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LARGE-SCALE WIND-TUNNEL TESTS OF A N AIRPLANE MODEL WITH FOUR PROPELLERS AND ROTATING CYLINDER FLAPS
by Jumes A. Weiberg and Stunley 0. Dickinson
Ames Reseurcb Center Moffett Field CuZzF
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. APRIL 1970
I -
https://ntrs.nasa.gov/search.jsp?R=19700013803 2020-04-05T02:12:33+00:00Z
TECH LIBRARY KAFB, NM
0131521
1. Report No. 2. G o v e r n m e n t A c c e s s i o n N o ,
NASA TN D-5142
4. T i t l e a n d S u b t i t l e LARGE-SCALE WIND-TUNNEL TESTS OF AN AIRPLANE MODEL WITH FOUR
PROPELLERS AND ROTATING CYLINDER FLAPS
7. A u t h o d s ) James A. Wejbergand Stanley 0. Dickinson
9. Pe r fo rming Organ iza t ion Name and Addres s
NASA Ames Research Center Moffett Field, CaliI., 94035
~ ~
2 . Sponsor ing Agency Name and Addres s
National Aeronautics and Space Washington, D. C. 20546
Administration
5. S u p p l e m e n t a r y N o t e s
3. R e c i p i e n t ' s C a t a l o g No. ~ . . " ~ ~ .
5. R e p o r t D a t e April 1910
6. P e r f o r m i n g D r g a n i z a t i o n C o d e
0 . P e r f o r m i n g O r g a n i z a t i o n R e p o r t No. A-2169 ,
I O . Work Uni t No. 721-01-00-13-00-21
11. C o n t r a c t or Gron t No.
13. T y p e of R e p o r t a n d P e r i o d C o v e r e d
Technical Note
14. S p o n s o r i n g A g e n c y C o d e
7. Key Words Suggested by Author(s) ~ ~~
Rotating cylinder flaps 18. D i s t r i b u t i o n S t a t e m e n t
Unclassified - Unlinlited
~ .-
9. S e c u r i t y C l a s s i f . ( a f t h i s r e p o r t )
$3 .00 11 Unclassified Unclassified
22. P r i c e * 21- No. of P a g e s 20. Secur i ty Classif. (of t h i s p a g e )
'For sale by the Clear inghouse for Federal Scient i f ic and Technical Informat ion Springfield, Virginia 22151
r
NOTATION
b wing span, 26.17 ft
BLC boundary-layer control
CD drag coefficient, - drag qs
chf flap hinge-moment Coefficient, - HM
qsf Cf
CL lift coefficient, - lift qs aCL a, CL, lift-curve slope, -
C lift due to flap deflection, - aCL L6 f
Cm pitching-moment coefficient, - M qs c
P pn3D
CP power coefficient, ~
CT thrust coefficient, - T pn2D4
CY side-force coefficient, - Y qs
c wing chord, 4.89 ft
cf flap chord aft of cylinder center line, 2.75 ft
CH horizontal-tail chord, 4.15 ft
D propeller diameter, 4.77 ft
HM flap hinge moment, ft-lb
J propeller advance ratio, - V nD
M pitching moment about 0 . 3 5 c, ft-lb
n propeller angular velocity, rps
N number of propellers
A-2769 iii
P power, f t - lb/sec
q free-stream dynamic p r e s s u r e , p s f
r p r o p e l l e r b l a d e s t a t i o n , f t
R p r o p e l l e r b l a d e r a d i u s , f t , and r e s u l t a n t f o r c e , l b
RCF r o t a t i n g c y l i n d e r f l a p
wing a rea , 128 sq f t
f l a p a r e a , 6 0 . 5 sq f t
t h r u s t , l b
t h r u s t c o e f f i c i e n t , - qs T
c y l i n d e r p e r i p h e r a l v e l o c i t y , f p s
f r ee - s t r eam ve loc i ty , fp s
d e s c e n t v e l o c i t y , fpm
s i d e f o r c e , l b
fuse l age ang le o f a t t ack , deg
l i f t e f f ec t iveness pa rame te r , - c L 6 ( r e f . 1 ) cL c1
prope l l e r b l ade ang le a t 0.75R, deg
f l i gh t -pa th ang le , deg
a f t f l a p d e f l e c t i o n , d e g
f l a p d e f l e c t i o n , d e g
t h r u s t t i l t , deg
wing t i l t , deg
downwash angle , deg
s l ip s t r eam tu rn ing ang le , deg
mass d e n s i t y of a i r , s l u g s / c u f t
i v
I-
LARGE-SCALE WIND-TUNNEL TESTS OF AN AIRPLANE MODEL WITH
FOUR PROPELLERS AND ROTATING CYLINDER FLAPS
By James A. Weiberg and S tan ley 0. Dickinson
Ames Research Center
SUMMARY
Wind-tunnel tes ts were made of a model equipped with rotat ing cyl inder f l a p s and fou r p rope l l e r s . Compar i son w i th t he r e su l t s ob ta ined w i th t he same f l a p on a model with two p r o p e l l e r s o f larger diameter but with approx- imate ly the same s l ips t ream coverage on t h e wing showed t h a t l a r g e g a i n s were ob ta ined i n t he l i f t due t o t h r u s t and s l i p s t r e a m t u r n i n g a s a r e s u l t o f t h e i n c r e a s e d f l a p c h o r d r e l a t i v e t o p r o p e l l e r d i a m e t e r . The s l i p s t r e a m t u r n i n g ang le s fo r t he fou r -p rope l l e r model were approximately equal to the f lap d e f l e c t i o n .
