Document
i
LONGITUDINAL STABILITY A N D CONTROL
CHARACTERISTICS OF A POWERED MODEL OF
A T WIN-PROPELLER DEFLECTED-SLIPSTREAM
STOL AIRPLANE CONFIGURATION
by Richard J. Margason, Alexander D. Hammond, . . , '.. $ I 1
and Garl L. Gentry
\\ ,:
Ldngley Research Center
LangZey Station, Hampton, Va.
,1 z .
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 W A S H I N G T O N , D. C . J U L Y 1 9 6 6 i F - TECH LIBRARY KAFB, NM LONGITUDINAL STABILITY AND CONTROL CHARACTERISTICS OF A POWERED MODEL O F A TWIN-PROPELLER DEFLECTED-SLIPSTREAM STOL AIRPLANE CONFIGURATION By Richard J. Margason, Alexander D. Hammond, and Gar1 L. Gentry Langley Research Center Langley Station, Hampton, Va.
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sole by the Clearinghouse for Federol Scientific and Technical Information Springfield, Virginia 22151 - Price $4.00 LONGITUDINAL STABILITY AND CONTROL CHARACTERISTICS OF A POWERED MODEL OF A TWIN-PROPELLEE DEFLECTED-SLIPSTREAM STOL AIRPLANE CONFIGURATION By Richard J. Margason, Alexander D. H m o n d , and Gar1 L. Gentry Langley Research Center SUMMARY Results are presented of a wind-tunnel investigation of the static longi- tudinal stability and control capabilities of a twin-propeller deflected- slipstream STOL airplane in the take-off and landing speed range through the post-stall region at angles of attack up to 44O.
The results of this investigation show that the magnitudes of the pitching moments of the wing-body combination for the flaps-retracted ( 0 ' flap deflection) configuration were small. The tail-on data show that any of the tail configu- rations with the flaps-retracted configuration provide an adequate stability contribution and are capable of trimming the airplane. The wing-body combina- tion for the flaps-deflected (451' flap deflection) configuration had a large tail lift requirement for longitudinal trim, particularly for the highest power setting of the investigation. The small tail in the high position had the capability of trimming the airplane for the low power conditions (thrust coeffi- cients of 0 and 0.70). At higher power conditions the tail stalled before trim was achieved. The large tail in either position had the capability of trimming the airplane up to the angle of attack corresponding to the maximum lift coeffi- cient for all but the highest power condition (thrust coefficient of 2.42).
This tail stalled before trim was achieved for the highest power condition.
INTRODUCTION Recent interest in developing a small deflected-slipstream short take-off and landing (STOL) airplane has led to a need for stability and control data on this type of configuration. A static wind-tunnel investigation of a powered model of a twin-propeller deflected-slipstream STOL aircraft configuration was conducted to provide some of this information. The lateral control character- istics of this model have been presented in reference 1. The longitudinal stability and control characteristics are presented in the present report.
This investigation was undertaken to determine the longitudinal stability and control characteristics through the angle-of-attack range from -bo into the post-stall region (to 44O).
The investigation was conducted in the 17-foot (5.18 meter) test section
of the Langley 3OO-MPH 7- by 10-foot tunnel and covered two flap deflections
and several power conditions.
SYMBOLS The units used for the physical quantities defined in this paper are given both in the U.S.
Customary Units and in the International System of Units (SI).
Factors relating these two systems of units are presented in reference 2. The symbols used are defined as follows: wing chord, 1.29 feet (0.39 meter) drag coefficient, Drag qs Lift
lift coefficient, -
qs maximum trimmed lift coefficient lift-curve slope, per degree pitching-moment coefficient referred to model moment, center at wing quarter-chord (c/4), Pitching moment ,(see fig. 1) qsc propeller thrust coefficient based on free-stream velocity and wing T
area, - (often designated in literature as Tc)
qs propeller thrust coefficient based on slipstream velocity and T
propeller disk area, -
WJSp propeller diameter, feet (meters) height of the horizontal-tail chord above the wing chord, feet (meters ) tail incidence, degrees tail length measured horizontally from the wing quarter-chord to the horizontal-tail quarter-chord, feet (meters) number of propellers
e?, pounds/foot2 (newtons/meter 2 )
dynamic pressure,
slipstream dynamic pressure, q + -, T pounds/foot2 (newtons/mete$)
N s p
propeller disk area, d, foot2 (mete2)
tail area, f o o t ' (meter21
wing area, 9.04 foot2 (0.84 meter2)
total propeller thrus5, pounds (newtons) free-stream velocity, feet/second (meters/second)
- St - It
nondimensional horizontal-tail volume, sw distance measured along airfoil chord line from the leading edge, feet (meters ) distance measured perpendicular from airfoil chord line to airfoil lower surface, feet (meters) distance measured perpendicular from airfoil chord line to airfoil upper surface, feet (meters) angle of attack, degrees deflection of movable surface (with subscript to denote surface
deflected) , degrees
downwash angle at the horizontal tail, degrees downwash angle at the horizontal tail when wing angle of attack is zero, degrees air density, slugs/foot3 (kilogram/meter3) Subscripts: f flap (see fig. 3 ) t tail V vane (see fig. 3 ) MODEL AND APPARATUS A three-view drawing of t h e model i s presented i n figure 1 and photo- graphs are presented i n f i g u r e 2. The wing had an NACA 44-13 a i r f o i l section, a l5.3O-inch (0.39 meter) chord, w a s unswept, and had a span of 7.00 feet The wing contour w a s formed with (2.13 meters) with an aspect r a t i o of 5.42.
f a i r e d wooden blocks fastened t o a metal spar which supported t h e two motor nacelles and t h e fuselage strongback as w e l l as t h e brackets which held t h e f l a p system.
a 20-percent-wing- The double-slotted h i g h - l i f t f l a p system consisted of chord vane with a St. Cyr 156 a i r f o i l section and a 40-percent-wing-chord f l a p with a modified Rhode St. Genese 35 a i r f o i l section over t h e forward 30 percent of i t s chord f a i r e d i n t o t h e wing a i r f o i l section over t h e rear 70 percent of its chord. The flap and vane ordinates, as w e l l as t h e f l a p and vane positions when deflected, a r e given i n figure 3 .
Two d i f f e r e n t horizontal t a i l s were tested. Both had an aspect r a t i o of 3-15 and an NACA 4415 a i r f o i l section whose p r o f i l e w a s modified t o give a 9 percent maximum thickness and were mounted inverted t o provide an inverse camber. The two t a i l s had different areas, spans, and chords. (See f i g . 1.)
The s m a l l t a i l w a s t e s t e d i n a high position (ht = 0 . 9 4 ~ )only; t h e large t a i l w a s t e s t e d i n both t h e high and the low position ( h t = 0 . 1 5 ~ ) . Both t a i l posi- tions were above t h e wing-chord plane. For t h e t a i l configurations tested, t h e nondimensional horizontal-tail volumes V , a r e t h e following: s m a l l t a i l i n t h e high position, 0.85; large t a i l i n t h e high position, 1.15; large t a i l i n t h e low position, 1.04. Additional data on t h e geometric characteristics are also presented i n f i g u r e 1.
