Force-Test Investigation of the Stability and Control Characteristics of a 1/4-Scale Model of a Tilt-Wing Vertical-Take-Off-and-Landing Aircraft

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NASA MEMO 11-3-58L N L_ I

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MEMORANDUM

FORCE-TEST INVESTIGATION OF THE STABILITY AND CONTROL CHARACTEIKISTICS OF A I/4-SCALE MODEL OF A TILT-WING VERTICAL-TAKE-OFF-AND-LANDING AIRCRAFT By William A. Newsom, Jr., and Louis P. Tosti Langley Research Center Langley Field, Va.

NATIONAL AERONAUTICS AND

SPACE ADMINISTRATION

WASH I NGTON January 1959 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MEMORm_UM 11-5-58L FORCE-TEST INVESTIGATION OF _{E STABILITY AND CONTROL CHARACTERISTICS OF A I/4-SCALE MODEL OF A TILT-WING VERTICAL-TAKE-OFF-AND-LANDING AIRCRAFT By William A. Newsom, Jr., and Louis P. Tosti SUMMARY A wind-tunnel investigation has been made to determine the aerody- namic characteristics of a i/4-scale model of a tilt-wing vertical-take- off-and-landing aircraft. The model had two 3-blade single-rotation propellers with hinged (flapping) blades mounted on the wing, which could be tilted from an incidence of 4 ° for forward flight to 86 ° for hovering flight.

The investigation included measurements of both the longitudinal and lateral stability and control characteristics in both the normal forward flight and the transition ranges. Tests in the forward-flight condition were made for several values of thrust coefficient_ and tests in the transition condition were made at several values of wing incidence with the power varied to cover a range of flight conditions from forward- acceleration (or climb) conditions to deceleration (or descent) conditions The control effectiveness of the all-movable horizontal tail_ the ailerons and the differential propeller pitch control was also determined. The data are presented without analysis.

INTRODUCTION An investigation has been made of the stability and control charac- teristics of a i/4-scale model of the Vertol 76 vertical-take-off-and- landing (VTOL) aircraft. The results of the free-flight tests of the model are reported in reference i_ and the results of the force tests are presented in the present paper.

The force tests included measurementof both longitudinal and lateral

stability characteristics for the transition and normal-forward-flight

conditions. The tests in the forward-flight condition were madeat wing

incidences of 4° and 14° for thrust coefficients from 0 through 0.5. The

tests in transition flight were madefor wing incidences from 20° through

80° with various power settings to represent conditions of steady level

flight, forward acceleration (or climb), and deceleration (or descent).

The control effectiveness of the all-movable horizontal tail was deter-

mined for both the forward-flight and transition conditions, and the

effectiveness of the ailerons and the differential propeller pitch con-

trol was determined for the transition conditions.

SYMBOLS

The forces and momentsare based on the stability-axis system, which

is an orthogonal system with the origin at the airplane center of gravity.

The Z axis is in the plane of symmetry and perpendicular to the relative

wind, the X-axis is in the plane of symmetry and perpendicular to the

Z-axis, and the Y-axis is perpendicular to the plane of symmetry.

FL lift, lb

FD drag, lb

pitching moment, ft-lb

My

Fy side force, ib

rolling moment, ft-lb

MX

M Z yawing moment, ft-lb

C L llft coefficient, FL

qS C D drag coefficient, FD qS pitching-moment coefficient, My C m qSc side-force coefficient, F__y Cy qS rolling-moment coefficient, C_ qSb Cn yawing-moment coefficient, MZ qSb wing span, ft > hJ P 0q _q i (I] H kj4 H H D k_ k_ oo c+ c_ c_ edge down, deg (9 < c+ ng edge down, deg ci- c-h c+ FJ _--_ (I) o 0 0 b--' _-_ I-'- l.J.

c'_ r'0 r'o 2-- _dius, deg o o 9_ I ',D k.N {nots !