INTRODUCTION
The sma l l - sca l e and two-dimensional t e s t s o f r e f e r e n c e 2 i n d i c a t e d t h a t a r o t a t i n g c y l i n d e r b u i l t i n t o t h e l e a d i n g e d g e o f a f lap could p rovide l a rge g a i n s i n f l a p l i f t with low ro ta t ing cy l inde r power , low l o n g i t u d i n a l trim requirements, and low f l a p h i n g e moments. Based on t h e s e d a t a , t e s t s were made of a large-scale , three-dimensional , two-propel lered model with a r o t a t - i n g c y l i n d e r f l a p (RCF) . The. r e s u l t s a r e p r e s e n t e d i n r e f e r e n c e 3 . Addi- t i o n a l t e s t s were made o f t h e r o t a t i n g c y l i n d e r f l a p on a model having four propel le rs o f smal le r d iameter bu t p rovid ing approximate ly the same s l i p - stream coverage on the wing. For this model with reduced propel ler d iameter t he r a t io o f f l ap chord t o p rope l l e r d i ame te r was g r e a t e r t h a n f o r t h e model of r e fe rence 3 and hence would provide greater s l ipstream turning as indicated i n r e f e r e n c e 4 . The t e s t s were made i n t h e Ames 40- by 80-Foot Wind Tunnel. The inves t iga t ion i nc luded an examinat ion of the e f fec ts o f cy l inder speed , p r o p e l l e r t h r u s t , n a c e l l e s p a c i n g , wing leading-edge s l a t s , wing and n a c e l l e t i l t , f u s e l a g e s t r a k e s , and h o r i z o n t a l - t a i l l o c a t i o n on f l a p e f f e c t i v e n e s s , s t a l l c h a r a c t e r i s t i c s , l o n g i t u d i n a l s t a b i l i t y , s l i p s t r e a m t u r n i n g , and descent c h a r a c t e r i s t i c s .
MODEL
The model f o r t h e s e t e s t s shown i n f i g u r e 1 i n c o r p o r a t e d t h e r o t a t i n g c y l i n d e r f l a p d e s c r i b e d i n r e f e r e n c e 3 . The geometry and dimensions of the model a r e g i v e n i n f i g u r e 2 . A s shown i n t h i s f i g u r e , two n a c e l l e l o c a t i o n s
. _-._ " "
' P ro fes so r A. Alvarez-Calderon was a c o n s u l t a n t f o r t h e d e s i g n o f t h e r o t a t i n g c y l i n d e r f l a p u s e d on both the two and the four -propel le red models .
were t e s t e d on t h e model. The wing could be t i l t e d , 20" leading edge up and t h e n a c e l l e s c o u l d b e t i l t e d 15" nose down. The pivot axes are shown i n f i g u r e 2 .
Details o f t h e f l a p are shown i n f i g u r e s l (c) and 2(b) . A 10.9-inch- diameter machined aluminum cylinder (0.25-inch wall) was b u i l t i n t o t h e l e a d - i ng edge o f t he f l ap . The c y l i n d e r was in four segments , each 65 .1 inches long and dr iven by an e lectr ic motor . Each segment had 18.2-inch-diameter d i s k s on the ends a t t h e wing t i p s and f u s e l a g e . The cylinder segments were sepa ra t ed 0 .25 i nch and f i t t ed wi th 13 .4- inch-d iameter d i sks a t mid-semispan. The f l a p i n c l u d e d a s l o t t e d a f t p o r t i o n w i t h a chord t ha t was 28 p e r c e n t o f t h e wing chord,
The h o r i z o n t a l t a i l cou ld be l oca t ed i n two p o s i t i o n s r e p r e s e n t i n g a low and a tee t a i l . These pos i t ions are shown i n f i g u r e 2 ( a ) .
Details o f t h e wing leading-edge s l a t , e n d p l a t e s , n a c e l l e f a i r i n g s , and f u s e l a g e s t r a k e s a r e shown i n figure 2 .
The model had 4 three-b laded propel le rs o f 4 .77- foot d iameter d r iven by e l e c t r i c m o t o r s . The geometry of the propel ler b lades i s shown i n f i g u r e 3 ( a ) . The b lade angle a t 0 .75 b lade rad ius was 2 1 . S o .
TESTS AND CORRECTIONS
Tests were made a t free-stream dynamic pressures f rom 0 t o 2 .6 psf (Reynolds number = 1 . 5 m i l l i o n ) . The d a t a f r o m t h e s e t e s t s i n c l u d e t h e d i r e c t p r o p e l l e r f o r c e s as well as the aerodynamic forces . The p r o p e l l e r c h a r a c t e r - i s t i c s a r e g i v e n i n f i g u r e s 3 ( b ) and 3(c) . Forces and moments a r e p r e s e n t e d about the wind axes for a moment c e n t e r l o c a t e d as shown i n f i g u r e 2 ( b ) .
Tunnel -wal l cor rec t ions were no t appl ied to the da ta because the r e l a t i v e s i z e o f t h e model and t h e wind tunnel was wi th in t he boundar i e s i nd i - c a t e d i n r e f e r e n c e 5 fo r bes t co r re l a t ion be tween w ind- tunne l and f l i g h t - t e s t r e s u l t s . The convent iona l tunnel -wal l cor rec t ions a re :
c1 = flu + 0 . 3 CL
CD = CD, + 0.0052 C L ~
where t h e s u b s c r i p t u s t ands fo r uncor rec t ed da t a .
No c o r r e c t i o n s were made f o r e f f e c t s of t h e model suppor t s as t h e s e e f f e c t s were e s t i m a t e d t o b e small ( C D ~ ~ ~ ~ = 0.027) .
C y l i n d e r r o t a t i o n a l power inpu t was determined from measurements of e l e c t r i c a l power inpu t t o t he d r ive mo to r s and co r rec t ed fo r mo to r e f f i c i ency obtained from a dynamometer c a l i b r a t i o n of the motors .