Since no directional s t a b i l i t y tests w e r e included i n t h e investigation, The v e r t i - t h e v e r t i c a l t a i l served only as a support f o r t h e horizontal t a i l .
c a l surface consisted of a sheet of 1/2-inch (1.27 centimeters) aluminum with a rounded leading edge and a beveled t r a i l i n g edge.
The three-blade propellers were made of balsa covered with glass-fiber cloth and w e r e driven by water-cooled variable-frequency e l e c t r i c motors oper- ated i n p a r a l l e l from a variable-frequency power supply, which kept t h e motor The speed of rotation of each speeds matched within 20 revolutions per minute.
propeller w a s determined by a stroboscopic indicator which received t h e output For a l l t h e frequency of s m a l l a l t e r n a t o r s connected t o each motor shaft.
tests t h e r i g h t propeller rotated i n a clockwise direction and t h e l e f t pro- peller rotated i n a counterclockwise direction when viewed from the r e a r of t h e model. The During t h e tests t h e speed o f rotation w a s maintained a t 6000 r p m .
t h r u s t coefficient w a s varied by changing t h e wind-tunnel speed.
The motors were mounted inside aluminum-alloy nacelles by means of s t r a i n - gage beams so t h a t t h e propeller t h r u s t could be measured. The t o t a l l i f t , longitudinal force, pitching moment, r o l l i n g moment, yawing moment, and side force were measured by a strain-gage balance mounted t o t h e fuselage a t t h e Only longitudinal components of t h e data are presented i n wing quarter-chord.
t h i s report. The r e s u l t s of l a t e r a l control t e s t s of this model a r e presented i n reference 1.
TESTS AND CORRECTIONS The investigation was made i n the 17-foot (5.18 meter) t e s t section of the Langley 3OO-MPH 7- by 10-foot tunnel. For the powered t e s t s t h e free- stream dynamic pressure was varied from about 1.3 t o 5.3 pounds/foot2 (72 t o 254 newtons/meter2), depending on the desired t h r u s t coefficient. The s l i p - stream dynamic pressure was r e l a t i v e l y constant a t about 7.5 pounds/foot2 (359 newt;ons/meter2) f o r a l l t h r u s t coefficients. A free-stream dynamic pres- sure of about 6.0 pounds/foot2 (287 newtons/meterZ) w a s used f o r the propeller- For the powered t e s t s the Reynolds number, based on wing chord, of off t e s t s .
the flow i n the slipstream averaged about 0.65 X lo6; f o r t h e propeller-off
t e s t s the Reynolds number i n the f r e e stream averaged about 0.58 X lo6. Since e r r o r s due t o blockage, slipstream contraction, and tunnel-wall e f f e c t s have been found t o be small f o r models of t h i s s i z e i n the l7-foot t e s t section 3), no corrections f o r these types of e r r o r have been applied t o the ( r e f .
data.
The propeller thrust data have been presented as the conventional t h r u s t coefficient, t h a t is, thrust nondimensionalized by free-stream dynamic pressure times wing area I n a l l cases a thrust coefficient of zero was (CT = T/qS).
obtained by removing the propellers from the model. For the propeller-on data the t h r u s t w a s measured by strain-gage beams a t t h e motors.
The t h r u s t coeffi- c i e n t s based on these measurements a r e presented with the basic data.
Although the motor rotation speed was held constant, t h e thrust increased as the angle of attack of the model increased; as a r e s u l t , the t h r u s t coefficients a r e not constant f o r a p a r t i c u l a r t e s t . For convenience the average values of t h r u s t coefficients near zero angle of attack f o r the data presented i n t h i s report (used as reference values throughout the report) a r e l i s t e d i n the following t a b l e :
Reference value of - I
'T,s I 0 0 .20 .44 -39 .50
::is 1
.64 2.42 .78 It is often desirable to use the propeller thrust coefficient based on slip-
stream velocity and propeller disk area. Figure 4 is a plot of the relation
between these two thrust coefficients for the model tested.
E U L T S AND DISCUSSION The results of a wind-tunnel investigation of the longitudinal control and stability characteristics of a model of a twin-propeller deflected-slipstream STOL airplane are presented in the following figures: Figure Basic data: Flaps retracted, 6f = 0':
T a i l o f f . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5
Small tail, high position . . . . . . . . . . . . . . . . . . . . 6 to 8
Large tail, high position . . . . . . . . . . . . . . . . . . . . 9 to 11
Large tail, low position . . . . . . . . . . . . . . . . . . . . . 12 to 1 4
Flaps deflected, 6f = 45': . . . . . . . . . . . . . . . . . . . . .
T a i l o f f . . . . . . . . 15
Small tail, high position . . . . . . . . . . . . . . . . . . . . 1 6 to 1 . 9
Large tail, high position . . . . . . . . . . . . . . . . . . . . 20 to 23
Large tail, low position . . . . . . . . . . . . . . . . . . . . . 24 to 27
Comparisons :
Effect of tail area . . . . . . . . . . . . . . . . . . . . . . . . 28 to 30
Effect of tail height . . . . . . . . . . . . . . . . . . . . . . . 3lto 32
Basic Data The basic data figures present the variation of the lift and pitching- moment coefficients with angle of attack and the variation of drag and pitching- moment coefficients with lift coefficient. In addition, the variation of thrust coefficient with angle of attack is presented for the propeller-on tests. The pitching-moment coefficients for all the data are presented about the wing quarter-chord line. The angle of attack used in this investigation ranged from - 4 O to 44O.
The basic data for the configuration with theohorizontal tail off are pre- sented in figure 5 for the flaps-retracted ( t i f = 0 ) configuration and in fig-
15 for the flaps-deflected (6f = 4 5 ' ) configuration. The lift-curve slope
ure The following as well as the maximum lift increases with increasing power.
Cs, table gives a summary of the lift-curve slope and of the maximum lift coeffi- cients (model with the horizontal tail off) for both of the flap deflections: ..
.........
. . I..,.,.- . ......-.._.. .
I; CL ,max .- 0.067 1.12 1.83 2.11 ..
2.47 0.068 4.40 5.35 .l50 6.48 The t a i l - o f f pitching moments f o r t h e flaps-retracted (6f = 0 0 ) configura- t i o n are unstabl-e f o r angles of a t t a c k up t o wing stall and are generally neu- t r a l beyond t h a t angle of a t t a c k ( f i g . 5). The s m a l l magnitude of the pitching moments for t h i s configuration shows t h a t t h e horizontal t a i l is required configurations.
mainly t o provide s t a b i l i t y for the flaps-retracted (6f = 00
)
6 t o 14) show t h a t any of the t a i l configurations The tail-on data ( f i g s .
t e s t e d with t h e flaps-retracted configuration provide an adequate s t a b i l i t y contribution and, i n addition, are capable of trimming t h e airplane throughout t h e l i f t - c o e f f i c i e n t range of t h e investigation.
