Document
NASA TP 15 01 c .1 1
NASA Technical Paper 1501
Response Technique Including
Correlations With Analytical
and Flight Test Results
L. Tracy Redd, Perry W. Hanson, and Eleanor C. Wynne NOVEMBER 1979 TECH LIBRARY KAFB, N M
NASA Technical Paper 1501
c
Evaluation of a Wind-Tunnel Gust
Response Technique Including
Correlations With Analytical
and Flight Test Results
L. Tracy Redd, Perry W. Hanson, and Eleanor C. Wynne Langley Reseu rch Cetzter Hamptoiz, Virginia National Aeronautics and Space Administration Scientific and Technical Information Branch SUMMARY A wind-tunnel technique f o r o b t a i n i n g g u s t frequency-response f u n c t i o n s for u s e i n p r e d i c t i n g t h e response of f l e x i b l e a i r c r a f t to atmospheric turbu- l e n c e is e v a l u a t e d by comparing t h e t u n n e l test r e s u l t s f o r a dynamically R s c a l e d cable-supported aeroelastic model with a n a l y t i c a l and f l i g h t d a t a .
The technique, which employs o s c i l l a t i n g vanes i n t h e t u n n e l throat s e c t i o n c to g e n e r a t e a s i n u s o i d a l l y varying flow f i e l d around t h e model, w a s e v a l u a t e d by use of a 1/30-scale model of t h e B-52E a i r p l a n e , f o r which c o n s i d e r a b l e f l i g h t g u s t response d a t a were a v a i l a b l e . The s t u d i e s show good c o r r e l a t i o n between t h e wind-tunnel r e s u l t s , f l i g h t test r e s u l t s , and a n a l y t i c a l predic- t i o n s f o r response i n t h e s h o r t - p e r i o d and wing f i r s t elastic modes of motion, which are t h e modes of primary s i g n i f i c a n c e f o r response of f l e x i b l e a i r c r a f t t o atmospheric t u r b u l e n c e .
INTRODUCTION The response of a i r c r a f t to atmospheric t u r b u l e n c e is an important d e s i g n c o n s i d e r a t i o n from t h e s t a n d p o i n t of g u s t l o a d s , s t r u c t u r a l f a t i g u e , and r i d e A method commonly used to determine t h e response of a i r c r a f t to random q u a l i t y .
g u s t s is based on random process t h e o r y or t h e so-called power spectral a n a l y s i s technique (see refs. 1 and 2, f o r example). I n t h i s method t h e response of a f l e x i b l e a i r p l a n e is determined f o r e x c i t a t i o n by s i n u s o i d a l g u s t s of varying frequency. T h i s response f u n c t i o n is commonly r e f e r r e d t o as t h e frequency- response f u n c t i o n .
Airplane frequency-response f u n c t i o n s are g e n e r a l l y determined from power s p e c t r a l data measured d u r i n g f l i g h t tests ( r e f . 3 ) or by a n a l y t i c a l methods ( r e f s . 4 t o 7 ) . The procedures f o r determining response f u n c t i o n s from f l i g h t tests are c o s t l y ; furthermore, t h e y cannot provide e a r l y d e s i g n d a t a f o r spe- c i f i c c o n f i g u r a t i o n s . A n a l y t i c a l methods, on t h e other hand, may prove t o be inadequate, p a r t i c u l a r l y i n t h e t r a n s o n i c speed range, where accurate d e f i n i - is l a c k i n g . T h e r e f o r e , a wind-tunnel method t i o n of unsteady aerodynamic l o a d s is d e s i r a b l e for e v a l u a t i n g frequency-response f u n c t i o n s e x p e r i m e n t a l l y under a method would allow v e r i - c o n t r o l l e d c o n d i t i o n s during t h e d e s i g n phase. Such f i c a t i o n of t h e a n a l y s i s i n a t i m e l y manner. The need f o r a c o n t r o l l e d e x p e r i - mental method l e d t o t h e development of a unique wind-tunnel g u s t response technique developed f o r u s e i n t h e Langley t r a n s o n i c dynamics t u n n e l .
The wind-tunnel technique c o n s i s t s of measuring t h e response of an aero- e l a s t i c a l l y s c a l e d model to an o s c i l l a t i n g v e r t i c a l g u s t f i e l d g e n e r a t e d i n t h e t u n n e l by o s c i l l a t i n g vanes l o c a t e d upstream of t h e test s e c t i o n . The model is flown on a two-cable support system (ref. 8 ) which permits s i m u l a t i o n of f r e e - f l i g h t modes of motion. Two p r e l i m i n a r y m o d e l s t u d i e s ( r e f s . 9 and 1 0 ) u s i n g
I
t h e technique showed good c o r r e l a t i o n between a n a l y s i s and wind-tunnel response measurements f o r r i g i d body modes of motion.
L
The next logical step in the verification of the method was to conduct tests using a dynamically scaled aeroelastic model and to compare the results with analytical predictions and flight data. Because considerable gust response data were available for the B-52E airplane, it was selected as the test vehicle for a comparative analysis of the wind-tunnel, flight, and analytical test meth- ods. The study was conducted through an effort carried out jointly by NASA, the U . S . Air Force, and the Boeing Company (Wichita Division). This paper describes the B-52E model wind-tunnel tests involved in the study and compares those test results with analytical and flight data.
AND ABBREVIATIONS SYMBOLS cb wing vertical bending-moment coefficient, Mb/qSz Cn coefficent of normal acceleration, mz/qS
dynamic pitching-moment coefficient , eEW/2VqS
cq
-
C mean aerodynamic chord, m c.g. center of gravity distance from oscillating vane quarter chord to gust probe head dV ( 1 4 . 9 m) F.S. model body station measured alonq fuselage center line (F.S. 0 is 4.83 cm forward of fuselage nose), cm f frequency, Hz g gravitational constant, 9.80 m/sec2
k reduced frequency , cW/2V
L.W.S. left wing station 1 characteristic length, m M Mach number wing vertical bending moment, N-m Mb m mass, kg P.R.G. pitch-rate gyro 9 dynamic pressure, Pa R.W.S. right wing station wing area, m 2 V v e l o c i t y , m/sec W weight, N W . S . wing s t a t i o n , measured p a r a l l e l to wing t r a i l i n g edge ( W . S . 0 is a t i n t e r s e c t i o n of wing l e a d i n g edge and f u s e l a g e c e n t e r l i n e ) , cm * ..
z v e r t i c a l a c c e l e r a t i o n normal to f u s e l a g e c e n t e r l i n e , m/sec2 a b s o l u t e v a l u e of measured g u s t a m p l i t u d e , deg 1Egl a b s o l u t e averaged v a l u e of g u s t amplitude over wing span, deg IEgI o s c i l l a t i n g amplitude of g u s t vanes, deg atmospheric a i r or t u n n e l test-medium d e n s i t y , kg/m3 phase a n g l e between model response and g u s t , deg t h e o r e t i c a l phase a n g l e between g u s t a t model c . g . and g u s t vane, deg measured phase a n g l e between g u s t a t model c.g. and g u s t vane, deg averaged v a l u e of g u s t phase a n g l e along model wing span, deg w c i r c u l a r frequency, 2nf, rad/sec S u b s c r i p t s : A f u l l - s c a l e a i r p l a n e M model APPARATUS The study w a s conducted i n t h e Langley t r a n s o n i c dynamics t u n n e l which has a 4.88-m s l o t t e d test s e c t i o n and is c a p a b l e of t e s t i n g a t Mach numbers up to 1.2 i n a i r or Freon-12, over a wide range of a i r or test-medium d e n s i t i e s .
The t u n n e l is c a p a b l e of g e n e r a t i n g a simulated ( s i n u s o i d a l ) g u s t f i e l d i n t h e test s e c t i o n .