2
Flap hinge moments were ob ta ined from measurements with s t r a i n gages on t h e f l a p a c t u a t o r arms.
RESULTS AND DISCUSSION
The d a t a are p r e s e n t e d i n f i g u r e s 4 t o 27 as l i s t e d i n t a b l e 1.
F lap E f fec t iveness
The e f f e c t o f c y l i n d e r r o t a t i o n on l i f t is shown i n f i g u r e 4 . Cylinder speed is expressed as a r a t i o o f c y l i n d e r s u r f a c e s p e e d t o free-stream veloc- i t y U / V . A t low va lues o f U/V, l i f t i nc reases r ap id ly w i th i nc reas ing U/V. u n t i l t h e f l o w on t h e f l a p is a t tached , cor responding to the knees o f the cu rves i n f i gu re 4 . The r a t i o U / V f o r a t t a c h e d f l o w f o r a g i v e n f l a p d e f l e c t i o n was t h e same as fo r t he two-p rope l l e red model o f r e f e rence 3 . The r a t i o U / V fo r a t t ached f l ow was dependent only on f l a p d e f l e c t i o n and was independen t o f ang le o f a t t ack , p rope l l e r s l i p s t r eam effects , wing t i l t , n a c e l l e t ilt , and n a c e l l e s p a c i n g .
For t he r ema inde r o f t he f i gu res where in U / V i s cons t an t , t he va lue o f U / V given on t h e f i g u r e s a n d i n t a b l e 1 i s a t or above t h e v a l u e f o r a t tached f low. The power r e q u i r e d t o r o t a t e t h e c y l i n d e r i s shown i n f i g u r e 5 ( a ) and is t h e same as fo r t he two-p rope l l e r model of r e f e r e n c e 3 . Figures 5(b) and 5 (c ) show the r e l a t ionsh ip be tween cy l inde r speed , a i r speed , and U/V.
L i f t and drag due t o f l a p d e f l e c t i o n and t h r u s t a r e shown i n f i g u r e 6 . The d a t a i n t h i s f i g u r e a r e f o r a va lue o f U / V a t o r above t h a t r e q u i r e d f o r a t tached f low on t h e f l a p . A t z e r o t h r u s t , t h e l i f t due t o f l a p d e f l e c t i o n shows reasonable agreement with the f lap theory of reference 1 f o r f l a p d e f l e c - t i o n s up t o 70". The a p p a r e n t d e t e r i o r a t i o n o f f l a p e f f e c t i v e n e s s a b o v e 70' may b e due t o w i n g - f u s e l a g e j u n c t u r e e f f e c t s s i n c e t u f t s showed at tached f low on t h e f l a p e x c e p t n e a r t h e f u s e l a g e , and the p re sence o f l a rge ex te rna l v o r t i c e s on the fu se l age and n a c e l l e s .
The l i f t due t o t h r u s t f o r 60" f l a p d e f l e c t i o n i s compared i n f i g u r e 7 with the two-propel ler data f rom reference 3 . The l i f t due t o t h r u s t was g r e a t e r f o r t h e model w i t h t h e f o u r p r o p e l l e r s b e c a u s e o f t h e g r e a t e r t u r n i n g o f t h e s l i p s t r e a m f o r t h e l a r g e r r a t i o o f f l a p c h o r d t o p r o p e l l e r d i a m e t e r . Inc luded i n f i gu re 7 are es t imated va lues f rom re ference 4 and comparisons with a blowing BLC f l a p from reference 6 . For the same f l a p c h o r d t o p r o - p e l l e r d i a m e t e r r a t i o , t h e l i f t due t o t h r u s t i s t h e same f o r t h e RCF and t h e blowing BLC f l a p and i s es t imated reasonably well by the theory o f re fe rence 4 .
The a b i l i t y o f t h e a f t f l a p t o p r o v i d e l i f t c o n t r o l i s shown i n f i g u r e 8 f o r a 60" main f l a p d e f l e c t i o n . An a f t f l a p d e f l e c t i o n o f 20" provides a cap- ab i l i t y o f abou t 0 .13 g normal acce le ra t ion . Above 20" d e f l e c t i o n , t h e e f f e c - t iveness decreased . A t h i g h t h r u s t c o e f f i c i e n t s , t h e a f t f l a p is more e f f e c t i v e f o r d r a g c o n t r o l t h a n l i f t c o n t r o l .
3
" .,."."".. _........-... .. ..._.... ....._.
S t a l l C h a r a c t e r i s t i c s
Observa t ions o f the f low about the wing , f laps , and fu se l age w i th . t u f t s and streamers i n d i c a t e d t h a t m a x i m u m l i f t appea red t o be l imi t ed by fu se l age f l o w d e t e r i o r a t i o n r a t h e r t h a n f l a p o r wing s t a l l . Various a t tempts were made t o c o n t a i n o r d e l a y t h i s f l o w i n t e r f e r e n c e t o h i g h e r a n g l e s o f a t t a c k b y t h e use of s la ts , r educed nace l l e spac ing , wing tilt , and fu se l age s t r akes .
To inc rease t he p rope l l e r s l i p s t r eam cove rage nea r t he fu se l age , t he n a c e l l e s were moved inboa rd (nace l l e spac ing B, f i g . 2 ( a ) ) . The e f fec t o f t h e reduced nacel le spacing with 70" f l a p d e f l e c t i o n i s shown i n f i g u r e 9 . The comparison shown i n t h i s f i g u r e a l s o i n c l u d e s t h e e f fec t o f p r o p e l l e r r o t a t i o n ; however, f i g u r e 1 3 ( a ) shows t h a t , a t least f o r 80" f l a p d e f l e c t i o n , t h i s e f fec t would be small. Figure 9 t h u s i n d i c a t e d t h a t , a t a t h r u s t c o e f f i c i e n t o f 4 , the c loser nace l le spac ing improved the f low over the wing and f l a p n e a r t h e f u s e l a g e and inc reased maximum l i f t .
