = + 7 5 Degrees 1.

- - --- Figure VII-3 shows a t y p i c a l d i s t r i b u t i o n of spanwise o s c i l l a - t o r y b e ~ d i n g moments f o r a r e l a t i v e l y tiigh t h r u s t , hign wind- speed condition. For comparison, t h e c a l c u l a t e d endurince l i m i t is a l s o shown, When t h e r a t i o o f t h e measured Load t o t h e 0s- c i l l a t o r y l o a d i s considered, t h e s t a t i o n 52.5 beamwise bending aoment and t h e s p i n d l e chordwise bending moment a r e t h e most c r i t i c a l . ( T h i s is a l s o t r u e f o r conversion and a i r p l a n e mode.)

Consequently, t h e l o a d s d a t a presented i n t h i s r e p o r t p e r t a i n t o t h e s t a t i o n 52.5 beamwise and t h e s p i n d l e c h o d i s e l o a d s (and t h e p i t c h f i n k load, which i n d i c a t e s blade t o r s i o n a l l o a d s and c o n t r o l system loads), The v a r i a t i o n i n o s c i l l a t o r y l o a d s v e r s u s t h r u s t f o r t h r e e wind speeds i s shown i n F i g u r e s VII-4 through VII-6 f o r s t a t i o n 52 - 5 beamwise bending moment, t h e s p i n d l e ctlordwise bending m o m e n t , and t h e p i t c h l i n k loads. Figure 711-7 shows t h e v a r i a t i o n i n blade and c o n t r o l l o a d s w i t h a i r s p e e d f o r a c o n s t a n t thrust.

2. Conversion Mode, o llAm = +60 Degrees. +30 Degrees, and +15 Deprees The design e i p speed i n conversion mode is 700 f p s , but most of t h e test d a t a f o r conversion mode were taken a t 740 f p s s i n c e a test s t a n d frequency w a s i n resonance w i t h two per rev a t t h e T e s t r p m corresponding t o 700 fps !see S e c t i o n s 111 and VIII).

s t a n d o s c i l l a t o r y load c o n s i d e r a t i o a s d i c t a t e d l i m i t e d u s e of t h e r p corresponding t o a t i p speed of 700 f p s . Camparison of t h e small amount of blade load d a t a taken a t 700 f p s with t h o s e taken a t 740 i p s r e v e r l s a n e g l i g i b l e d i f f e r e n c e i n magnitude and

t r e n d -

Figures VII-8 through VII-10 show t h e measured blade and yoke berding moments and p i t c h l i n k loads v e r s u s t h r u s t f o r s e v e r a l conversion angles. The most n o t i c e a b l e t r e n d w i t h conversion a n g l e i s t h e steady reduction i n s e n s i t i v i t y t o t h r u s t as t h e s h a f t s n g l e of a t t a c k is reduced.

3. Airplane Mode, a MAST = 0 Degrees -- - - The a i r p l a n e mode d a t a were taken with zero c y c l i c p i t c h , i n con- trast t o ' . e l i c o p t e r and conversion mode where l o n g i t u d i n a l c y c l i c was used t o z e r o t h e 1or.gitudinal f l a p p i n g .

The var:atioa i n b l a d e and c o n t r o l oscLlLatory l o a d s w i t h a i r - Tkis w a s speed a n d t h r c s t was small, a s shown i n Figure VII-11.

expected, because t h e o s c i l l a t i n g e x c i t a t i o n 5n t h e a x i a l flow condition is p r i m a r i l y due t o g r a v i t y .

The v a r i a t i o n w i t t : mast angle of a t t a c k is shown i n Figure VII-12 This same v a r i a t i o n f o r 500 r p and a wind speed of 185 knots.

a t t h e i d e n t i c a l test c o n d i t i o n s was determined during t h e dynamic test w i t h t h e p r o p r o t o r mounted a t t h e wivg t i p . P comparison of t h e d a t a was made t o d e t e r m i n e t h e i n f l u e n c e of wing-proprotor aerodynamic i n t e r f e r e n c e on b l a d e l o a d s . n , a t i n f l u e n c e w a s found t o be s m a l l .

The s t e a d y p i t c h l i n k l o a d s arc h i g h e r i n a i r p l a n e mode than t h e y are i n h e l i c o p t e r and conversion modes because t h e c o l l e c t i v e p i t c h is h i g h e r . F i g u r e VII-13 shows t h e v a r i a t i o n i n pitch-1ik.k s t e a d y l o a d s w i t h a i r s p e e d i n a i r p l a n e mode. The i n f l u e n c e of t h r u s t on t h e s t e a d y l o a d is s m a l l . R e p r e s e n t a t i v e p i t c h l i n k s t e a d y Loads f o r h e l i c o p t e r and conversion mode are a l s o shown f o r comparison.

CORREIATION O F THEORY WITH MEASWED OSCILLATORY IDADS T h e o r e t i c a l b l a d e l o a d s w e r e c a l c u l a t e d by means of a h y b r i d computer v e r s i o n of t h e BHC P r o p r o t o r A e r o e l a s t i c A n a l y s i s , Program DYNS, d e s c r i b e d i n S e c t i o n V. The hybrid v e r s i o n h a s been developed s p e c i f i c a l l y f o r t h e purpose of computing b i a d e Loads f o r wind t u n n e l test c o n d i t i o n s . While t h e r e w e r e some d i f f e r e n c e s between t h e l o a d s p r e d i c t e d b e f o r e t h e test and t h e measured Loads, t h e s e have been t r a c e d t o t h e swashplate phasing e r r o r d i s c u s s e d i n Subsection VI1.A.

With t h e i n p u t e r r o r s c o r r e c t e d , t h e c o r r e l a t i o n is e x c e l l e n t .

(The t h e o r y a p p e a r s t o be s l i g h t l y c o n s e r v a t i v e . ) The p r e d i c t e d o s c i l l a t o r y Loads are p l o t t e d i n F i g u r e s VII-3 through VII-13 w i t h t h e measured d a t a shown f o r comparison.

N o a t t e m p t h a s been made t o p r e d i c t t h e o s c i l l a t o r y p i t c h l i n k Loads, hence c o r r e l a t i v e d a t a are n o t shown. The d e s i g n o s c i l - l a t o r y l o a d s w e r e e s t a b l i s h e d by means of an e m p i r i c a l method which BHC u s e s f o r s e m i r i g i d r o t o r s . The s t e a d y p i t c h i n g moment o f t h e b l a d e h a s been c a l c u l a t e d f o r a i r p l a n e mode and i s com- The c o r - pared with t h e measured s t e a d y l o a d i n F i g u r e VII-13.

r e l a t i o n is e x c e l l e n t .

@ BELL HEUC'X>PTER CORA-NV

@ SPINDLE BEAM @ BLADE BEAM 2 2 . 8 @ BLADE BEAM 52.5 @ BLADE BEAU 7 5 . 0 @ SPINDLE CHORD @ BLADE CHORD 52.5 @ BLADE CHORD 75.0 @ BLADE CHORD 1 1 2 @ PITCH LINK @ MAST TORQUE Figure V I I - 2 . B l a d e a n d C o n t r o l Loads 2;; FOUX)UT FRAME

I

$ = 90° is the azimuth of a = 30" amST = 1 5 .