The main f e a t u r e s of t h e o s c i l l a t i n g vane system used t o g e n e r a t e t h e g u s t f i e l d are shown i n f i g u r e 1 . Two sets of vanes, i n a b i p l a n e arrangement, are l o c a t e d on t h e s i d e w a l l s of t h e e n t r a n c e s e c t i o n of t h e t u n n e l . The vanes have a span of 1.07 m, a taper ratio of 0.5, and a p a n e l aspect ratio of 1.2. The b i p l a n e vanes on each w a l l are o s c i l l a t e d about t h e quarter chord by a h y d r a u l i c motor and a l a r g e f l y w h e e l by means of l i n k a g e s , which produce n e a r l y s i n u s o i d a l vane o s c i l l a t i o n s about t h e mean angle-of-attack p o s i t i o n . Under normal circum- s t a n c e s , t h e vane amplitudes are mechanically a d j u s t a b l e from Oo to &12O, and t h e frequency is remotely a d j u s t a b l e from 0 t o 20 Hz by means of an electrical control system which also synchronizes t h e motions of t h e t w o sets of vanes. By means o f t h i s system, t h e phasing of t h e t w o s i d e s can be v a r i e d from " i n sync" to 180° o u t o f s y n c h r o n i z a t i o n ; however, t h i s c a p a b i l i t y w a s n o t used i n t h i s i n v e s t i g a t i o n .
The f l o w a n g l e of t h e airstream w a s measured w i t h f o u r d i f f e r e n t i a l pres- s u r e probes mounted 0.61 m apart on a h o r i z o n t a l bar as shown i n f i g u r e 2. The bar w a s a t t a c h e d to a s t a n d which could b e . l o c a t e d a t any l o n g i t u d i n a l s t a t i o n i n t h e tunnel. The bar could be a d j u s t e d to any vertical s t a t i o n w i t h i n a d i s - t a n c e o f k0.61 m from t h e c e n t e r of t h e test s e c t i o n and could be remotely tra- versed l a t e r a l l y on t h e s u p p o r t , k0.30 m from t h e t u n n e l c e n t e r l i n e .
The p r e s s u r e probes ( f i g . 3) c o n s i s t e d of t w o (0.178 cm i n n e r diameter) steel t u b e s or "claws," a t t a c h e d to opposite s i d e s of a t w o - a r m v a r i a b l e - r e l u c t a n c e d i f f e r e n t i a l p r e s s u r e t r a n s d u c e r . The claws o f t h e probes were o r i e n t e d to measure flow a n g u l a r i t y i n t h e v e r t i c a l p l a n e by means of t h e a t t e n d a n t pressure d i f f e r e n t i a l between t h e upper and lower claw t u b e s .
Flow a n g l e measurements i n t h e test s e c t i o n were made i n t h e v i c i n i t y of t h e model c e n t e r of g r a v i t y ( c . g . ) , which was l o c a t e d approximately 1 4 . 9 m down- stream of t h e vane quarter chord.
F L O W FIELD Measurements of t h e flow f i e l d g e n e r a t e d by t h e airstream oscillator are shown i n f i g u r e s 4 and 5 . F i g u r e 4 p r e s e n t s a three-dimensional view of t h e v a r i a t i o n of g u s t amplitude leg[ and phase a n g l e @g as a f u n c t i o n of l a t - eral p o s i t i o n i n t h e t u n n e l and vane o s c i l l a t i o n reduced frequency. These plots were made by f i t t i n g a s u r f a c e s p l i n e ( r e f . 1 1 ) through t h e measured d a t a p o i n t s (shown as circles i n t h e f i g u r e ) . These f i g u r e s i n d i c a t e t h a t t h e amplitude d e c r e a s e s and t h e phase l a g d e c r e a s e s l i n e a r l y w i t h i n c r e a s i n g frequency. A l l t h e d a t a shown a r e based on 1-minute averages of t h e o u t p u t of an e l e c t r o n i c sine-cosine r e s o l v e r system t h a t g i v e s t h e in-phase and out-of-phase components of t h e f i r s t harmonic of stream angle with respect t o vane p o s i t i o n . T h i s pro- cedure w a s used to average o u t t h e e f f e c t of random t u n n e l t u r b u l e n c e .
The d a t a shown i n t h e f i g u r e s are f o r a v e l o c i t y of 35.4 m/sec, which is r e q u i r e d to s i m u l a t e t h e v e l o c i t y of t h e f u l l - s c a l e airplane. The vane ampli- t u d e ( k 6 O ) w a s determined from t h e practical c o n s i d e r a t i o n t h a t t h e g u s t a n g l e a t t h e model c.g. must be l a r g e enough to e x c i t e t h e s h o r t - p e r i o d and lower elastic modes, but n o t so l a r g e t h a t t h e model c o u l d be damaged. These d a t a show a c o n s i d e r a b l e v a r i a t i o n i n g u s t amplitude i n t h e spanwise d i r e c t i o n a t l o w f r e q u e n c i e s . The phase d a t a are f a i r l y uniform i n t h e spanwise d i r e c t i o n , e x c e p t a t t h e higher f r e q u e n c i e s . Although t h e r e is c o n s i d e r a b l e v a r i a t i o n i n spanwise g u s t amplitude a t l o w f r e q u e n c i e s , t h e g u s t is s u f f i c i e n t l y uniform t o produce model responses similar to t h o s e p r e d i c t e d f o r a g u s t uniformly d i s t r i b - uted over t h e wing span. This f a c t is shown i n t h e d a t a p r e s e n t e d i n following s e c t i o n s . Consequently, t h e spanwise v a r i a t i o n s i n t h e g u s t amplitude and phase a n g l e were averaged across t h e span f o r each v a l u e of reduced frequency k i n analyzing t h e model g u s t response data. The r e s u l t s of t h i s averaging process are shown i n f i g u r e s 5 ( a ) and 5 ( b ) . The format used i n f i g u r e 5 ( b ) c o n f i n e s v a r i a t i o n s of the phase a n g l e d a t a to 51 80° ( i n s t e a d of accumulating t o 21 400° as i n f i g . 4 ( b ) ) . A t l o w reduced f r e q u e n c i e s t h e measured phase a n g l e $g approximates t h a t p r e d i c t e d by t h e e q u a t i o n where is t h e phase d i f f e r e n c e caused by t h e t i m e r e q u i r e d f o r a p o i n t on $d a t r a v e l i n g wave a t a frequency f to move downstream t h e d i s t a n c e from t h e vane q u a r t e r chord to t h e probe head dv a t a v e l o c i t y V. The d i f f e r e n c e between t h e measured and t h e t h e o r e t i c a l phase l a g a t high reduced f r e q u e n c i e s can be a t t r i b u t e d t o unsteady f l a w e f f e c t s on t h e vanes.
MODEL PROPERTIES AND INSTRUMENTATION Model S c a l i n g C o n s i d e r a t i o n s The 1/30-geometric-scale model w a s designed t o s i m u l a t e t h e B-52E a i r p l a n e dynamically and a e r o e l a s t i c a l l y a t a g r o s s mass of 189 964 kg a t a f l i g h t a l t i - t u d e of 1646 m with e x t e r n a l tanks a t t a c h e d . Froude number, reduced frequency, and mass r a t i o of t h e model and a i r p l a n e were matched. The model s t i f f n e s s p r o p e r t i e s , i n e r t i a p r o p e r t i e s ( d i s t r i b u t e d masses and moments of i n e r t i a ) , aerodynamic p r o p e r t i e s , and c e n t e r - o f - g r a v i t y l o c a t i o n also matched t h a t of t h e f u l l - s c a l e B-52E a i r p l a n e . The model s c a l i n g f a c t o r s d e r i v e d from t h e preceding c o n d i t i o n s are t a b u l a t e d i n t a b l e I.