The e f f ec t o f f u l l - s p a n , wing leading-edge s l a t s i s shown i n f i g u r e 10 . With 60" f l a p d e f l e c t i o n , t h e s l a t s inc reased maximum l i f t a t t h r u s t c o e f f i - c i e n t s o f 2 o r less . With a h i g h e r t h r u s t c o e f f i c i e n t o r h i g h e r f l a p d e f l e c - t i o n t h e f u l l - s p a n s l a t was de t r imen ta l . I t is n o t known i f t h e s l a t would be more e f f e c t i v e i f t a i l o r e d t o t h e h i g h e r t h r u s t c o e f f i c i e n t s . A l s o , t e s t e d were p a r t i a l s p a n s l a t s between the inboard nacel les and the fuselage on t h e model w i th t he r educed nace l l e spac ing (nace l l e spac ing B , f i g . 2 ( a ) ) a n d f l a p s 70". This s l a t conf igura t ion d id no t improve l i f t o r s t a l l cha rac t e r - i s t i c s ( f i g . l O ( c ) ) .
The effect o f wing tilt is shown i n f i g u r e 11. The d a t a i n t h i s f i g u r e show t h a t t h e wing a t t i t u d e r e l a t i v e t o t h e f u s e l a g e d i d n o t a p p e a r t o a f f e c t t h e wing m a x i m u m l i f t c a p a b i l i t y w i t h f l a p s d e f l e c t e d .
An a t tempt was made t o improve the f low around the large angular fuselage by large strakes on the forward lower par t o f the fuse lage ( f ig . 2 (a ) ) . These s t r a k e s improved the fuse lage f low and increased maximum l i f t as shown i n f i g u r e 1 2 .
The effect o f d i r e c t i o n o f p r o p e l l e r r o t a t i o n on s t a l l c h a r a c t e r i s t i c s was a l s o i n v e s t i g a t e d . Two modes o f r o t a t i o n were t e s t e d : a l l l i k e r o t a t i o n , and with a l l inboard blades down going. The r e s u l t s a r e shown i n f i g u r e 1 3 . The counterrotat ing arrangement with down-going b l a d e s n e x t t o t h e f u s e l a g e r educed f l ow sepa ra t ion i n t ha t r eg ion bu t d id no t affect t h e l o n g i t u d i n a l f o r c e c h a r a c t e r i s t i c s . The counter ro ta t ing p ropel le r a r rangement , however , g r e a t l y r e d u c e d t h e s i d e - f o r c e v a r i a t i o n w i t h a n g l e o f a t t a c k ( f i g . 1 3 ( b ) ) s i m i l a r t o t h e r e s u l t s shown i n r e f e r e n c e 7 .
Long i tud ina l S t ab i l i t y and Con t ro l
The effect o f h o r i z o n t a l - t a i l l o c a t i o n on s t a b i l i t y i s shown i n f i g u r e 1 4 f o r a f l a p d e f l e c t i o n o f 80". L a r g e v a r i a t i o n s i n t h e t a i l - o n s t a b i l i t y a C m / a C L are caused by changes i n t h e downwash parameter ae/act with Tc ' ( s i m i l a r t o t h e r e s u l t s shown i n ref . 6 ) .
4
The model w i th t he low t a i l shows l a r g e r v a r i a t i o n s o f aC / X L with Tcl and becomes u n s t a b l e a t h i g h Tc' (moment c e n t e r a t 0.35 cy. The magni- t ude o f t he t a i l pitching-moment contribution was approximate ly the same f o r both t a i l p o s i t i o n s . F o r f l a p d e f l e c t i o n s o f 6 0 " o r less a t low angles of a t t a c k ( f i g s . 1 7 a n d 1 8 ) , t h e model w i th t he low t a i l was uns tab le , even wi th power o f f (Tcl = 0 ) . The r e s u l t s o f r e f e r e n c e 8 i n d i c a t e t h a t t h i s may b e t h e r e su l t o f t he w ing- fuse l age wake impinging on t h e t a i l . The tee t a i l provided a p o s i t i v e c o n t r i b u t i o n t o s t a b i l i t y ( f i g . 1 4 ( b ) ) .
Zero Ai rspeed Charac te r i s t ics
A measu re o f t he e f f ec t iveness o f a f l a p i s i t s a b i l i t y t o c o n v e r t t h r u s t i n t o l i f t by de f l ec t ion o f t he s l i p s t r eam. The s l i p s t r eam tu rn ing e f f ec t ive - nes s a t zero forward speed i s shown i n f i g u r e 1 5 . I n c l u d e d i n t h i s f i g u r e are es t imated va lues f rom re ference 4 , comparisons with the two-propeller RCF da t a o f r e f e rence 3 , and d a t a f o r a four -propel le r b lowing BLC f l a p from r e f e r e n c e 6 . The RCF o n t h e f o u r - p r o p e l l e r model w i t h a l a r g e r f l a p c h o r d r e l a t i v e t o t h e p r o p e l l e r d i a m e t e r p r o v i d e d g r e a t e r s l i p s t r e a m t u r n i n g f o r a g i v e n f l a p d e f l e c t i o n t h a n e i t h e r t h e RCF on the two-propel le r model o r t h e blowing BLC f l a p . The d a t a i n f i g u r e 15 a r e p l o t t e d w i t h f l a p d e f l e c t i o n r e f e r e n c e d t o t h r u s t a x i s a n d show t h a t s l i p s t r e a m t u r n i n g a t zero a i r speed i s a f u n c t i o n o f t h e f l a p d e f l e c t i o n r e l a t i v e t o t h e t h r u s t a x i s . On t h e f o u r - p r o p e l l e r RCF model, 74O s l ip s t r eam tu rn ing a t 87 -pe rcen t t h rus t r ecove ry was obtained with 70" f l a p d e f l e c t i o n ( a f t f l a p d e f l e c t e d 20').
