Summary of Flight-Test Results of the VZ-2 Tilt-Wing Aircraft

NASA TN

D-989

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TECHNICAL

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SUMMARY OF FLIGHT-TEST RESULTS OF

THE VZ-2 TILT-WING AIRCRAFT

By Robert J. Pegg

Langley Research Center

Langley Air Force Base, Va.

ADMINISTRATION

AND SPACE

NATIONAL AERONAUTICS

WASHINGTON

February 1962

IN

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

TEC_NZCAL NOTE D-989

SUMMARY OF FLIGHT-TEST RESULTS OF

THE VZ-2 TILT-WING AIRCRAFT

By Robert J. Pegg

_Y

L

Flight-test information gained from a tilt-wing research aircraft

tested at the Langley Research Center has shown that design problems

exist in such fields as low-speed stability and control, handling

qualities, and flow separation during transition. The control power in

the near-hoverlng configuration was considered by the pilots to be

inadequate in yaw, marginal in pitch, and excessive in roll.

Solutions for some of the design problems are indicated; for

example, the addition of a leading-edge droop to the wing in an attempt

to delay flow separation resulted in such significantly improved handling

qualities in the transition range that an additional descent capability

of 1,100 feet per minute was obtained.

INTRODUC TI ON

The research program for vertical-take-off-and-landing (VTOL) air-

craft being carried on by the National Aeronautics and Space Administration

includes phases of simulator, wind-tunnel, and flight testing. Previous

work attempted in some of these areas and pertaining to the VZ-2 aircraft

includes: (a) studies of control response characteristics made with a

variable-stability helicopter (ref. l) which provide the basis for the

control-power criteria of reference 2; (b) pilot opinion correlated to

problems encountered in flying the various VTOL test beds (ref. 3); (c)

time-history information of the aircraft making complete transitions and

specific problem areas studied from an operational standpoint (refs. 4 to

6); (d) i/4-scale-model free-flight and force-test data (refs. 7 to i0).

Examination of existing flight data showed a lack of satisfactory

full-scale flight information on a tilt-wingVTOL-type aircraft. The

tests reported herein were therefore carried out to gather documented

data and pilot opinion on ground-effect characteristics, dynamic and

static stability, maneuver stability, control response, and rate-of-

descent limitations of the test-bed aircraft.

The investigation presented herein is aL[med at bridging the gap

between pilot experience and the wind-tunnel or theoretical results by

presenting flight measurementsof aerodynami; characteristics of the

VZ-2 aircraft and interpreting the results oi_ these measurementsin

terms of meansfor improvement for future de 3igns.

SYMBOLS

b L

wing span, ft

C

chord of rotor blade, in. 5

acceleration due to gravity, 32.2 ft/sec 2

g

moment of inertia about roll axis, slug-ft 2

Ix

moment of inertia about pitch axis, slug-ft 2

Iy

moment of inertia about yaw axis, 31ug-ft 2

Iz

wing angle (measured from a line parallel to the upper

iw

longeron), deg

M

pitching moment, ft-lb

AM ft-lb

speed stability,

AV knot

number of rotor blades

n

P

power, hp

P angular rolling velocity, radians/_ec

R

rotor-blade radius, in.

V

airspeed, knots

W

weight of aircraft, lb

fuselage angle of attack, deg

_f

TEST EQUIPMENT

A drawing of the test aircraft including modifications is shown in

figure l, and its principal dimensions and physical characteristicsare

listed in table I. Figure 2 shows the test aircraft in transition

flight.

Control of the aircraft in forward flight is obtained longitudinally

from an all-movable horizontal-tail surface, with lateral and directional

control obtained from the ailerons and rudder, respectively. In hovering

flight, pitch and yaw control are obtained by varying the thrust of small

L

fans which are located at the aft end of the fuselage. Differential

collective pitch on the main rotor blades produces roll control. A

phasing mechanism within the aircraft automatically varies the amount of

4 differential collective pitch, aileron, and pitch-fan blade pitch as the

wing angle is changed. The rudder and horizontal tail, however, were not

connected to this phasing mechanism during this investigation and their

deflections did not vary with wing angle.