,ol Loads Waveforms in Helicopter, conversion. and Airplane Modes.

VII-s h CXMAsT = 7 5 DEGREES RPM = 5 6 5

-

alm = 0 DEGREES VT = 1 2 0 KNOTS THRUST = 5 5 6 5 POUNDS

--- -

MEASURED

0 BEAMWISE

-

A CHORDWISE

THEOR'I' BEAMWISE

-

A CHORDWISE

ENDURANCE LIMITS .

-

-

CHORDWISE \ i

---

B M I S E

*\\ (INDEPENDENT O F LOAD I N

-

, THE OTHER DIRECTION)

\

I I

\

\ \

' 1

I \

\ Q

1 , \ \ A

SPANWISE STATION - r / R

Figure V I I - 3 . S p a n v i s e O s c i l l a t o r y B e n d i n g Moment D i s t r i b u t i o n i.n H e l i c o p t e r Mode.

I +

I I I

I

-

- 75 DEGREES -

a - ~ ~ ~ RPM = 565 aim = 0 DEGldES

-

MEASURED

0 80 KNOTS

0 120 KNOTS -

A 140 KNOTS

THEORY t r i 140 KNOTS I

DESIGN 1

-ENDURANCE LIMIT I

--

I I 1

THRUST - LB x 1 0 '

Figure VII-4. Blade Station 52.5 Beam Oscillatory Bending Moment versus Thrust in Helicopter Mode.

2 3 4 5 6 7

THRUST - LB x 10'~

F i g w e V I I - 5 .

Spindle Chord Oscillatory Bending Moment Versus Thrust i n Helicopter Mo6e.

(eb BELL HEUCOPTER CCS~A-NV

THRUST - LB x 10'~

Figure VII-6. Blade P i tch-Link OsziJ-latory Load Versus Thrust i n Helicopter Mode.

I u

THKiST - 4590 POUNDS

6 0 0 BLADE BEAM STA 5 2 . 5

0 SPINDLE CHORD

A PITCH LINI; BLADE BEAM STA 5 2 . 5 2 0 2 . 0 0 8 0 100 120 140

SIMULATED TRUE AIRSPEED - KNOTS

Figur :e VII-7. oscillator^ roads Versus Airspeed, Helicopter Mode.

Blade S t a t i o n 5 2 . 5 Oscillatory ~eamwise Figure VII-8.

Bending Moment Versus Thrust i n Conversion Mode.

Figure V I I - 9 . Yoke Spindle Chord O s c i l l a t o r y Bending Moment - Conversion Mode.

0 1 2 3 4 5

PROPROTOR THRUST - LB x l o o 3

Pitch-Link O s c i l l a t o r y Load Versus Figure VII-10.

Thrust i n Conversion Mode.

0 50C 1000 I5 CO 2COO

PROPROTOR THRUST - POUNDS

Glade S t a t i o n 5 2 . 5 Eem Bending Moment Figure VII-11.

c.cd Pitch Link Loads versus Proprotor T k r s s t i r , Airplane Mode.

0 2 4 6 8 10 12 14 16

MAST ANGLE OF A T T A C K - DEGREES

Figure VII-12. Blade and Pitch Link Loads Versus Angle of Attack i n Airplane Mode.

700. 4 #' 30 0 POWERED PROPROTOR TEST 2 0 0

0 a m S T = 0 DEG, RPM = 4 5 8

a m ~ T = 30 DEG, RPM = 565

0 aMAST = 6 0 DEG, RPM = 565

A a , m S T = 75 DEG, RP?II = 565

0 STANDARD STIFFNESS TEST

100 1 5 0 2 0 0 2 5 0 30 0

AIRSPEED - KNOTS

Pitch Link Steady L o a d V e r s u s igure V I I - 1 3 .

Airspeed i n Airplane Mode, V I l I . SOISE A N D VIBRATION A , NOISE During powered t e s t i n g a microphone 68 f e e t upstreaa of the node1 and 5 f e e t above the f l o o r on the c e n t e r l i s r of the tunr,el monitored the noise Le-:eL. !This Locatiov i c the same a s t h a t f o r a past program which neasured the r:oise of conven- tionaL rotors. See References 2 9 and 39 t The output of the microphone was recorded f o r a a s t t i l t angles fron zero t o 7 5 degrees, f o r several t i p speeds, and a t various t h r u s t and power s e t t i n g s .

Figure VIII-1 shows the coaparison between proprotor and con- ventional rotor noise. For tunnel v e l o c i t i e s above about 80 knots, the noise of the proprotor i n takeoff mode is Lower than t h a t of a square-tipped conventional rotor. Also, the r a t e of increase i n proprotor noise. as tunnel v e l o c i t y increases. i s somewhat l e s s than f o r conventional r o t o r > . The proprotor's noise i n cruise mode, a t tilt angles between zero and 30 degrees, i s a t l e a s t 7 decibels Lower than t h a t f o r the takeoff =ode.

(Predictions show a reduction on the order of 15 decibels.)

The proprotor sounds more l t k e a propeller th&n a rotor. lacking the Loud blade s l a p which is c h a r a c t e r i s t i c of conventional r o t o r s a t high speeds. I n cruise mode, the proprotor i s reason- ably quiet--quiet enough, i n f a c t , t h a t the sound of the w i n d tunnel masks it a l n o s t coaple t e l y .

The noise measurements shown i n Figure V I I I - 1 can be extrapolated t o give estimated noise l e v e l s :or the Model 300. Figure V I I I - 2 shows the r e s u l t s of such an extrapolation, and shows the noise c h a r a c t e r i s t i c s of a v a r i z t y of other a i r and surface vehicles for p::rposes of coaparison. A s the figure shows, the noise of the Model 300 during takeoff w i l l be about the saae a s '-hat of a medium h e l i c o p t e r , whereas i t s noise i n cruise f l i g h t w-11 be about the same as t h a t of a Light helicopter. The observer on a busy s t r e e t corner would be unable t o hear the kodel 300 pass- ing over i n crxise f l i g h t a t an a l t i t u d e of 1030 f e e t .

Vibration l e v e l s were measured a t two s t a t i o n s on the dynamic t e s t stand pylon: t1.e intersection of the c e n t e r l i n e s of the conversion spindle and a a s t (Pylon Station Q), and a t Pylon Station 36. A t S t a t i c n 9 the vibration was measured along the s h a f t a x i s , i n the wing beam sense, and i n the pylon yaw sense.

A t Station 36, the vibration was measr~red i n the wing beam and pylon yaw senses.

The doainant vibration was a t the blade passage frequency ( t h r e e per r e v ) and was due t o aerodynamic interference between the wing and the proprotor. I t s amplitude increased with both airspeed and mast a n g l e of a t t a c k . The amplitude was e x t r e m e l y s e . l s i t i v c t o rpxi i n t h e c a s e of t h e d e s i g n - s t i f f n e s s t e s t s t a n d s i n c e t h e l a t e r a l be:.ding n a t u r a l frequency of t h e rnast was i n resonance w i t h t h r e e - p e r - r e v e x c i t a t i o n s a t 470 rpm.