The r o t a t i n g c y l i n d e r f l a p model provided a g i v e n s l i p s t r e a m t u r n i n g w i t h l e s s p i t c h i n g moments than the blowing BLC f l a p . With t h e f o u r - p r o p e l l e r RCF model, these moments were zero for a moment c e n t e r l o c a t e d a t 0.35 c.
The d a t a i n f i g u r e 1 5 i n d i c a t e t h a t t h e model w i th 80" f l ap de f l ec t ion can hover with approximately 20° ang le o f a t t ack o r w ing tilt . A t forward f l i g h t f o r t h e model w i t h t h e w i n g t i l t e d , minimum drag occurs a t a large neg- a t i v e f u s e l a g e a n g l e o f a t t a c k ( f i g . 2 3 ( b ) ) . Less nega t ive fu se l age ang le o f a t t a c k i s r e q u i r e d w i t h n e g a t i v e f l a p d e f l e c t i o n .
Descen t Cha rac t e r i s t i c s
The e f f e c t s of f l a p s , s la t s , and wing and t h r u s t tilt on the descent c h a r a c t e r i s t i c s a r e shown i n f i g u r e 1 6 . I n t h i s f i g u r e , t h e d e s c e n t a n g l e and descent veloci ty Vs at C L ~ ~ ~ a r e shown as f u n c t i o n s o f f l a p d e f l e c t i o n
f o r an assumed wing loading W/S of 40 ps f . F igu re 16 shows t h a t t h e d e s c e n t angle a t QmaX i nc reased w i th bo th main f l a p and a f t f l a p d e f l e c t i o n .
Add i t iona l ga ins i n descen t capab i l i t y can be ob ta ined i f the inboard and ou tboa rd p rope l l e r s are ope ra t ed a t d i f f e r e n t t h r u s t s e t t i n g s t o v a r y t h e s p a n - wise d i s t r i b u t i o n o f l i f t and drag as shown i n r e f e r e n c e 8 . The spanwise dis- t r i b u t i o n o f t h r u s t was v a r i e d b y o p e r a t i n g t h e p r o p e l l e r s a t d i f f e r e n t r o t a - t iona l speeds wi th the b lade angle he ld cons tan t . The e f f e c t o f d i f f e r e n t i a l
5
p r o p e l l e r t h r u s t on the descent angles and descent ve loc i ty a t C h a x i s shown i n f i g u r e 1 6 ( c ) . The d i f f e r e n t i a l p r o p e l l e r t h r u s t , f o r a p p r o x i m a t e l y t h e same a v e r a g e t o t a l t h r u s t , r e s u l t e d i n a l o s s i n l i f t ( f i g . 27 ) . Drag a t maximum l i f t , however, was inc reased g iv ing an i nc rease i n descen t ang le o f about 4" ( f ig . 16 (c ) ) . Descen t ang le s o f ove r 20" were ob ta ined a t maximum l i f t .
CONCLUDING REMARKS
The r e s u l t s h a v e shown t h a t t h e r o t a t i n g c y l i n d e r f l a p c a n p r o v i d e t h e la rge va lues o f s l ips t ream turn ing and l i f t due t o t h r u s t i n d i c a t e d by t h e l a r g e f l a p c h o r d r e l a t i v e t o p r o p e l l e r d i a m e t e r . Hover could be achieved with moderate amounts of angle of a t t a c k o r wing inc idence . A t f o r w a r d v e l o c i t i e s , the d rag due to wing incidence could be reduced by small amounts o f nega t ive f l a p d e f l e c t i o n . F u s e l a g e f l o w d e t e r i o r a t i o n a p p e a r e d t o limit maximum l i f t and could reduce the e f fec t iveness o f the hor izonta l t a i l f o r some v e r t i c a l l o c a t i o n s . The e f f ec t s o f t h i s f l ow de t e r io ra t ion cou ld be r educed by fuse l age s t r akes and by the spac ing be tween the p rope l le rs and fuse lage .
Ames Research Center National Aeronautics and Space Administration
Mof fe t t F i e ld , Calif. 94035, Dec. 10, 1969
6
I
REFERENCES
1. DeYoung, John: Theoretical Symmetric Span Loading Due t o F l a p D e f l e c t i o n f o r Wings of A r b i t r a r y Plan Form at Subsonic Speeds. NACA Rep. 1071, 1952. (Supersedes NACA TN 2278.)
2 . Alvarez-Calderon, Alberto; and Arnold, Frank R. : A Study of t h e Aero- dynamic C h a r a c t e r i s t i c s o f a High-Lift Device Based on a Ro ta t ing Cylinder and Flap. Stanford Universi ty Tech. Rep. RCF-1, 1961.
4 . Kuhn, Richard E . : Semiempir ical Procedure for Est imat ing L i f t and Drag Character is t ics of Propel ler-Wing-Flap Configurat ions. f o r Ver t ica l - and Short- Take-Off-and-Landing Airplanes. NASA MEMO 1-16-59L, 1959.
5 . Cook, Woodrow L . ; and Hickey, David H . : Comparison of Wind-Tunnel and F l igh t -Tes t Aerodynamic Data i n t h e T r a n s i t i o n - F l i g h t S p e e d Range f o r Five V/STOL Aircraf t . Paper 26 , NASA SP-116, 1966.
