Document
h
NASA C o n t r a c t o r R e p o r t 179458
Propfan Test Assessment
Testbed Aircraft Flutter
I. Y
Model- Test Report
C.M.J. Jenness
Lockheed-Georgia Company
Marietta, Georgia
- P r e p a r e d f o r L e w i s R e s e a r c h C e n t e r u n d e r C o n t r a c t N A S 3 - 2 4 3 3 9
NASA
N a t i o n a l A e r o n a u t i c s a n d S p a c e A d m i n i s t r a t i o n L e w i s R e s e a r c h C e n t e r C l e v e l a n d , O h i o 4 4 1 35 u n t i l 19 June 1987 FOREWORD The f l u t t e r model t e s t program described h e r e i n was conducted by t h e Propfan Test Assessment under Lockheed Georgia Company as a p a r t o f t h e c o n t r a c t NAS3-24339 with t h e NASA-Lewi s Research Center. M r . C. M. Jenness
was t h e Lockheed p r o j e c t engineer and M r . J . F . Lubomski was t h e NASA tech-
n i c a l m o n i t o r f o r t h e f l u t t e r model program.
The wind t u n n e l t e s t s were performed i n t h e Transonic Dynamics Tunnel a t t h e NASA Langley Research Center. M r . C. H. R u h l i n was t h e NASA Langley p r o j e c t engineer and M r . 0. J. Crooks was t h e Lockheed t e s t en- g i neer.
The f l u t t e r model d e t a i l design and f a b r i c a t i o n were performed by Calcu- search, I n c o r p o r a t e d ( A t l a n t a , Georgia) under t h e d i r e c t i o n o f M r . D. C. Cone.
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LIST O F FIGURES F i q u r e
- T i t l e Page
M o d i f i e d Gulfstream I1 Propfan Testbed A i r c r a f t Propfan Testbed A i r c r a f t Design F l i g h t Test Envelope and F l u t t e r Safety Margin Bare Wing Model i n NASA Langley Transonic Dynamics Tunnel Twin Propfan Model w i t h W i ndmi 11 i ng Props S i n g l e Propfan Model with ( a ) W i n d m i l l i n g Prop (b) Weighted Spinner S i n g l e Propfan Model General Arrangement Model Propfan Powerplant S i m u l a t i o n 30 Model Cable Mount System Schematic
Model Test Envelope 32
Model Propfan W i n d m i l l i n g Speed f o r 52.5' Blade P i t c h Angle 11 34 Test C o n d i t i o n s f o r Test Number 1 T e s t Conditions f o r Test Number 2 13 Test Conditions f o r Test Number 3 T e s t Conditions f o r Test Number 4 Test Conditions f o r Test Number 5 T e s t Conditions f o r Test Number 6 T e s t Conditions f o r Test Number 7 18 Test Conditions f o r Test Number 8 T e s t Conditions f o r Test Number 9 20 Test Conditions f o r Test Number 10 T e s t Conditions f o r Test Number 11 22 Test Conditions f o r Test Number 12
LIST OF FIGURES (CONT'D)
T i t l e Page
F i q u r e Test C o n d i t i o n s f o r T e s t Number 13 23.
24 Test C o n d i t i o n s f o r T e s t Number 14 25 Test C o n d i t i o n s f o r T e s t Number 15 3c; ua , b u l a t e d .._-_.._ M _ _ _ _ _ . _ _ -I r i . . A A - - D - . --I-.- r I , . .
v e l s u a i.lea>ur eu r I U L L ~ I ouuriuar-y -
L V Twin Propfan C o n f i g u r a t i o n w i t h D e s t a b i l i z i n g Booms
27 Cal c u l ated versus Measured F1 u t t e r Boundary -
S i n g l e Propfan C o n f i g u r a t i o n w i t h D e s t a b i l i z i n g Booms and Spinner
28 C a l c u l a t e d versus Measured F l u t t e r Boundary - S i n g l e
Propfan C o n f i g u r a t i o n w i t h D e s t a b i l i z i n g Boom and Propf an Wing Geometry Fuselage Geometry V e r t i c a l S t a b i l i z e r Geometry 32 H o r i z o n t a l S t a b i 1 i z e r Geometry Bare Wing V e r t i c a l Bending S t i f f n e s s 34 Bare Wing T o r s i o n a l S t i f f n e s s 35 Bare Wing Fore-and-Aft S t i f f n e s s Propfan Wing V e r t i c a l Bending S t i f f n e s s 37 Propfan Wing T o r s i o n a l S t i f f n e s s Propfan Wing Fore-and-Aft Bending S t i f f n e s s 39 Fuselage V e r t i c a l Bending S t i f f n e s s
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40 Fuselage L a t e r a l Bending S t i f f n e s s r 41 Fuselage T o r s i o n a l S t i f f n e s s \ -
Measured Normal Modes - Bare Wing C o n f i g u r a t i o n
. & -
- Twin Propfan C o n f i g u r a t i o n
43 Measured Normal Modes _ -
Measured Normal Modes - S i n g l e Propfan C o n f i g u r a t i o n
- r v i i i LIST OF SYMBOLS Symbo 1 Def i n i ti on b Reference l e n g t h BL B u t t o c k l i n e - L a t e r a l d i s t a n c e from f u s e l a g e p l a n e o f symmetry (Fig. 1) C F l e x i b i l i t y i n f l u e n c e c o e f f i c i e n t ( d e f i n e d i n Tables 21 and 22) D Propfan diameter E I Beam bending s t i f f n e s s FS Fuselage s t a t i o n - D i s t a n c e a l o n g f u s e l a g e r e f e r e n c e l i n e from a p o i n t .20m (8 in.) f o r w a r d o f nose radome (Fig. 1)
GJ
Beam t o r s i o n a l s t i f f n e s s C e n t r o i d a l mass moments of i n e r t i a about x, y, and t axes, r e s p e c t i ve 1 y in.
Inches K i lograms kg l b Pounds M Mach number Design d i v e Mach number MD m Meters N Newtons n Propfan r o t a t i o n speed Dynamic pressure S Seconds V Ve 1 o c i t y Design d i v e v e l o c i t y "0 V/nD Propfan advance r a t i o
W L Water l i n e - V e r t i c a l d i s t a n c e from s t a t i c ground
l i n e a t 1) nose gear (Fig.
LIST OF SYMBOLS (CONT'D)
Def i ni ti on Symbol
w s Wing station - Spanwise distance i n wing reference plane
Distances along longitudinal, lateral, and vertical axes, respecti ve 1 y
- - -
Center of gravity locations along x, y, and z axes, re- X I Y, 2 spectively Atmospheric density P Frequency o f oscillation c3 xi i SUMMARY The Propfan Test Assessment ( P T A ) program i n c l u d e s f l i g h t t e s t s o f a propfan powerplant mounted on the l e f t wing o f a m o d i f i e d G u l f s t r e a m I 1 t e s t b e d a i r c r a f t . A s t a t i c balance boom i s mounted on t h e r i g h t wing t i p f o r l a t e r a l balance. F l u t t e r analyses i n d i c a t e t h a t these i n s t a l 1- a t i o n s reduce t h e wing f l u t t e r speed and t h a t t o r s i o n a l s t i f f e n i n g and t h e i n s t a l l a t i o n o f a f l u t t e r s t a b i l i z i n g t i p boom a r e r e q u i r e d on t h e l e f t wing f o r adequate f l u t t e r safety margins.
t e s t s o f a 1 / 9 t h scale h i g h speed f l u t t e r model o f t h e t e s t - Wind t u n n e l bed a i r c r a f t were conducted i n t h e NASA Langley Transonic Dynamics Tunnel d u r i n g August 1985. One o b j e c t i v e o f t h e t e s t s was t o substan- t i a t e t h e a n a l y t i c a l l y p r e d i c t e d wing f l u t t e r s a f e t y of t h e s i n g l e propfan t e s t b e d p r e l i m i n a r y design and a s i m i l a r design w i t h p r o p f a n powerplants on b o t h wings. A second o b j e c t i v e was t o o b t a i n d a t a w i t h which t o v a l i d a t e t h e f l u t t e r a n a l y s i s methods b e i n g used i n t h e a i r - c r a f t f i n a l design. The t e s t program i n c l u d e d t h e design, f a b r i c a t i o n , and t e s t i n g o f t h e f l u t t e r model and t h e c o r r e l a t i o n o f t h e f l u t t e r t e s t d a t a w i t h a n a l y s i s r e s u l t s .
was designed t o s i m u l a t e t h e o p e r a t i o n o f t h e t e s t b e d a i r c r a f t The model t h r o u g h o u t i t s f l i g h t t e s t envelope and t o demonstrate a 20 p e r c e n t f l u t t e r speed margin above l i m i t d i v e speed. It was d y n a m i c a l l y s c a l e d f o r t e s t i n g a t f u l l s c a l e Mach numbers i n Freon 12, which l i m i t e d t h e maximum t e s t dynamic pressure t o about 9000 N/m' (188 l b / f t ' ) .
The model wings and f u s e l a g e u t i l i z e d s i n g l e spar, segmented s h e l l con- s t r u c t i o n . Two complete wings were f a b r i c a t e d , one o f which r e p r e s e n t e d I 1 w i t h s t a t i c balance booms on each t i p and t h e unmodified G u l f s t r e a m t h e o t h e r o f which r e p r e s e n t e d the t o r s i o n a l l y s t i f f e n e d t e s t b e d d e s i g n w i t h a p r o p f a n powerplant and f l u t t e r s t a b i l i z i n g boom on each side.
The unsymmetrical s i n g l e propfan c o n f i g u r a t i o n was represented by i n - s t a l l i n g one-half o f each wing design. The unpowered p r o p f a n powerplants i n c l u d e d d y n a m i c a l l y scaled power sections, gearboxes, and propfans. The 0.30111 (12 i n ) diameter propfans had g r a p h i t e r e i n f o r c e d epoxy blades which c o u l d be s e t a t d i f f e r e n t p i t c h angles t o v a r y t h e w i n d m i l l i n g speed. E q u i v a l e n t weight n o n r o t a t i n g s p i n n e r s c o u l d be sub- s t i t u t e d f o r t h e propfans.
The model was t e s t e d on a v e r y c o m p l i a n t two-cable mount system which produced minimal e f f e c t s on t h e wing f l u t t e r s t a b i l i t y . It was i n s t r u - mented w i t h a combination o f s t r a i n gage b r i d g e s and m i n i a t u r e accel- erometers t o measure i t s loads and dynamic response.
The t e s t procedure c o n s i s t e d o f speed b u i l d u p s a t s e v e r a l t u n n e l t o t a l pressures u n t i l f l u t t e r occurred o r t h e t e s t envelope l i m i t s (1.zvg, M = 0.90) were reached. The propfan b l a d e p i t c h was s e t a t 52.5' i n o r d e r t o achieve t h e nominal r o t a t i o n speed a t a p p r o x i m a t e l y M = 0.9.