_D LJ1 II ii II O ro 4_ O o O o c_ d- o O ct- II o II H _n .g-- < 0 O II --.1 !

ale flying model o light test inves- r+ !

ul kD rawing of the kO s scaled up to the ropellers with ric motor which

drove the propellers through shafting and right angle gear boxes. The

speed of the motor was Changedto vary the thrust of the propellers.

The propeller blade angle was set at 12° except during the tests to

determine the lateral control effectiveness of differential pitch of the

two propellers.

The wing was pivoted at the 37-percent mean-aerodynamic-chord sta-

tion and could be rotated to provide incidence of 4° to 86° . The model

had an all-movable horizontal tall and conventional aileron and rudder

controls for forward flight. Roll control in hovering flight was pro-

vided by varying the pitch of the propellers differentially. For pitch

and yaw control in hovering flight, the model had Jet reaction controls

in the rear of the fuselage instead of the recessed tail "fans" in the

horizontal and vertical tails which are used on the airplane, and th_

tail fans were not represented in the model tests.

TESTS

The tests were made in the Langley full-scale tunnel with the model

support strut mounted near the lower edge of the entrance cone and about

5 feet above a ground board. Electric strain-gage balances were used to

measure the forces and momentson the model and an electric tachometer

was used to set the various model propeller speeds needed in the tests.

Although during someof the tests, another model was left in the tunnel

approximately 15 feet behind and slightly to the left of the present

model, no corrections for flow angularity or blockage due to its presence

have been applied to the data, since the blockage and interference effects

were believed to be very small.

For the forward-flight condition, tests were madeto determine the

longitudinal stability and control characteristics and the lateral sta-

bility characteristics for thrust coefficients from 0 to a value of 0.5

which represented full power at a lift coefficient of 1.0. These tests

were madewith a wing incidence of 14° as well as with the design inci-

dence of 4°. The longitudinal stability tests covered a range of angles

of attack from -5 ° to 20° and tail incidences from 0° to -15°. The

lateral stability tests covered a range of sideslip from -20° to 20° and

angles of attack from 0° to 20°.

For the investigation of the transition-flight condition, tests were

made for a range of power settings from that required for a forward

acceleration of i/2g or a rate of climb of 500 feet per minute for the

full-scale aircraft (whichever was the greater) to that required for a

deceleration of i/2g or a rate of descent of 500 feet per minute (which-

ever was the less). For tests at wing incidences of 20° , 40°, and 60°_

the forward acceleration or deceleration proved to be the determining

condition and the tests were madewith power settings which 3 with the

fuselage at zero angle of attack, gave forward accelerations of i/2g and I/4g, zero acceleration, and decelerations of i/2g and i/4g. With these power settings, the angle of attack was varied for longitudinal stability and control tests from -15 ° to 20 ° with the stabilizer off and with the stabilizer set at various angles of incidence from 0 ° to 15 ° .

Lateral stability and control tests at wing incidences of 20 °, 40 °, and 60 ° were made with power settings which gave forward accelerations of i/4g, O, and -i/4g with the fuselage at angles of attack of 0 ° and i0 °.

These tests covered a range of sideslip angles from 20 ° to _200, deflec- tions of the right aileron from 30 ° to -30 ° , and total differential pro- peller pitch from 0 ° to 6 ° (for the condition of zero acceleration at = 0° only).

For a wing incidence of 80 ° , the condition for rate of climb of.

500 feet per minute required the greater power variation, and the tests were consequently set up to represent the climb and descent conditions.

Tests were made with the power settings for steady level flight with the fuselage at 0 ° angle of attack, for a 45 ° climb or descent with the fuse- lage level (angle of attack, -45 ° or 45 ° , respectively), and for a 26._ ° climb or descent with the fuselage level. For the longitudinal stability and control tests, the fuselage angle of attack was varied approximately ±I0 ° from these conditions with the horizontal tail off and with the tail on at angles of incidence of 0° and 15 °. Tests were made at sideslip of 20 ° to -20 ° , and aileron-effectiveness tests were made with the fuselage level for a range of deflection of the right aileron from 30 ° to -30 ° for the level-flight, the 26.5 ° climb, and the 26._ ° descent power conditions. The effectiveness of the differential propeller pitch control was determined only for the level flight condition.