To simulate t h e a i r p l a n e a l t i t u d e ( a i r d e n s i t y ) and to allow t h e model t o be heavy enough f o r i n s t r u m e n t a t i o n and b a l l a s t i n g , t h e tests were c a r r i e d o u t i n Freon-12 gas. The r e s u l t i n g d e n s i t y r a t i o of 4.07 between model and a i r - plane (see t a b l e I) provided a model mass of 28.7 kg. Comparisons of t h e model a i r p l a n e test c o n d i t i o n s and other parameters are given i n t a b l e 11.
Model P h y s i c a l C h a r a c t e r i s t i c s The model is shown mounted i n t h e Langley t r a n s o n i c dynamics t u n n e l i n f i g - u r e 6, and a schematic diagram of t h e mount system is shown i n f i g u r e 7. The system ( r e f . 8) c o n s i s t s of t w o " f l y i n g " c a b l e s on which t h e model is suspended.
The forward f l y i n g c a b l e loop is i n t h e v e r t i c a l plane; t h e rear f l y i n g c a b l e l o o p is i n t h e h o r i z o n t a l plane. Four "snubber c a b l e s , " normally s l a c k d u r i n g t e s t i n g , can be remotely a c t i v a t e d to snub or r e s t r a i n t h e model i n case a n i n s t a b i l i t y occurs.
T o achieve r e a s o n a b l e s i m u l a t i o n of t h e short-period mode of t h e B-52E air- plane, t h e b a s i c two-cable-mount system of r e f e r e n c e 8 w a s modified as shown i n case, t h e c a b l e s were a t t a c h e d through p i v o t s t o t h e model f i g u r e 7 . I n t h i s near t h e c e n t e r of g r a v i t y , and t h e p u l l e y s were mounted a t t h e t u n n e l w a l l i n s t e a d of f o l l o w i n g t h e normal p o s i t i o n i n g of t h e p u l l e y s i n s i d e t h e contour of t h e model f u s e l a g e . T h i s m o d i f i c a t i o n r e s u l t e d i n a mount c o n f i g u r a t i o n which had a r o t a t i o n a l s t i f f n e s s low enough i n p i t c h t o allow adequate s i m u l a t i o n of t h e s h o r t - p e r i o d f r e e - f l i g h t mode.
Except f o r t h e v e r t i c a l t a i l , t h e model s t i f f n e s s w a s provided by alumi- num a l l o y spars covered w i t h separate r i g i d f a i r i n g s e c t i o n s which provided t h e aerodynamic contour of t h e a i r p l a n e . Each s e c t i o n w a s a t t a c h e d to t h e spars independently so t h a t s t i f f n e s s of t h e spars would n o t be a l t e r e d .
Nacelle s t i f f n e s s w a s provided by a s i n g l e mounting s t r u t a t t a c h e d t o t h e wing spar. The v e r t i c a l t a i l had an aluminum plate core w i t h b a l s a c o v e r i n g to pro- v i d e t h e a i r f o i l shape.
The model geometry is shown i n f i g u r e 8. The a i l e r o n s and all-movable h o r i z o n t a l s t a b i l i z e r s provided r o l l and p i t c h t r i m c o n t r o l , r e s p e c t i v e l y , and both c o n t r o l s could be o p e r a t e d remotely from t h e t u n n e l c o n t r o l room.
Model Dynamic C h a r a c t e r i s t i c s Before t h e wind-tunnel t e s t i n g was begun, t h e model w a s mounted on t h e f l y i n g c a b l e s i n t h e t u n n e l test s e c t i o n , and model ground v i b r a t i o n tests were conducted i n t h e wind-off c o n d i t i o n . The suspension system used allowed proper s i m u l a t i o n of t h e a i r p l a n e s h o r t - p e r i o d mode. A s o f t s p r i n g a t t a c h e d t o t h e model a t t h e c.g. was used t o suspend t h e model v e r t i c a l l y i n t h e tun- n e l f o r t h e s e v i b r a t i o n tests. A comparison of t h e symmetric modal f r e q u e n c i e s measured i n t h e s e tests and t h o s e c a l c u l a t e d f o r t h e u n r e s t r a i n e d model are shown i n t a b l e s I11 and I V . The d a t a show t h a t c a l c u l a t e d and measured frequen- cies of t h e f l e x i b l e modes ( t a b l e I I I ( a ) ) are i n good agreement. The agreement between c a l c u l a t e d and measured r i g i d body modes ( t a b l e I I I ( b ) ) and n a c e l l e s t r u t modes ( t a b l e I V ) is considered adequate.
The modal node l i n e s o b t a i n e d i n t h e model v i b r a t i o n tests are shown i n f i g u r e 9 f o r t h e f i r s t f o u r symmetric elastic modes. These are t h e modes which were most r e a d i l y e x c i t e d by t h e g u s t v e l o c i t y g e n e r a t e d i n t h e t u n n e l . As shown i n f i g u r e 4, t h e g u s t amplitude is almost z e r o above a reduced frequency of 0.20 (which corresponds to a frequency of about 1 0 Hz) .
I n s t r ument a t i o n Locations of t h e onboard i n s t r u m e n t a t i o n and equipnent f o r measuring t h e model response and f o r powering t h e h o r i z o n t a l s t a b i l i z e r and a i l e r o n c o n t r o l s u r f a c e s are shown i n f i g u r e 1 0 . The s e n s o r s c o n s i s t e d of (1) f o u r sets of s t r a i n - g a g e b r i d g e s mounted on t h e wing spar, o r i e n t e d and c a l i b r a t e d t o measure bending moments; ( 2 ) t h r e e v e r t i c a l accelerometers mounted on t h e f u s e l a g e spar; and ( 3 ) a p i t c h - r a t e gyro mounted on t h e f u s e l a g e spar a t t h e model c.g. A i l e - ron and s t a b i l i z e r c o n t r o l s u r f a c e s were d r i v e n by independent dc torque motors
r
through crank-pushrod l i n k a g e s . The pushrods were connected by f l e x i b l e bellows couplings. The a i l e r o n motors each provided a maximum t o r q u e of k0.381 N-m, and through t h e l i n k a g e s , c o n t r o l s u r f a c e o s c i l l a t i o n amplitude of k25O w a s provided i n t h e frequency range from 0 to 20 Hz. The h o r i z o n t a l s t a b i l i z e r motor, used for t r i m o n l y , produced a maximum torque of k3.884 N-m and d e f l e c t e d t h e s u r f a c e a t a rate of 0.06 deg/sec. The s t a b i l i z e r w a s capable of t r a v e l from 100 t r a i l - i n g edge down to 5O t r a i l i n g edge up w i t h respect t o t h e f u s e l a g e w a t e r l i n e .
For cable-mounted models, an electrical "umbilical" c a b l e is r e q u i r e d t o connect model s e n s o r s to r e a d o u t i n s t r u m e n t a t i o n . I n o r d e r t o keep t h e umbili- cal c a b l e small and l i g h t , t h e r e b y minimizing its e f f e c t on model response, t h e power t o t h e onboard i n s t r u m e n t a t i o n and c o n t r o l actuator motors w a s s u p p l i e d t o t h e model through f o u r copper-clad snubber c a b l e s .