A s noted i n S e c t i o n 111. t h e dynanic t e s t s t a n d s do not respond e x a c t l y a s t h e Mcdel 300 wing-pylon system rlould i n t h e n o r a a l o p e r a t i n g r p n , t h r e e - p e r - r e v frequency rznge because t h e i r s t r u c t u r a l d e t a i l s ( s u c h a s t h e t r a n s n i s s i o n and pylon c a s e , engine i n s t a l l a t i o n , and t h e v a s t ) a r e not dynamically s i r n i l a r .

D e t a i l e d s c a l i n g of t h e s e f e a c u r e s was n o t n e c e s s a r y t o d u p l i c a t e dynamic s t a b i l i t y c h a r a c t e r i s t i c s . N a t u r a l f r e q u e n c i e s of both t e s t s t a n d s were c l o s e t o resonance with t h r e e per r e v , producing r e l a t i v e l y h i g 5 t h r e e - p e r - r e v v i b r a t i o n s . Even s o . t h e t e s t o p e r a t i o n w a s not s i g n i f i c a n t l y r e s t r i c t e d .

Five a c c e l e r o m e t e r s were i n s t a l l e d on t h e powered t e s t s t a n d , p r i m a r i l y f o r t h e purpose of a o n i t o r i n g t e s t s t a n d o s c i l l a t o r y Loads. The torquemeter b o l t r i n g w a s e s t i m a t e d t o be t h e c r i t i c a l aember o f t h e t e s t s t a n d , and t h e r e f o r e a c c e l e r a t i o n l i m i t s were e s t a b l i s h e d which would preclude exceeding i t s endurance l i m i t . During t h e conversion and h e l i c o p t e r mode tests, o p e r a t i o n of t h e p r o p r o t o r a t 535 rpm (nR = 709 f t 'set) w a s avoided because t h e two-per-rev v i b r a t i o n l e v e l s exceeded t h e e s t a b l i s h e d L i m i t . (These r e s u l t e d from a t e s t s t a n d aode being i n resonance a t 535 rpm, with t h e two-per-rev torque generated by t h e gimbal when t h e r o t o r f l a p p e d . ) V i b r a t i o n a t o t h e r p r o p r o t o r speeds was n o t a probl.za.

1. V i b r a t i o n C h a r a c t e r i s t i c s of t h e T e s t Stands During Nornal P r o p r o t o r Operation Figure V I I I - 3 shows how t h e t h r e e - p e r - r e v v i b r a t i o n of t h e d e s i g n - s t i f f n e s s dynamic t e s t s t a n d v a r i e d with r p v . The i n s t a l l a t i o n of t h e pylon yaw l i n k simulated f u l l y converted a i r p l a n e n ~ d e f l i g h t (pylon o n t h e doc~nstop) The l a r g e a a p l i - t u d e i n t h e 450-470 rpa range i s due t o resonance of t h e q a s t L a t e r a l bending mode with t h r e e p e r r e v Figure V I I I - 4 shows the v a r i a t i o n i n a n p l i t u d e with a i r s p e e d a t c o n s t a n t 458 rpm.

There a r e no d a t a i n t h e 170-190-knot range because of t h e mast l a t e r a l bending aode resonance a t 458 r p a i n t h i s speed range.

Without t h e yaw L i n k t h e amplitude of t h e t h r e e p e r r e v was much lower ( a s shown i n Figure V I I I - 5 ) . I n t h i s c a s e t h e i n t r o d u c t i o n of t h e pylon yaw mode ( 1 4 . 1 c p s , o r 1.85 per r e v ) f o r c e s t h e rnast l a t e r a l bending mode t o over 53 cps ( g r e a t e r t h a n e i g h t per rev a t 458 rpm). For t h a t c o n d i t i o n t h e r e a r e no modes n e a r t h r e e p e r rev ( s e e S e c t i o n 1 1 1 ) .

Figure V I I I - 6 shows how t h e v e r t i c a l (beamwise) t h r e e - p e r - r . ? v v i b r a t i o n a t Pylon S t a t i o n 36 v a r i e d w i t h s i n u l a t e d a i r s p e e d or both t h e o n e - f o u r t h and design s t i f f n e s s t e s t s t a n d s . Note tk.at t h e o n e - f o u r t h - d e s i g n - s t i f f n e s s t e s t s t a n d d a t a a r e not

@ BELL HELICOPTER - - N I

a t constanr rpm because of t h e c o l l e c t i v e p i t c h l i m i t a t i o n s discussed i t ; Section V . A t 390 knots t h e v i b r a t i o n l e v e l (1.2g) is l e s s than 50 percent of the Model 300 design L i m i t ( + 2 . 5 g ) .

Figure V I I I - 7 shows t h e i n f l u e n c e of angle of a t t a c k on the v e r t i c a l v i b r a t i o n a t Pylon S t a t i o n 36. The i n c r e a s e i n a n p l i - tude with angle of a t t a c k is due t o the i n c r e a s i n g wing upwash.

Vibration d a t a from t h e powered t e s t a r e surnvarized i n Figure VIII-8. I t i l l u s t r a t e s the t r e n d i n v i b r a t i o n l e v e l during a conversion a t c o n s t a n t a i r s p e e d and rpm, Since t h e response of t h e powered t e s t stand i a t h e t h r e e - p e r - r e v frequency range i s e s s e n t i a l l y i n v a r i a n t with conversion angle ( s e e S e c t i o n 1111, the reduction i n v i b r a t i o n l e v e l a s the pylon is converted i s due t o reduced e x c i t a t i o n .

2. Vibration P u r i n e S t o p - S t a r t Operation Fxploratory t e s t s t o simulate the s t o p - s t a r t phase of t k e f o l d - i n g proprotor concept %ere made on both the standard and one- q u a r t e r - s t i f f n e s s dynamic t e s t s t a n d s . Of prlmary i n t e r e s t were t h e v i b r a t i o n l e v e l s t h a t would be encountered a s a s t a r t i n g o r stopping o p e r a t i o n passed through the v a r i o u s resonant frequen- c i e s of t h e b l a d e s , wings, and pylons. The 25-foot proprotor of t h e s e t e s t s w a s not designed t o be folded i n f l i g h t , t h e r e f o r e d i d not have a f l a p p i n g lockout o r a h i g h - r a t e c o l l e c t i v e p i t c h arrangement, f e a t u r e s which would improve i t s f e a t h e r i n g charac- t e r i s t i c s . Neverthe l e s s , t h e proproror f e a t h e r e d without d i f f i c u l t y a t a i r s p e e d s up t o 265 knots ( s i m u l a t e d ) and a n g l e s of a t t a c k up t o 6 degrees.

The s t o p - s t a r t t e s t s on t h e d e s i g n - s t i f f n e s s t e s t s t a n d were conducted a t a n a i r s p e e d of 136 k n o t s . The c o l l e c t i v e p i t c h range r e s t r i c t e d the rpm sweeps from the low p i t c h l i m i t (600 rp3) t o t h e h i q h p i t c k l i m i t (250 rpm). S e v e r a l resonances of t h e blade and wing were t r z r i s i t e d d u r i n g t h e s e sweeps without any s i g n i f i c a n t i n c r e a s e i n loads o r v i b r a t i o n . The time t o change from 600 t o 250 rprn was 2.75 seconds; and t o r e t u r n t o 600 r p a , 2.48 seconds.