I n contrast t o t h e flaps-retracted (6f = Oo) configuration, t h e flaps- deflected (6f = 4 5 O ) configuration requires a large increment of pitching moment f o r t r i m . The t a i l - o f f data of figure 15 show t h a t nose-up increments of pitching-mom2nt coefficient ranging from approximately 0.5 to 1.1 a r e therefore, a large down load i s needed, depending on t h e power condition; required a t t h e t a i l position. I n order t o satisfy this requirement, a normal- force coefficient as large a s -0.95 m u s t be developed by t h e large horizontal t a i l i n t h e high position, assuming no l o s s i n dynamic pressure a t t h e t a i l .
The This value approaches t h e maximum normal-force coefficient f o r t h i s t a i l .
requirement f o r tail normal-force coefficient i s even more severe f o r t h e other two t a i l configurations t e s t e d because of t h e i r smaller t a i l volumes.
The data f o r t h e flaps-deflected (6f = 45O) configuration with each of t h e several horizontal t a i l s are presented i n figures 16 t o 27. The configuration with the small t a i l i n t h e high position ( f i g s . 16 t o 19) i s s t a b l e (at & m / & , = -0.32) and can be trimmed a t low power s e t t i n g s (CT = 0 CT = 0 , and CT = 0.70) through t h e angle of a t t a c k f o r wing s t a l l . Some t a i l s t a l l i s present a t low angles of a t t a c k f o r t h e t a i l incidence of -10.4O. A t higher power s e t t i n g s (CT = 1.25 and CT = 2.42) t h e t a i l i s beginning t o s t a l l a t more positive t a i l incidences. When CT = 2.42, t h e t a i l with an incidence of
-10.4O i s completely s t a l l e d . The configuration can no longer be stably
trimmed f o r any of t h e t a i l incidences used for these t e s t s with the center- of-gravity position corresponding t o the wing quarter-chord.
The data f o r t h e flaps-deflected configuration with t h e large t a i l i n t h e high position ( f i g s . 20 t o 23) a r e similar t o those f o r the configuration w i t h the s m a l l t a i l but with a l a r g e r t r i m increment due t o the additional area.
A t CT = 1.25 the model i s trimmed a t an angle of a t t a c k of 80 with the t a i l incidence of - 5 . 5 O but the s t a b i l i t y i s reduced from t h a t with more positive t a i l incidences. This indicates t h a t the t a i l i s p a r t i a l l y s t a l l e d a t t h i s incidence. The t a i l with an incidence of -10.4O i s completely s t a l l e d i n t h e A t CT = 2.42 t h e model i s not trimmed below low range of angle of attack.
the angle of a t t a c k f o r wing s t a l l .
The large t a i l i n t h e low position ( f i g s . 24 t o 27) w i l l trim t h e f l a p - deflected model for a l l but t h e highest power s e t t i n g (CT = 2.42). This tail, however, i s i n a higher f i e l d of downwash and is s t a l l e d f o r more of the t a i l incidences than it i s i n t h e high position.
Downwash Analysis The data f o r a l l the configurations t e s t e d were analyzed t o determine the downwash angle and t h e dynamic pressure a t the horizontal t a i l .
For the cases where both the wing and t h e t a i l a r e unstalled it was possible t o obtain the
variation of downwash angle (eo and &/&) and an average value of q/qt
over the range of angles of a t t a c k where the variation of the parameters with angle of a t t a c k i s l i n e a r . These r e s u l t s a r e shown i n the j llowing t a b l e f o r a l l t h e t a i l configurations t e s t e d with the flaps-retracted 6f = 00) configuration.
Horizontal-tail
a€/&
€ 0 9 deg cT configuration Small t a i l , 0 1.02 0.29 1 . 1 6 1.05 high position 1.16 17 -33 1.09 1.15 .44 ... .. _.___ -- _ _ _ ~ I 0.28
Large t a i l , 0 0.60
high position -17 * 65 9 30 .44 .80 . _ _ Large t a i l , 0 1.06 0.28 low position 1.10 .48 .44 1.58 I For the configurations with the f l a p s deflected t o 4 5 O it was not pos- s i b l e t o determine corresponding information f o r t h e highest power s e t t i n g (CT = 2.23) or f o r t h e large t a i l i n the low position because the t a i l was s t a l l e d f o r most of the t a i l incidences used i n these t e s t s . The r e s u l t s f o r the other t a i l configurations a r e a s follows: Horizontal-tail configuration 0 1.08 5.66 Small tail, 0.37 high position 1 . 1 0 8 . 3 0 9 63
7 0
1 . 1 6 9.28 1 . 2 5 I
1.17
Large t a i l , 0 5.67 0.39
high position .66 1.17
7.55 7 0
1.25 1.16
8.37 7 8
Effect of T a i l Area The effect of t a i l area on a s t a b i l i t y contribution f o r the t a i l a t 0 ' incidence i n t h e high position i s presented i n figure 28 f o r 6f = 00 and i n figure 29 f o r 6f = 4 5 O . Over the e n t i r e range of angle of a t t a c k f o r t h e flaps-retracted (6f = 00) configuration, the data f o r the large t a i l show the expected increase i n the t a i l contribution t o s t a b i l i t y due t o the additional area. However, the wing angle of a t t a c k f o r zero l i f t on t h e t a i l i s d i f f e r - ent f o r each t a i l because t h e average downwash angle i s d i f f e r e n t . The data f o r t h e flaps-deflected (6f = 45O) configuration a r e similar t o those f o r the flaps-retracted (6f = O o ) configuration up t o t h e wing angle of a t t a c k cor- l i f t f o r a l l but the highest power s e t t i n g . Beyond responding t o zero t a i l t h i s wing angle of a t t a c k t h e t a i l l i f t becomes nonlinear f o r both horizontal t a i l areas.
When the f l a p s are r e t r a c t e d (6f = Oo) both horizontal t a i l s a r e capable of trimming the model up t o t h e angle of attack of wing s t a l l f o r the thrust coefficients of t h e t e s t s . A s shown i n the basic data, t h e flaps-deflected The e f f e c t of t a i l area on (6f = 450) configuration presents a t r i m problem.
the maximum trimmed l i f t coefficient i s presented i n figure 30 as a f'unction of t h r u s t coefficient. This figure shows the increased t r i m capability of the large t a i l due t o the additional area. The data f o r t h e small t a i l i n the high position a r e based on t h e same moment center, O.25c, as the r e s t of the data i n t h i s report. The data f o r the large t a i l have t h e moment center adjusted t o 0.31~ so t h a t t h e s t a t i c s t a b i l i t y a t zero t h r u s t coefficient i s the same ( C ~ C L = 0.36) a s t h a t f o r the small t a i l . A l s o presented a s the dashed curve i n figure 30 is t h e m a x i m u m attainable l i f t coefficient f o r t h e this curve with the maximum-trimmed- t a i l - o f f configuration. Comparison of l i f t - c o e f f i c i e n t curves f o r the two t a i l s indicates that t h e small t a i l i s capable of trimming up t o the maximum a t t a i n a b l e l i f t coefficient ( t a i l - o f f
values of C L , - ) f o r thrust coefficients approaching 0.70 and t h a t t h e large
t a i l i s adequate f o r thrust coefficients up t o 1 . 2 5 . Neither of the tails i s adequate i n s i z e t o t r i m t h e airplane up t o t h e m a x i m u m a t t a i n a b l e l i f t coef- f i c i e n t s ( t a i l o f f ) a t t h e highest t h r u s t coefficient (CT = 2 . 4 2 ) .