A m i c r o e l e c t r o n i c m u l t i p l e x i n g system w a s used t o t r a n s m i t electrical com- mands to t h e onboard i n s t r u m e n t a t i o n and to r e c e i v e d a t a s i g n a l s f r u n t h e model s e n s o r s simultaneously through t h e 0.159-cm-diameter c o a x i a l u m b i l i c a l c a b l e .
are given i n t h e appendix. T h i s sys- The d e t a i l s of t h e m i c r o e l e c t r o n i c system t e m w a s c a p a b l e of t r a n s m i t t i n g 1 0 c h a n n e l s of analog d a t a s i m u l t a n e o u s l y from t h e t r a n s d u c e r s w i t h i n t h e model through t h e c o a x i a l u m b i l i c a l c a b l e t o t h e con- t r o l room o u t s i d e t h e test s e c t i o n . I n a d d i t i o n , through t h i s c o a x i a l c a b l e , t h e system could simultaneously code g a i n and o f f s e t commands to t h e onboard s i g n a l c o n d i t i o n i n g u n i t s , operate commands t o t h e t h r e e c o n t r o l motors, and g e n e r a t e a timing s i g n a l f o r d a t a s y n c h r o n i z a t i o n .
ANALYTICAL MODEL RESPONSE STUDIES Model Mount E f f e c t s A n a l y t i c a l s t u d i e s were conducted by t h e Boeing Company (Wichita D i v i s i o n ) t o determine t h e cable-mount e f f e c t s on t h e model motions. The e q u a t i o n s of motion and methods of a n a l y s i s of r e f e r e n c e 1 2 were used i n t h e response s t u d - ies. These s t u d i e s c o n s i s t e d of comparisons of t h e response c h a r a c t e r i s t i c s of t h e model when mounted on t h e f l y i n g c a b l e s and i n free f l i g h t . I n both cases, t h e measured v e r t i c a l g u s t data shown i n f i g u r e 4 were used i n t h e a n a l y s i s t o g e n e r a t e spanwise g u s t l o a d s on t h e model. I n each case, t h e a m p l i t u d e of t h e model response d a t a w a s normalized by t h e averaged spanwise v e r t i c a l g u s t a n g l e legl i n d i c a t e d i n f i g u r e 5 ( a ) . Typical r e s u l t s of t h i s a n a l y s i s are shown i n f i g u r e 11 i n terms of t h e v a r i a t i o n w i t h t h e reduced frequency k of t h e ratios IC,/eg I , I Cq/Eg I , I cb/Eg I , and t h e a s s o c i a t e d phase a n g l e s . (See f i g . 10 f o r measurement l o c a t i o n s . ) These results i n d i c a t e t h a t t h e mount has a small e f f e c t on t h e model response a t t h e reduced f r e q u e n c i e s correspond- o n l y i n g t o t h e s h o r t - p e r i o d ( k = 0.04) and f i r s t elastic (wing-bending, k = 0.072) A s i g n i f i c a n t e f f e c t is i n d i c a t e d on t h e wing fore-and-aft bending mode modes.
(second elastic mode, k = 0.14). The s h a r p peak i n t h e cable-mount data a t low v a l u e s of k (k = 0.005) is a s s o c i a t e d w i t h t h e v e r t i c a l t r a n s l a t i o n mode of t h e model on t h e mount system and, hence, does n o t appear i n t h e f r e e - f l i g h t a n a l y s i s .
Effect of Spanwise Variation of Gust Field on Analytically Derived Frequency Responses of Cable Mounted Model Ideally, the gust field generated in the wind tunnel should be nearly uniform, but the tunnel measurements in figure 4 show that the gust amplitude apparently has significant spanwise variation. Therefore, it was decided to determine the difference between the theoretical model response caused by the measured tunnel gust distribution and that resulting from a uniform gust obtained by averaging the measurements across the span. The power spectral density approach (ref. 1 ) was used in the response calculations. This approach provides statistical descriptions of the dynamic response from a combination of power spectral description of the turbulent velocities with solutions of linear equations of motion for the airplane. Typical comparisons of results of the analysis using the measured tunnel gust distribution (replotted from
I % I
fig. 1 1 ) and those caused by a uniform gust distribution are shown in figure 1 2 . This comparison shows that the spanwise gust variations in the tun- nel have only a small effect (about 4 percent maximum) on the model response except at reduced frequencies above k = 0 . 1 4 . The phase angles for the uni- form gust field were calculated by using the cross-spectra between the turbu- lence input at the tunnel streamwise station corresponding to the location of the model c.g. and the response of the model. The data from figure 1 2 showing model response indicate that the gust in the tunnel can be considered uniform along the span. The insensitivity of the model response to the spanwise vari- able gust amplitude is probably a result of the nearly constant spanwise phase difference (fig. 4 ( b ) ) .
Modeling of Atmospheric Turbulence Atmospheric turbulence has been found to be essentially homogeneous and, therefore, with respect to the axes of an airplane in level flight, the turbu- lence gust velocities would be expected to vary along the span of the airplane as well as along the direction of flight. However, the most commonly used for- mulation of the spectral approach to calculating airplane response to turbu- lence has assumed a one-dimensional or "uniform" gust field. In this formula- tion the gust velocities are considered to vary randomly in the direction of flight but are uniform along the span. For an airplane with a large wing span (greater than about 5 percent of the turbulence scale), such as the B-52E, the one-dimensional analysis may not be sufficient. For this case, a more adequate gust model is obtained by assuming the turbulence to be a random two-dimensional isotropic gust (refs. 1 3 and 1 4 ) . Furthermore, the analyses of reference 1 5 indicate that the two-dimensional gust analysis produces calculated airplane responses which more nearly agree with flight test results than does the one- dimensional analysis. Because the model response in the tunnel approximates
that from a one-dimensional gust field (see fig. 1 2 ) , it was desirable to
determine the differences between the model response to one-dimensional and two-dimensional gust inputs. (The two-dimensional gust input of refs. 13 and 1 4 was used for these calculations.) If these differences are sufficiently large, then corrections which include these differences, plus corrections for the slight differences between the one-dimensional analysis and the response result- i n g from t u n n e l g u s t s ( f i g . 12) and t h e cable-mount e f f e c t s ( f i g . 11) , should be a p p l i e d to t h e wind-tunnel d a t a b e f o r e comparisons w i t h t h e f l i g h t test d a t a can be made. Typical comparisons between t h e one- and two-dimensional random g u s t i n p u t s f o r t h e model are shown i n f i g u r e 13.. These d a t a show t h a t t h e two-dimensional g u s t a n a l y s i s g i v e s peak v a l u e s f o r t h e f i r s t elastic mode (i.e., k ZJ 0.07) which are about 10 p e r c e n t lower t h a n t h o s e from t h e uniform a n a l y s i s . This d i f f e r e n c e , when added to t h e d i f f e r e n c e i n t h e r e s p o n s e s of t h e model caused by t h e t u n n e l g u s t and t h e uniform (one-dimensional) a n a l y s i s ( f i g . 12) and to t h e d i f f e r e n c e caused by cable-mount e f f e c t s ( f i g . 11 1 , g i v e s maximum c o r r e c t i o n factors of approximately 12 p e r c e n t f o r t h e reduced frequency range o f i n t e r e s t (i.e., 0.01 6 k 5 0.14). A comparison of t h e phase a n g l e s i n f i g u r e s 11 t o 1 3 i n d i c a t e t h a t no c o r r e c t i o n s i n phase a n g l e are appar- e n t l y n e c e s s a r y , e x c e p t f o r t h e cable-mount e f f e c t s a t l o w reduced frequency ( k < 0.01) i n f i g u r e 11 and f o r cable and t u n n e l spanwise v a r i a t i o n e f f e c t s a t t h e second elastic mode frequency (k = 0.14).
RESULTS AND DISCUSSION Measured and C a l c u l a t e d Model Response The model response t o t h e g u s t s generated by o s c i l l a t i n g t h e g u s t vanes a range of f r e q u e n c i e s a t a t u n n e l f l o w v e l o c i t y o f 35.4 m/sec w a s mea- over sured i n terms of a c c e l e r a t i o n s and moments experienced by t h e model. A com- p a r i s o n of t h e uncorrected measured and t h e c a l c u l a t e d frequency responses ( f i g . 12) of t h e model are shown i n f i g u r e 14. ( N o t e scale change r e l a t i v e to p r e v i o u s f i g u r e s . ) The nondimensional response c o e f f i c i e n t s o f a c c e l e r a t i o n , p i t c h i n g moment, and bending moment a t v a r i o u s s t a t i o n s on t h e model are p l o t t e d i n t h i s f i g u r e . For t h e l o c a t i o n s of t h e response s t a t i o n s , r e f e r to f i g u r e 10.