On t h e o n e - f o u r t h - d e s i g n - s t i f f n e s s s t a n d , s t o p - s t a r t t e s t s were a t simulated speeds of 185 and 265 knots. F u l l s t o p s t o zero rprn were vade i n both c a s e s . A t 185 knots t h e angle of a t t a c k was varieti f r o % 0 degrees t o +6 degrees. The time t o s t o p o r s t a r t was approximately 6 seconds ( e q u i v a l e n t t o 3 seconds i n r e a l t i m e ) . The optim~rn r e a l time required f o r t h e s t o p - s t a r t operatiorl has been shown by model t e s t s t o be 2 t o 3 seronds.

Figure VIII-9 shows a time h i s t o r y of p r o p r o t o r and t e s t stand response during s t o p p i n g a t 185 knots a n d z e r o angle of a t t a c k .

A s the blade passage frequency t r a n s i t s each of the s t a n d n a t u r a l modes, a buildup a p p e a r s . Model t e s t s and t h e o r y ( s e e References 22 and 28) have e s t a b l i s h e d t h a t t h e aerodynamic i n t e r f e r e n c e between wing and r o t o r c a u s e s hub s h e a r s a t t h e blade passage f r e q u e n c y and g e n e r a t e s a s m a l l avount of e x c i t a t i o n a t t w i c e t h e b l a d e p a s s a g e f r e q u e n c y . T h e s h o r t e r t h e t i a e r e q u i r e d f o r s t o p p i n g , t h e smaller t h e a a p l i t u d e of t h e buildup--down t o t h e t i m e f o r mlnimum b u i l d u p , g i v e n a s h t = 2 . 7 5 , r D '38.5 seconds ( R e f e r e n c e 26) o r a b o u t 2 . 2 s e c o n d s i n t h e c a s e of a 2 5 - f o o t - d i a m e t e r p r o p r o t o r .

As F i g u r e V I I I - 1 0 shows, t h e a m p l i t u d e of t h e response i n c r e a s e s w i t h a n g l e o f a t t a c k . The a n g l e of a t t a c k i l l u s t r a t e d ( 4 d e g r e e s ) c o r r e s p o n d s t o a 1.5g maneuver. The a m p l i t u d e a l s o i n c r e a s e s w i t h a i r s p e e d a s shown i n F i g u r e VIII-11.

The l i r n i t e d number o f d a t a p o i n t s keeps t h e r e s u l t s of t h e s t o p - s t a r t t e s t s e s s e n t i a l l y q u a l i t a t i v e . b u t t h e tests d i d show t h a t f e a t h e r i n g t h e b l a d e s and s t o p p i n g t h e r o t o r i s f e a s i b l e . Loads a n d a c c e l e r a t i o n s were w e l l w i t h i n a l l o v a b l e li-its and t h e r e m s no e v i d e n c e of i n s t a b i l i t y . A m o d i f i e d v e r s i o n of t h e 2 5 - f o o t p r o p r o t o r which i n c o r p o r a t e s a f l a p p i n g l o c k o u t and b l a d e f o l d i n g w i l l be t e s t e d i n t h e f a l l of 1971.

3. C o r r e l a t i e n of Theory w i t h Measured V i b r a t i o n Under c o n t r a c t t o t h e U. S. A i r F o r c e F l i g h t Dynamics L a b o r a t o r y , t h e BHC has d e v e l o p e d a h i g h l y r e f i n e d t h e o r y which p r e d i c t s aerodynamic i n t e r f e r e n c e between wing and p r o p r o t o r ( R e f e r e n c e 2 8 ) .

Good c o r r e l a t i - o n wkth measured v i b r a t i o n d a t a from d y n a z i c a l l y s c a l e d models h a s been e s t a b l i s h e d and is zhown i n t h a t r e p o r t .

A s t u d y of t h e c o r r e l a t i o n of BHC v i b r a t i o n t h e o r y w i t h t h e v i b r a t i o n measured d u r i n g t h e dynamic s t a b i l i t y t e s t h a s c o t been made. A l a r g e amount of d i g i t a l computer t i v e would be r e q u i r e d f o r s u c h a s t u d y and because a c c u r a c y of t h e t h e o r y h a s a l r e a d y been i n v e s t i g a t e d , t n e expense w a s c o n s i d e r e d unwar- r a n t e d . F u r t h e r m o r e , t h e d a t a i n d i c a t e t h e v i b r a t i o n problem i - .

e s s e n t i a l l y one of a v o i d i n g resonance w i t h t h r e e per r e v .

CHANGE I N NOISE LEVEL I N 75-150 HERTZ OCTAVG - DECIBELS

(ARBITRARY COMMON REFERENCE LFl'EL)

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50 100 200 500 1000 2000

DISTANCE TO VEHICLE - FEET

Figure V I I I - 2 . Estimated 5 x t e r . 7 a l Noise L e v e l s .

300-099-034 V I I I -6

@ BELL HELICOPTER coup.Nr

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YAW LINK I N

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- A Pylon S t a t i o n 0 Yaw

0 Pylon S t a t i o n 0 Beam 0 Pylon S t a t i o n 3 6 . 0 Beam

-

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PYLON YAW MODE RESONANCE - 6 .

1 . 2 350 40G 450 500 550 600 PROP2OTOR RPM 1-3. Pylon Three-Per-Rev V i b r a t i o n L e v e l Versus Figure VII Proprotor RPM f o r Standard S t i f f n e s s Test Stand (Airplane Mode).

@ eELL HELICOPTER c o u p n l v

TUNNEL VELOCITY - KNOTS

Figure VI 11-4 Pylon T h r e e - P e r - R e v Vibration L e v e l V e r s v ~ s Tunnel Velocity for Standard Xifffness T e s t Stand ( A i r p l a n e Mode) .

I I I I * 1 I MAST = 0 DEGREES

-

RPM = 4 5 8 Y A W LINK OUT

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ACCELEROMETER LOCATIOFIS 0 PYLON AXIAL

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A PYLON STATION 0 Y A W

a PYLON STATION 0 BEAM

0 PYLON STATION 3 6 . 0 BEAM

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F i g u r e VIII-5. Pylon Three-Per-Rev V i b r a t i o n Level Versus T u n n e l V e l o c i t y for Standard S t i f f n e s s Test S t a n d (Airplane Mode - Y a w Link O u t ) .

VIII -9 * L a m ~ ~ = 0 DEGREES

- -

YAW L I N K I N

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0 DESIGN STIFFNESS TEST STAND ( 4 5 8 RPM) 0 1/4 DESIGN STIFFNESS TEST STAND - -

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Three-Per-Rev B e m i s e Vibration at Figure ~111-6, Pylon S t a t i o n 36 Versus Simulated Airspeed (Airplane Mode) .