Effect of T a i l Height The e f f e c t of t a i l height f o r t h e large t a i l a t zero incidence i s pre- sented i n figure 31 f o r 6f = O0 and i n figure 32 f o r 6f = 4 5 O . For t h e flaps-retracted (6f = Oo) configuration t h e r e s u l t s with the t a i l i n t h e two The t a i l i n the high position genera1.Q positions a r e generally similar.
produces a greater change of pitching moment with change of angle of a t t a c k than it does i n t h e low position. For t h e f l a p s a e f l e c t e d (6f = 4 3 O ) config- uration ( f i g . 32) t h e t a i l i n the low position produces a l a r g e r increment of pitching moment. A t the highest thrust coefficient ( C T = 2.42) t h e t a i l i n t h e low position i s s t a l l e d throughout t h e angle-of-attack range and t h e t a i l i n t h e high position i s s t a l l e d above 80 angle of attack. The model with t h e t a i l i n the high position shows a tendency t o p i t c h up a t low power s e t t i n g s a t angles of a t t a c k above about 20°.
CONCLUSIONS The r e s u l t s of a wind-tunnel investigation of t h e s t a t i c longitudinal s t a b i l i t y and control c h a r a c t e r i s t i c s of a model of a twin-propeller deflected- slipstream STOL airplane configuration indicate t h e following conclusions: 1. The magnitudes of the pitching mments of t h e wing-body combination The f o r t h e flaps-retracted (Oo f l a p deflection) configuration were small.
tail-on data show t h a t any of the t a i l configurations with the flaps-retracted configuration provide adequate s t a b i l i t y contribution and a r e capable of trimming t h e airplane.
2. The wing-body combination f o r t h e flaps-deflected ( f l a p deflection 4 3 O ) configuration had a l a r g e t a i l l i f t requirement f o r longitudinal trim, p a r t i c - u l a r l y f o r t h e highest power s e t t i n g of the investigation.
3 . The small t a i l i n the high position had t h e capability of trimming t h e A t airplane f o r the low power conditions ( t h r u s t coefficients of 0 and 0.70).
higher power conditions the t a i l s t a l l e d before t r i m was achieved.
4. The large t a i l i n e i t h e r position had t h e capability of trimming t h e airplane up t o the angle of a t t a c k corresponding t o t h e maximum l i f t coeffi- c i e n t for a l l but the highest power condition (thrust coefficient 2.42). This t a i l s t a l l e d before t r i m was achieved f o r t h e highest power condition.
Langley Research Center, National Aeronautics and Space Administration, Langley Station, Hampton, Va., February 16, 1966.
REFERENCES 1. Margason, Richard J.; and Hammond, Alexander D.: Lateral Control Charac- t e r i s t i c s of a Powered Model of a Twin-Propeller Deflected Slipstream N A S A TN D-1585, 1964.
STOL Airplane Configuration.
2. Mechtly, E. A.: The International System of Units - Physical Constants and
Conversion Factors. NASA SP-7012, 1964.
3 . Staff of Powered-Lift Aerodynamics Section, NASA Langley Res. Center: Wall Effects and Scale Effects i n V/STOL Model Testing. AIAA Aerodynamic Testing Conf., Mar. 1964, pp. 8-16.
I
t5.50 -
f39.371
--f
'-g
42m ft06.68/ Wing: NACA 4415 A i r f o i l secfion I29 (39 Chord, ti (mJ 9.04 184) Areo, f t 2 (m2) ZOO(Zt3) S p o n , t t ( m ) Aspect rofio Small toil Horizonfot foil Lorge t a i l 288 f.27) Areo, f t ' l mz) 386 (36) 3.00 1-91) S p o n , f f ( m J 3.48 f106) Chord.tf l m ) / / I 1.34) .96 (.29) 3. t3 Aspect rotio 3.t3 2.67 It,wing chord High position 270 Lowposition 244 2.4 t .94 ht,wing chord H i g h posifion .W .i5 Lowposition .I5
: I
, 1 1 Propellers: Sta 6/92 2.00 (61) D iometer, tf f m ) (t5Z8) .33 1/01 Nacelle diamefer, f f fm) 1 Number o t blades (each) I . S t a 66.79 ft69.65) Sfa t8.58 j 5 / 7 5 % Sfa 0 f42t9) (0) 75% a4 Sta 62.79 Sto, 24.40 f6L981 .
f3Zt6) 2.30 Moment center fV4) Figure 1.- Three-view drawing of 1/5-scale model and table of geometric characteristics. All dimensions are in inches (centimeters) u n l e s s otherwise noted.
(a) Top quarter f r o n t view.
L-63-9676 Model i n wind tunnel.
Figure 2.- L- 63-9677 ( 0 ) Lower quarter f r o n t view.
Figure 2.- Concluded.
~ Sta. 730 S la. 0 StL? 0 I (bJ F l a p deflected 45* (a) P l a i n w i n g ( N A C A 4415 airfai1).
F L A P ORDINATES I- X/C 1 u,/c v,/c 0 0 0 ,0125 0 0 ' 0 ,0250 .OG5 ,0460 -.0290 .0500 -0387 ,0750 ,0905 -.0446 -0435 . IO00 ,1039 -.0448 ,0460 ,1269 ,1500 -.0409 -.0468 .zoo0 ,1440 -.0300 -.0444 ,1630 ,3000 -.Ut40 -.042U ,4000 ,1660 .oo/o 70373 ,5000 . 1600 .Ut80 -.0312 6000 ,1440 .0300 -.ozm . 1 170 -.02/7 ,7000 ,0320 ,0830 ,8000 .0300 : -0159 ,9000 .0484 .0180 ,8750 ,0435 -.o/oo
,0274 .0107 1 /.oooo -0040
,9500 ,0040 -- ,0065 - 0 I I Figure 3.- Geometric characteristics of wing section and flap deflection. A l l dimensions given in fraction of wing chord unless otherwise noted.
-0 . 4 .8 /.2 1.6 20 24 28 3.2 3.6 4.0 4.4 4.8 52 56 6 . 0 CT Figure 4.- Slipstream t h r u s t coefficient plotted as a f u n c t i o n of free-stream t h r u s t coefficient.
(a) Variation of CL with a.