The tests at wing incidences of 4 ° and 14 ° were made at an airspeed of about 29 knots, which gave an effective Reynolds number based on the wing chord and free-stream velocity of about 400,000. For the tests at higher angles of wing incidence it was necessary to reduce the tunnel airspeed below 29 knots to avoid exceeding the model motor limitations.

The Reynolds number based on the wing chord and slipstream velocities varied between 200,000 and 790;000.

PRESENTATION OF RESULTS The results of the force test investigation to determine the aero- dynamic characteristics of a i/4-scale model of the Vertol 76 VTOL air- craft are presented in figures 2 to 14. The data for the normal forward flight tests (i w = 4 ° and 14 °) are presented in coefficient form, but since the coefficients approach infinity and become essentially meaning- less as the velocity approaches zero, the data for the transition flight tests (iw = 20 ° , 40 ° , 60 ° , and 80 ° ) have been scaled up to the weight

and center-of-gravlty locations of the full-scale airplane listed in

table II. It should be noted, however, that although the data have been

scaled up to correspond to the weight values of table II for tests made

at power settings that gave zero net drag and acceleration or deceleration

of 1/4g and 1/2g at _ = 0°, the data can be interpolated and rescaled in

terms of other conditions such as climb or glide or trim at other angles

of attack. If the data are rescaled, all forces and momentsare simply

multiplied by the factor required to make the lift equal to the desired

value for the desired condition, and the velocity is multiplied by the

square root of this factor. All tests were madewith a meanblade angle

of 12o instead of a blade angle adjusted to the proper value for each

condition, since the variation of rotor speed and blade angle for the

airplane was not known. Instead of adjusting the blade angle, the pro-

peller speed was adjusted to give the proper thrust and, consequently, the proper slipstream velocity and position for each test condition.

All the data of figures 2 and 3 were obtained at the normal forward

flight conditions (i w = 4° and 14°, Tc = 03 0.25, and 0.50). Figure 2

shows the variation of lift coefficient, drag coefficient, and pitching-

momentcoefficient with angle of attack_ and figure 3 showsthe variation

of rolling-moment coefficient, yawing-momentcoefficient, and side-force

coefficient with sideslip angle. During each test, horlzontal-tail

deflections of 0°, -10°3 and -15° were used, and data were also obtained

with the horizontal tail off.

The data obtained in the tests through the transition flight range

are presented in figures 4 to ll. In figures 4 to 8 are plots presenting

the variation of lift, drag, and pitching momentwith angle of attack at

several horizontal tail deflections, the separate figures representing

the different flight conditions of zero acceleration and of both accelera-

tion and deceleration of 1/4g and 1/2g. Figure 9 shows data obtained for

the variation of lift, drag, and pitching momentwith angle of attack for

the climbing and descending flight conditions with 80° wing incidence.

For figure 9(a) the model was set at _ = -26.5 ° and the 26.5° climb

(or forward acceleration of 1/2g) condition was established. The angle

of attack was then varied up and down from -26._ ° for the tests. The

data of figure 9(b) represent a 45° climb (or acceleration of 1 g) and

the test conditions were set up in a manner similar to those discussed

for figure 9(a). The condition for a 26.5° descent angle was represented

by establishing 1/2g of deceleration at _ = 26.5° and a 45° descent

angle was represented by a deceleration of 1 g at _ = 45° (figs. 9(c)

and 9(d)). Figure lO showsthe variation of rolling moment, yawing

moment, and side force with sideslip angle for wing incidences of 20°_

40° , 60° , and 80° . The tests were for the conditions of zero accelera-

tion and for accelerating or decelerating flight of 1/4g at both _ = 0o

and _ = lO°. The lateral stability data for a wing incidence of 80°

and a 45° climb (acceleration of 1 g) are presented in figure II. The

results of the tests of the lateral control effectiveness through the

transition flight range are presented in figures 12 to 14. In figures 12

and 13, the effect of deflection of the right aileron on the rolling

moment, yawing moment, and side force is presented, and figure 14 shows

the effect of differential propeller pitch.