6 . Weiberg, James A . ; and Page, V . Robert: Large-Scale Wind-Tunnel Tests of an Airplane Model With an Unswept, Aspect-Ratio-io Wing, Four Propel le rs , and Blowing Flaps. NASA TN D-25, 1959.
7 . Weiberg, James A . ; and G i u l i a n e t t i , Demo J . : Large-Scale Wind-Tunnel Tes t s o f an Airplane Model With an Unswept T i l t Wing of Aspect Ratio 5 . 5 , and With Various S ta l l Cont ro l Devices . NASA TN D-2133, 1964.
8 . Deckert, Wallace H . ; Page, V . Robert; and Dickinson, Stanley 0 . : Large- Sca le Wind-Tunnel Tes t s o f Descent Performance o f an Airplane Model With a T i l t Wing and D i f f e r e n t i a l P r o p e l l e r T h r u s t . NASA TN D-1857, 1964.
7
TABLE 1.- FIGURE INDEX
i n f :
L
-
6T -
- 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 -
Data 6 f T a i 1 P r o p e l l e r r o t a t i o n
CL vs. u/v Power r e q u i r e d ACL VS. ACL VS. T, 6 f l
ACL VS. 6A Effect o f :
Nacelle spac ing S l a t s Wing t i l t S t r a k e s P r o p e l l e r r o t a t i o n Hor i zon ta l t a i l
S l i p s t r e a m t u r n i n g D e s c e n t c h a r a c t e r i s t i c s
i gu ra t i c 5 6 .6 6 . 6 6 .6 6 . 6 6 . 6 6 . 6 6 . 6 6 . 6 6 . 6 6 . 6 6 . 6 6 . 6 6 . 6 6.6 0 6.6 6 . 6 6 .6 6 .6 6.6 6 . 6
A low, off
o f f o f f o f f o f f low 1 ow o f f o f f low o f f t e e o f f o f f t ee
o f f low
o f f t ee t ee low o f f
COl
0 10 10 0 0 0 0
20 30 30
0 0
20 20 20
0 0 0 0 0 0
10, 20
___- , vs . C L - 0-4 0-4 0-4 0-6 0-4 0-4 0-4 0-6 0-4 0-4 0-6 0-4
0-6 0-6
0-4 0-8 0-6 0-4 0-4 0-4 0-4 0-4 -
- c c c c C C C C C 0 C 0 0 0 0 0 0 0 0 0 0 0 -
o f f o f f
on o f f
on o f f
on o f f
m, off o f f o f f o f f o f f
on o f f o f f o f f
on o f f o f f o f f o f f
40 40 40 60 60 60 60 60 60 60 70 70 70 70 70 80 80 80 80 80 80
10, 70
Conf igu ra t ion B CD, a , Cm VS. C L
o f f o f f 6 . 6 20 70 o f f o f f 6 .6 20 70 o f f o f f 6 . 6 20 70 o f f o f f 6.6 20 55 o f f o f f 4 10 25 o f f o f f 0 -10 - 10 o f f o f f 0 0 0 o f f o f f 0 0 0
I -
0-2 0-2 0-2 0-4 0-6 0-6 0-6 (a)
- 0
20 20 20 20
0 20 20 -
- 0 0 0
15 15 0 0 0
a . D i f f e r e n t i a l t h r u s t T,' = 8 inboa rd , 0 outboard and
T,' = 4 inboa rd , 0 outboard
8
A-40375.1
(a) The model with wing tilted and flaps Oo.
Figure 1.- The model installed in the Ames 40- by 80-Foot Wind Tunnel.
9
(b) Front detail of wing and cylinder.
Figure 1.- Continued.
A-39265.1
I
(c) Rear detail of cylinder and flap.
Figure 1.- Concluded.
Wing Area 128.0 Span 26. I 7 Chord 4.89 Aspect ratio 5.35 Section NACA 5321 (mod.)
Horizontal tail Area 54 .O Span 13.0 Chord 4.1 5 Aspect ratio 3.26 Section NACA 63-216 (inv.) Tail length
1 Low I 5.50 I
Tee I 8.75 1 1
Location of cH/4 Sta. WL
Dimensions in feet Tee 33.6 13.5 Low 30.3 5.5 ,
Nacelle spacing LI L2 ' A 10.92 21-40
B 9.25 19.73
1- 36.4
(a) General dimensions.
Figure 2.- Geometry of the model.
I
50c 7 F y f l a p hinge
Thrust axis - - WL 2.85 f t .
Sta., f t WL, f t 14.35 2.54 Nacelle pivot 14.29 2.20 Wing pivot 14.83 2.56 Moment center
Aft flap hinge
SI
--I .19c I Aft flap /
.17c - 7- - L .33c
End pla te ' (b) F l a p , s l a t , and end-plate geometry.
F igure 2 . - Concluded. 13
8
6
.- c
f .s 4 'D
73 Q)
0 - m
2
0
.6
- g .4 .- 3
'D al
0 R -
Q) U 0 - !ij .2
. I
0
(a) Blade geometry.
Figure 3.- Propeller characteristics.
14
2 4 6 8 IO 12 Free -stream dynamic pressure, psf
(b) Thrust.
Figure 3.- Continued.
15
7
6
5
4
3
2
I
0
.7
.6
.5
.4
. 3
.2
. I
0
Figure 3 . - Concluded.
16
8f, 8 ~ , Td a, Tail deg deg deg
0 40 0 4.2 0 Off 0 70 0 4.3 0 Off 0 80 0 0 0 Low
14
12
IO
8
CL
6
8fl TA a, 8 W ~ Tail deg deg deg deg deq
A 25 I O 4.1 -24 20 -15 Of f h 55 20 2.1 -24 20 -15 Off n 70 20 4.0 0 0 0 Tee n70 20 4.5 o o o off
0 2 4 6 0 IO U v
( b 1 Nacelle spacing B
-
Figure 4.- Effect of cylinder rotation on lift.