F i f t e e n c o n f i g u r a t i o n s were t e s t e d . Ten o f t h e s e r e p r e s e n t e d t h e nominal a i r c r a f t p r e l i m i n a r y d e s i g n s and were i n t e n d e d t o v e r i f y t h e i r p r e d i c t e d f l u t t e r s a f e t y margins. I n c l u d e d were t e s t s of t h e s i n g l e and t w i n prop- and a t e s t o f t h e s i n g l e propfan f a n c o n f i g u r a t i o n s w i t h o u t f l u t t e r booms, c o n f i g u r a t i o n w i t h a s i m u l a t e d f a i l u r e o f t h e gearbox-to-power s e c t i o n connections. No f l u t t e r o r n e a r - f l u t t e r c o n d i t i o n s o c c u r r e d i n any o f t h e s e t e s t s .
The o t h e r f i v e t e s t s were made w i t h d e s t a b i l i z i n g wing booms t o o b t a i n f l u t t e r d a t a f o r a n a l y s i s v a l i d a t i o n . F l u t t e r o r n e a r - f l u t t e r p o i n t s were obtained i n t h r e e o f t h e s e t e s t s . The t w i n p r o p f a n c o n f i g u r a t i o n f l u t t e r e d i n a 16 Hz symmetric mode a t M = .77 and 8857 N/m' (185 l b l f t ' ) dynamic pressure, The s i n g l e p r o p f a n c o n f i g u r a t i o n f l u t t e r e d i n a 15 Hz unsymmetric mode a t M = . 7 9 and 9193 N/m' (188 l b / f t ' ) w i t h a s p i n n e r , and was near f l u t t e r a t M = .77 and 9000 N/m' (188 l b / f t ' ) w i t h a wind- m i l l i n g propfan.
E x c e l l e n t c o r r e l a t i o n s w i t h t h e t e s t d a t a were achieved i n p o s t - t e s t f l u t t e r a n a l y s i s u s i n g a c t u a l model p r o p e r t i e s . It was concluded t h a t t h e f l u t t e r a n a l y s i s method used was capable o f a c c u r a t e f l u t t e r p r e d i c - t i o n s f o r b o t h t h e (symmetric) t w i n p r o p f a n c o n f i g u r a t i o n and t h e (unsymmetric) s i n g l e p r o p f a n c o n f i g u r a t i o n . The same method w i 11 be used f o r t h e f i n a l f l u t t e r a n a l y s i s o f t h e a i r c r a f t .
The f l u t t e r a n a l y s i s a l s o r e v e a l e d t h a t t h e d i f f e r e n c e s between t h e t e s t e d model c o n f i g u r a t i o n s and t h e c u r r e n t a i r c r a f t d e s i g n caused t h e ( s c a l e d ) model f l u t t e r speed t o be s i g n i f i c a n t l y h i g h e r t h a n t h a t o f t h e a i r c r a f t , a t l e a s t f o r t h e s i n g l e p r o p f a n c o n f i g u r a t i o n w i t h o u t a f l u t t e r boom. Thus, it cannot be concluded f r o m t h e model t e s t s a l o n e t h a t t h e c u r r e n t ( o r f i n a l ) a i r c r a f t d e s i g n s n e c e s s a r i l y have adequate f l u t t e r s a f e t y margins. V e r i f i c a t i o n o f t h e a i r c r a f t f i n a l d e s i g n should, t h e r e - f o r e , be based on f l u t t e r p r e d i c t i o n s made w i t h t h e t e s t v a l i d a t e d a n a l y s i s method.
INTRODUCTION The Propfan Test Assessment (PTA) i s a NASA-sponsored program t o e v a l u a t e by f l i g h t t e s t s t h e s t r u c t u r a l i n t e g r i t y and n o i s e c h a r a c t e r i s t i c s o f an e f f i c i e n t , high-speed p r o p e l l e r known as a propfan. The PTA Program i s being c a r r i e d o u t by t h e Lockheed-Georgia Company under C o n t r a c t NAS3-24339 w i t h t h e NASA-Lewis Research Center. It was i n i t i a t e d i n August 1984, f o l l o w i n q t h e e v a l u a t i o n o f t h e r e s u l t s o f svstem s t u d i e s performed by t h e Lockheed-Georgia and Douglas A i r c r a f t Companies (Ref.
1 and 2 ) .
The PTA Program w i l l u t i l i z e a 2.74 meter (9-foot) diameter, 8-blade p r o p f a n developed by t h e Hami 1 t o n Standard D i v i s i o n o f U n i t e d Technologies
i n t h e Large-Scale Advanced Prop-Fan (LAP) Program. The p r o p f a n d r i v e -
system c o n s i s t s o f a m o d i f i e d A l l i s o n Model 570 i n d u s t r i a l engine and m o d i f i e d A 1 1 i s o n T56 r e d u c t i o n gearbox. The propfan powerplant w i 11 be mounted on t h e l e f t wing of a m n r l i f j ~ d Gu!fstream American GI1 testbed a i r c r a f t ( F i g u r e 1). A s t a t i c balance boom w i l l be i n s t a l l e d on t h e r i g h t wing t i p t o counterbalance t h e w e i g h t o f t h e propfan powerplant.
The program a l s o i n c l u d e s t h e p r e l i m i n a r y design o f a t w i n propfan G I 1 t e s t b e d c o n f i g u r a t i o n .
P r e l i m i n a r y f l u t t e r analyses i n d i c a t e d t h a t t h e propfan i n s t a l l a t i o n a d v e r s e l y a f f e c t s t h e G I 1 wing f l u t t e r s t a b i l i t y and t h a t m o d i f i c a t i o n s a r e r e q u i r e d t o p r o v i d e adequate f l u t t e r s a f e t y margins d u r i n g t h e prop- f a n f l i g h t t e s t i n g . The f l u t t e r m o d i f i c a t i o n s i n c l u d e t h e a d d i t i o n o f e x t e r n a l doublers on t h e l e f t wing upper and lower surfaces t o p r o v i d e increased t o r s i o n a l s t i f f n e s s inboard of t h e propfan n a c e l l e and a f l u t t e r s t a b i l i z i n g "dynamic balance'' boom on t h e l e f t wing t i p . The same modi- f i c a t i o n s a r e made t o b o t h wings of t h e t w i n p r o p f a n t e s t b e d configuration.
The t e s t b e d f u s e l a g e and empennage a r e s t r u c t u r a l l y unmodified from t h e G I 1 design.
The f l u t t e r p r e v e n t i o n program c o n s i s t s of f l u t t e r analyses, a f l u t t e r model t e s t , a ground v i b r a t i o n (resonance) t e s t , and a f l i g h t f l u t t e r t e s t . The f l u t t e r model t e s t was i n c l u d e d i n t h e program p r i m a r i l y be- cause o f t h e u n c e r t a i n e f f e c t s o f t h e p r o p f a n a t h i g h subsonic Mach numbers and o f t h e c o n s i d e r a b l e unsymmetry o f t h e s i n g l e p r o p f a n c o n f i g - u r a t i on.
The o b j e c t i v e s o f t h e model t e s t a r e t o s u b s t a n t i a t e t h e a n a l y t i c a l l y p r e d i c t e d f l u t t e r s a f e t y margins o f t h e s i n g l e and t w i n p r o p f a n t e s t b e d p r e l i m i n a r y designs and t o v a l i d a t e t h e methods used i n t h e a i r c r a f t f i n a l f l u t t e r analyses. To meet these o b j e c t i v e s , a 1/9-scale high-speed f l u t t e r model capable of r e p r e s e n t i n g t h e s i n g l e and t w i n p r o p f a n , t e s t b e d c o n f i g u r a t i o n s was designed and f a b r i c a t e d . The model i n c l u d e s dynam- i c a l l y scaled p r o p f a n powerplants w i t h windmi 11 i n g propfans t o r e p r e s e n t The model t h e e s s e n t i a l aerodynamic and gyroscopic e f f e c t s o f t h e props.
was t e s t e d a t f u l l - s c a l e Mach numbers i n t h e NASA Transonic Dynamics Tunnel (TDT) a t t h e Langley Research Center from 12 t o 30 August 1985.
The t e s t s a r e b e l i e v e d t o be unique i n t h a t p r o p e l l e r e f f e c t s were i n v e s t i g a t e d a t Mach numbers up t o 0.90.
T h i s report describes the design, construction, a n d testing o f the flutter model. The results o f post-test analysis correlations with the test data are also included.
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MODEL DESCRIPTION
SCALING
The f l u t t e r models were designed t o s i m u l a t e t h e o p e r a t i o n of t h e t e s t b e d a i r c r a f t t h r o u g h o u t i t s f l i g h t t e s t envelope and t o demonstrate a 20 per- c e n t f l u t t e r speed s a f e t y margin above l i m i t d i v e speed ( F i g u r e 2 ) . The models were designed f o r t e s t i n g i n Freon 12 i n t h e NASA Langley Tran- s o n i c Dynamics Tunnel (TDT). The Freon 12 t e s t medium was necessary because i t s low s o n i c v e l o c i t y reduces t h e dynamic pressure t o about one- f o u r t h of t h a t i n a i r a t t h e same Mach number. T h i s reduces t h e r e q u i r e d s t r e n g t h o f t h e models t o achievable l e v e l s .
The models were d y n a m i c a l l y scaled, w i t h t h e e x c e p t i o n o f t h e p r o p f a n b l a d e s and empennage, which have higher-than-scaled s t i f f n e s s e s and ( a p p r o x i m a t e l y ) scaled mass p r o p e r t i e s . T h i s approach reduced the risk.
o f encountering f l u t t e r o f these components, which m i g h t have prevented t e s t i n g o f t h e wing.
t h e s u c c e s s f u l The model d e s i g n scales a r e shown i n Table 1. A geometric s c a l e o f 1/9 was s e l e c t e d because i t p e r m i t t e d t h e use o f t h e s m e p r ~ p f a n blades iised on t h e PTA s t a b i l i t y and c o n t r o l model, and i t was c o m p a t i b l e w i t h t h e TOT t e s t s e c t i o n dimensions (4.88111 x 4.88m o r 16 f t x 16 ft).
The Mach, d e n s i t y , v e l o c i t y and dynamic p r e s s u r e s c a l e s were e s t a b l i s h e d b y r a t i o j n g t h e "design p o i n t " c o n d i t i o n s f o r t h e a i r c r a f t and model.
e
The most a c c u r a t e dynamic s i m u l a t i o n o f t h e a i r c r a f t occurs i n t h e v i c i n i t y o f t h e d e s i g n p o i n t . The design p o i n t was 1 . 2 V D a t M-1.0 The corresponding model d e s i g n p o i n t was l o c a t e d w i t h i n t h e ( F i g u r e 2).