Figure 5.- Effect of propeller t h r u s t coefficient (in terms of reference CT) on longitudinal aerodynamic characteristics of flaps-retracted (6f = 00) configuration w i t h tail off, A 0 12 A 4 16
0 .2 . 4 . 6 .8 -.6 - 4 -.2 -.8
CD (b) Variation of C L with CD.
Figure 5.- Continued.
. 6 CT o Ofprop off1 0.17 .4 0 0.44 .2 -2 cm -.4 -.6
- .8
-10 -1.2 -1. 4
-
(c) Variation of Cm with 01.
Figure 5.- Continued.
-1.4 0 4 .8 /.2 /.6 2.0 24 28 32 3 . 6 -.8 -4 ‘ L (d) V a r i a t i o n of C, with C k F i g u r e 5.- C o n t i n u e d .
(e) Variation of CT with a.
Figure 5.- Concluded.
(a) Variation of CL with a.
Figure 6.- Effect of tail incidence on longitudinal aerodynamic characteristics of configuration with small tail i n high position. b f = 00; CT = 0 .
3 . 2 2.0 l. 6 CL 1 . 2 .8 .4 -.4 '48 -.6 -.4 -.2 0 .2 I .6 .8 l.0 /. 2 1.4 1 . 6 CD (b) V a r i a t i o n of CL w i t h CD.
Figure 6.- Continued.
ro w . 6 .4 .2 -.2 Cm -.4 -.6 -.8 -LO -1.2 ( c ) Variation of Cm with a.
Figure 6.- Continued.
ci
(d) Variation o f C , with CL.
Figure 6 . - Concluded.
(a) Variation of CL w i t h a.
Figure 7.- Effect of tail incidence on longitudinal aerodynamic characteristics of configuration w i t h small tail i n h i g h position. 6f = 00; reference CT = 0.17.
-?8 -.6 -.4 -. 2 0 4 .6 .8 1.0 1 . 2 1.4 1.6 CD ( b ) Variation of CL with CD.
Figure 7.- Continued.
. 6 .4 .2 -2 -.4 -.6 -.8 -10
- 1.2
0 4 8 12 16 20 24 28 32 36 40 44 48 -4 Q, deg (c) V a r i a t i o n of C m with a.
F i g u r e 7.- C o n t i n u e d .
,. I -8 -4 0 4 .8 12 1 . 6 2.0 2.4 2.8 3.2 3.6 CL ( d ) Variation of C m with CL.
Figure 7.- Continued.
w 4.0 3 . 6 3.2 1.6 1.2 . 8 .4 (e) Variation of CT w i t h a.
Figure 7.- Concluded.
-9-4 0 4 8 I2 I6 20 24 28 32 36 40 44 48 a, deq (a) Variation of C L w i t h a.
Figure 8.- Effect of t a i l incidence o n longitudinal aercdynamic characteristics of configuration w i t h small tail in h i g h position. 6f = Oo; reference CT = 0 . 4 4 .
w P w Iu 3 . 2 2.4 /.6
c , 1 . 2
.8 .4 -.4 -.8 U 1.0 1 . 2 l.4 L 6
- .4 -.2 0 .2 1 .6 .8
CD (b) Variation of CL with CD.
Figure 8.- Continued.
. 6 .4 .2 -.4 -.6 -.8 -LO -1.2 -14 4 0 4 8 12 I6 20 24 28 32 36 40 44 48 Q, deg (c) Variation of C, w i t h a.
w Figure 8.- Continued.
w w .t= -,.7 0 .8 /.2 1 . 6 20 24 2.8 3.2 3 . 6 -.8 -4 CL (d) Variation of C, w i t h CL.
Figure 8.- Continued.
4.0 3.6 3.2 2.4 2.0 1.6 1 . 2 . 8 .4 - 4 0 4 8 12 16 20 24 28 32 36 40 44 48 0, deg (e) Variation o f CT with a.
w wl Figure 8 . - Concluded.
(a) V a r i a t i o n of CL w i t h a.
F i g u r e 9.- Effect of t a i l i n c i d e n c e o n l o n g i t u d i n a l aerodynamic c h a r a c t e r i s t i c s of c o n f i g u r a t i o n w i t h large t a i l in h i g h position. Bf = 00; CT = 0 .
3.2 2.4 /.6 .4 -.4 -.0 -.8 -.6 -.4 -. 2 0 .2 1 .6 .8 l.0 /.2 /.4 /.6 CD (b) Variation of CL with CD.
Figure 9.- Continued.
w ( 1 , . 6 .4 .2 -.6 -.8 -10 -14
-
( c ) Variation of Cm with a.
Figure 9.- Continued.
' I . 7 / 1 2 1 . 6 2.0 2.4 2.8 3.2 3.6 I .8 -.8 -.4 CL (d) Variation of C , w i t h CL.
Figure 9.- Concluded.
(a) Variation of CL w i t h a.
Figure 10.- Effect of t a i l incidence o n longitudinal aerodynamic characteristics of configuration w i t h large tail in h i g h position. 6f = 00; reference CT = 0.17.
-.0 -.8 -.6 -.4 -. 2 0 .2 4 .6 .8 /.O /.2 /.4 1 . 6 CD ( b l V a r i a t i o n of C L w i t h CO.
Figure 10.- Continued.
. 6 .4 .2 -2 Cm -.4 -.6 -.8 -10 -1.2 I6 20 24 28 32 36 40 44 48 -4 0 4 8 /2 a, deg (c) Variation of C , with a.
Figure 10.- Continued.
0 .8 / 1 2 1 . 6 2.0 2.4 2.8 3.2 3.6 CL (d) Variation of C , w i t h CL.
Figure 10.- Continued.
4.0 3.6 3.2 1 . 6 I. 2 .8 .4 (e) Variation of Cy with a.
Figure 10.- Concluded.
(a) Variation of CL w i t h a.
Figure 11.- Effect of t a i l incidence o n longitudinal aerodynamic characteristics of c o n f i g u r a t i o n w i t h large t a i l in h i g h position. 6f = 00; reference CT = 0.44.
3.2 / . 6
c, 1.2
.8 .4
- .4
-.8 -8 -.6 -.4 -. 2 .2 I .6 .8 l.0 /.2 /.4 / . 6 CD (b) Variation of CL with CD.
Figure 11.- Continued.
(c) Variation of C m with a.
Figure 11.- Continued.
- 1 . 7 0 .8 /2 1 . 6 2.0 2.4 2.8 3.2 3.6 -.8 -.4 CL (d) Variation of C m with CL.
Figure 11.- Continued.
4.0 3 . 6 3.2 2 . 8 2.4 2 . 0 1 . 6 1.2 .8 .4 n W 0 4 8 I2 I6 20 24 28 32 36 40 44 48 -4 0, deg (e) Variation of CT with a.
F i g u r e 11.- Concluded.
0 4 8 I2 /6 20 24 28 32 36 40 44 48 0, deq (a) Variation of CL with a.
CT = 0.
Figure 12.- Effect of tail incidence on longitudinal aerodynamic characteristics of configuration with large tail i n low position. 6f = 00; (b) Variation of CL with CD.