All d a t a are f o r symmetric response w i t h t h e model on t h e cable-mount system.
I n g e n e r a l , t h e measured r e s u l t s a g r e e w e l l w i t h c a l c u l a t e d d a t a , except a t t h e higher reduced f r e q u e n c i e s k > 0.12. The l o w g u s t l e v e l produced a t t h e higher v a l u e s o f k undoubtedly r e s u l t e d i n i n a c c u r a c i e s i n t h e measure- ment both of t h e response and t h e g u s t amplitudes. The a n a l y s i s t e n d s to pre- d i c t a higher response than w a s measured f o r t h e f i r s t elastic mode ( k = 0.072) a t t h e f u s e l a g e a f t end ( f i g . 1 4 ( c ) ) , t h e p i t c h i n g moment f o r t h e a c c e l e r a t i o n a t t h e c.g. ( f i g . 14 ( d ) ) , and t h e bending moments along t h e wing ( f i g s . 1 4 (e) and 1 4 ( f ) ) .
The measured r e s u l t s shown i n f i g u r e s 1 4 ( b ) and 1 4 ( d ) show t h e repeatabil- i t y of t h e g u s t response tests f o r t w o d i f f e r e n t r u n s (i.e., r u n s 1 and 2 ) . For v a l u e s o f k below 0.16, t h e r e p e a t a b i l i t y is e x c e l l e n t .
Model and F l i g h t T e s t R e s u l t s The measured model and a i r p l a n e f l i g h t test r e s u l t s are compared i n f i g - u r e 15. Two d i f f e r e n t sets of model d a t a are p r e s e n t e d - t h e "uncorrected" d a t a t o account f o r from f i g u r e 1 4 and t h e " c o r r e c t e d " d a t a , which have been a d j u s t e d t h e cable-mount and two-dimensional g u s t e f f e c t s .
L
The model and f l i g h t test d a t a shown i n f i g u r e 15 g e n e r a l l y a g r e e w e l l for values of reduced f r e q u e n c i e s from 0.03 t o 0 . 1 4 , which i n c l u d e t h e s h o r t - p e r i o d and f i r s t e l a s t i c m o d e s . However, t h e pitching-moment d a t a ( f i g . 1 5 ( d ) ) do not a g r e e w e l l , and much scatter appears i n t h e f l i g h t data for k > 0.08. It should be noted t h a t a t l a w reduced f r e q u e n c i e s (k < 0.03), a l l t h e a i r p l a n e d a t a were a f f e c t e d by s p u r i o u s pilot-induced motions ( r e f . 1 5 ) . A l s o , a t t h e higher reduced f r e q u e n c i e s ( k > 0 . 1 4 ) , the low g u s t l e v e l produced by t h e air- stream oscillator l e d to measurement errors i n t h e model data. The short-period mode of motion (k - 0.04) d i d not show up v e r y w e l l i n e i t h e r t h e f l i g h t test or model response data. A t b e s t , it can b a r e l y be d e t e c t e d i n t h e model d a t a of f i g u r e 1 5 . On t h e whole, it is completely overpowered by t h e s t r o n g f i r s t elas- t i c mode. This illustrates t h e high degree of f l e x i b i l i t y i n t h e B-52E a i r p l a n e s t r u c t u r e .
The major d i f f e r e n c e noted between c o r r e c t e d and uncorrected model data is t h a t c o r r e c t i o n s reduce t h e p e a k amplitude of t h e f i r s t elastic mode of t h e model ( k = 0 . 0 7 2 ) by a b o u t 1 0 p e r c e n t . The r e s u l t is better agreement w i t h t h e f l i g h t data. The c o r r e c t i o n improved t h e phase a n g l e and, i n most cases, t h e c o e f f i c i e n t amplitudes a s s o c i a t e d w i t h t h e low-frequency c a b l e mode ( k = 0.01); however, i n g e n e r a l , t h e c o r r e c t i o n s produced o n l y minor changes i n t h e results.
CONCLUDING REMARKS A wind-tunnel technique f o r o b t a i n i n g g u s t f requency-response f u n c t i o n s for u s e i n p r e d i c t i n g t h e response of f l e x i b l e aircraft to atmospheric turbu- l e n c e has been evaluated by comparing t h e t u n n e l test r e s u l t s f o r a dynamically s c a l e d cable-supported aeroelastic m o d e l of t h e B-52E a i r p l a n e w i t h a n a l y t i c a l and f l i g h t d a t a .
These s t u d i e s show good c o r r e l a t i o n between wind-tunnel, f l i g h t test, and a n a l y t i c a l p r e d i c t i o n s f o r t h e short-period and f i r s t elastic modes of motion.
Since t h e s e t w o modes are g e n e r a l l y t h e ones p r i m a r i l y associated with a i r c r a f t g u s t response, t h e good c o r r e l a t i o n of results f o r t h e s e modes i n t h e p r e s e n t s t u d y i n d i c a t e s t h a t t h e wind-tunnel/airstream oscillator technique should be a u s e f u l and v a l i d tool f o r e v a l u a t i n g g u s t response c h a r a c t e r i s t i c s of a i r p l a n e s .
Langley Research Center N a t i o n a l Aeronautics and Space Administration Hampton, VA 23665 September 24, 1979 1 0 APPENDIX MICROELECTRONIC INSTRUMENTATION SYSTEM A block diagram of t h e m i c r o e l e c t r o n i c system used i n t h e B-52E model wind-tunnel g u s t tests is p r e s e n t e d i n f i g u r e 1 6 . Both t h e i n s t r u m e n t a t i o n components l o c a t e d onboard t h e model and t h o s e l o c a t e d i n t h e c o n t r o l room are shown. The connecting l i n k s between t h e t w o l o c a t i o n s a r e t h e c o a x i a l u m b i l i c a l c a b l e , which c a r r i e d s i g n a l s from t h e t r a n s d u c e r s and t r a n s m i t t e d commands to t h e t r i m c o n t r o l motors, and t h e four copper$lad snubber c a b l e s which c a r r i e d t h e power t o t h e onboard i n s t r u m e n t a t i o n .
The instrument components l o c a t e d i n t h e model (see f i g . 1 6 ) c o n s i s t e d of 1 0 t r a n s d u c e r s , a t r a n s d u c e r selector switch, 2 FM'data modulators, a s w i t c h c o n t r o l u n i t , a r e c e i v e r decoder u n i t , and 6 dc-dc v o l t a g e c o n v e r t e r s . Each of t h e t w o d a t a modulator u n i t s contained f i v e v o l t a g e - c o n t r o l l e d FM oscilla- tors; t h u s , each could handle f i v e incoming transducer s i g n a l s simultaneously.
I n a d d i t i o n , a remotely o p e r a t e d transducer selector u n i t (see f i g . 1 6 ) , which allawed time-sharing s i g n a l s , was used with d a t a modulator B; t h u s , t h i s modula- tor u n i t could handle s e l e c t i v e l y a t o t a l of 1 0 t r a n s d u c e r s i g n a l s , 5 a t a t i m e .
The transducer selector u n i t and t h e c o n t r o l s u r f a c e motors were c o n t r o l l e d by t h e switch c o n t r o l u n i t . Regulated power w a s s u p p l i e d to each of t h e t r a n s - ducers and o t h e r onboard i n s t r u m e n t a t i o n by t h e onboard power c o n v e r t e r s . The arrangement of onboard i n s t r u m e n t a t i o n i n t h e model is i n d i c a t e d i n f i g u r e 1 7 .