@ BELL HEUCOPTER COM-"

0 DESIGN STIFFNESS TEST STAND (100 KNOTS, 458 RPM)

A DESIGN STIFFNESS TEST STAND

(185 KNOTS, 500 RPM)

I

h A

ANGLE OF ATTACK - DEGREES

Figure VIII-7, Three-Per-Rev Beamwise Vibration a t Pylon S t a t i o n 36 Versus A n g l ~ of Attack (Airplane i%lde).

d BELL HELICOPTER w r

MAST ANGLE OF ATTACK - DEGREES

Mast Case Three-Per-Rev Vibratior Levels Figure VIXI-8, Versus Mast Azgle of Attack, Helicopter and Conversion Modes.

T IPE-SECONDS STA FEATH 10 3,000 IN-LB/I:l WING BEAM BENDING

- -

STATION 3 8 . 0

2 3 , 200 IN-LB/IN

WIYG TORSION STATION 3 8 . 0 TRANSIENT ' 199,000 IN-LB/IN WISG CHORD WING CHORD BENDING STATION 38.0 CONVERSION LINK AXIAL LOAD YAW LINK AXIAL LOAD FLAPPING " 1 , b FLAPPING 1 , BLADE BEAM BENDING STATION 52.5 YOKE CHORD BENDING AZIMUTH POSITION RED BLADE Time History of eath her-Stop Figure VIII-9.

Test Stand, Zero Degrees Ang Tunnel Speed = 9 2 . 5 Knots.

F O ~ * FRAME

V I I I - 1 3 I I STAM FEATHERING

3/REV BUILD-UP AT WING / )/REV BUILD-UP AT WItlG

TORSION FREQUENCY 7 BEAM FREQUENCY

TRANSIENT RESPONSE AT 3jREV BUILD-L'P AT WING WING CHORD FREQUENCY CHORD FREQUENCY Feather-Stop for Quarter Stiffness o Degrees Angle-of -At-:ack , 92.5 Knots*

I

STAWr FEATHERING WING BEAM BENDING STATION 3 8 . 0 WING TORSION STATION 3 8 . 0 I WING CHORD BENDING STATION 3 8 . 0 CONVERSION LINK AXIAL rnAD YAW LINK AXIAL LOAD a- FLAPPING b FLAPPING l m BLADE BEAM 3ENDING STATION 5 2 . 5 YOKE CHORD BENDING AZIMUTH POSITION RED BLADE F i g u r e V I I I - 1 0 . Time H i s t o r y of F e a t h e r - T e s t Stand, Four Degrees T u n n e 1 Speed = 9 2 .5 Knot I; 6 7 8 3 13 11 12 A A A I L A I A -UP AT WING 3/REV BUILD-UP AT WING UENC Y BEAM FREQUENCY

-

- - - I REV BUILD-UP AT WING CHORD FREQUENCY J JC I * - \ & L I f Feather-Stop for Quarter Stiffness Jr Degrees Angle-of-Attack, 92.5 Knots.

1-14 1 2 3 4 I a a START . - % FEATHERING 1 0 3 , 0 0 0 IN-LB/IN WING BEMI BEtIDING STATION 3 8 . 0 2 3 , 2 0 0 IN-LB/IN WING TORSION STATION 38.0 199,000 IN-LB/IN WING CHORD BENDING STATION 38.0 CONVERSION LINK AXIAL LOAD YAW LINK AXIAL LOAD PYL9N YAW MODE NEAR 3 f R E V RESONANCE PRIOR TO FEATHERI:;; 1 4 . 1 DEG/IN FLAPPING BLADE BEAM BENDING STATION 5 2 . 5 YOKE CHORD BENDING I AZIMUTH P O S I T I O N

i

RED BLADE F i g u r e V I I I - 1 1 . T i m e History o?

T e s t Stand, Zero Degrees T u n n e l Speed = 132 Knots.

4 5 6 7 8 9 10 & 1 A A L A

a BEAM 3/REV FREQUENCY BUILD-UP AT WING

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f F e a t h e r - S t o p for Q u a r t e r S t i f f n e s s ro Degi-ees A n g l e - o f - A t t a c k , 1 3 2 Knots.

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0 B E u H E m c o P E R coM-Nr

I X . CONCLUSIONS A . GENERAL 1. F u l l - s c a l e wind-tunnel t e s t s have v e r i f i e d the proprotor technology developed dirring t h e Army Composite A i r c r a f t Program.

2. Theory and model t e s t s form a r e l i a b l e b a s i s f o r t h e design of f u l l -scale p r o p r o t o r s .

3. The advanced design 25-foot-diameter proprotor w i l l provide good performance i n h e l i c o p t e r f l i g h t , w i l l a t t a i n high propul- s i v e e f f i c i e n c y i n a i r p l a n e f l i g h t , and w i l l have p o s i t i v e dynamic s t a b i l i t y and a c c e p t a b l e t l a d e loads.

4 . These tunnel t e s t s , by e s t a b l i s h i n g t h a t t h e p r e d i c t e d Levels of performance, s t a b i l i t y , and load margins a r e a t t a i n a b l e , pro- vide a f i r m base f o r the next Logical s t e p i n t h e development of p r o p r o t o r technology--a r e s e a r c h a i r c r a f t f l i g h t program.

1. P r o p r o t o r s can provide l i f t e f f i c i e n t l y i n hover ( a t a t y p i c a l o p e r a t i n g c o n d i t i o n , t h r i l s t was measured a t 8 . 5 pounds p e r h o r s e - power and t h e f i g u r e of merit was 0.78).

2 . Measured s t a t i c t h r u s t exceeded t h e maximum p r e d i c t e d t h r u s t by 1 5 percent (maximum t h r u s t was not defined i n the r e s t due t o t e s t stand power l i m i t a t i o n s a t t h e t e s t rpm) .

3. I n the blade element momentum a n a l y s i s . t h e use of t h e cLas- s i c a l h e l i c ~ p t e r t i p l o s s f a c t o r B = 1 - ~ ! q / b , l e a d s t o under p r e d i c t i o n of t h e maximum a t t a i n a b l e s t a t i c t h r u s t . I n a d d i t i o n t o a decreased t i p l o s s , maximum l i f t c o e f f i c i e n t s h i g h e r than measured during the two-dimensional a i r f o i l t e s t s m u s t be used i n t h e c a l c u l a t i o n s t o match t h e h i g h e s t t h r u s t measured.

4 . The proprotor provides a broad range of L i f t and propulsive f o r c e f o r conversion.

5 . For a given l e v e l of L i f t and propulsi17e f o r c e , power can be p r e d i c t e d a c c u r a t e l y . P r o p r o t o r t i p path plane o r i e n t a t i o n pre- d i c t i o n s f o r t h e given f o r c e combinations a r e good except f o r mast a n g l e s g r e a t e r than about 60 degrees a t which point t h e p r e d i c t i o n s begin t o be l e s s a c c u r a t e and over p r e d i c t t h e required angle by 2 t o 3 d e g r e e s .

6. Propulsive e f f i c i e n c i e s were measured i n excess of 90 per- c e n t and were g e n e r a l l y s l i g h t l y h i g h e r than predicted by t h e o r y .