TDT o p e r a t i n g envelope a t M=l.O and a dynamic p r e s s u r e o f 8426 N/m' (173
p s f ) . The Freon d e n s i t y and sonic v e l o c i t y a t t h i s p o i n t were determined f r o m t h e TDT o p e r a t i n g d a t a i n Reference 3 and were used t o e s t a b l i s h t h e d e n s i t y and v e l o c i t y scales. The remaining s c a l e s (frequency, weight, mass moment o f i n e r t i a , and s t i f f n e s s ) were d e r i v e d f r o m t h e o t h e r s by means o f standard dynamic s i m i l a r i t y laws.
WING CONFIGURATIONS
Three wing c o n f i g u r a t i o n s were tested. They were t h e b a r e wing, s i n g l e propfan, and t w i n p r o p f a n c o n f i g u r a t i o n s . Photographs o f t h e t h r e e a r e 3'-5. The bare wing c o n f i g u r a t i o n ( F i g u r e 3) r e p r e s e n t e d shown i n F i g u r e s an unmodified G u l f s t r e a m I 1 ( G I I ) wing w i t h a 1134kg (2500 l b ) s t a t i c balance boom on each t i p . I t s purpose was t o s u b s t a n t i a t e t h e p r e d i c t e d f l u t t e r s t a b i l i t y o f t h e r i g h t wing of t h e s i n g l e propfan t e s t b e d a i r - c r a f t , w i t h o u t t h e c o m p l i c a t i n g e f f e c t s o f an unsymmetrical l e Y t wing.
The t w i n p r o p f a n c o n f i g u r a t i o n ( F i g u r e 4 ) r e p r e s e n t e d t h e s t i f f e n e d GI1 wing design w i t h p r o p f a n powerplants and 136kg (300 l b ) f l u t t e r s t a - b i l i z i n g booms on b o t h sides. I t s purpose was t o s u b s t a n t i a t e t h e p r e d i c t e d f l u t t e r s t a b i l i t y o f t h e t w i n p r o p f a n t e s t b e d a i r c r a f t pre- l i m i n a r y design. It was a l s o used t o o b t a i n symmetrical f l u t t e r d a t a f o r a n a l y s i s v a l i d a t i o n and f o r comparison w i t h t h e unsymmetrical f l u t t e r d a t a o b t a i n e d w i t h t h e s i n g l e propfan c o n f i g u r a t i o n .
The single propfan configuration (Figure 5 j represented the unsymmetrical flight test aircraft (preliminary) design, with its propfan powerplant and flutter boom on the left side and static balance boom on the right.
Its left wing semispan was the same as that of the twin propfan config- uration and its right semispan the same as that of the bare wing configuration, The same fuselage and empennage were used with all three wing configurations. The single propfan configuration was used to sub- stantiate the predicted flutter safety of the highly unsymmetrical testbed aircraft design. It was also used to obtain unsymmetrical flutter data for analysis validation and comparison with the symmetrical twin propfan flutter data. For these tests, destabilizing, aft slung wing tip booms were substituted for the flutter stabilizing booms in order to induce well defined flutter instabilities within the model test envelope. Tests of the single and twin propfan configurations were made with windmilling propfans (Figure 5a) and with equivalent weight non- 'rotating spinners (Figure 5 . b ) .
CONSTRUCTION The single propfan model general arrangement i s shown i n Figure 6.
The model wings and fuselage are constructed with hollow aluminum spars and segmented, f i berg1 ass-rei nforced, wooden aerodynamic fai rings which are attached to the spars with aluminum 'lbridges". The spars provide the scaled stiffness and part of the mass and inertia properties of the corresponding aircraft components. The fairings provide the external shape and remai ni ng mass and inertia properties. Fuel mass properties are represented by removable metal weights which attach directly to the wing spars. Two complete wings are provided, one representing the twin propfan configuration and the other representing the bare wing config- uration. The single propfan configuration is represented by installing the twin propfan left wing semispan and the bare wing right semispan.
The wings are built without twist to alleviate the excessive outer wing down loading which would otherwise occur at high dynamic pressures. A 25% chord roll trim tab is located on each right wing semispan. It is operated through a worm gear and torque tube by an electric motor located in the fuselage.
The propfan powerplants (Figure 7 ) consist of masses representing the power section, gearbox, and propfan, which are supported by springs re- presenting the engine mounts and a built-up aluminum truss representing the nacelle structure. Each propfan consists of 8 graphite-epoxy blades mounted i n an aluminum hub. The hub is bolted to a steel shaft which turns on two ball bearings i n the simulated gearbox. The blade pitch is adjustable so that the windmilling speed can be varied. The blades are identical to those on the stability and control model and are designed for rotation speeds u p to 18,000 rpm. The maximum rotation speed during flutter model testing was approximately 7500 rpm.
The fuselage spar has a wing attachment fitting which simulates the roll and yaw flexibilities of 'aircraft wing attachment. Pulley brackets for the cable mount system are bolted to the fuselage spar forward and aft of the wing fitting. A Y-shaped aluminum box structure attaches the sim- - ulated jet engines to the aft fuselage spar. Flexures at its ends s i m u l a t e t h e v e r t i c a l and p i t c h f l e x i b i l i t i e s o f t h e a i r c r a f t engine p y l ons.
Because t h e y a r e n o t s t i f f n e s s scaled, t h e f i n and s t a b i l i z e r a r e unsegmented monocoque surfaces. They c o n s i s t o f aluminum r o o t and t i p r i b s and spars w i t h bonded fiberglass-cloth-reinforced epoxy skins.
P l a s t i c foam cores s t a b i l i z e t h e skins. The f i n s k i n s extend o v e r and a r e b o l t e d t o t h e a f t fuselage spar, A r u d d e r o f s i m i l a r c o n s t r u c t i o n p r o v i d e s yaw t r i m . It i s operated through a worm gear d r i v e and t o r q u e t u b e by an e l e c t r i c motor l o c a t e d i n t h e fuselage. The s t a b i l i z e r i s a t t a c h e d t o t h e f i n through a p i v o t f i t t i n g and jackscrew, which p r o v i d e s
+/- 5 degrees p i t c h t r i m , The jackscrew i s operated through a j o i n t e d
t o r q u e tube by an e l e c t r i c motor l o c a t e d i n t h e fuselage,
PRODUCT ASSURANCE
t h e . The model d e s i g n and f s b r i c z t i s n were c a r r i e d o u t i n compliance w i t h q u a i i t y assurance c r i t e r i a s p e c i f i e d i n t h e NASA-Langley Wind-Tunnel Model Systems C r i t e r i a handbook (Ref. 4), as summarized below. C a l i b r a - t i o n of t h e t e s t i n s t r u m e n t a t i o n i s d e s c r i b e d i n t h e I n s t r u m e n t a t i o n s e c t i o n o f t h i s r e p o r t .
The model d e t a i 1 d e s i g n and f a b r i c a t i o n were vendor supplied. Q u a l i t y assurance c o n t r o l s were i n i t i a l l y e s t a b l i s h e d and i n c o r p o r a t e d i n t o t h e "Propfan Test A i r p l a n e F l u t t e r Model S p e c i f i c a t i o n " (Ref. 5), which was a p a r t o f t h e vendor purchase agreement. Requirements f o r design reviews, i n s p e c t i o n , acceptance, and d e l i v e r y were s p e c i f i e d . A design r e v i e w o f each major model component was accomplished p r i o r t o i t s f a b r i c a t i o n .
P e r i o d i c v i s i t s t o t h e vendor's f a c i l i t y were made t o ensure s p e c i f i c a t i o n compliance, r e s o l v e problems, and approve necessary d e s i g n changes.
The c r i t i c a l wing and fuselage spars were 100-percent X-ray i n s p e c t e d a t Lockheed a f t e r w e l d i n g and h e a t treatment. C e r t i f i c a t i o n s o f t h e m a t e r i a l s and h e a t t r e a t m e n t processes used were p r o v i d e d by t h e vendor.
These documents, a l o n g w i t h the i n t e r p r e t e d X-ray f i l m s were s u b m i t t e d t o t h e NASA-Langley TDT F a c i l i t y S a f e t y O f f i c e r w i t h t h e "Propfan T e s t Assessment t e s t b e d A i r c r a f t F l u t t e r Model System S a f e t y " r e p o r t (Ref. 6 ) f o r approval p r i o r t o t h e wind tunnel t e s t .
were accomplished a t t h e ven- The i n s p e c t i o n and acceptance of t h e model d o r ' s f a c i l i t y . A combination o f s t a t i c loads, s t i f f n e s s , mass p r o p e r t y , and resonance t e s t s was performed t o v e r i f y vendor-supplied d a t a and compliance w i t h s p e c i f i c a t i o n s . Shipment t o t h e Lockheed f a c i l i t y was accomplished by Lockheed personnel, i _- _ - CABLE MOUNT SYSTEM
The model was restrained in the TDT test section by the cable mount
system shown schematically in Figure 8 . The flying cable mount system consisted of a forward cable oriented in a vertical plane and a rear cable oriented in a horizontal plane. The ends o f the forward cable attached to the test section floor and ceiling at tunnel station 13.54 (533) and passed over pulleys located in the model forward fuselage at
FS 0.804 (31.64). The rear cable loop passed over pulleys located in
the model aft fuselage at FS 1.686 (66.36) to pulleys mounted on the test section walls at tunnel station 29.01 (1142) to a remotely controlled tensioning device located outside the test section. The rear cable ten- sion was normally maintained at 667N (150 l b ) by a soft spring in this device.
The flying cable system was very compliant in the vertical, lateral, and angular degrees o f freedom and therefnre d.:d n c t signSficmtiy affect t h e wing flutter stability. For example, the maximum rigid body mode frequency (pitching) at zero airspeed was approximately one-tenth o f the wing f 1 utter frequency.
Emergency restraint was s u p p l i e d by t h e f ~ u r - t a b l e snubber system shown i n Figure 8 . The cables attached to the model fuselage, at FS 1.273 (50.11, passed over pulleys located i n the test section wall slots and then t o a tensioning and damping device located outside the test section.
The normally slack snubber cables could be quickly tensioned to 356N (80 lb) by a pneumatic cylinder i n this device. A spring-damper cartridge maintained tension and provided damping for each cable.
A dynamic stability analysis of the model and cable mount system was per- formed to ensure adequate stability of the model with and without the snubber cables tensioned.
INSTRUMENTATION
MODEL SENSORS
The model was instrumented w i t h a combination o f s t r a i n gage b r i d g e s and m i n i a t u r e accelerometers t o measure t h e 1oads and dynamic response.
H a l l - e f f e c t p u l s e t r a n s d u c e r s and frequency counters were used t o moni- t o r t h e p r o p f a n r o t a t i o n speeds. S i x t e e n o f these channels were d i s p l a y e d and recorded on two 8-channel pen recorders. T h i r t e e n chan- n e l s were a l s o recorded on an F M t a p e recorder, a l o n g w i t h a t i m e code and v o i c e recorded i n f o r m a t i o n . The recorded channels v a r i e d w i t h t h e wing c o n f i g u r a t i o n b e i n g tested, as shown i n Tables 2-4.