Figure 12.- Continued.
. 6 .4 .2 -2 cm -.4 - . 6 -.8 -10
- /.2
- I R /6 20 24 28 32 36 40 44 48 -f.7 -4 0 4 8 /2 0, deq (c) Variation of C , with a.
Figure 12.- Continued.
.6 . 4 .2
- .2
cm -.4
-. 6
-.8 - 1 . 7 -.8 -.4 0 .8 12 1 . 6 2 . 0 2.4 2.8 3.2 3.6 CL (d) Variation o f C m with CL.
u Figure 12.- Concluded, w LJl c (a) V a r i a t i o n of C L w i t h a.
F i g u r e 13.- Effect of t a i l i n c i d e n c e o n l o n g i t u d i n a l aerodynamic c h a r a c t e r i s t i c s of c o n f i g u r a t i o n w i t h large t a i l in low position. b f = 00; reference CT = 0.17.
(b) V a r i a t i o n of CL w i t h CD.
F i g u r e 13.- Continued.
i -1.4 0 4 8 I2 I6 20 24 28 32 36 40 44 48 - 4 a, deg C , with a.
(c) Variation of Figure 13.- Continued.
L 4.0 3 . 6 3.2 2 . 8 2 . 4
c, 2.0
1 . 6 1.2 .8 .4 G -4 0 4 8 f2 1 6 20 24 28 32 36 40 44 48 a, deg (e) Variation of CT w i t h a.
Figure 13.- Concluded.
(a) Variation of CL w i t h a.
Figure 14.- Effect of tail incidence on longitudinal aerodynamic characteristics of configuration with large tail in low position. bf = Oo; reference CT = 0 . 4 4 .
-.U -.8 -.6 -.4 -. 2 0 .2 4 .6 . 8 l.0 /.2 /.4 /.6 CD (b) Variation of CL with CD.
Figure 14.- Continued.
. 6 .4 .2 -.2 cm -.4 -.6 -.8 -10 (c) Variation of C , with a.
Figure 14.- Continued.
.. I
-.8 - 4 0 .8 62 1 . 6 2 1 0 2 1 4 2.8 3.2 3 . 6 CL (dl Variation of C, with CL.
Figure 14.- Continued.
4 . 0 3 . 6 3.2 2 . 8 2 . 4 C r 2 . 0 1 . 6 1 . 2 . 8 .4 - -4 0 4 8 I2 I6 20 24 28 32 36 40 44 48 a, de7 (e) V a r i a t i o n of CT w i t h a.
Figure 14.- Concluded.
(a) Variation of CL w i t h a.
Figure 15.- Effect of propeller t h r u s t coefficient (in terms of reference Cy) o n longitudinal aerodynamic characteristics of flaps-deflected (ijf = 450) configuration w i t h t a i l off.
C L 4 / -/
-
l.6 -/12 -8 - 4 0 4 .8 1.2 1 . 6 2.0 24 2.8 CD (b) Variation of CL with CD.
Figure 15.- Continued.
(c) Variation of C , with a.
Figure 15.- Continued.
6.
I CJ 0 0 /prop o f f ) A 0.70 h L25 b 242 -/ 0 / 2 3 4 5 6 7 8 CL (d) Variation of C , w i t h Ck F i g u r e 15.- Continued.
f6 20 24 28 32 36 40 44 48 -4 0 4 8 f 2 a, deg (e) Variation of CT with a.
Figure 15.- Concluded.
/ - / (a) Variation of C L with a.
Figure 16.- Effect of tail incidence on longitudinal aerodynamic characteristics of configuration with small tail i n high position. 6f = 450; Cy = 0 .
I - a C L 4 / -I /.2 1 . 6 2.0 24 2.8 -l6 -12 -.8 -4 0 4 .8 CU (b) Variation o f CL w i t h CD.
Figure 16.- Continued.
. 6 .4 .2 -.2
cm -.4
-.6 -.8 -10 -L?
4.4 -4 0 4 8 /2 /6 20 24 28 32 36 40 44 48 Q, deg (c) Variation of C , with a.
Figure 16.- Continued.
. 6 .4 .2 -.2
cm -.4
-.6
- . 8
4.2 4.4 -1 0 I 2 3 4 5 6 7 8 CL (d) Variation of C , with CL.
Figure 16.- Concluded.
I C L 4 / - / -4 0 9 8 I2 I6 20 24 28 32 36 40 44 48 (a) Variation of CL with a.
Figure 17.- Effect of tail incidence on longitudinal aerodynamic characteristics of configuration with small tail in high position. bf = 45O; reference CT = 0.70.
C L 4 I -1 -/.6 -12 -.8 -4 0 4 .8 1.2 1 . 6 2.0 2 . 4 2.8 CD (b) Variation of C L with CD.
Figure 17.- Continued.
. 6 .4 .2 -2 Gin -.4 - . 6 -.8 -10 - 1 . 2 -L4 -4 0 4 8 12 I6 20 24 28 32 36 40 44 48 a,deg (c) V a r i a t i o n of Cm with a.
Figure 17.- Continued.
. 6 .4 .2 I -.2 Cm -.4 -.6
- .8
I -LO 4.2 4.4 -/ 0 / 2 3 4 5 6 7 8 CL (d) Variation of C , w i t h CL.
Figure 17.- Continued.
4.0 3.6 3.2 2 . 8 2.4
c, 2.0
1.6 1.2 .8 .4 -4 16 20 24 28 32 36 40 44 48 0 4 8 12 a, de7 (e) Variation of CT with a, Figure 17.- Concluded.
CL 4 I -1 /6 20 24 28 32 36 40 44 48 -4 0 4 8 Q, deg (a) Variation of C L with a.
Figure 18.- Effect of tail incidence on longitudinal aerodynamic characteristics of configuration with small tail i n high position. 6f = 45O; reference CT = 1.25.
C L 4 / -/ 1 . 2 /.6 2.0 2 1 4 2.8 -L6 - L 2 -.8 - 4 0 4 .8 CD (b) Variation of CL with CD.
Figure 18.- Continued.
. 6 .4 .2 -.2 cm -.4 -.6 -.8 -LO -1.2 4 4 -4 0 4 8 I2 I6 20 24 28 32 36 40 44 48 0 3 deg (cl Variation of C , with a.
Figure 18.- Continued.
. 6 .4 .2 -.2 Cm -.4 -.6
- .B
-LO -1.2 -1. 4 -1 0 1 2 3 4 5 6 7 8 CL (d) Variation of C , w i t h CL.
Figure 18.- Continued.
4.0 3 . 6 3.2 2 . 8 2.4 2.0 1 . 6 1 . 2 .8 .4 I6 20 24 28 32 36 40 44 48 -4 0 4 8 I2 a, deg (e) V a r i a t i o n of Cy w i t h a.
F i g u r e 18.- Concluded.
C L 4 I - i -4 0 4 8 /2 /6 20 24 28 32 36 40 44 48 Q, deg (a) Variation of CL with a.
m Figure 19.- Effect of tail incidence on longitudinal aerodynamic characteristics of configuration with small tail i n high position. 6f = 45O; reference CT = 2.42.