1111111
c CL
.c 0 c a a CL
0
c
- c 0
I-"
.-
.I- d e .I- 2, .- 0 0 - 9
60
40
20
n ., ( a ) Cylinder power
IO
8
6
4
2
n " (b) Velocity ratio
0) 0 O
a .I- L
200
0 2 4 6 8 IO Cylinder rpm/1000
( c ) Cylinder surface speed
Figure 5.- Cylinder power requirements.
18
? c
m
0 a
b 0
16
12
?
4
0
Figure 6 . - Flap effectiveness; a = -So. 19
I
ac,
IO
8
6
4
2
0
IO
8
6
4
2
0 2 4
7 0 RCF 0 Blowing BLC (Ref. 6)
Theorv (Ref. 4)
6 Tc '
8 IO 12
Figure 7.- Lift due t o thrust.
20
0 IO 20 30 40 50 S A * deg
Figure 8.- Effect of a f t f l a p d e f l e c t i o n ; cSf = 60°, - = 6.6, a = -10". U v
2 1
N N
0 2 4 6 -20 -10 0 IO X) 2 0 -2 - 4 CD a , deg c m
Figure 9.- Effec t o f nace l le spac ing; 6 = 70°, = 20°, - U = 6.6, t a i l o f f . f v
I
- 2 - I 0 I 2 CD
20 IO 0 IO 20 2 I 0 - I a, deg cnl
(a) 6f = 60’; 6 = O’, t a i l o f f . A
Figure 10.- Effect of s la t s ; nace l le spac ing A, - = 6 . 6 . U v
12
I I
IO
9
8
CL 7
6
5
4
3
2 - I 0 I 2 3 4 -20 -10 0 I O I 0 -I
CD a , deg Cm
(b) 6 f = 70'; 6 = 20°, tail off. A
Figure 10.- Continued.
-10 0 IO 20
(c) Inboard slats; 6f = 70°, = 2 0 ° , tee tail.
Figure 10.- Concluded.
I
Cb 1 ' I I! ! ' u-"l
-2 -I 0 I
CD
.20 -10 0 IO 20 awing, deQ
Figure 11.- Effect of wing tilt; t a i l o f f , nace l l e spac ing B .
(b) 6 f = 70"; 6 A = 20", = 6 .6
Figure 11.- Concluded.
2 I 0 Cm
Figure 1 2 . - E f fec t o f fu se l age s t r akes ; 6 = 7 0 ° , 6 = 20°, - = 6.6 , t e e t a i l , i n b o a r d slats, nacelle spacing A .
U A v
8
7
CL
6
5
4
3
2 0 I 2 3
Figure 13.- Effect of propel ler rotat ion; 6 = 8 0 ° , 6 = O o , f A
3 2 I 0 c m
- = 6 . 6 , t e e t a i l , nacel le spacing A . U V
d 0 2.0 4.3
Propeller rotation
C 'Y
@I cy Figure 13.- Concluded.
30
8
7 CL
6
CD -20 -10 0 IO 20
a , deg 3 2 I 0 - I
Cm
(a) b f = 80"; 6 = 0 " , - - " - 6.6 A V
Figure 14 . - Effect of a horizontal t a i l ; nacel le spacing A .
w N
0 2 4 6 -20 -10 0 IO 4 2 0 -2
(b) 6f = 70'; 6 = 20°, - U = 6 . 6 A v
Figure 14. - Concluded.
0 SA= 18, Hinge I, Props qP -"-Ref.4
A
TD M O
- .4
-
1.0 R T -
.8
0 20 40 60 80 sf , deg
2 Prop RCF; cf/D = 0 . 3 4 (ref. 5)
0 20 40 60 80 100 sf, deg
0 20 40 60 80 sf, deg
(Relative to thrust axis) 4 Prop blowing BLC; cf/D = 0 . 3 8 4 Prop RCF; cf/D = 0.67 (ref. 6)
Figure 15.- Zero airspeed characteristics.
0 0 0 0
> VI
- - I
-3
+ C a, 0 VI a, 0
(a) Effect of flap deflection; = 0'.
Figure 16.- Descent characteristics.
34
I
0
P, a, -0
< -10
- a, P, c 0
c 5 - 2 0
D M
- 30
T 1
1 I
(b) Effect of aft flap; 6f = 60°.
Figure 16.- Continued.
35
I
0 0 0
.4- E,
>* e .- 2.
0 0 a, > - c C a, 0 u)
E
0
- I
- 2
-3
0 I Uniform thrust 0 2 Differential thrust
CY U a,
x
- a, CY
0 c
+ C a, 0 u)
E
0
-10
- 20
4"
0 I 2 3 4 5 T; Average
(c) Effect of differential thrust; 6f = 70°, 6 = 20'. A
Figure 16.- Concluded.
36
Figure 17.- Aerodynamic W
-20 -10 0 IO 20 30 a , deg
(a) = O o , - - V ” - 5.0
characteristics; b f = 40°, nacelle
2 I 0 - I Cm
spacing A.
L
-3 -2 -I 0 I 2 CD
2 I 0 -I C m
(b) = 10'; - - V " - 6.6 , t a i l o f f .
Figure 1 7 . - Continued.
-3 -2 -I 0 I CD
.20 -10 0 I O 20 2 I 0 -I a, deg C m
Figure 17 . - Concluded.