CONCLUDING Pd_IARKS

Data have been presented for a i/4-scale model of a vertical-take-

off-and-landing aircraft with a wing capable of being tilted from 4°

incidence for forward flight to 86° incidence for hovering flight.

Included are longitudinal and lateral stability data covering a range

of conditions simulating zero forward acceleration and accelerating and

decelerating flight in level, climbing, and descending flight.

Langley Research Center,

National Aeronautics and Space Administration_

Langley Field, Va., October 2, 1958.

REFERENCE

i. Tosti, Louis P.: Flight Investigation of the Stability and Control

Characteristics of a i/4-Scale Model of a Tilt-Wing Vertical-Take-

Off-and-Landing Aircraft. NASA M_O II-4-58L , 1959.

TABLE I .- SCALED-UP GEOMETRIC CHARACTERISTICS OF THE MODEL Propellers (5 blades each rotor): Diameter, ft ........................ 9.33 Solidity .......................... 0.239 Chord, ft ......................... 1.0 Wing: Airfoil section ..................... NACA 4415 Aspect ratio ........................ 5.42 Chord, ft ......................... 4.75 Taper ratio ......................... 1.0 Area 3 sq ft ......................... 118.2 Span, ft ......................... 24.88 Ailerons (each): Chord, ft ........................ 1.22 Span, ft ......... ................ 4.83 Hinge line, percent chord ................ 74.1 Vertical Tail: Airfoil section ..................... NACA 0012 Aspect ratio ........................ 1.25 Chord, ft ......................... 4.0 Taper ratio ........................ 1.0 Span, ft .......................... 5.0 Rudder (hinge line perpendicular to fuselage center line): Chord, ft ........................ 1.25 Span, ft ......................... 5.0 Horizontal Tail: Airfoil section ..................... NACA 0012 Aspect ratio ........................ 3.10 Chord, ft ......................... 3.0 Center section chord, ft .................. 4.21 Area (including center body), sq ft ............ 29.70 Span, ft .......................... 9.90 ) 9 TABLE II.- WEIGHT OF THE FULL-SCALE AIRCRAFT WITH THE CENTER- OF-GRAVITY LOCATIONS FOR VARIOUS WING INCIDENCE ANGLES Weight, 3,139 lb] Center-of-gravity position (from wing pivot), ft deg Vertical Horizontal (below) (forward) 0.490 i. 310 .477 i.25o I.228 .460 2O 40 1.12O .394 6o .297 1.o57 8O .180 1.023 i0 t_ I l-J.

3.1 C) od .,-I o5 r--- .M k D_ r,,o o ,-I ,.-I O.

od ¢J o ,._ 4._ I

i

i Od !

I

J i t deg off D_ o -I0 .4 -15 Cm

X

"2 1.6

J

i 1.4 //

/

[] # 1.2 1.0 .8 CL .6 x/'/x_' ._j / /, C D -5 0 5 I0 15 20 a,de_ (a) T' = O; i = 4 ° .

C w Figure 2.- Longitudinal stablllty and control characteristics in normal forward flight.

't, deg .6 © [] O_ /x .4 .2 Cm J ]_ -2 1.8 1.6

/

f

1.4 CL // 1.2 _ / 1.0 / // .8 _'y

/

.6 CD .4 .2 0 5 10 15 20 -5 a,deg ' = O} i = 14 °

(b) To ,,, •

Figure 2.- Continued.

,4 .2 C m 0 :r --[ -.2 2.0 jl

, 4_._ "

tB / , CL It,d_ 1.6 /# O-- off rn____ 0 0 I0 A-- ---I5 1.2 Z/_/_/_ _j/

,o /,,,7

/

.8 .4 C o

_r_ .'--;'- '_

.2 J_ -.2 -5 0 5 I0 15 20 a,deg !

Figure 2.- Continued.