The c o n t r o l r o o m i n s t r u m e n t a t i o n used i n t h e model tests ( r e f e r t o f i g . 1 6 ) c o n s i s t e d of a two-way m u l t i c o u p l e r , a group carrier c o n v e r t e r , an analog tape r e c o r d e r , a demodulator, a d i s p l a y and c o n t r o l panel, a c o n t r o l code g e n e r a t o r , a c o n t r o l s i g n a l modulator. The d i s p l a y and c o n t r o l panel contained a s e t and of m i n i a t u r e oscilloscopes for continuous monitoring of each d a t a channel and a c o n t r o l panel f o r changing channel g a i n s and o f f s e t s . The model t r i m c o n t r o l s i g n a l s for t h e s t a b i l i z e r and a i l e r o n c o n t r o l s u r f a c e s were also routed through t h e d i s p l a y and c o n t r o l p a n e l . The s i g n a l s to and from t h e o s c i l l a t i n g vane system were also passed through t h e m i c r o e l e c t r o n i c system, as shown a t t h e t o p of f i g u r e 1 6 . This allowed t h e o s c i l l a t i n g vane d a t a t o be time-synchronized w i t h t h e model response data on t h e analog t a p e .
REFERENCES 1 . H o u b o l t , John C.; S t e i n e r , Roy; and P r a t t , K e r m i t G.: Dynamic Response of Airplanes to Atmospheric Turbulence I n c l u d i n g F l i g h t Data on I n p u t and Response. NASA TR R-199, 1 964.
2. P r a t t , K e r m i t G.: Response of F l e x i b l e A i r p l a n e s t o Atmospheric Turbulence.
Performance and Dynamics of Aerospace V e h i c l e s , NASA SP-258, 1971, pp. 439-503.
3. Coleman, Thomas L.; Murrow, Harold N.; and P r e s s , Harry: Some S t r u c t u r a l Response C h a r a c t e r i s t i c s of a Large F l e x i b l e Swept-Wing A i r p l a n e i n Rough A i r . J. Aero. S c i . , v o l . 25, no. 8 , Aug. 1958, pp. 515-521, 536.
4. Bennett, Floyd V.; and P r a t t , K e r m i t G.: C a l c u l a t e d Responses of a Large Sweptwing A i r p l a n e t o Continuous Turbulence With F l i g h t - T e s t Comparisons.
NASA TR R-69, 1960.
5. Coleman, Thomas L.; P r e s s , Harry; and Meadows, May T.: An E v a l u a t i o n of E f f e c t s of F l e x i b i l i t y on Wing S t r a i n s i n Rough A i r f o r a Large Swept- Wing Airplane by Means of Experimentally Determined Frequency-Response Functions With an Assessment of Random-Process Techniques Employed.
N A S A TR R-70, 1960. (Supersedes NACA TN 4291 .)
6. Chernoff, Max; and Rothman, Herbert L.: Unsteady Frequency Response Func- t i o n s f o r U s e i n Power S p e c t r a l A n a l y s i s . J. Aerosp. S c i . , v o l . 29, no. 2, Feb. 1962, pp. 121-129.
7. Zbrozek, J. K.: L o n g i t u d i n a l Response of A i r c r a f t t o O s c i l l a t o r y Vertical Gusts (Frequency A n a l y s i s I n c l u d i n g t h e E f f e c t of Unsteady Aerodynamics).
Rep. N o . Aero.2559, B r i t i s h R.A.E., Nov. 1955.
8. Reed, W i l l i a m H., 111; and Abbott, Frank T., Jr.: A New "Free-Flight" Mount System f o r High-speed Wind-Tunnel F l u t t e r Models. Proceedings o f Sympo- sium on Aeroelastic & D y n a m i c Modeling Technology, RTD-TDR-63-4197, P t . I, U.S. A i r Force, Mar. 1964, pp. 169-206.
9 . Gilman, Jean, Jr.; and Bennett, Robert M.: A Wind-Tunnel Technique f o r Measuring Frequency-Response Functions f o r G u s t Loads Analyses.
J. Aircr., v o l . 3, no. 6, Nov.-Dec. 1966, pp. 535-540.
10. Rainey, A. Gerald; and Abel, I r v i n g : Wind-Tunnel Techniques f o r t h e Study o f Aeroelastic E f f e c t s on A i r c r a f t S t a b i l i t y , C o n t r o l , and Loads. Aero- elastic E f f e c t s From a F l i g h t Mechanics S t a n d p o i n t , AGARD CP N o . 46, Mar. 1970, pp. 18-1 - 18-15.
11. Harder, Robert L.; and Desmarais, Robert N.: I n t e r p o l a t i o n Using Surface S p l i n e s . J. Aircr., v o l . 9, no. 2, Feb. 1972, pp. 189-191.
A.: B-52 Aeroelastic Model - Summary R e p o r t . B3-7763-1
12. G i l l e y , T.
(Contract N o . F33615-67-C-1264), Boeing Co., A p r . 25, 1968.
1 3 . Sawdy, David T.: bn t h e Two-Dimensional Atmospheric Turbulence Response of an A i r p l a n e . Ph. D. D i s s . , Univ. of Kansas, 1 9 6 6 . ( A v a i l a b l e as NASA CR-91116 .)
1 4 . P r a t t , K e r m i t G.: E f f e c t o f Spanwise V a r i a t i o n of Turbulence on t h e N o r m a l Acceleration of A i r p l a n e s With Small Span R e l a t i v e to Turbulence S c a l e .
NASA TM X-72748, 1975.
1 5 . G i l l e y , T. A.; and C a s t , Rudy D.: B-52C-F Dynamic Responses and Load
Survey ( V o l u m e 11) - WFT 1293 ( V o l u m e I1 - T h e o r e t i c a l and Experi-
mental Frequency Response Function Comparisons). Doc. N o . D3-7060-2
(Contract N o . AF34(601)-17947) , Boeing Co. , Sept. 27, 1 9 6 6 . (Avail-
a b l e from DDC as AD 904 598 .)
1 3 TABLE 1 . - MODEL SCALE FACTORS
I Scaled q u a n t i t y
Formula Factor Reduced frequency 1 .o Froude number - I - - I 1 .o wM PA/zA\3 Mass r a t i o 1 .o Dimension S e l e c t e d 1 /30 Tunnel = 0 . 0 0 8 ~ ~ F l u i d d e n s i t y 4 . 0 7 A i r p l a n e = 0.001 9634 V e l o c i t y 0.183
"("r
Dynamic p r e s s u r e 0.136 PA vA Frequency 5.48 Weight 0 . 0 0 0 1 51 1 4 TABLE 11.- COMPARISON B m E N M3DEL AND AIRPLANE TEST CONDITIONS Parameter A i r plane Model ~
. . . . 1646 NA
A l t i t u d e . m . . . . . .
. . . . 193 3 5 . 4
Velocity. m/sec . . . .
. . . . 0.569 0.226
Mach number . . . . . .
. . . . 18 826 2557
Dynamic p r e s s u r e . Pa .
3 0 0
Temperature. K . . . . . . . . 2 86
28.7
Mass. kg . . . . . . . . . . . 189 964
Density. kg/m3 . . . . . . . . 1 . 0 1 2 4.1 23
Speed of sound. m/sec . . . . . 339 1 5 6
Freon- 1 2
F l i g h t environment . . . . . . A i r
0 . 2 3 3
Mean aerodynamic chord. m . . . 6 . 9 9
1 . 8 8
Wing span. m . . . . . . . . . 56.39
0.41 3
. . . . 371 . 6 7
Wing area. m2 . . . . .