P 0
CL
F i g
14
13
12
I I
IO
8
-3 -2 -I 0 I 2 CD
(a> 6*
;ure 18.- Aerodynamic c h a r a c t e r i
I I I =! "
-20 -10 0 IO 20 "
a, deg
; t a i l o f f .
.s; 6 = 60°, - = 6.6, nac U f v :e l 1
I Cm
.e spacing A .
-3 -2 - I 0 I 2 3 CD
,20 - IO 0 IO 20 30 2 I 0 - I Crn
(b) b A = 0'; tail o f f , slats on.
Figure 18.- Continued.
-20 -10 0 IO 20
b A = oo; low tail.
Figure 18.- Continued.
I
C L
12
I I
IO
9
0
7
6
5
4
3
2 -2 - I 0 I 2 3
CD 2 I 0
Cm
(d) = 0'; low t a i l , slats on.
Figure 18.- Continued. P w
P P
-I 0 I 2 3
(e) = 20'; tail off.
Figure 18.- Continued.
12 - I k i l B i E E E Z
0 4.2 4.2 On
L
-I 0 I 2 3 4
CD
-20 -10 0 IO 20 a, deg
I - 1-
I 0 -I Cm
(f) = 30'; tail o f f .
Figure 18.- Continued. P ul
(g) = 30'; low t a i l .
Figure 18.- Concluded.
T I
13
12
I I
IO
I I
I I I I I
I
L
-2 -I 0 I 2 3 CD
-20 -10 0 IO 20 a, deg
(a) 6A = 0'; tail off.
Figure 19. - Aerodynamic characteristics; 6 = 70°, - = 6.6, U
f v
* '; . n .
2 I 0 -I Crn
nacelle spacing A.
P
4
3
2 -I 0 I 2 3 -20 -10 0 IO 20 2 I 0
CD a , deg Cm
(b) = Oo; tee tail.
Figure 19.- Continued.
14
IO
9
CL
8
7
6
5
4
3~ I 2 3 4 -20 -10 0 IO 20 CD a, deg
(c) = 20"; t a i l o f f .
2 I 0 -I
Figure 19 . - Continued.
0 4.2 A 6.1
Propeller rotation
7 - -2 -I 0 I 2 3
CD
(d) = 20'; tail off, slats on.
I I I ' I
3 2 I 0 -I
Figure 19. - Continued.
C 'L
.20 -10 0 IO Q, deg
(e) = 20'; tee tail.
Figure 19.- Concluded.
2 I 0 -I Crn
U V (a) - = 0; low tail.
Figure 20.- Aerodynamic characteristics; 6 = 80°, 6 = Oo, nacelle spacing f A
-I
A.
C 'L
I I f I
I
I
I
-I 0 I 2 3 4 -20 -10 0 IO 2 I 0 -I CD a, deg Cm
+
(b) - = 6.6; t a i l o f f . U v
Figure 20.- Continued.
Cn P
-I 0 I 2 3 -20 -10 0 IO 20 I 2 0 -I CD a, deg cm
( c ) - U = 6.6; t a i l o f f , s la ts on. v Figure 20.- Continued.
C
I I
IO
9
0
'L
7
6
5 0 0 0
0 2 .o 4.3
Propeller rotation rn 4
I
I 2 3 -20 -10 0 IO 2 I 0 - I CD a, deg Cm
(d) - = 6.6; tee tail.
Figure 20.- Continued.
U V
12
I1
IO
9
8
7
6
5
4
3
2 0 I 2 3
CD
-20 -10 0 I O a, deg
U v (e) - = 6.6; tee taiI
Figure 20.- Continued
3 2 I 0
U (f) v = 6 . 6 ; low tail.
Figure 20 1 . - Concluded.
Ch f
.4
.2
0
..2
.4
.6
f
58
4
3
2
CL
I
0
-I -2 -I 0 I - 20
C D
Figure 22.- Aerodynamic characteristics;
-10 0 IO 20 I 0 -I
(a) = 0"
6 = o o , 6 = oo, - = 0 , U f A V
Lrn
tail o f f , nacelle spacing B.
C L
5
4
3
2
I
0
-I
-3 -2 -I 0 I -40 -30 -20 -10 0 I O I 0 - I , c D a , deg Cm
(b) 6w = 20'
Figure 2 2 . - Concluded.
-3 -2 - I 0 I -30 -20 -10 0 IO 20 I 0 - I CD a, deg Cm
Figure 2 3 . - Concluded.
I
I
-40 -30 -20 -10 0 IO a, deg
Figure 24.- Aerodynamic characteristics; 6 = 2S0, 6 = IOo, - - f A " - 4.0, v nacelle spacing B.
6T = -15', tail o f f ,
" -4 -2 0 2 4
~
-40 -30 -20 -10 0
Figure 25.- Aerodynamic characteristics; 6 = 55", 6 = 2 0 ° , " = 6.6, 6 = 20° , 6 = -15', tail off, f A v W T nacelle spacing B.
Figure 26. -
CL
16
14
12
I O
8
6
4
2
0 2
Aerodynamic
4 6 CD
characteristics;
-20 -10 0 I O 4 2 0 -2 Cm
6f = 7 0 ° , 6 = 2o", v - A " - 6.6, tail o f f , nacelle spacing B.
12
IO
... ... ......... ..... ". ...... * ~ , . . ....... " .. " . . , .......
. _ & _ _
..... - ..
0 2 4 6 ~
.40 -30 -20 -10 0 a
(b) 6w = 20"
Figure 26. - Concluded.
Inboard Outboard 1
2.2 0
0 8.0 0
2
0 2 4 6 40 -30 -20 -10 0 a , deg
Figure 27.- Effect of differential propeller thrust; 6 = 7 0 ° , 6 = 2 0 ° , - = 6.6, tail off, 6w = 20°,
nacelle spacing B.
U f A V