.4 .2 "

Cm _

0 • " //< 1.8

/Y

1.6

/

1.4 / C/

/

t?_ It, deg 0 off

//

1.0 [] ---- 0 ..... I /' --- -I 5 / .8 .6 i CD I0 15 20 (d) T' = 0.29_ iw = 14 ° • c Figure 2.- Continued.

c_ . " "t_ _z _X _2

I

-4 _2

A/

J.!

,Z

C L LO / it,decj off 8 __ -I0 -15 .2 % "5 0 2-O a,deg ' = 0._; L = 4° .

(e) _c

Figure 2.- Concluded.

D 17

J a, deg -I o--0 - I n-_- --I0 <> .... 20 -- I I I l Cy -J :2

\

-.3 -4 O8 ,,_[I .O4 (> J J Cn 0 / :04 v I --:08 .08 .04 Cl 0 :04 I -08 -20 -15 -I0 -5 0 5 I0 15 20 ,8, deg (a) T O = O; i w = 4 ° .

Figure 3.- Lateral stability characteristics in normal forward flight.

a_deg 0--0 D----IO

\

Cy \ \ -.I i \ -.2 \\i -.3 -.4 .04 J -.O8 .04 _L "_ _'_--" - CI 0 _. _<> .O4 _06 I0 15 20 -20 -15 -I0 -5 0 5 /9, deg (b) T' = O; i = 14 °.

c Figure 3.- Continued.

a,deg tJ 0 D----IO

\

Cy =2 \ \\ -3

\

-4 ;5 ,04 J j_ f Cn 0 t -D4 /J -08 .08 .04 Cl 0 =04 -:08 15 2O -2O -15 -I 0 -5 0 5 I0 ,8, d eg (c) T' = 0.25; ± = 4°.

c 1¢ Figure 3.- Continued.

a,decj .3 c)_ 0 D----IO .2

\

C,{ \

-2 \ -3 -.4 .04 L Cn 0 /

ii

_08 .O8 ct o .-04_ -:O8 -20 -_5 -_0 -5 0 5 _0 _5 2O B,d_ I (a) TO = 0.29_ iw = 14°" Figure 3.- Continued.

Cy Cn 0 -:04 .08 ,!

El. 0 -- -O4 :08 I -20 -15 -I0 -5 0 5 I0 15 20 (e) T' = 0.90} i = 4° Figure 3-- Concluded.

.L

r_ >- C -I 000 E O

\

E C "K/-- -2000 "_,_ n- It,deg o_ off -3000 [] 5 <>.... I0 A_-15 _D

I I

LL C_ -I000 4000 _ _ -_ 3000 2000 _ i I000 -15 -I0 -5 0 5 I0 15 20 a, deg (a) iw = 20 ° ) V = 68.5 knots.

Figure 4.- Longitudinal stability and control characteristics in the transition range with zero forward acceleration at _ = O.

It, deg off [] Z_ ioao E 0 E c- 2 - I000 if_ -2 000 I000 _Q & LI_ c_ e C3 -I000

I

300O ._I la..

2OO0 _I I000 5 I0 15 20 -15 40 V = 43.0 knots, (b) Continued.

_t,deg off Q A _ 2OOO ) r- E o E c r- n- 2OOO J_ J I000 J_ & e g] -I000 4OO0 e_ 3OOO 2OOO _J I000 2O -15 .-I0 -5 0 5 I0 15 a, deg (c) iw = 600; V = 29.8 knots.

Figure 4.- Continued.

D 25 O El A >- c- (D E O - 1000 E ¢- c- g_ I000 c5 e -I000 4000- 300O ..CI .._I [I 20O0 w-- .._I I000 40 -5 0 5 I0 a, deg (d) i_ -- 8O°; V = 8.53 knots.

Figure 4.- Concluded.

ft, deg _>............._f >- -I000 <-

_ -

-- ._ -<>.

IlJ E E -2OOO c if- --3000 I000 & o LL -2030 5OOO

I

J 4OOO -I0 -5 0 5 I0 15 20 a, deg (a) iw : 20o; V = 57.8 knots.

Figure 5-- Longitudinal stability and control characteristics in the 0 ° transition range with forward acceleration of i/4g at _ = .