S i x t o e i g h t o f t h e s t r a i n gage channels were c a l i b r a t e d and monitored t o ensure t h a t t h e model maximum d e s i g n loads were n o t exceeded. I n - c l u d e d were t h e wing v e r t i c a l bending, f u s e l a g e v e r t i c a l and l a t e r a l bending, left j e t a i g i n e pylon bending and i e f t p r o p f a n n a c e l l e v e r t i - c a l and l a t e r a l bending s t r a i n gage channels. These o u t p u t s were low- pass f i l t e r e d t o y i e l d ' ' s t a t i c ' ' loads. Generally, t h e wing bending and t o r s i o n and a f t f u s e l a g e v e r t i c a l bending and tors,ion s t r a i n gages and t h e f u s e l age nose and propfan gearbox accelerometers (when appl i cab1e) wsrs m o n i t o r e a (ana recorded] f o r i n d i c a t i o n s o f approaching f l u t t e r .
The wind t u n n e l parameters were o b t a i n e d v i a t h e TDT f a c i l i t y d a t a a c q u i s i t i o n system. S t a t i c and s t a g n a t i o n pressure, s t a g n a t i o n temper- a t u r e , Mach number and dynamic pressure were c o n t i n u o u s l y d i s p l a y e d and were p r i n t e d o u t f o r each t a b p o i n t . High speed (128 f t / s e c ) movie cameras l o c a t e d on e i t h e r s i d e and downstream o f t h e model were used t o r e c o r d s i g n i f i c a n t model responses.
CALIBRATION A N D ACCURACY
C a l i b r a t i o n s o f t h e s t r a i n gage channels used t o m o n i t o r t h e model s t a t i c l o a d s were made i n t h e wind t u n n e l b y a p p l y i n g approximate d e s i g n l o a d s t o t h e model and a d j u s t i n g t h e s e n s i t i v i t y so t h a t f u l l s c a l e de- The f l e c t i o n o f t h e pen r e c o r d e r would occur a t t h e d e s i g n l o a d l e v e l .
r e s u l t i n g accuracy o f t h e s t a t i c loads measurements was e s t i m a t e d t o be f5 percent. P r e c i s e measurements were n o t r e q u i r e d because o f t h e l a r g e (IOOX) s a f e t y margins b u i l t i n t o t h e model.
The model dynamic response channels were n o t c a l i b r a t e d because o n l y frequency and r e l a t i v e amplitude d a t a were required. The e s t i m a t e d accuracy o f t h e f l u t t e r frequency measurements i s -+2%. The p r o p f a n ro- t a t i o n speed measurements a r e estimated t o be w i t h i n * l X .
The p r i n c i p a l q u a n t i t a t i v e d a t a r e q u i r e d from t h e t e s t were t h e wind t u n n e l parameters a t t h e f l u t t e r p o i n t s and a t t h e maximum c o n d i t i o n s reached. The measurement accuracy of these d a t a i s discussed i n Ref- erence 1, and a r e g e n e r a l l y as follows: Dynamic pressure 21 % Mach number f. 002 TEST PROCEDURE MODEL TEST ENVELOPE The model t e s t envelope ( F i g u r e 9) was e s t a b l i s h e d by a p p l y i n g t h e design s c a l e s t o t h e t e s t b e d a i r c r a f t dynamic pressures a t 1.2 VD. The maximum model t e s t Mach number was l i m i t e d t o 0.9 by t h e TDT f a c i l i t y s a f e t y requirements. T h i s l i m i t a t i o n was considered unimportant i n t h a t t h e minimum f l u t t e r dynamic p r e s s u r e was p r e d i c t e d t o o c c u r a t M = .865.
F o r each c o n f i g u r a t i o n t e s t e d , b u i l d u p s i n dynamic pressure and Mach number were made along s e v e r a l l i n e s o f a p p r o x i m a t e l y c o n s t a n t t o t a l p r e s s u r e u n t i l f l u t t e r o c c u r r e d o r t h e t e s t envelope c o n d i t i o n s were reached. Generally, t h e t u n n e l was pumped down t o t h e minimum t e s t pres- s u r e ( u s u a l l y 7180-9575 N/m' (150-200 l b / f t ' ) b e f o r e b e g i n n i n g a t e s t .
A speed b u i l d u p was made at. t h i s prosssre, a f t e r which t h e a i r s p e e d was reduced t o zero, and Freon was bled i n t o t h e t u n n e l t o i n c r e a s e t h e pres- s u r e f o r t h e n e x t pass.
PROPFAN WINDMILLING SPEED A t a g i v e n b l a d e p i t c h angle, t h e propfans w i n d m i l l e d a t an a p p r o x i m a t e l y
c o n s t a n t advance r a t i o (VhD). Since t h e speed o f sound remained almost
c o n s t a n t d u r i n g t h e t e s t , t h e p r o p f a n r o t a t i o n speed v a r i e d almost l i n e a r l y w i t h Mach number, as shown i n F i g u r e 10. Thus, t h e scaled r o t a - t i o n speed o f 7560 RPM c o u l d be achieved a t o n l y one Mach number. The b l a d e p i t c h was s e t a t 52.5" so t h a t t h i s r o t a t i o n speed was reached a t appoximately M = .90. F l u t t e r analyses performed p r i o r t o t h e t e s t i n d i - c a t e d t h a t t h e wing f l u t t e r speed was i n s e n s i t i v e t o p r o p f a n advance r a t i o . T h i s b e i n g t h e case, i t was n o t w o r t h w h i l e t o t e s t each c o n f i g - u r a t i o n a t s e v e r a l d i f f e r e n t blade p i t c h angles i n o r d e r t o achieve t h e d e s i g n r o t a t i o n speed a t s e v e r a l d i f f e r e n t Mach numbers.
_-
SUMMARY OF PROBLEMS AND CORRECTIVE ACTION
CABLE MOUNT STAB I LITY
A model pitch-plunge instability was encountered when the snubber system was engaged during the initial testing of the bare wing configuration.
The 2 . 6 to 2.8 Hz instability occurred at dynamic pressures of 2400 to 3100 N/m2 (50-65 l b / f t ' ) , with rear flying cable tensions ranging from 445 to 890N (100 to 200 l b ) . The instability was eliminated by moving the model center-of-gravity froward 0.0761~1 (3 in.) by adding 2.96 kg to the nose ballast. The same nose ballast increment was used for the single and twin propfan configuration tests. A l s o , the snubber cable pulleys were moved forward by 0.23111 (9 i n . ) .
ROLL TRIM
Difficulties were experienced in maintaining rsll t r i m each time the wing confjguration was changed. The roll trim flap located on the right side of the wing was less effective than expected and was inadequate to com- pensate for the rolling moments caused by the lift unsymmetry.
The problem was alleviated by: ( 1 ) shimming t h e outer wing sections to reduce the lift unsymmetry, and ( 2 ) adding a chord extension to the trim tab to increase i t s effectiveness.
MODEL LOADS
stabilizer were higher The loads experienced by the wing and horizontal than predicted. Because the facility safety requirements prohibit testing with loads in excess of the maximum design values, this situation threatened to restrict the model test envelope. The problem was alle- viated in two ways. The wing root bending moments were reduced by shimming amount of washout to the untwisted the wing sections to introduce a small wing. The horizontal stabilizer design down load was increased by 35 percent, and a static test was performed to demonstrate a 50-percent margi n of safety.
TEST RESULTS
SUMMARY
A summary of the test configurations and results is given in Table 5.
For each test (configuration) the run and tab point numbers, wing con- figuration, fuel condition, prop rotation direction and wing boom configuration are given along with comments indicating the maximum dy- namic pressure reached and whether or not flutter occurred. Plots of dynamic pressure versus Mach number for each test are shown in Figures 1 1 through 25. For most configurations, tests were made at several total pressures. The variation o f dynamic pressure with Mach number for each of these passes is shown in the figures by dashed lines which lead to tab points indicating the maximum conditions reached. Where no flutter occurred, an open symbol and a tab point number are shown. Where flutter occurred, a solid symbol is used, and the flutter frequency is also in- ciicat-ec!. T a b l e 6 ! i s t s t h e Mach iiiiiiibei-, d;yilaiiiic pressure, veiocity,
density, Reynolds number, prop RPM and model response description for
each tab point shown i n the figures.
DESIGN VERIFICATION TESTS
Tests 1 through 9 and 13 were made to substantiate the predicted flutter safety of the single and twin propfan testbed aircraft preliminary de- signs. Nominal 1134 kg (2500 lb) balance booms and/or 136 kg (300 lb) flutter stabilizing booms were simulated for all except tests 9 and 13, i n which the flutter booms were removed. The minimum (12f%) fuel condi- tion was simulated for all except tests 4 through 6, in which the most flutter critical (by analysis) intermediate fuel condition of 1250 kg (10,000 lb) was simulated. No flutter or near flutter conditions oc- curred for these tests, including those made without flutter booms.
Test 6 simulated a severe failure condition in which the torquemeter and struts connecting the propfan gearbox to the engine power section were assumed to have failed. Because it represented a failure condition, this test was intended to reach only the V D boundary. The test was terminated at .97 V D when a spring representing the lower gearbox (engine) mount buckled and allowed the gearbox and prop to pitch down several degrees.
Because the strength of the full scale mount was not simulated on the model, this failure did not imply that a similar failure would be likely on the aircraft. No wing or whirl flutter conditions were observed on this test prior to the mount failure.
ANALYSIS VALIDATION TESTS
Tests 10 through 12, 14, and 15 were made with destabilizing wing tip booms to obtain data for validation of the method being used for the air- craft flutter analysis, Three destabilizing boom configurations were tested on the twin propfan configuration in an attempt to induce a flut- ter instability within the model test envelope. They were: (1) the nominal 136 kg (300 lb) flutter booms installed in reverse (aft slung): ( 2 ) the same booms modified by increasing their mass to simulate 235 kg (519 l b ) , and ( 3 ) l a r g e d e s t a b i l i z i n g booms s i m u l a t i n g 264 kg (582 l b ) each and having a p p r o x i m a t e l y t h r e e t i m e s t h e c e n t r o i d a l p i t c h i n g moment o f i n e r t i a o f t h e f l u t t e r booms. Mass d a t a f o r t h e boom c o n f i g u r a t i o n s a r e g i v e n i n Table 17.
A l t h o u g h h i g h l e v e l s o f dynamic response were observed w i t h a l l t h r e e d e s t a b i l i z i n g boom c o n f i g u r a t i o n s , s u s t a i n e d f l u t t e r o c c u r r e d o n l y w i t h t h e l a r g e d e s t a b i l i z i n g booms. F l u t t e r o c c u r r e d i n a 16 Hz symmetric wing bending, t o r s i o n mode a t M = .77 and 8857 N / m 2 (185 l b / f t 2 ) dynamic p r e s s u r e ( F i g u r e 22). When t h e same boom was t e s t e d on t h e s i n g l e prop- f a n c o n f i g u r a t i o n ( t e s t 1 4 ) , f l u t t e r o c c u r r e d i n a 1 5 Hz unsymmetric mode a t M = .79 and 9193 N/m' (192 l b / f t 2 ) ( F i g u r e 24). The same c o n f i g - u r a t i o n w i t h a w i n d m i l l i n g p r o p f a n i n s t e a d o f a n o n r o t a t i n g s p i n n e r was v e r y c l o s e t o s u s t a i n e d f l u t t e r a t M = .77 and 9000 N / m 2 (188 l b / f t ' ) when t h e t e s t was t e r m i n a t e d t o a v o i d unnecessary r i s k o f model damage ( F i g u r e 25).