F.S. 0 . 6 8 5 8
Center of g r a v i t y . m . . . . . F . S . 20.57
~ 1 5 1111111 TABLE 111.- SUMMARY OF MODEL FREQUENCIES (a) Wind-off elastic symmetric modal f r e q u e n c i e s ~ Frequency, Hz M o d e C a l c u l a t e d Measured (a) 1 3 . 4 5 3.12 2 6 . 9 0 6 . 7 5 3 8 . 8 2 8 . 7 5 4 10.68 10.65 11.45 5 11.06 6 11.12 11.70 7 14.13 14.65 17.35 8 16.43 9 17.85 17.69 a M o d e l mounted on soft s p r i n g and f l y i n g cables.
(b) Wind-on r i g i d body f r e q u e n c i e s -. - - ._ -_ . - Frequency, Hz
I
C a l c u l a t e d Measured . - aModel mounted on f l y i n g cables; v = 35.4 m/sec.
TABLE 1V.- WIND-OFF CANTILEVERED NACELLE STRUT MODES N a t u r a l frequency, Hz I n b o a r d s t r u t and nacelle M o d e , s t r u t b e n d i n g D e s i g n A c t u a l 11.50 11.48 S i d e 11.17 11.12
I
22.02 21 .E6 22.29 21.62 V e r t i c a l
-----
32.92
T o r s i o n 32.92 I -----
Left a n d r i g h t nacelle a A c t u a l f r e q u e n c i e s are average of s t r u t f r e q u e n c ies .
Figure 1.- View oE airstream o s c i l l a t o r vane system i n Langley transonic dynamics t u n n e l showing cutaway of mechanism.
L-79-310 Figure 2.- O s c i l l a t i n g flow measurement probes mounted on motor d r i v e n t r a v e r s a b l e bar.
53.45 kPa differential pressure transducer Transducer lead wires 0.178 I.D. steel tube Figure 3 . - Flow angle probe geometry. (Linear dimensions i n c m ; I.D. r e p r e s e n t s inner diameter ; O.D., outer diameter .)
/I-
1 . 5 , ’
1.0 ’
, / . . L L 1 , 1 # /
0 .04 - 0 8 . 1 2 .16 - 2 0 Reduced frequency, k (a) Gust amplitude.
Figure 4.- Variation of amplitude and phase angle of gust flow with reduced frequency and lateral position. 6, = +6O; V = 35.4 m/sec; and E = 0.233 m.
I : -1400 / / /- / / / / / 0 .04 .08 . 1 2 .16 ,20 Reduced frequency, k (b) Gust phase a n g l e .
Figure 4 .- Concluded.
1.4 1.2 1.0 .8 , deg
I %I
.6 .4 .2 0 .04 .08 .12 .16 .20 Reduced frequency,k ( a ) G u s t amplitude.
-60 -120 -180 0 .04 .08 .12 .16 .20 Reduced frequency, k (b) Gust phase a n g l e .
F i g u r e 5.- V a r i a t i o n of averaged spanwise g u s t amplitude and phase a n g l e w i t h reduced frequency.
Ib
Unibilical car I;-70-3779.1 Figure 6 . - Model of B-52E a i r p l a n e mounted i n Langley t r a n s o n i c dynamics t u n n e l .
Figure 7.- Modified two-cable-mount system.
.
rd line Center of gravity 25% chord line of gravity Center Figure 8.- Model geometry. ( A l l linear dimensions are in cm.)
F i r s t e l a s t i c mode Second elastic mode (wing b e n d i n g ) ; (wing f o r e a n d a f t ) ; f = 6 . 7 5 Hz f = 3 . 4 5 HZ T h i r d e l a s t i c mode; F o u r t h e l a s t i c mode; f = 8 . 7 5 Hz f = 1 0 . 6 5 Hz F i g u r e 9.- Measured node l i n e s for i m p o r t a n t symmetric s t r u c t u r a l modes of model.
I
I
Model on cable mount
f r e e f l i g h t - - -
Model i n I'
I, p e r deg
I %
Reduced frequency, k (b) Pitching moment at fuselage station 68.58.
Figure 11.- Continued.
.10 Model on cable mount .09
Model i n free f l i g h t - - -
.08 .06 . 0 3 .02 .01
200 r
-50 -100 -150 -200 0 - 0 2 - 0 4 - 0 6 .08 .10 .12 .14 .16 .18 .20 Reduced frequency, k (c) Wing bending moment at wing station 18.80.
Figure 1 1 .- Concluded.
Gust d i s t r i b u t i o n Measured
---
.14 Uniform
'16 F
I-
150 c
Reduced frequency, k (a) Normal acceleration at fuselage 68.58.
Figure 12.- Comparison of analytically derived frequency-response characteristics of cable-mounted model for measured and uniform gust e xc i tat ion.
Gust distribution Measured
Uniform - - -
@, deg o -50 -100 -150 I ' -200 0 - 0 2 - 0 4 .06 - 0 8 . I O . I 2 .14 -16 .18 - 2 0 , .
Reduced frequency, k (b) Pitching moment at fuselage station 68.58.
Figure 12 .- Continued.
Gust d i s t r i b u t i o n .10 - Measured
Uniform - - -
.09 .08 .07 ‘ b
-
, per deg .05
-
E g .04 . 0 3 .02 .01 100 1 5 0 L 1 .
Reduced frequency, k (c) Wing bending moment a t wing s t a t i o n 18.80.
F i g u r e 1 2 .- Concluded.
- 1 6 Gust distribution
r
-
.14
t Two dimensional Two dimensional
Uniform ---
.12 .10 .08 / .06 .02
I I I I I
@, deg 0 -50 -100 -150 C -200 Reduced frequency, k (a) Normal acceleration at fuselage station 68.58.
Figure 13.- Comparison of analytically derived frequency-response characteristics for two-dimensional and uniform gust excitation.
Gust d i s t r i b u t i o n Gust d i s t r i b u t i o n r Two dimensional Two dimensional
Unif o m --- Unif o m ---
r
25 -
20 -
I . I I I 1
a Reduced frequency, k (b) Pitching moment at fuselage station 68.58.
Figure 13.- Continued.
.10 -
Gust distribution
-
.09 Two dimensional
---
Unif o m -50 -100 -150 b -200 0 .02 .04 .06 .08 .10 .12 .14 .16 .18 .20 Reduced frequency, k (c) Wing bending moment a t wing s t a t i o n 18.80.
F i g u r e 1 3 .- Concluded.
... ... ,. .. ,_..-_________..._
.30 .25
Analytical -
.20 Experimental 0 0 0 Cn .15
li,l ’ per deg
.10 .05 -50
- 1 5 0 1 , : 8 f
f
-200 I I 1 1 1 - I 1 -J
0 .02 .04 .06 .08 .10 .12 .14 .16 .18 .20 Reduced frequency, k (a) Normal acceleration at fuselage station 9.65.
Figure 14.- Comparison of model uncorrected experimental and analytical frequency-response characteristics.
.30 Analytical Experimental .25 Run 1 0 Run 2 0 .20 .10 .05
2oo 150 L
-50 -100 -150 -200
0 .02 .04 .06 .08 . 10 .12 . 1 4 .16 .18 .20
Reduced frequency, k (b) Normal acceleration at fuselage station 68.58.
Figure 14 .- Continued.
, . -._ .-. ..- . - _.
I , ..I I I.
Analytical
.30 r
Experimental 0 .25 .20 , per deg .15 .10 .05
I I 1 1 1 1 1 1 1 1
4 , deg 0 -50
- 100
-150 \ -200 0 .02 .04 .06 .08 .10 .12 .14 .16 .18 .20 Reduced frequency, k (c) N o r m a l acceleration at fuselage station 135.89.
Figure 14 .- Continued.