't, deg off O [] _ A _Q I000 4-- >- E O E c -I000 __ o -2000 ( I000 c_ & U_ o C_ -I000 -2000 4OOO I 3O0O .J LL 2OOO ..J I000 45 -I0 -5 0 5 I0 15 20 a, deg (b) iw = 40o; V = 35-9 knots.

Figure 5.- Continued.

It, deg o off I000 >- c- ..._ ,-_--I I_ '--_ ....... ' "-- -_ E E -I000 c- -C o n- ICX:X:) .,C:I & LL e PI -I000 -2000 4OOO _..A I,_,- .,-----I_-.-..-i _ _- "- E v I," --3 LI.

w-- h 2O0O .--I I000 -15 -10 -5 0 5 I0 15 2O a, deg (c) iw : 60°; V : 14.5 knots.

Figure 5.- Concluded.

it, deg of O o

° II

A -IO00 -2000

I

-3000 n- .a -I000 & U,.

f e -2000 ...._, o llp "_ -3OOO 5OOO 4OOO 3C00 f LL

/

J

w .J A 20O0

//

ioool -15 -I0 -5 0 5 I0 15 20 a, deg (a) iw : 200; V : 49.5 knots.

Figure 6.- Longitudinal stability and control characteristics in the transition range with forward acceleration of I/2g at _ = 0 °.

_t,deg o off a__ 5 0 I0 a 1,5 o I .!

% >- -Io0o r- i E o f -2000 E ;f j-_-f / C 4:: -3000 n- -I000 a_ -20OO a -3OOO 4OOO 3OOO 2OOO Is_ _J I000 -15 -I0 -5 0 5 I0 15 20 e, deg (b) iw = 40o; V = 25.1 knots.

Figure 6.- Concluded.

\

,_..__ ..._..._ )_......_ --

\f

I000

\

E f \

'\

E -I000 c n- -2000 -300O 2OOO & _p IO00 T , _...,t : 5OOO 4OO0 3OOO J

I I I

2OOO ft, deg off -6 40 -5 0 5 I0 15 20 a, deg (a) iw = 20o; V = 82.5 knots.

Figure 7.- Longitudinal stability and control characteristics in the transition range with deceleration of i/4g at _ = 0 °.

_t, deg o off [] 5 0 .............. I0 +- 15 '...,.

a_ x ).- I000 %..

E E t C t- O n'- -I000 2OOO

d

& I(XX) e 4OOO ,._fq r-- 3OOO ..Q _.J U..

._1 I000 -15 -I0 -5 0 5 I0 IG 2O a, deg (b) iw = 40°; V = 63.7 knots.

Figure 7.- Continued.

D 33

't, deg o off 0 rO 3OO0 >- E E o I000 E

§

n- 2OOO & I000 U_ _b 4OOO I 2000 L -J I000 J5 -K) -5 0 5 I0 15 20 cl; de g (c) iw = 60o; V = 38.9 knots.

Figure 7.- Concluded.

3OOO

\

__n 2000 o x

, \

I000

q

\

tla \ [] E A \ \ \ E o \ a, \ c \ -/ l \ \ \ /

J

if_ \ r

-I000 r\

/ / / / \ \ -2OO0 20OO ..Q I.L _ I000 CD 5OOO -_ 30OO ^ .L_..

.j ZOO0 _t, deg o off [] 5 I000 ............. I0 " 15

I I111 I

45 40 5 I0 15 20 (a) iw = 20 °} V = 105 knots.

Figure 8.- Longitudinal stability and control characteristics in the transition range with deceleration of i/2g at _ = 0 °.

4OOO

\

El I0 <> \ 30O0 >-

I I I

_, X, ¢- 20O0 E E f C

(-

o \

"- ,._. \

n- _. I

"_ ',,, \

/ / J -IO00 \ , _/L 30OO ZOO(?

2-" - & U._ e IO00 ..o _..1 2OOO _J IO00 45 40 -5 0 5 I0 15 2O a,decj

(b)

iw = 40°; V = 241 knots.

Figure 8.- Continued.