DISCUSSION OF RESULTS
MODEL FLUTTER ANALYSIS
Post-test flutter analyses of the flutter model were performed for cor- relation with the data from tests 12, 14, and 15. The analysis method was the same as that being used for the full-scale testbed aircraft.
The lumped mass, beam-type structural representation utilized actual model mass and stiffness data and was adjusted to achieve good normal mode correlation with resonance test data.
A comparison o f the caculated and measured normal mode frequencies is
given in Table 7 for the twin propfan configuration and i n Table 8 for the single propfan configuration with nominal booms. The agreement was within 3 percent for all except two single propfan configuration modes which differed by 5 and 6 percent. The measured modes are further de- scribed in Figures 42, 43, and 44 of the Appendix, The unsteady aerodyamic derivatives used i n the analyses were predicted with a subsonic doublet lattice method. The wing derivatives were ad- justed to obtain agreement with the steady lift-curve-slope variation
with Mach number which was measured on ths Gulfstream I ! serodynamic
model. Doublet lattice aerodynamic derivatives were also used on the vertical and horizontal stabilizer surfaces. Quasi-steady propfan aero- dynamic derivatives calculated by Hamilton Standard were used for correlations with test 15.
The unsymmetry of the single propfan configuration required the represen- tation of the complete model (both sides) i n the analysis. Although the twin propfan configuration was symmetrical and could have been represented by a half-model, a complete model was also used for it for expediency.
The resulting math models contained 49 (single) and 52 (twin) component modes which were used to compute 35 normal modes for the flutter analyses.
Two percent structural damping (g = .02) was used for all modes.
ANALYSIS - TEST DATA CORRELATION
The flutter boundaries predicted for the twin and single propfan model configurations with large destabilizing booms are shown i n Figures 27, 28, and 29. A l s o shown in the figures are the data from tests 12, 13, and 15. Excellent correlation with the test data was achieved for all three cases. The flutter frequency was somewhat overpredicted for the twin propfan configuration, but was very close for the single propfan configuration. Because only one flutter (or near-flutter) test point was obtained for each case, the predicted flutter boundaries could .not be fully validated. However, the lack of flutter at tab point 376 (test 12) suggests that the shape of the predicted boundary for the twin propfan configuration is reasonable.
Although the predicted single propfan flutter boundary is higher than that of the twin propfan at Mach numbers below 0.8, it has a more pro- nounced "dip" and is about 7 percent lower in dynamic pressure (3.5% lower i n speed) a t M = 0.865. T h i s was unexpected, because t h e f l u t t e r speed f o r an unsymmetrical c o n f i g u r a t i o n i s g e n e r a l l y h i g h e r t h a n t h a t o f a summetrical c o n f i g u r a t i o n r e p r e s e n t i n g i t s more c r i t i c a l wing.
The more pronounced d i p seen h e r e was caused by t h e g r e a t e r s e n s i t i v i t y o f t h e s i n g l e p r o p f a n f l u t t e r i n s t a b i l i t y t o t h e d e n s i t y v a r i a t on w i t h Mach number i n t h e TDT.
on w i t h The p r e d i c t e d f l u t t e r boundary f o r t h e s i n g l e p r o p f a n c o n f i g u r a t a w i n d m i l l i n g p r o p f a n was n e g l i g i b l y h i g h e r t h a n t h a t w i t h o u t p r o p f a n e f f e c t s . This i n s e n s i t i v i t y t o p r o p aerodynamic and g y r o s c o p i c e f f e c t s i s c o n s i s t e n t w i t h t h e t e s t d a t a and o b s e r v a t i o n s o f s u b - c r i t i c a l response made d u r i n g the t e s t s .
AIRCRAFT DESIGN VERIFICATION No f l u t t e r or n e a r - f l u t t e r c o n d i t i o n s o c c u r r e d i n t h e t e s t s r e p r e s e n t i n g t h e a i r c r a f t ( p r e l i m i n a r y ) d e s i g n c o n f i g u r a t i o n s , i n c l u d i n g t h o s e r e p r e - s e n t i n g a propfan p o w e r p l a n t f a i l u r e and t h e l o s s o f t h e f l u t t e r boom.
These r e s u l t s a r e c o n s i s t e n t w i t h t h e a i r c r a f t p r e l i m i n a r y f l u t t e r a n a l - yses, which p r e d i c t f l u t t e r boundaries above 1.2 VD1 e x c e p t f o r t h e cases w i t h o u t f l u t t e r booms. I n t h e s e cases, f l u t t e r i s p r e d i c t e d t o o c c u r w i t h i n t h e a i r c r a f t f l i g h t t e s t envelope. However, t h e s t a b i l i t y ( n e t damping) of t h e p r e d i c t e d f l u t t e r mode v a r i e s v e r y g r a d u a l l y w i t h a i r - speed, and small changes i n damping r e s u l t i n l a r g e f l u t t e r speed changes.
A l s o , t h e s t a b i l i t y o f t h i s mode has been shown by a n a l y s i s t o be s e n s i - t i v e t o o t h e r parameters, i n c l u d i n g o u t e r wing mass and p r o p f a n n a c e l l e f l e x i b i 1 i t y .
The model d i f f e r s f r o m t h e c u r r e n t ( s c a l e d ) a i r c r a f t d e s i g n i n s e v e r a l r e s p e c t s . The most c r i t i c a l wing f u e l c o n d i t i o n s a r e empty and 4535 kg (10,000 l b ) . No f u e l i s l o c a t e d i n t h e o u t e r wing f o r e i t h e r o f t h e s e c o n d i t i o n s . The model wing c o u l d n o t be made l i g h t enough t o r e p r e s e n t t h e empty f u e l c o n d i t i o n , i n s p i t e o f i t s e f f i c i e n t t h i n w a l l aluminum spar. The excess mass i s d i s t r i b u t e d across t h e span and i s e q u i v a l e n t t o 1 2 i p e r c e n t of t h e G u l f s t r e a m I 1 c a p a c i t y f u e l mass, o r about 1135 kg (2500 l b ) . The a i r c r a f t f l u t t e r analyses i n d i c a t e t h a t i n c r e a s e d o u t e r wing mass increases t h e s t a b i l i t y o f t h e s e n s i t i v e boom-off f l u t t e r mode.
The model propfan n a c e l l e f l e x i b i l i t i e s a l s o d i f f e r from t h e ( s c a l e d ) a i r c r a f t design. The f r e q u e n c i e s o f t h e model e n g i n e - n a c e l l e normal modes a r e g e n e r a l l y l o w e r t h a n t h e s c a l e d ( p r e d i c t e d ) a i r c r a f t modes, which i n d i c a t e s t h a t t h e model engine mount and/or n a c e l l e s t r u c t u r a l f l e x i b i l i t i e s a r e g r e a t e r t h a n d e s i r e d . A i r c r a f t f l u t t e r analyses i n d i - c a t e t h a t increased n a c e l l e f l e x i b i l i t y i n c r e a s e s t h e s t a b i l i t y of t h e s e n s i t i v e boom-off wing f l u t t e r mode.
A t h i r d d i f f e r e n c e between t h e model and t h e ( s c a l e d ) a i r c r a f t i s t h e nose b a l l a s t added t o t h e model f o r c a b l e mount dynamic s t a b i l i t y . The b a l l a s t represents 3227 kg (7100 l b ) on t h e a i r c r a f t . I t s e f f e c t on t h e c r i t i c a l f l u t t e r mode was n o t s e p a r a t e l y determined.
An analysis was performed to determine the net e f f e c t of these dif- the s t a b i l i t y of the single propfan model without a f l u t t e r ferences on
boom. T h e predicted f l u t t e r speed was well above 1.2 Vo, which i s con-
s i s t e n t w i t h the t e s t resultsr b u t indicates t h a t the model i s more s t a b l e t h a n the a i r c r a f t f o r t h i s configuration. Therefore, the absence of f l u t t e r i n the t e s t of t h i s configuration does not indicate t h a t the f l u t t e r s t a b i l i z i n g boom i s unnecessary f o r the testbed a i r c r a f t .
T h e e f f e c t s of these differences on the f l u t t e r s t a b i l i t y o f the other t e s t configurations a r e probably less significant, b u t were not deter- mined by comparative f l u t t e r analyses. Therefore, the t e s t r e s u l t s alone cannot be used t o verify t h a t the f i n a l a i r c r a f t design w i l l have adequate f l u t t e r safety margins. They do indicate, however, t h a t no serious f l u t t e r i n s t a b i l i t i e s of the a i r c r a f t have been overlooked i n the analysis.
CONCLUSIONS AND RECOMMENDATIONS
The r e s u l t s o f t h e design v e r i f i c a t i o n t e s t s g e n e r a l l y c o n f i r m t h e pre- d i c t e d wing and w h i r l f l u t t e r s t a b i l i t y o f t h e p r e l i m i n a r y d e s i g n c o n f i g u r a t i o n s tested. However, t h e model t e s t s d i d n o t a c c u r a t e l y sim- u l a t e t h e most c r i t i c a l c u r r e n t c o n f i g u r a t i o n s , and t h e d i f f e r e n c e s were shown by a n a l y s i s t o i n c r e a s e the s t a b i l i t y and f l u t t e r speed o f t h e model r e l a t i v e t o t h e c u r r e n t a i r c r a f t design, a t l e a s t f o r t h e s i n g l e p r o p f a n c o n f i g u r a t i o n w i t h o u t a f l u t t e r boom.
T h i s leads t o t h e f o l l o w i n g conclusions: o No unexpected f l u t t e r i n s t a b i l i t i e s a r e caused by t h e r o t a t i n g p r o p f a n o r t h e unsymmetry o f t h e s i n g l e p r o p f a n t e s t b e d c o n f i g - u r a t i o n .
0 The wing f l u t t e r s t s h i ! i t y i s e s s e n t i a l ! y t h e same w i t h rntatin9 p r o p f a n s as w i t h e q u i v a l e n t w e i g h t n o n - r o t a t i n g spinners.
o The f l u t t e r s t a b i l i t y o f t h e symmetrical t w i n p r o p f a n c o n f i g - u r a t i o n i s a p p r o x i m a t e l y t h e same as t h a t o f t h e unsymmetrical s f n g l e propfan c o n f i g u r a t i o n .