7 . The t e s t r e s u l t s confirmed t h a t t y p i c a l c r u i s e e f f i c i e n c y of approximately 75 percent a r e p o s s i b l e with a proprotor a s p r e - d i c t e d i n t h e Task 1 Design Study.

0 BELL HEUCOPTER c o u p x ~ v

C. DkNAMIC STABILITY 1. ALL proprotor/pylon/wing modes e x h i b i t e d good s t a b i l i t y through a simulated speed of 408 knots.

2 . The c r i t i c a l mode, t h e fundamental wing bending mode, e x h i b i t e d increased damping with a i r s p e e d ( a t c o n s t a n t prop- r o t o r rpm).

3. There was c l o s e agreement between t h e 200-knot simulated speed and t h e 200-knot a c t u a l speed.

4 . The f u l l - s c a l e t e s t r e s u l t s were i n e x c e l l e n t agreement with t h e o r y and one -f i f t h - s c a l e a e r o e l a s t i c model t e s t s .

D BLADE FLAPPING 1. Blade f l a p p i n g produced no r e s t r i c t i o n s or, t h e operation of t h e p r o p r o t o r .

2 . A simulated 3.0g maneuver a t 265 knots produced l e s s than 7 degrees of f l a p p i n g a s compared with t h e 12-degree design allow- a b l e .

3 . Measured f l a p p i n g d e r i v a t i v e s a g r e e with t h e o r y .

4. Evaluation of an e l e c t r o n i c f l a p p i n g c o n t r o l l e r showed s t e a d y f l a p p i n g reduction of 100 percent with reductions of t r a n s i e n t f l a p p i n g of 40 t o 40 percent.

5 . The f l a p p i n g c o n t r o l l e r increased the wir;gts response t o simulated g u s t s 6 . I n c r e a s i n g t h e response r a t e of t h e f l a p p i n g c o n t r o l l e r t o reduce t r a n s i e n t f l a p p i n g reduced t h e dynamic s t a b i l i t y of t h e proprot or/pylon/wing system.

7 . The e f f e c t of t h e f l a p p i n g c o n t r o l l e r on f l a p p i n g and dynamic s t a b i l i t y agreed with t h e o r y .

E , BLADE AND CONTROL SYSTEM LOADS 1. Measured bladz and c o n t r o l system Loads were acceptable i n a l l f l i g h t modes.

2. O s c i l l a t o r y loads were l e s s thar, t h e c a l c u l a t e d f a t i g u e endurance l i m i t i n h e l i c o p t e r and conversion modes even though l i t t l e time i s spent i n those modes and complete conversion can be made i n 11 seconds. O s c i l l s t o r y Loads vere verV low i n a i r - plane mode.

3. O s c i l l a t o r y blade loads were i n good agreement with p r e d i c - t i o n s i n h e l i c o p t e r mode but were i n poor agreement with o r i g i n a l p r e d i c t i o n s a t i n t e r m e d i a t e conversion a n g l e s .

300-099-004 I X - 2 4 . The o r i g i n a l p r e d i c t i o n s were found t o have e r r o r s i n t h e i n p u t d a t a ( r e l a t i n g t o s w a s h p l a t e phasing and hub s p r i n g a o m e n t s ) . When t h e s e e r r o r s were c o r r e c t e d , t h e t h e o r y p r e - d i c t e d o s c i l l a t o r y Loads s l i g h t l y h i g h e r than t h o s e measured.

5 . The t e s t r e s u l t s and t h e o r y show t h a t t h e f l a p p i n g r e s t r a i n t produced by t h e hub s p r i n g i n c r e a s e s t h e o s c i l l a t o r y bl2de l o a d s .

Only a moderate l e v e l of f l a p p i n g r e s t r a i n t , such a s t h a t i n c o r - porated i n t h e system t e s t e d , a p p e a r p r a c t i c a l .

F . N O I S E AND VIBRATION 1. Measured p r o p r o t o r n o i s e i n h e l i c o p t e r mode was comparable t o n o i s e measured from i s o l a t e d h e l i c o p t e r main r o t o r s i n t h e tunnel..

I t can t h e r e f o r e be concluded t h a t t h e n o i s e f r o 9 a p r o p r o t o r a i r - c r a f t w i l l be lower t h a n t h a t from a h e l i c o p t e r o p e r a t i n g a t e q u a l t i p speed and g r o s s w e i g h t , because of t h e absence of a t a i l r o t o r o r t h e o v e r l a p e f f e c t s of a tandem r o t o r h e l i c o p t e r .

2 . I n a i r p l a n e f l i g h t , n o i s e l e v e l s w i l l be e x t r e m e l y low.

P r o p r o t o r t?oise l e v e l s i n p r o p e l l e r o p e r a t i o n were ' >o low t o be d i s t i n g u i s h e d f r o a background t u n n e l n o i s e .

3 . During t h e powered t e s t t h e t e s t s t a n d t n r e e - p e r - r e v v i b r a t i o n Level d e c r e a s e d r a p i d l y a s t h e s t a n d was c o n v e r t e d from h e l i c o p t e r t o a i r p l a n e mast a n g l e of a t t a c k , r e f l e c t i n g t h e r e d u c t i o n i n o s c i l l a t o r y a i r l o a d i n g .

4 . I n a i r p l a n e mode t h e b l a d e passage f r e q u e n c y v i b r a t i o n i s dominant. Aerodynamic i n t e r f e r e n c e between t h e wing and prop- r o t o r is t h e s o u r c e of v i b r a t i o n i n a i r p l a n e q o d e .

5. Crew s t a t i o n and cabin v i b r a t i o n l e v e l s i n a i r p l a n e mode w i l l be v e r y l o w , provided t h a t wing and pylon n a t u r a l f r e q u e n c i e s a r e n o t i n resonance w i t h t h e b l a d e passage f r e q u e n c y .

6. S t a r t - s t o p s s i m u l a t i n g t h i s o p e r a t i o n f o r a s t o p - f o l d prop- r o t o r were performed w i t h o u t d i f f i c u l t y a t s i m u l a t e d speeds up t o 265 k n o t s .

"Advancement of P r o p r o t o r Technology Task I - Design Study Summary, NASA C o n t r a c t o r Report C R 114363, September 1969.

S c h e r r e r , Richard, e t a1 . , "NASA-Lockhead Short-Haul T r a . 1 ~ - p o r t S t u d y , " N A ~ ~ T e p o r t SP-116, A p r i l 1966.

Armstrong, LTC M a r s h ~ l l B . , " T a c t i c a l Uses and F u t u r e Concepts f o r Ti1 t i n g - P r o p r o t o r Low-Disc-Loading VTOL A i r c r a f t i n Marine Corps O p e r a t i o n , I 1 p r e s e n t e d a t AHS/U. 3 . Army Symposium on Ope r a t i o n t l C h a r a c t e r i s t i c s and T a c t i c a l Uses of V e r t i c a l L i f t A i r c r a f t , Noveinber 1967.

tfModeL 266 Composite A i r c r a f t Program: Explor- ~ f i n i - t i o n F i n a l Report," i3ell H e l i c o p t z r Ccmpar- '. : 266-049-217, J u l y 1967.