't, deg J o off [3 4OOO

\

• " - 15 \ \\

I

2000 \ +- \! "-._ . < O E \ ."

--- I000 " -- _'_ n- 3OOO 2OOO & Ix.

1:3 4OOO 3OO0 _A 2(?OO h I000 W5 -I0 -5 0 5 I0 15 20 a, deg (c) iw = 600; V = 66.8 knots.

Figure 8.- Concluded.

E] 0 >- o t c E O E -I000 c c- n- ..Q & b_ - IOOO T v C3 - 2o00 .Q ..J 2OOO ._I IO0O -35 -50 -25 -20 -15 a, deg (a) 26.5 ° climb at _ = -26.5°_ V = 8.58 knots.

Figure 9.- Longitudinal stability and control characteristics in climb and descent conditions at iw = 80 ° • it,de cJ off O- >- c- j_

E -,ooo

E C _- -2000 n- .Q -- -I000 & b.

e n .200(3 -3000 _Q _r" _A " 2000 • ._ i_._ _ _A I000 o -55 -50 -45 -40 -35 cl, deg (b) 45 ° climb at = -45°_ V = 9.27 knots.

Figure 9-- Continued.

I

Figure 9-- Continued.

4o it ,deg 0 off rl <> .Q iooo i >- i ]

/

c- _f E O E -I000 o- C_ 3OOO J_ - 2000 _J LL --] I000 4O 45 50 55 a,deg (d) 4_ ° descent at = 450; V = 31.25 knots.

Figure 9.- Concluded.

D 41 tw, deg V, knots o 20 6Z3 a -40 4:5.0 v ......... 60- _1.1 M0oo,

I I 1

u_ A 80- 8.4 ,,f

_I i i I

" 0 o -I000 o _I 2OO0 ¢: / ...f! _._ 7" ........ %> 1- I I oE E - I000 /- /:

/

- 2OOO 4OOO

\\

-_ 2OOO I000 _.---- - ._.._ ..__ E E 0 nr -I000 -2000 2O -15 -Io -5 o 5 Io 15 ,8,deg (a) Zero forward acceleration at _ = 0 °.

Figure i0.- Lateral stability characteristics in the transition range.

'w, deg V, knots o 2.0 67.3 n ---40 4::5.0 O-- -60 3 I.I J_ I000 LL .--- __ o - IO00 J 2OOO / .K" / / I000 ........., .---'-J E £ E ¢- -I000 7, / / >-- -2000 I000 x ¢,- _' 0 E E _ _ i_ _L -- -I000 n_ -2OOO lO 15 -20 -15 -I0 -5 0 5 2O /3, deg = i0 ° .

(b) Zero forward acceleration at Figure i0.- Continued.

Iw, deg V,knots o 20 58.0 -_ I000 a _u 57.2 <>_ _60 14.7 -I000 i:1 ._1 I000 N i t E E -I000 j

• O

>- -2000 3000 ¸ J_ I000 x >-- C I1J E 0 o E c- -- -I000 £E -2000 -3000 2O -15 -I0 -5 0 5 I0 15 -20 B, deg 0 ° (e) Acceleration of i/4g at Figure i0.- Continued.

'w, deg V,knots 20 58.0 rI_ -40 37.2 -60 14.7 .Q I000 h O -I000 .J ,_ 2000 It; ]-- -- - N I000 E E o 0 E ¢.- /

f

__ -_ooo

..Q I000

2"-

(a3 x ¢.-- Z E E - 1000 L--'_ c- n-- -2000 -20 - 15 -I 0 -5 0 5 I0 15 /9, deg (d) Acceleration of i/4g at c_ = i0 °.

Figure i0.- Continued.

Iw, deg V, knots 2ooo_i ! T -- --40 61,5 i0 ........ 60 38.4 iooo 5 o o -I000 2000 r t .-" j ,d'_, = 0 --" -I000 -_

i / "

/ j _ -2ooo / j :p. /" -3OOO ,/ -4O0O 4OOO 3000 < _" _ _ -...,.