It i s a l s o reasonable t o conclude f r o m t h e model a n a l y s i s / t e s t d a t a cor- r e l a t i o n t h a t t h e f l u t t e r a n a l y s i s methods b e i n g used i n t h e a i r c r a f t d e s i g n a r e capable o f a c c u r a t e l y p r e d i c t i n g i t s wing f l u t t e r charac- t e r i s t i c s , i n c l u d i n g t h e e f f e c t s o f mass and s t i f f n e s s unsymrnetry and p r o p f a n aerodynamic and gyroscopic c o u p l i n g . The v e r i f i c a t i o n o f t h e a i r c r a f t f l u t t e r s a f e t y margins would, t h e r e f o r e , be based on t e s t v a l i d a t e d a n a l y t i c a l p r e d i c t i o n s r a t h e r than t h e model t e s t r e s u l t s .
The s t r u c t u r a l r e p r e s e n t a t i o n used i n t h e a i r c r a f t f l u t t e r a n a l y s i s should be v a l i d a t e d by normal mode c o r r e l a t i o n s w i t h ground v i b r a t i o n t e s t data.
B L FS 0 . 0 0 BL 0.00 BL
7 . 2 1 3 m ( 2 8 4 . 0 in.) --:-\I
10 e 287m ( 4 0 5 . 0 in.)
~ _ _ _ _ _ _ _ _ _ __ ____ 21.971rn 0.00- - ’ - FS ( 8 6 4 . 0 in.)
0.00
B i 0.00
0.00 __._ Modified Gulfstream I I Propfan Testbed Aircraft Figure 1 , - .
-
Q) U E lu eo Q) n
-
L Q)
>
> N C w r - c W a 0 ' O Z I U
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S/W - a33dSliIV I N 3 l V A I n b 3
I I I I I 0 0 0 0
0 0 -
0 0 I n m N
S I O N X - 0 3 3 d S t l l V I N 3 l V A l n b 3
Figure 3 . Bare Wing Model in NASA Langley Transonic Dynamics Tunnel Figure 4 . Twin Propfan Model with Windmilling Props I b) Single Propfan Model with ( a ) Windmilling Prop Figure 5 .
(b) Weighted Spinner 0 . 9 1 4 m ( 3 6 . 0 0 in.)
FS 0.851m ( 3 3 . 4 9 in.)
FS FS 0.023111 2.438m ( 0 . 8 9 in.)
( 9 6 . 0 0 in.)
2.337m ( 9 2 . 0 0 in.)
I
-
1 3 I
I I BL
I I 0.467m
'b , , - e , , , ( A 8 . 3 8 in.)
. 2 6 in.)
( 8 111 c 7m in.)
- . .- - - Figure 6 . Single Propfan Model General Arrangment Model Propfan Powerplant Simulation Figure 7.
C
.-
m m In Y E =r In M
-
- I
i
I- v) /
t
c
m
.-
L i
\
1 . 2 VD I - LL \ m -I w 120 2 0 . 2 . 3 . 4 . s .6 . 7 . 8 .9 MACH NO.
Figure 9. Model Test Envelope
-
-
w 6000 -
I- Z
-
E
-
U w a .
v) Z
4000 -
-
I- z,
>
w
/
U
30001
2ooi
0 1 I I I I I I I I I . 8 1 . O I . 2 .4 . 6 MACH NUMBER - -..
Figure 10. Model Propfan Windmilling Speed for Pitch Angle 52.5' Blade T A B P T 69
P
1 .2VD / T A B P T 37
/ / P
-
. 2 . 3 .4 . 5 . 6 .7 . 8 .9 MACH NO.
Figure 1 1 . Test Conditions for Test Number 1 2 O( TAB PT 93 18(1 N
c
L L \ m 120 -I
W ' 1
az ui v) W az a
u
-
E U
z
>.
P Figure 1 2 .
Test Conditions f o r Test Number 2 T A B PT 135 TAB PT 143 1 . 2 ~ ~
/
/ /
/
12c 0 - 0 . 2 : 3 . 4 . 5 . 6 . 7 . 8 .9 M A C H NO.
Figure 1 3 , T e s t Conditions for T e s t Number 3
-
TAB PT 178 1 . 2 ~ ~ T A B PT 167
-
F
I
/'
-
/
/
1 4 0 - m
-
X > 6
- 2
W ct 3 s 1 0 0 - W w
-
a - r .
E
z
60-
40 -
20 -
o b .2 . 3 . 4 . 5 . 6 . 7 . 8 .9 MACH NO.
Figure 14, T e s t Conditions for T e s t Number 4 B PT
-
. 2 . 3 .4 . 5 .6 . 7 .8 . 9 MACH NO.
Figure 15. Test Conditions for Test Number 5 16( N
k 14(
\ m -0- I --\ 12( T A B PT p?
2 2 1 , o v) u,
/
w 100 p?
/
L W
/
5 80
a
z
>.
P 2 0 -0 . 2 . 3 . 4 . 5 . 6 . 7 . a .9 MACH NO.
Figure 16. T e s t Conditions for T e s t N u m b e r 6 a m N I- L L r4 x ' 6 \ a 12c E A \
z
I T A B PT W W 5 o?
' 231
5 I00
cn cn VI VI W W e
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a 80 W W
-
E I
a
a
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z 60
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n n - 1 2 ( ( -0 .2 . 3 .4 .5 . 6 .7 .8 .9 MACH NO.
Test Conditions for Test Number 7 Figure 17.
n c U X
--
T A B P l 2 4 1 /
2L 0 I I
" I I . 2 . 3 .4 . s . 6 . 7 . 8 . 9 MACH NO.
Figure 18.
Test Conditions for Test Number 8
2oo r
T A B PT 1 . 2 V D I T A B P T b 265 . 2 . 3 .4 .5 .6 . 7 . 8 . 9 MACH NO.
Figure 1 9 . T e s t Conditions for T e s t Number 9
2 o o r
!
1 .2VD
/
’ T A B PT
. 2 . 3 .4 . s - 6 . 7 . 8 . 9 MACH NO.
Figure 20.
Test Conditions for Test Number 10 1 . 2 V D T A B PT T A B PT N 140 n I- c LL
- X
H
rn
E -I 120 \ I Z T A B PT w w l Y Q 1 326
2 100
m m m w w C Y Q a a W U
c -
E I
a a
Z Z / > >. 60 n P
/ #'
4 ( 2( I I I 1 1 ( .2 .3 .4 . 5 . 6 . 7 . 8 .9 M A C H NO, T e s t Conditions for T e s t Number 1 1 Figure 21.
20c 1 6 H z SYM 18( TAB PT a 16( 14( c X q q n N I L U E \
z
W 5 &
1 oa
L!
VI W & a .
a0 W
-
E
a
z
>.
a
-
0 0 ;2 . 3 . 4 . 5 . 6 . 7 . 8 .9 MACH NO.
Figure 2 2 .
Test Conditions for T e s t N u m b e r 12
-
T A B PT 423 9 '
, l.7r7
180 -
8 -
160 -
7 -
140 -
m 6 - N
t 120- -
X \ N rn E -I \
z 5 -
' 1 0 0 -
W c1: a !
v) v) u , 4 - 8 0 - w a !
p?
a n . 2 . 3 .4 .5 .6 . 7 .8 . 9 MACH NO.
Figure 23. Test Conditions for Test Number 13 TAB P T 435 FLUTTER l 5 H z UNSYM
2 o o r
-
-
w 1 4 0 - , . , I - 0 c LL \ X N m E
-I 1 2 0 - -
K p!
a n 8 O - u
u
-
-
E
a
z
z
> 6 0 - n
-
2oL 0
. 2 . 3 . 4 . 5 . 6 . 7 . 8 . 9 M A C H NO.
Figure 2 4 , Test Conditions f o r Test Number 14 T A B P T 447 NEAR FLUTTER 1 5 Hz UNSYM
1.2VD /”
\
/
I
N I - L \ m 1 2 0 J I t u E 100 3 ) u , VI w
u
p /
-
I
a
z 60
/ >.
/ P
/ 0 0 /
- 0 2(1 t I I I I I I I C . 2 . 3 . 4 .5 . 6 . 7 . 8 .9 MACH NO.
Test Conditions for Test Number 15 Figure 2 5 .
C A LC U L A T ED 18-19Hz SYM MODE TEST_ 12 T A B P T 389
20( \
16Hz SYM MODE 18( 16( (Y 14C i- U \ m -I 120 I w o !
2 100
ul w o !
n
/
W 80
/
-
/
z
2 60
. 2 . 3 .4 . 5 . 6 . 7 . 8 . 9 MACH NO.
Figure 26.
Calculated versus Measured Flutter Boundary -
Twin Propfan Configuration with Destabilizing Booms
k 49
C A LC U LA T ED 15-16Hz UNSYM MODE
TEST 14 TAB P T 435 \
FLUTTER 15Hz UNSYM
\
I / n N 140 I - c L L X \ rn “ E 6 \
z
I w W a : p?
3 5 rr) m 100 m u , W w a : L U 80 U
z
. 2 . 3 .4 .5 .6 . 7 .8 .9 M A C H NO.
Calculated versus Measured Flutter Boundary - Single Propfan
Figure 27, Configuration with Destabilizing Booms and Spinner CALCULATED 1 15-16 H Z UNSYM MODE-
\
l o r
TEST 15 T A B P T 447
\
NEAR FLUTTER 1s”; UNSYM 7 ; . \ \ ; 1 .2VD .2 . 3 .4 . 5 . 6 .7 . 8 . 9 MACH NO.
Figure 28, Calculated versus Measured Flutter Boundary -
Single Propfan Configuration with Destabilizing Boom and Propfan TABLE 1.
MODEL DESIGN SCALES QUANTITY SYMBOL SCALE Geometry bm/ba 1 / 9 Mach Number k / M a 1/1 Density PmlPa 111 Ye l o c i t y VmIVa 1/2.02 Dynamic Pressure w / q a 1/4.08 Frequency wm/wa 4.455411 Weight (Mass) h / W a 1/729 Mass Moment o f I n e r t i a I m / I a 1/59.049 Stiffness (Beam) EImIEIa, GJmIGJa 1126.772 I TABLE 2.