W C L 4 i I i !
!
!
I I !
I -1 f.2 f.6 2.0 2 . 4 2.8 - L 6 -f.2 -.8 -4 0 4 .8 CD !
(b) Variation of CL with CD.
Figure 19.- Continued.
j . 6 .4 .2 -.2 cm -.4 -.6 -.8 -LO -/.2 4 4 /2 1 6 20 24 28 32 36 40 44 48 -4 0 4 8
Q, e?
(c) Variation of Cm with a.
Figure 19.- Continued.
. 6 .4 .2 -.2 -.4 -.6
- . 8
-LO -1.2 -I. 4 -I 0 I 2 3 4 5 6 7 8 CL (d) Variation of C , with Cb Figure 19.- Continued.
4.0 3 . 6 3.2 2 . 8 2.4 1 . 6 .8 .4 I6 20 24 28 32 36 40 44 48 -4 0 4 8 I2 Q, deq (e) Variation of Cy with a.
Figure 19.- Concluded.
/ - / /6 20 24 28 32 36 40 44 48 -4 0 4 8 /2 a,deg ( a ) Variation of C L with a.
Figure 20.- Effect of tail incidence on longitudinal aerodynamic characteristics of configuration with large tail i n high position. bf = 45O; CT = 0.
C L 4 I -I - L 6 -12 -.8 -4 0 4 .8 1 . 2 1 . 6 2.0 24 2.8 CD (b) V a r i a t i o n of CL w i t h CO.
Figure 20.- Continued.
. 6 .4 .2 -.2 cm -.4 -.6 -.8 -10 - L 2 4.4 f6 20 24 28 32 36 40 44 48 - 4 0 4 8 12 Q, deq (c) Variation of C m with a.
Figure 20.- Continued.
. 6 .4 .2 -.2 -.4 -.6
- . 8
- 10
-1.2 - I n - 1 . 7 -1 0 1 2 3 4 5 6 7 8 (d) Variation of C , with CL.
Figure 20.- Concluded.
CL 4 t - f 1 2 16 20 24 28 32 36 90 94 98 -4 0 4 8 Q, deg (a) Variation of CL with a.
Figure 21.- Effect of tail incidence on longitudinal aerodynamic characteristics of configuration with large tail i n high position. bf = 45O; reference CT = 0.70.
C L 4 / -/ 1.2 i.6 2.0 2.4 2.8 - L 6 -/.2 -.8 -4 0 4 .8 CD (b) Variation of CL with CD.
Figure 21.- Continued.
/6 20 24 28 32 36 40 44 48 -4 0 4 8 1 2 Q, deq (c) Variation of Cm with a.
Figure 21.- Continued.
. 6 .4 .2 -.6
- . 8
-LO 4.2 1.7 -/ 0 / 2 3 4 5 6 7 8 CL (d) V a r i a t i o n of C m w i t h CL.
F i g u r e 21.- Continued.
4 0 3.6 3.2 2 . 8 2.4
cr 2.0
1.6 1.2 .8 .4 I2 I6 20 24 28 32 36 40 44 48 -4 0 4 8 a, deg (e) Variation of CT with a.
Figure 21.- Concluded.
f - / -4 0 4 8 /2 1 6 20 24 28 32 36 40 44 48 Q, deg (a) Variation of CL with a.
Figure 22.- Effect of tail incidence o n longitudinal aerodynamic characteristics of configuration w i t h large t a i l in h i g h position. bf = 45O; reference CT = 1.25.
C L 4 / -/ 1.2 1 . 6 2.0 2.4 2.8 -l6 -L2 -.8 - 4 0 4 .8 CD (b) Variation of C L with CD.
Figure 22.- Continued.
. 6 . 4 .2 -.2 Cm -.4 -.6 -.8 -10 -/.2 -4 0 4 8 I6 20 24 28 32 36 40 44 48 Q, (c) Variation of C , with a.
Figure 22.- Continued.
. 6 .4 .2 -.6
- .8
- 1 2 .I A - 1 . 7 -/ 0 I 2 3 4 5 6 7 8 CL (d) Variation of Cm with CL.
Figure 22.- Continued.
4.0 3.2 2 . 8 2.4
cr 2 . 0
1 . 6 1 . 2 .8 .4 - /6 20 24 28 32 36 40 44 48 - 4 0 4 8 /2 @ , dep (e) Variation of Cy with a.
Figure 22.- Concluded.
C L 4 f - f f.2 1 6 20 24 28 32 36 40 44 48 -4 0 4 8 0,dW (a) Variation of CL with a.
Figure 23.- Effect of tail incidence on longitudinal aerodynamic characteristics of configuration with large tail in high position. 6f = 45O; reference CT = 2.42.
C L 4 / -/ -/.6 -f.2 -.8 -4 0 4 .8 1 . 2 1 . 6 2.0 2.4 2.8 CD (b) V a r i a t i o n of CL w i t h CD.
w Figure 23.- Continued.
. 6 .4 .2 -.2 -.6
- .8
- /.Q
4.2 4.4
-
I I 2 3 4 5 6 7 8 CL (c) Variation of C , with CL.
Figure 23.- Continued.
. 6 . 4 .2 -.2 cm -.4 - . 6 -.8 -LO - 1 . 2 -4 0 4 8 1 2 I6 20 24 28 32 36 40 44 48 a,dep (d) Variation o f C , with a.
Figure 23.- Continued.
4.0 3.6 3.2 2.8 2.4 2.0 1.6 1.2 .8 .4 -4 0 4 8 I2 I6 20 24 28 32 36 40 44 48 Q, deq (e) V a r i a t i o n of CT w i t h a.
F i g u r e 23.- Concluded.
C L 4 / - / - 4 0 4 8 I2 16 20 24 28 32 36 40 44 48 Q, deg (a) Variation of CL with a.
Figure 24.- Effect of tail incidence on longitudinal aerodynamic characteristics of configuration with large tail i n low position. 6f = 450; CT = 0 .
- ._ C L 4 / -I -l6 -12 -.8 -4 0 4 .8 /.2 /.6 2.0 2.4 2.8 CD (b) Variation of CL with CD.
Figure 24.- Continued.
. 6 .4 .2 -2 Cm -.4 - . 6 -.8 -10 -1.2 4 4 I6 20 24 28 32 36 40 44 48 -4 0 4 8 12 a, deg (c) Variation of C, w i t h a.
F i g u r e 24.- Continued.
. 6 .4 .2 -.2 cm -.4 -.6
- . 8
- 1.0
-1.2 - 1 . 7 -I 0 I 2 3 4 5 6 7 8 CL (d) Variation of C , with CL.
Figure 24.- Concluded.
C L 4 / - / -4 0 4 8 I2 I6 20 24 28 32 36 40 44 48 a, deg la) Variation of CL with a.