H a f c e r , R . , "The Domain of t h e C o n v e r t i b l e R o ~ G ~ , " AGARD V/STOL A i r c r a f t , September 1964.

L i c h t e n , R. L., -- e t a l . , "A Surve;- of Low-Disc-Loading VTGL

A i r c r a f t Designs , I t AIAA P a p e r 67-756, November 1965.

L i c h t e n , Robert L. , "An A n a l y s i s of L o w - ~ i s c - L o a d i n g VTOL A i r c r a f t Types," p r e s e n t e d a t AGARD, P a r i s , J a n u a r y 1966.

F i s c h e r , J . N i l e , -- e t a l . , "VTOL v e r s u s A l t e r n a t f v e I n t r a -

T h e a t e r A i r T r a n s p o r t Systems - A Cost E f f e c t i v e n e s s

Comparison , I t p r e s e n t e d a t AIAA M i l i t a r y A i r c r a f t Sys terns Meeting, October 1966.

S t e p n i e w s k i , W. Z. and P r a g e r , P. C . , I1VfOL - New F r o n t i e r

of F l i g h t , I 1 V e r t i - F l i t e , A p r i l 1967.

Brown, E. L . and F i s c h e r , J. N . , "Comparative p r o j e c t i o n s of H e l i c o p t e r s , Compound H s l i c o p t e r s and T i l t i n g - P r o p r o t o r

Low-Disc-Loading VTOL A i r c r a f t f o r C i v i l A p p l i c a t i o n s ,"

AIAA Paper 67-939, p r e s e n t e d a t t h e 4 t h Annual Meeting, October 1967.

D e c k e r t , W. H . and F e r r y , R . G . , ['Limited F l i g h t Evaluation

of t h e AT-3 A i r c r a f t ," TR-60-4, A i r Force F l i g h t T e s t

C e n t e r , May 1960.

" P i l o t E v a l u a t i o n o f t h e B e l l Model XV-3 V e r t i c a l Takeoff and Landing A i r c r a f t , " Report ATO-'I'R-62-1, U . S. Army T r a t ~ s p o r t a t l o n M a t e r i a l Command, F e b r u a r y 1962.

Reeder , John -2. , ''Hand: i n g Qua1 i t i e s Experiences t A ~ i t h S e v e r a l VTOL Resea?:ch A i r c r a f t , I t NASA TN-D-735, March 1961.

H a l l , W. E a r l , ' P r o p r o t o r S t a b i l i t y a t High A d v ~ n c e R a t i o s , " Journal of the American H e l i c o p t e r S o c i e t y , ~ ~ l n c 1966.

''Rotor/Pylon S t a b i l i t y a t High Advance Ratios , I 1 Bell Helicopter Company Report 599-063-903.

Brown, E . L . , e t a 1 . , "Results of T i l t i n g - R o t o r Dyna!sic S t a b i l i t y ~ o Z l T e s t s , November 1965-March 1966 ," & L l ).l.elicapter Company Report 599-063-904, A p r i l 2 8 , 196t.

"I1266 Composite Research A i r c r a f t Design Concept. I' B e ll Helicoptei' Company Report D266-099-101.

Mzy 1965.

I'D266 ComposLte Research A i r c r a f t R e s u l t s of A n a l y t i c a l Studies ,I' E e l l H e l i c o p t e r Coapany Report D266-099-110.

May 1966.

"Model 266 C ~ a p o s i t e A i r c r a f t Program: Design Concepts ," Bell He1 i c o p t e r Company Report 266-099-201, ''Model 266 Composite 4 i r c r a f t Program: Dynamic fjlodel T e s t s and Artalytical S t u d i e s of high-Risk Areas ," B e l l H e l i c o p t e r Coapany R e p rt 266-099-212.

Wernicke , K. C - . , "Tilt-Proprotor Coaposits A i r c r a f t , Design S t a t e of t h e A r t , " presented a t t h e 24th Annual National Forua of t h e American H e l i c o p t e r S o c i e t y , May 1968.

Gaffey, T . M., et aL., "Analysis and Model T e s t s gf the

.-

Proprotor Dynamlcs of a T i l t - P r o p r o t o r VTOL A i r c r a f t ," pres:nted a t t h e A i r Force V/STOL Technology and Planr,i?g Conference , September 1969.

"Advancement of P r o p r o t o r Techwlogy ,I' NASA Contract NAS2-5386, Yay 7 , 1969.

Wernicke, K, G . : and Edenborough, H. K,, "Full-Scale Prop- r o t o r Development, '' presented a t t h e 27+h Annual Yational V/STOL Forum af t h e Americaa H e l i c o p t e r S o c i e t y , May 1971.

"PIASTRAY Users Maoual , I ' ?Jat i o n a l Aeronautics and Space A d a i n i s t r a t i o n Report NASA SP-222, Washington. D . C., October 1969.

Livingston, C . L.. lzRotor Aerodynamic C h a r a c t e r i s t i c s Prcgrarn F3S( . ! ) , I 1 Bell H e l i c o p t e r Company Report 599-004-900, 24 June 1967.

Neal, G. T . , ''Proprotor Dynaaic S t a b i l i t y Analysis - Program

DYN-4 ," Belt H e l i c o p t e r Company Report 599-099-011, t o be

published.

2 8 , Yen, J i n g G . , e t a t - , "A S t u d y of F o l d i n g P r o p r o t o r VTOL Dynamics , " AFFDL-TR-71-7, F e b r u a r y 1 9 7 1 .

2 9 . Cox, C. R . , "Rotor Noise Measurements i n Wind T u n n e l s , " P r o c e e d i n g s T h i r d Cal/AVLABS S p p o s i u r n , B u f f a l o , New York, Aerodynamics of R o t a r y Wing and V/!XOL A i r c r a f t , Vol. 1 , June 1 9 6 9 .

SO. Cox, C. R, , " F u l l - S c a l e H e l i c o p t e r Rotor Xoise Measurements i r i 4aes 40- by 8 0 - F o o t Wind Tunttel," B e l l H e l i c o p t e r Company Report 576-099-052, J u l y 1967, APPENDIX Nomenclature Powered Test Data Listings Drawings NOMENCLATURE Corncute r SpboL Output Description

-

Rotor a c t i v i t y f a c t o r = 218 (based on a constant blade chord of 14 inches).

FLAP Fore a ~ d aft f i ~ p o i n g angle with respect to the haft (8eg) ( see Figurr A - 1 ) .

CYCLIC Fore and a f t c y c l i c angle with respect t o the s h a f ~ (deg' isee Figure A - 1 1 .

C Speed of socnd ( f ps 1 .

CL L i f t c o e f f i c i e n t = L/( 16pn 2 R 4 ).

2 3 CLH Helicopter l i f t c o e f f i c i e n t = L .'(p?;fl R ) .

CPB Airplane power c o e f f i c i e n t based on the m a s t torque power data = 550 ( H ~ ) / ( 3 2 p n 3 ~ 5 > . This value used i n the t e x t and figures - CPF Propulsive force c o e f f i c i e n t = ~ ~ ' 1 6 p n ~ R ~ .