INSTRUMENTATION FOR BARE WING CONFIGURATION PEN RECORDER CHANNELS SEISED QUANTITY CHANNEL
1 Left wing root vertical bending moment - static
2 Right wing root vertical bending m n e n t - static
3 Left stabilizer root vertical bending moment - static
4 Left jet engine pylon vertical load - static
5 F i r ! !nter;! bandSng imaeiit - s t a t i c
6 Aft fuselage (FS 1.93 r75.81) vertical bending moment - static
7 Aft fuselage (FS 1.93 r75.81) lateral bending moment - static
8 Fuselage nose vertical acceleration - dynamic
9 Left wing root vertical bending moment - dynamic
1 0 Right wing root vertical bending moment - dynamic
1 1 Left wing root torsion moment - dynamic
12 Right wing root torsion moment - dynamic
13 Fin root lateral bending moment - dynamic
14 Left stabilizer root vertical bending moment - dynamic
15 Aft fuselage (FS 1.93 [75.8]) torsion moment - dynamic
16 Fuselage nose lateral acceleration - dynamic
TAPE RECORDER CHANNELS CHANNEL SENSED QUANTITY
1 Left wing root vertical bending moment - dynamic
2 Right wing root vetical bending moment - dynamic
Left wing root torsion moment - dynamic
4 Right wing root torsion moment - dynamic
5 Fin root lateral bending moment - dynamic
6 Left stabilizer root vertical bending moment - dynamic
7 Aft fuselage (FS 1.93 E75.81) torsion moment - dynamic
8 Fuselage nose lateral acceleration - dynamic
9 Fuselage nose vertical acceleration - dynamic
1 0 Aft fuselage (FS 1.93 r75.81) vertical bending moment - dynamic
1 1
Aft fuselage (FS 1.93 [75.8]) lateral bending moment - dynamic
12 Left wing root vertical bending moment - static
13 Right wing root vertical bending ronent - static
14 Tina code TABLE 3.
INSTRUMENTATION FOR SINGLE PROPFAN CONFIGURATION PEN RECORDER CHANNELS SENSED QUANTITY CHANNEL
1 Left wing root vertical bending moment - static
Right wing root vertical bending moment - static
Propfan nacelle vertical bending moment - static
4 - static
Propfan nacelle lateral bending moment
Fin root lateral bending moment - static
6 Aft fuselage (FS 1.93 [75.8]) vertical bending moment - static
7 Aft fuselage (FS 1.93 175.81) lateral bending moment - static
8 Left jet engine pylon vertical bending moment - static
9 Left wing root vertical bending moment - dynamic
10 Right wing root vertical bending moment - dynamic
1 1 Left wing root torsion moment - dynamic
12 Right wing root torsion moment - dynamic
Fin root lateral bending moment - dynamic
14 Aft fuselage (FS 1.93 [75.8]) vertical bending moment - dynamic
15 Propfan gearbox lateral acceleration - dynamic
16 Propfan gearbox vertical acceleration - dynamic
TAPE RECORDER CHANNELS CHANNEL SENSED QUANTITY
Left wing root vertical bending moment - dynamic
Right wing root vertical bending moment - dynamic
3 Left wing root torsion moment - dynamic
4 Right wing root torsion moment - dynamic
Fin root lateral bending moment - dynamic
6 Propfan rotation speed
7 Aft fuselage (FS 1.93 [75.8]) torsion moment - dynamic
8 Propfan nacelle vertical bending moment - dynamic
9 Propfan nacelle lateral bending moment - dynamic
1 0 Aft fuselage ( F S 1.93 [75.8]) vertical bending moment - dynamic
1 1 Aft fuselage (FS 1.93 [75.8]) lateral bending M m e n t - dynamic
Left wing root vertical bending moment - static
Right wing root vertical bending moment - static
Time code TABLE 4.
INSTRUMENTATION FOR TWIN PROPFAN CONFIGURATION PEN RECOROER CHANNELS SENSED QUANTITY CHANNEL
L e f t wing r o o t v e r t i c a l bending moment - s t a t i c
2 Right wing r o o t v e r t i c a l bending moment - s t a t i c
3 L e f t propfan n a c e l l e v e r t i c a l bending moment - s t a t i c
4 L e f t propfan n a c e l l e l a t e r a l bending moment - s t a t i c
L e f t wino rant tortinn m m n t - c t r t i c
A f t fuselage (FS 1.93 (75.81) v e r t i c a l bending moment s t a t i c
7 A f t fuselage (FS 1.93 [75.8]) l a t e r a l bending moment - s t a t i c
8 L e f t j e t engine pylon v e r t i c a l bending moment - s t a t i c
9 L e f t wing r o o t v e r t i c a l bending moment - dynamic
10 Right wing r o o t v e r t i c a l bending moment - dynamic
11 L e f t wing root t o r s i o n rnonent - dynamic
12 Right wing r o o t t o r s i o n moment - dynamic
13 A f t fuselage (FS 1.93 [75.8]) l a t e r a l bending moment - dynamic
14 A f t fuselage (FS 1.93 [75.8]) v e r t i c a l bending moment - dynamic
Right propfan gearbox v e r t i c a l acceleration - dynamic
16 L e f t propfan gearbox v e r t i c a l acceleration - dynamic
TAPE RECORDER CHANNELS SENSED QUANTITY CHANNEL 1 L e f t wing r o o t v e r t i c a l bending - dynamic
2 Right wing r o o t v e r t i c a l bending - dynamic
L e f t wing r o o t t o r s i o n nmment - dynamic
4 Right wing r o o t t o r s i o n moment - dynamic
Right propfan r o t a t i o n speed 6 Left propfan r o t a t i o n speed 7 A f t fuselage (FS 1.93 [75.8]) t o r s i o n
8 L e f t propfan n a c e l l e v e r t i c a l bending - dynamic
9 L e f t propfan nacelle l a t e r a l bending - dynamic
10 A f t fuselage (FS 1.93 [75.8]) v e r t i c a l bending moment - dynamic
'
11 A f t fuselage (FS 1.93 I75.81) l a t e r a l bending moment - dynamic
L e f t propfan gearbox v e r t i c a l acceleration - dynamic
Right propfan gearbox v e r t i c a l acceleration - dynamic
14 Time code I L L n E P % \ z
c -
c -
E E \ z \ 2
B
P
I
a
ie"
c P Y Y Y Y rl c 4 L P P c
-
I- a I c I-
E
-
0,
E
;-" 10
B
i i i
$
: : : I a a a ID 4 e I. 4 D
[ c Y 4
Y Y Y Y Y Y Y Y I.
4 L Y Y r( L r( L 4 L 4 b.
P P P P TABLE 7.
COMPARISON OF CALCULATED AND MEASURED NORMAL MODE FREQUENCIES -
TWIN PROPFAN CONFIGURATION SYMMETRIC MODES C a l c u l a t e d Measured Cal.c/Meas.
Freq. Mode D e s c r i p t i o n
Freq. (Hz) Freq. (Hzl
14.01 1 .oo Wing 1 s t v e r t i c a l bending
14.04
21.50 21.48 1 .oo Wing 1 s t t o r s i o n , p r o p f a n n a c e l l e
v e r t i c a l bending 1.01 Propfan n a c e l l e l a t e r a l bending 22.96 22.70 0.98 Fuselage v e r t . bend., j e t engine 28.88 29.38 v e r t i c a l
30.47 30.50 1 .oo Propfan powerplant yawinq
31.50 0.99 Propfan powerplant p i t c h i n g 31.34 36.99 36.28 1.02 J e t engines v e r t i c a l ANTISYMMETRIC MODES Cal c u l ated Measured Ca 1 c/Meas.
Freq. (Hz) Freq. Mode D e s c r i p t i o n Freq. (Hz) 14.94 15.10 .99 A f t f u s e l a g e l a t e r a l bending, t o r s i o n .98 Wing 1 s t v e r t i c a l bending, a f t 17.65 18.09 f u s e l a g e l a t e r a l bending 22.42 22.70 .99 Propfan n a c e l l e l a t e r a l bending 24.57 24.41 1.01 J e t engine v e r t i c a l Propfan powerplant p i t c h i n g , j e t 28.82 engi ne v e r t i c a l 30.31 30.64 .99 Propfan powerplant yawing 32.11 A f t f u s e l a g e t o r s i o n , f i n t o r s i o n 34.35 33.85 1.01 F i n t o r s i o n , fwd. f u s e l a g e l a t e r a l bending 39.03 Propfan powerplant p i t c h i n g , wind 2nd t o r s i o n 39.20 40.50 .97 Fwd. f u s e l a g e l a t e r a l bend., f i n 1 a t e r a l bending c L
-
Y m t m u .r c Y U al w .r .r 4 c L VI C t L al m > c t * .C I - n c m 4 d rc U Y E C a a c c m .I- .r al U L m VI U u .
n 3 L .C LL c 0 c m 3 : I- m
s
u Y Y m z 4 .r c -c al m 'c .C 3 a Y L W 2 W .C I - al -I 01 3 c > m 9) 5 m m c , a c 01 rc u u - .r m c 0) m v) Y y.
.r c , -I a w u W OI al -c u a c .TI -1 c , al .r Y w > Y Y 0 Y c w
P C r
d al CI c , m al .C m
> 0 4 -
v) c al U al > VI al
PI
I al e 3 1 - n a .r Y r c
-
OI Y Y VI 0 I- 0 al L 0 c VI al 0 LL C w c u 0 u 0 0 m > m m m c , n 0- z z F m al .r .r a l
-
L C L 3 5 .C C C 3 : 3 OI m 0 m m 3 : m 'c icI c
-
I- Y Y Q- y . c , Y a c -c 4 a a a 0 c -c C m m c m CII r
.r .- L L e .- .r
al L 3 L Q nz n n 4. U . a dc a u
APPENDIX - MODEL B A S I C DATA
Model geometric, mass, s t i f f n e s s , and resonance d a t a a r e presented i n F i g u r e s 29 t h r o u g h 44 and Tables 9 t h r o u g h 2 2 . The model " a c t u a l " d a t a were determined almost e n t i r e l y by t e s t s . The ''desired" d a t a r e p r e s e n t as o f June 1985. The d e s i r e d wing mass t h e a i r c r a f t p r e l i m i n a r y d e s i g n o f t h e Gulfstream I 1 c a p a c i t y f u e l d i s t r i b u t i o n d a t a i n c l u d e 1 2 i p e r c e n t because t h e model wings c o u l d n o t be made l i g h t enough t o s i m u l a t e t h e empty a i r c r a f t wings. I n most cases, t h e " a c t u a l " d a t a agree v e r y c l o s e l y w i t h t h e d e s i r e d c h a r a c t e r i s t i c s . Some o f t h e a c t u a l spar s t i f f n e s s d a t a d i f f e r s i g n i f i c a n t l y f r o m t h e d e s i r e d d a t a o v e r s h o r t d i s t a n c e s . However, t h e n e t e f f e c t o f t h e s e l o c a l d i f f e r e n c e s on t h e model normal modes i s much l e s s s i g n i f i c a n t because o f t h e e f f e c t i v e I' i nt e g r a t i on s It i nvo 1ved .
Measured normal modes f o r t h e bare wing, t w i n propfan, and s i n g l e p r o p f a n c o n f i g u r a t i o n s a r e presented i n Figilrps 42; 41, 2nd 44, r e s p e c t i v e l y .
The r e s o n a n t frequency and node ( z e r o m o t i o n ) l i n e s f o r each mode a r e shown on an i s o m e t r i c view o f t h e model.