Figure 25.- Effect of tail incidence on longitudinal aerodynamic characteristics of configuration with large tail i n low position. 6f = 45O; reference CT = 0.70.
C L 4 I -1 1.2 1.6 2.0 2.4 2.8 -I6 -/.2 -.8 -4 0 I .8 CD (b) Variation of C L with CD.
Figure 25.- Continued.
. 6 .4 .2 -.2 Cm -.4 - . 6 -.8 -LO - 1 . 2 4 0 4 8 f2 1 6 20 24 28 32 36 40 44 48
Q,m
(c) Variation of C , with a.
P P Figure 25.- Continued.
w .
. 6 .4 .2 -.2 Cm -.4 -.6
- .8
-LO 4.2 4.4 -/ 0 / 2 3 4 5 6 7 8 CL ( d ) Variation of C , with CL.
Figure 25.- Continued.
4.0 3 . 6 3.2 2 . 8 2.4 C r 2.0 1.6 1.2 .8 .4 - -4 0 4 8 12 1 6 20 24 28 32 36 40 44 48 Q, de7 l e ) V a r i a t i o n of CT with a.
Figure 25.- Concluded.
C L 4 / -/ I6 20 24 28 32 36 40 44 48 -4 0 4 8 I2 Q, deg (a) Variation of CL with a.
Figure 26.- Effect of tail incidence on longitudinal aerodynamic characteristics of configuration with large tail i n low position. 6f = 450; reference CT = 1.25.
C L 4 I -1 - L 6 -/2 -.8 -4 0 4 .8 1 . 2 1 . 6 2.0 2.4 2.8 CD (bi Variation of CL w i t h CD.
Figure 26.- Continued.
. 6 .4 .2 -2 Cm -.4 - . 6 -.8 -10 -f.2 (c) Variation of C m with a.
Figure 26.- Continued.
. 6 .4 .2 -2 Cm -.4 -.6
- .8
- 1.0
-1.2 -/. 4 -/ 0 / 2 3 4 5 6 7 8 CL (d) Variation o f Cm with CL.
Figure 26.- Continued.
!
E
4.0 3 . 6 3.2 2 . 8 2.4
cr 2.0
1 . 6 1 . 2 .8 .4 /6 20 24 28 32 36 40 44 48 -4 0 4 8 /2 a, deg ( e ) Variation of CT with a.
Figure 26.- Concluded.
C L 4 - 1 -, -4 0 4 8 I2 f6 20 24 28 32 36 40 44 48 Q, deg ( a ) Variation of CL with a.
Figure 2 7 . - Effect of tail incidence on longitudinal aerodynamic characteristics of configuration with large tail i n low position. 6f = 45O; reference CT = 2.42.
- L 6 -f.2 -.8 - I 0 I .8 1.2 1.6 2.0 2.4 2.8 CD (b) Variation of CL with CD.
Figure 27.- Continued.
. 6 .4 .2 -.6 -.8 -10 (cl Variation of Cm with a.
r Iu Figure 27.- Continued.
w . 6 .4 .2 -.2 cm -.4 -.6
- .8
-lO -1.2 -I. 4 -/ 0 I 2 3 4 5 6 7 8 CL (d) Variation of C , with CL.
Figure 27.- Continued.
(e) Variation of CT with a.
Figure 27.- Concluded.
(a) CT = 0 (propeller off).
Figure 28.- Effect of tail area for configuration with horizontal tail i n high position. i t = Oo; 6f = go.
p
.6 .4 .2 -.2 Cm -.4 - . 6 -.8 -10
- 1 . 2
/6 20 24 28 32 36 40 49 48 - 4 0 4 8 /2 Q, deg (bl Reference CT = 0.17.
Figure 28.- Continued.
. 6 .4 .2 -.6
- . 8
- 1.0
-1.2 4.4 I6 20 24 28 32 36 40 44 48 -4 0 4 8 adeg (c) Reference CT = 0 . 4 4 .
Figure 28.- Concluded.
- 4 0 4 8 I2 16 20 24 28 32 36 40 44 48 Q, deg (a) CT = 0 (propeller off).
Figure 29.- Effect of tail area for configuration with horizontal tail i n high position. it = Oo; df = 45O.
: !
(b) Reference CT = 0.70.
Figure 2 9 . - Continued.
/6 20 24 28 32 36 40 44 48 -4 0 4 8 1 2 a, deg (c) Reference CT = 1.25.
Figure 29.- Continued.
. 6 .4 .2 -2 -.4 - . 6 -.8 -10 - L I6 20 24 28 32 36 40 44 48 c 0 4 8 f2 a,dw (d) R e f e r e n c e CT = 2.42, Figure 29.- Concluded.
I-’ w w Figure 30.- Effect of t a i l area on maximum trimmed l i f t coefficient.
. 6 .4 o f f .2 -2 Cm -.4 -.6 - .8 -LO -1.2 - 1 . 4 16 20 24 28 32 36 40 44 48 -4 0 4 8 12 a, deg (a) CT = 0 (propeller off).
Figure 31.- Effect of t a i l height for configuration w i t h large tail. i t = 00; bf = Oo.
. 6 .4 off .2 -2 -.6
- .8
-LO -1.2 -1. 4 I2 I6 20 24 28 32 36 40 44 48 -4 0 4 8 Q, deg (b) Reference CT = 0.17.
Figure 31.- Continued.
I2 I6 20 24 28 32 36 40 44 48 - 4 0 4 8 a, deg 1c) Reference CT = 0 . 4 4 .
F i g u r e 31.- Concluded.
. 6 .4 o f f .2 -.2
- .4
-.6
- .8
-LO -1.2 -1. 4 - 4 0 4 8 I2 I6 20 24 28 32 36 40 44 48 CJ, deg (a) CT = 0 (propeller off).
32.- Effect of tail height for configuration with large tail. it = Oo; df = 45O.
Figure . 6 .4 ff .2 -.2 Cm -.4 -.6 -.8 -LO -/.Z (b) Reference CT = 0.70.
Figure 32.- Continued.
. 6 .4 .2 -2 Cm -.4 -.6
- .8
4.2 4.4 -4 0 4 8 /2 /6 20 24 28 32 36 40 44 48 Q, deg (c) R e f e r e n c e CT = 1.25.
F i g u r e 32.- Continued.
. 6 .4 .2 -2
- .4
-.6
- .8
- 1.0
4.2 4.4 I2 /6 20 24 28 32 36 40 44 48 -4 0 4 8 0, deg (d) Reference Cy = 2.42.
F i g u r e 32.- Concluded.
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TECHNOLOGY UTILIZATION PUBLICATIONS: Information on tech- nology used by NASA that may be of particular interest in commercial and other nonaerospace applications. Publications include Tech Briefs; Technology Utilization Reports and Notes; and Technology Surveys.
Details on the availability o f these publications may be obtained from: SCIENTIFIC AND TECHNICAL INFORMATION DIVISION N AT1 0 N A L A E RO N A UTI CS A N D SPACE AD M I N I STR AT1 0 N Washington, D.C. PO546