\ 2000 _'_ \ I000

\

o I E -I000 n_ \ "2000 !

-3000 i i ] i

i

-4oc_ ] i 15 20 -20 -15 -I0 -5 o 5 I0 B, deg (e) Deceleration of i/4g at m = 0°.

Figure lO.- Continued.

'w, de_ V,knots o 20 78.5 I000 []-- _ .40 61.5 o ...... 60 38.4

o

-I000 c E t" E ¢- -,ooo _--

/

-2000 / -3000 / X I000 c" E E C n" ,...., - I000 - 2000 -20 - 15 -I 0 -5 0 5 10 15 20 _, deg (f) Deceleration of i/4g at _ = i0 °.

Figure i0.- Concluded.

lw, deg 8O O I000 u_

,! o

U 0 © (9 o o 4.-- o -I000 d I000 N o C _) E o E - I000 o >- -0- I000 x C E o E -- -I000 -15 -I0 -5 0 5 IO 5 2O -6 -20 n_ Angle -of s,desh p, _, deg Figure Ii.- Lateral stability characteristics in 45 ° climb. V = 9.27 knots at i w = 80o; _ = -45 °.

._ I w ,deg V,knots - o-- 20 673 _- []-- -- 40 43.0 LI..

1000 <_.... 60 ,.3I. I --- 8.4 o -6 0 !='-= ...... "--_ o 3000 _I -_ 2000 J J J / f j.__1 _ I000 i o _ E 0 _ 1_ # -,ooo _ _ _ - 2000 3000 _ .J_ x, \ rL

\

I000 E E

L]\

"6 n.- - I000 - 2000 -30 -20 4 0 0 I0 20 30 SOR,deg (a) Zero forward acceleration at _ = 0 °.

Figure 12.- Aileron effectiveness in the transition range.

o

c- (D _C_Jf E O E -IOCX3 _ _ _/ c- O ..Q \ x_ 2000 _,_

._ \

_- I000 --'--- " o _c _ n _ -I000 -2000 -30 -20 -I0 0 I0 20 30 aR,deg = 0° (b) Forward acceleration of i/4g at _ Figure 12.- Continued.

.1'

._J 2O00 ¢: I000

7;_ --E

i //_ E E _n t- I/ i I / i - I000 O >- -20OO

3ooo \\

\ \ 2000 \.

"_._ "_.,., C I000 _ E E n _

\

- I000 - 2000 -30 -20 40 0 I0 20 30 aR,deg (c) Deceleration of i/4g at _ = 0 °.

Figure 12.- Concluded.

Iw ,deg o-- 80 I000 o o .-, .)

O O J

J

N I000

J

E J E o 0 E J f.- o -I000 >- ,,', I000 (L'---...

x _ c ° 0 E E ¢.- = - I000 2_O 3O -6 -50 -20 -4O 0 I0 n_ f _o R, deg (a) 26.5 ° climb; _ = -26.5o; V = 9.4 knots.

Figure 13.- Aileron effectiveness in 26.5 ° climb and descent at i w = 80 ° .

Jw,deg 0 -- 8O La I000 o o o _J ¢:: ,,.) I000 .-.---I N E ,,., 0

E J

o J / E __...... J --= -I000 I000 X ._...-._ ) f f E ..---_ r---" Q.J E o E -I 000 c -30 -20 -00 0 to 2o 3o c_ _o R,deg (b) 26.5 ° descent at _ = 26.5 °} V = 6.7 knots.

Figure 13.- Concluded.

>: ,ooo h _ 0 ' w,deg V, knots 5 I000 ---- 40 43.0 o -- 20 673 ...1 , 60 3 II = -I000

\

E -2000 E

\

-3000

\

{3 >-

\

-4000 \ \ - 5000 I000 .,_- - I000 g E E - 2000 r_ - 3000 \\ -4O00

\

- 5000 - 60O0 left 15 14 13 12 II I0 9 Right 9 I0 II 12 13 14 15 Propeller blode ongle,/_p,deg Figure 14.- Effectiveness of differential propeller pitch control with zero forward acceleration at _ = 0 °.

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