1 . 5 8 4 FS( 6 2 . 3 8 ) 1 . 7 1 9 1 . 5 0 3 F S ( 6 7 . 6 6 ) F S ( 5 9 . 1 7 ) W S 0 . 5 6 4 4 W S 0 . 4 0 6 4 \ \ +.--_, \
B L O ~ --
1 I
. 8 7 0 1 . 0 8 4 1 . 4 3 4 FS ( 5 6 . 4 7 ) FS ( 4 2 . 6 9 ) F S ( 3 4 . 2 5 ) F i g u r e 29. Wing G e o m e t r y W ODm
m z m I
u . O D o n O D 0 - 0 - u l , .
u . b 0 - ul u .
ID- ?:
u
o\o m N cc) E urn
<
L L - W Lux m
e a
CL
a
W
\
Figure 32. Horizontal Stabilizer Geometry 6.0 \ ---DESIRED \
N \ 0 A C T U A L - R I G H T
z
-
N
a A C T U A L - LEFT
I
b
E m E A I -J Z \ I I \ L n W I c 0
k
-
X 1 .( x
-
-
w w I I \ v) v) v) v) w
b
w z \ z U U U
-
!A
-
I- I- v) v) \ U
u
z
- z
b
-
n
a
z 9
z
W w \ m m \ -I - J U \ U W
-
W
-
\
I- I - o? \ W w
>
> .o: . 2 .4 . 6 . 8 1 . o FRACTION OF SEMISPAN Figure 33. Bare Wing Vertical Bending Stiffness
- - D E S I R E D
0 A C T U A L - R I G H T
\
a A C T U A L - L E F T
\ \ 8 , N
z
\
-
I \ m -I \ I e l I c x I ul ul 1 . 0 - ; ul W - 2 ' L
z
, L u ,
-
L L
- - I -
\ ul I-
-
ul J
-I - a
a z
z 0 \
-
- - u l
\ a?
ul a ? 0 \ 0 - I - \ I- \ \
t
.03\ I I I I I .l-
0 . 2 .4 . 6 . 8 1 .o
F R A C T I O N O F SEMISPAN B a r e Wing Torsional Stiffness F i g u r e 3 4 , N N Z
-
I \ DESIRED
& 2 0 z
\ -I I ACTUAL R I G H T 0' I =t I
a ACTUAL LEFT
w
'El \
I c
- 10
X X , N
-
N w
-
W . I \ I . m m b , v) W
v) o \
' Z W LL z L L
. -
L I -
-
v, ! - v)
. u
z
W
-
\ Z n
-
z '0
n u
z
\ m W m l . C \ - I - ' L L I - \
L - a
- I
a \
n I \ ' 2 n
z - a
\ I
a
LLI I - a ?
w a 1 I I I I p?: 0 . l ' I LL 0 . a . 9 . 4 .6 . 8 1 .o L FRACTION OF SEMISPAN Figure 3 5 . Bare Wing Fore-and-Aft Stiffness 7 . 0 \
a
\ - - D E S I R E D \
0 A C T U A L - R I G H T
o ' a
A C T U A L - L E F T
\ \
b\
\
b
\ \
'4
I- 1 . 0 - v) \
-
q
- u
, z
\
-
, a
\
z
- - w
\
m
\
- - I .10
\
, u .
-
\ I- &
\
W \
- >
.03 .l- . 2 .4 . 6 . 8 1 . o FRACTION OF SEMISPAN Figure 36. Propfan Wing Vertical Bending Stiffness 6.0 \ \ --DESIRED
\ O A C T U A L - R I G H T
A C T U A L - L E F T
a \
10 [
\ \p N N
z
E &\
-
I I El
z
m -I \ I I 1 . o \ In I
b
c \ L \ E l x
‘,o
I
‘ a
I VI \ VI v)
v) 1 . 0 - 9
- g
W U \
z
‘ U L
\B
-
L L I -
-
\ - m I- El \ v) - - I
a
1 \
- z
a
\ Z .1
-
, V I v) \ c) e I- \ ! - \ I I 1 I I .o:
. 1 [
0 . 2 .4 . 6 . 8 1 .o
F R A C T I O N OF SEMISPAN F i g u r e 37. Propfan Wing Torsional Stiffness \ 10.0 --DESIRED N t \
z
-
0 ACTUAL - RIGHT
I m
A C T U A L - LEFT
-I I W I I
0 -
‘c X X i a N
-
N
-
W W I I VI \ VI VI VI W W
z
Z U U U \
1 .o
L L -
-
\a
I- I - VI VI U \
u
z
Z -
\
-
n
- a
\
z
z
u W m
g
m \ ! -
I- 1 .o
U LL
\
a
a
I \ I P n
z
\
z
a
a \
I W \ p!
.4 LL .1 0 . 2 . 4 . 6 . 8 1 .o FRACTION OF SEMISPAN Figure 38. Propfan Wing Fore-and-Aft Bending Stiffness
i
1 " .
I N / / /
I N
/
/ 6
.-
/ c, u , ul S
.-
u S Q) m
-
m I U
.-
Y L
j s r a
>
Q) I ul
I 0
m
-
I Q) I I 0 LL I Q, m
E
.-
I -
m I A - N
I =
Y ) o ~ o m o m o * = r = t m m N N r r O . . - . . . - . . . .
zu-N - s- 0 1 X ii
I I I I I I I I 1 r O = r N O O D i D ~ N O r r - C
N I - 8 1 - 9-OL X I3
Z n .
n a I In I- I O
-
I
I
/ 0 In m
0 ,c -
E
.-
t /' 0 In a 1 : t=i m
f
-
-0 c
o f
In I -12 I I 'In In \ ' N \ \ \ 0 In 0 In 0 In 0 m N N w c 0
I
z"'-N - s- 0 1 X 1 3
I
I
I I / I I
0 ,’
VI VI al
d C
f‘ I
-
I cp 0 C I
.-
I VI 1 0 I- I
f o
I I I .- a 0 m m 0 rn 0 N N
7-
p - ~ - O L x r3
I
S- ~ 1 I I I I I I I I I 1.
O m 0 0 r - U ) m a m N - O
-
N I - 8 1 - 9-OL X f3
B A R E m COh'FIGURATION 12.5 S D N E l K f C m X . T A ! l W N
Figure 42. Measured Normal Modes - Bare Wing Configuration
( 1 o f 2 )
a -
A
FIN OhTY SDRVEYED
Measured Normal Modes - Bare Wing Configuration ( 2 o f 2 )
Figure 4 2 .
Figure 43. Measured Normal Modes - Twin Propfan Configuration ( 1 o f 2 )
'IWLN PROPFAN CONFIGURPLPION 12.5 $ FUEL A W I ' 1 - C EXCITATION 18.09
Measured Normal Modes - Twin Propfan Configuration ( 2 of 2 )
Figure 4 3 .
SINGLE PROPPAN CONFIGURATION 12.5$ PliEL Figure 44.
Measured Normal Modes - Single Propfan Configuration
( 1 of 2 ) ( 2 o f 2 )
Measured Normal Modes - Single Propfan Configuration
F i g u r e 4 4 .
. . .
* . . . . . . . .
r c e . - . - N N . - r L 0 )
.I- -
8 0 B Q ) ' .C Q ) c n o m c 'C 0 .,- m Y v) m e c Q U c w 'C u c w u c c a n - .a cl - . a c . a m m
s 0
.r Y C L y . m 0 PI 'C I a m . . . . . . . . . .
e - c c N N r e m m N . . . . . e . . . .
e c r r - N N r r r
-
I - - w m m -I
P
7 c m m m
-
L 2 : A U t
e
-
Y = I
L al .I- C .C n c , v , e a u c
c
.C L I
-
c r . - c c . - c c I I I I I I I I I I I I I I Y E I I I I m c I l l 1 . . . .
I
I - -
¶
! ! 0
Y Y Y @ a 0 c 0 Y P'5 > 0 e a d Y a V .e
H .I
2 Y I a a . .
. . . . . . . . . . .
F e e l - - & & I I I I I I I I I I I I I C .C
Q
c d I- w
iw c c c
c c 4 4 O 3 3
. E y 2
UI 2 5 3 . .
Ln I-
s
z I .-. .. .. . .. , . - .
h A n n P m e m .. .o le.. b c c c x x * W c
P
. .
?
Y Y Y u) e- . .
X x * * W R
(v - N 5 . 8
. .
n * - m n n -lo c c W X w - 1 m
s
Y u 7 VI -- , I s ?
3 b
-? -.
. .
. ._ ..
w
i '
95:.
. .
. . - .
.”
t A t ; REFERENCES 1. Bradley, E. S . , e t al., "Advanced Turboprop Testbed Systems Study, I I NASA Lewis Research Center, CR 167928 (prepared by Lockheed-Georgi a Company and issued as LC81 ER0202) , Jul y e 1982.
2. Goldsmith, 1. M. , "Advanced Turboprop Testbed Systems Study," NASA
Lewis Research Center, CR 167895 (prepared by McDonnel 1 Douglas Corporation), J u l y 1982.
3. "The Langley Transonic Dynamics Tunnel , 11 Langley Research Center, Hampton, V i r g i n i a, LWP-799, September 1969.
4. "Wind Tunnel Model Systems Criteria,?" NASA L a n g l e y Research Center LHB 1710.15, September 1983, Revised January 1985.
5. Crooks, 0. J . , iiPropfan Test Airplane F l u t t e r Model Design Specifica- tion," Lockheed-Georgia, Company, PTA 41-001, September 1984.
6. Jenness, C. M. J., "Propfan T e s t Assessment Testbed A i r c r a f t F l u t t e r Model System Safety, I t Lockheed-Georgia Company, PTA 41-003, J u l y 1985.
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Loc kheed- Geo r g i a Company 86 South Cobb Drive NAS3- 24 339 Marietta, GA 30063 13. Typo of R e m and Pwlod Covered 2. dponrorlng &anw p(rrm urd Addma Contractor Report ~~ ~ NASA Lewis Research Center 114. Sponrorlng Agency Code 21000 Brookpark Road Cleveland, OH 44135 Project Yanager: E , J . Graber 1 8 . Ab.h.Et A 1/9-thscale h i g h speed f l u t t e r model o f the Gulfstream 11, modified t o the PTA testbed a i r c r a f t configurations, was tested d u r i n g August 1985 i n the NASA- Langley Transonic Dynamics Tunnel. The dynamically scaled propfan powerplants included w i n d m i l l i n g propfans. Tests o f the unsymmetric single a n d symmetric t w i n p r Q p f a n canfigurations revealed no w i n g o r propfan whirl f l u t t e r instabil- i t i e s w i t h i n t h e scaled a i r c r a f t f l i g h t envelope, expanded by a 20% f l u t t e r speed margin. Test d a t a were also acquired w i t h destabilizing w i n g t i p booms f o r the purpose o f analysis validation. Excellent correlations w i t h the d a t a were achieved i n post-test f l u t t e r analyses using actual model properties and modified doublet l a t t i c e aerodynamic derivatives. I t i s concluded t h a t the analysis method provides accurate f l u t t e r predictions for verification of the PTA testbed a i r c r a f t final design.
Prop- f a n
Unclassified - -
Wind Tunnel Testing A i rcraft F1 utter Model
unclassified I uncl assi f i ed
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