-4 Analytical - 3 . 0 x 10
-
Experimental Run 1 0 Run 2 2.5 2.0 1.0 .5
I I 1 I I I I I
9 , deg 0 -50 -100 -150
L
-200 0 .02 .04 .06 .08 .10 .12 .14 .16 .18 .20 Reduced frequency, k (d) Pitching moment at fuselage station 68.58.
Figure 14 .- Continued.
I .10 Analytical Experimental 0 .09 .08 .07 .04 .03 .02 .01 (e) Wing bending moment at wing station 18.80.
Figure 14 .- Continued.
Analytical Experimental 0
2oo r
.-
I
50 --
0, deg 0
I I I I I I
0 .02 .04 .06 .08 .10 . 1 2 . 1 4 .16 .18 .20 Reduced frequency, k (f) Wing bending moment a t wing s t a t i o n 6 9 . 4 3 .
F i g u r e 1 4 .- Concluded.
Model (uncorrected) 0 Model (corrected) A .25
Flight test +
030 1
150 c
B
-
100 B
-
A -50
- e
.d u Q -100 -0 -r-J u o u *r( M .d M O 0 Wing fore and aft I 0 -02 .04 .06 .08 .lO .12 .14 .16 .18 - 2 0 Reduced frequency, k (c) Normal acceleration at fuselage station 135.89.
Figure 15.- Continued.
Model (uncorrected) 0 Model (corrected) A
Flight test + +
4 c
A - -
Q A
-50
6 g @ Q e m &AA 4 4
M e .rl E O 3 M o u -0
e w w - 100
& a
u o u .rl M .d A M O M
a E2
e - 0 S Z Z 4 g g .d al 40 -4 c
1 ; F z l P s a
-150 sa
-200 I + I 4 1 I I I I I I 1
0 .02 .04 .06 .08 . IO - 1 2 .14 .16 .18 .20
Reduced frequency, k (d) Pitching moment at fuselage station 68.58.
Figure 1 5 .- Continued.
.10 Model (uncorrected) 0 .09 Model (corrected) A F l i g h t test
.08 1
8%0
O o A
100 150 E
A * ' A 0 A
-200 I 1 I I + u
-04 .06 .08 . l o .12 .14 .16 .18 -20 0 .02 Reduced frequency, k (e) Wing bending moment at wing station 18.80.
Figure 1 5 .- Continued.
Model (uncorrected) 0
r
Model (corrected) A
F l i g h t test +
A A
0 8
I 1 I I 4 1 1
loot
A n O a J o u w w M ! z Q C 3 m
I 1 'I I I I
0 .02 .04 .06 .08 .10 .12 .14 .16 .18 .20 Reduced frequency, k (f) Wing bending at wing station 69.43.
Figure 15.- Concluded.
Model instrumentation Transducer group 1
I O v -+TCG. 80 t o r s i o n
L. W. S. 18.80 bending Transducer group 2 28 V +14 V 5 V I- r Transducer selector Switch
10 v - R. W. S. 69.43
control c o n t r o l unit
- Blank
4 - n k Limit 5 V 2 8 V Coaxial 1 i decoder Transducer to model group 3 +14 v F. S. 135.89 accelerometer motor motor 28 v F. S. 68.58 accelerometer 5 s h a n n e l 5 signals
28 F. S. 9.65 accelerometer p' A b
~ +i--- - v 4 V -14 V 5 V 28 V +14 V Blank
- y r 28 V
inverter 28 V +14 V k 1 4 V converter converter converter converter converter converter 1 Copper-clad .
L-- -
snubber cables Figure 16.- B l o c k diagram of microelectronic data system.
Control room instrumentation I I
- Control data . -
_. . .
Clock 4 5-channel d r k s c i ! a t n g s v a n e
I Multiplexed modulator
I
c a r r i e r A U I Receiver decoder
I - ‘m
+14 V 5 V 28 V
I
b> address
I I A b Audio Time base I 4 V -14 V
I
I
c a r r i e r playback
I
I
I
I ZO-channel
Coaxial cable
ana log Demodulator .
m ulticoupler ‘-.i+q
to model output
I
Control w o r d
I
f
I 20-c ha n ne1 I t r i m generator I I
I
I I power I 4 v 28 v
I
I
lCOi3
converter converter I I
-
i F i g u r e 16.- Concluded.
, IC L-79-311 Figure 17.- Arrangement of microelectronic data system components in model.
._ _ _ 2. Government Accession No. 3. Recipient’s Catalog No.
1. Report No.
NASA TP-1501
I
5. Report Date 4. Title and Subtitle November 1979 EVALUATION OF A WINDTUNNEL GUST RESPONSE TECHNIQUE 6. Performing Organization Code INCLUDING CORRELATIONS W I T H ANALYTICAL AND FLIGHT TEST RESULTS 8. Performing Organization Report No.
7 Authorb) L-13137 L. Tracy Redd, P e r r y W. Hanson, and Eleanor C. Wynne 10. Work Unit No.
- 9. Performing Organization Name and Address 505-33-53-01 NASA Langley Research C e n t e r 11. Contract or Grant No.
Hampton, VA 23665 -~ 13. Type of Report and Period Covered 12. Sponsoring Agency Name and Address T e c h n i c a l Paper N a t i o n a l A e r o n a u t i c s and Space A d m i n i s t r a t i o n 14. Sponsoring Agency Code Washington, DC 20546 15. Supplementary Notes - - 16 Abstract A wind-tunnel t e c h n i q u e for o b t a i n i n g g u s t frequency-response f u n c t i o n s f o r u s e i n p r e d i c t i n g t h e r e s p o n s e of f l e x i b l e a i r c r a f t to a t m o s p h e r i c t u r b u l e n c e is e v a l u a t e d by comparing t h e t u n n e l t e s t r e s u l t s f o r a dynamically s c a l e d c a b l e - s u p p o r t e d aero- e l a s t i c model w i t h a n a l y t i c a l and f l i g h t d a t a . The wind-tunnel t e c h n i q u e , which employs o s c i l l a t i n g vanes i n t h e t u n n e l t h r o a t s e c t i o n to g e n e r a t e a s i n u s o i d a l l y v a r y i n g flow f i e l d around t h e model, w a s e v a l u a t e d by use of a 1/30-scale model of t h e €3-52E a i r p l a n e , for which c o n s i d e r a b l e f l i g h t g u s t r e s p o n s e d a t a were a v a i l a b l e .
The s t u d i e s show good c o r r e l a t i o n between t h e wind-tunnel r e s u l t s , f l i g h t t e s t r e s u l t s , and a n a l y t i c a l p r e d i c t i o n s for r e s p o n s e i n t h e s h o r t - p e r i o d and wing f i r s t e l a s t i c modes of motion, which are t h e modes of p r i m a r y s i g n i f i c a n c e for r e s p o n s e of f l e x i b l e a i r c r a f t to a t m o s p h e r i c t u r b u l e n c e .
. .
.- 18. Distribution Statement 17. Key Words (Suggested by Author(s)) U n c l a s s i f i e d - Unlimited G u s t Frequency r e s p o n s e Wind-tunnel t e c h n i q u e Atmospheric t u r b u l e n c e I S u b j e c t C a t e g o r y 01 Aeroelas t i c mode -~ 20. Security Classif. (of this page) 19. Security Classif. (of this report1 $5.25 U n c l a s s i f i e d Unc las si f i e d - - ..- - _ _ _ - _. . .
* For sale by the National Technical Information Service. Springfield, Virginia 22161 NASA-Langley, 1979 Postage and Fees Paid THIRD-CLASS BULK RATE National Aeronautics and National Aeronautics and Space Administration Space Administration NASA451 Washington, D.C.
20546 Official Business Penalty for Private Use, $300 1 1 7u,zL, 7 1 1 3 7 9 500903DS
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