CP BH Helicopter power c o e f f i c i e n t based on the q a s t torque power data = 550 ( H ~ ) / [ ~ I ~ R ~ ( ~ ~ R ) ~ 1 .

CPHSOL Helicopter power coef f i c i e n t / r o t o r s o l i d i t y r a t i o = CpH/'v.

CPl Airplane power c ~ e f f i c i e n t based on the w i n d - t u n n e l balance por-ler data = 550 (Hpl I/( 3 2 p n 3 ~ 5 ) .

CP 1 H Helicopter power c o e f f i c i e n t based on the wind-tunnel balance power data = 550 (lIpl ) / [ p a ~ 2 ( ~ ~ ) 3 ] .

CP2 Airplane power c o e f f i c i e n t based on the test-stand c e l l p o w r data = 5 5 0 (FiP2)/( 32pn3~5).

CPZH Helicopter power c o e f f i c i e n t based on the test-stand load c e l l power data = 550 ( H ~ ~ ) / [ ~ T R ~ ( R R ) ~ 1.

CT Airplane t h r u s t c o e f f i c i e n t = ~ / 1 6 p n ~ R ~ .

C o m p u t e r S ~ ~ O L o u r p u t D e s c r i p t i o n H e l i c o p t e r t h r u s t c o e f f i c i e n t ' r o t o r s o l i d i t y CTH SOL r a t i o = C T ~ / ' O .

D r a g ( 1 5 ) .

D Drag r e f e r r e d t o sea L e v e l s t a n d a r d DSIG c o n d i t i c n s = !3,/mt ! LS).

3 12

F i g u r e o f m e r i t = 0 . 7 9 7 %

CpH = 0.798 C33/*,/cP.

R o t o r "HI' force, p e r p e n d i c u l a r t o t h e s h a f t a x i s ( Lb) .

R o t o r "H" force r e f e r r e d t o sea L e v e l H S I G s t a n d a r d c o n d i t i o n s = H / u ' (Lb).

H o r s e p o w e r b a s e d o n t h e aast t o r q u e . T l i i s v a l u e u s e d i n t h e t e x t a n d f i g u r e s .

HPBSIG H o r s e p o e r b a s e d on t h e mast t o r a u e r e f e r r e d t o sea l e v e l s t a n d a r d c o n d i t i o n s = Hp,/ua .

H o r s e p o w e r b a s e d o n t h e w i n d - t u n n e l b a l a n c e d a t a .

H c r s e p o w r b a s e d o n t h o w i n d - t u n n e l b a l a n c e H P l S I G d a t a r e f e r r e d t o sea L e v e l s t a n d a r d c o n d i t i o n s = H p L / u t .

S.:c\rsepower b a s e d o n t h e t e s t s t a n d l o a d c e l l d a t a .

H o r s e p o w e r b a s e d o n the t e s t s t a n d l o a d c e l l d a t a r e f e r r e d t o sea L e v e l s t a n d a r d c o n d i t i o n s = Hp2/u1 .

A i r p l a n e a G v a n c e r a t i o = V/2nR.

A i r p l a n e a d v a n c e ratio c o r r e c t e d f o r s h a f t a o g l e o f a t t a c k = (V/2nR) c o s (amST).

A i r p l a n e a d v a n c e r a t i o c o r r e c t e d f o r t i p p a t h p l a n e a n g l e = (V/2nR) c o s (aTpp).

Computer O u t p u t L k s c r i p t i o n Symbol L i f t ( L b ) .

LSIG L i f t r e f e r r e d t o s e a - L e v e l s t a n d a r d c o r l d i t i o n s = L / u g ( l b ) .

A d v a n c i n g t i p Mach number

= ( l / C ) Jv' + ( f l ~ ) * + 2VRR c o s (aTpp)

RPS R o t o r r p s .

F F P r o p u l s i v e f o r c e = -D ( l b ) .

PFSIG P r o p u l s i v e f o r c e r e f e r e e d t o sea L e v e l s t a n d a r d c o r i d i t i o n s = PF/U@ ( L b ) .

D a t a p o i n t number.

Dynamic p r e s s u r e ( p s f 1 .

Mast t c r q i i e ( f t - l b ) .

hotor r a d i u s ( f t : .

RF R e s ~ l t a n t f o r c e = 4 - ( l b ) .

RFSIG R e s u l t a n t f o r c e r e f e r r e d t o sea l e v e l s t a n d a r d c o n d i t i c n s = %/ul ( l b ) .

R P M R o t o r rpm.

RUN Run number.

T T h r u s t a l o n g t h e s h a f t axis ( l b ) .

TSIG T h r u s t a l o n g t h e s h a f t a x i s r e f e r r e d t o s e a l e v e l s t a n d a r d c o n d i t i o n s = T / u g ( l b ) .

V e l o c i t y (f p s ) .

T u n n e l s p e e d ( k n o t s ) .

V ( KTS ALN Mast a n g l e o f a t t a c k ' d e g ) .

ALTPP T i p p a t h p l a n e a n g l e of a t t a c k - = aMST -- a l ~ - 90 ( d e g ) .

P r o p r o t o r b l a d e f l a p p i n g a n g l e , m e a s u r e d b e t w e e n t h e b l a d e - s p a n axis a n d t h e mast a x i s . P o s i t i v e f o r b l a d e f l a p p i n g u p .

@ BELL H E L A C O P T E R comprrur

Computer D e s c r i p t i o n SXI\bol O u t p u t Change ir. s p i n n e r b a s e p r e s s u r e r e l a t i v e t o f r e e s t r e a m s t a t i c p r e s s u r e .

'I ETC B P r o p u l s i v e e f f i c i e n c y b a s e d on t h e mast t o r q u e power d a t a = (TV) c o s (aMAST ) / ( 5 5 0 H p i .

T h i s v a l u e u s e d i n t h e t c x t a n d f i g u r e s .

P r o p u l s i v e e f f i c i e c c y b a s e d on t h e w i n d - t u n n e l b a l a n c e power d a t a = (TV) c o s i a m ~ ~ ) / ( 550 Hpl).

Propu 1s i v e e f f i c i e n c y b a s e d on t h e t e s t s t a n d Load c e l l power d a t a = (TV) cos (aMST /( 550 H p 2 .

THERF R e s u l t a n t force a n g l e = a r c t a n (L,/PF) ( d e g ; .

COLLECT T i p c o l l e c t i v e a n g l e ( d e g ) .

MU H e l i c 3 p t : e r a d v a n c e r a t i o c o r r e c t e d f o r t i p p a t n p l a n e a n g l e = (!!/OR) c o s ( a T p p ) .

H e l i c o p t e r a d v a n c e r a t i o = v/OR.

M L !

T. SPYYD R o t o r t i p speed = RR ( f p s ) .

RHO A i r d e n s i t y ( s l u g / f t 3 ) .

R o t o r s o l i d i t y = 0.089 ( b a s e d on a c o n s t a n t b l a d e c h o r d of l k i n c h e s 1, SIGMA -P A i r d e n s i t y r a t i o = pIPo.

R o t o r a n g u l a r v e l o c i t y ( r a d / s e c ) .

Direction of Angles and Forces.

Figure A - 1 .

@ BELL HEMCOPTER c--Nr

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