Power is supplied by an 850-horsepower gas-turbine engine and is

controlled by the pilot through the collective-pitch lever. Maximum

usable power has been limited by various dynamic components to 650 horse-

power. The aircraft has automatic stability augmentation equipment that

provides rate damping about the pitch and roll axes. This equipment may

be turned on or off by the pilot.

Instrumentation was provided to measure and record airspeed, altl-

tude, rotor speed, control and wing positions, longitudinal and normal

accelerations, angle of sideslip and angle of attack, engine torque, and

angular velocities about the body inertia axes. Vibratory-loads measure-

ments were also recorded during these flight tests and are presented in

reference ll.

A camera mounted on the vertical tail photographed wool tufts on the

_ight wing for visual indication of flow disturbances. A camera was also

mounted on the cockpit bulkhead to photograph the control-panel instru-

mentation for visual monitoring of the pilot's instrument display.

DISCUSSION OF DATA

Dynamic-Stability Characteristics

Lon6itudinal d_namic stability.- A measure of the dynamic-stability

characteristics of the aircraft is indicated by the resulting motion of

the aircraft when it is disturbed by a gust or a sudden angle-of-attack

changebrought on by the pilot. Figure 3 showsthese characteristics by

time-history plots of longitudinal pulse inputs (pitch dampers off) with

the accompanyingangular pitching velocities in the hovering configuration

in ground effect and three forward flight co_iditions of 40, 62, and

i00 knots. A longitudinal disturbance (puls,_ input) under near-hovering

conditions (5-knot wind) initiates an oscillation which expands at such a

rapid rate as to appear as a divergence on the first swing back through

trim. Figure 3(a) showsthis condition, where after approximately

3-5 seconds pitching velocity has increased to such a magnitude, with no

indication of peaking, that the pilot was compelled to apply corrective

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action. The pilot indicated that he felt that the aircraft had zero or

negative angular-velocity dampingwith a tendency toward simple diver-

gence. Also, the time histories of typical _)ulse inputs in the forward-

flight conditions, as shownin figures 3(b) t;o 3(d), indicated that as

the wing angle is decreased the period of the oscillation is decreased

and the damping of the oscillation is increased. Figure 4 shows the

variation of period of longitudinal oscillation with airspeed.

Lateral dynamic stability.- In the hovering configuration (i w = 80o),

the roll response to a pulse input is an oscillation which expands at such

a rapid rate as to appear as a divergence on the first swing through the

trim position (fig. 5(a)). At cruise speeds (iw = 9°), the oscillation is

well damped, but of short period (fig. 5(b)). At the intermediate air-

speeds, time-history traces were not obtainel because of the directional

instability of the aircraft.

Unstable Dutch roll oscillations were n)ted by pilots following

recovery at high power from a descent condition with a wing angle of 20 o.

These oscillations were believed to be due t) the downward inclination of

the principal inertia axes with respect to the flight path following the

recovery from the descent.

Static-Stability Characteristics

Static lateral-directional stability.-'_,_o flight conditions

(i w = 40o3 V = 40 knots; iw = 9 ° , V = lO0 kaots) were explored to

determine the static directional stability characteristics of the test

aircraft. Figure 6, which is a plot of pedal position as a function of

sideslip angle, shows that an instability exLsted in both flight

regimes. This instability is indicated by a reversal in the slope of

the curve between angles of sideslip of lO ° _nd -15 ° for iw = 40 ° and

between angles of sideslip of -3 ° to 2 ° for i w = 9 ° . Pilot comments

for other intermediate flight ranges indicate that this condition exists

throughout the forward-flight speed range. _t the lower airspeeds,

pilots say that this instability is objectionable as a result of the

deterioration of handling qualities because of increased effort and

diverted attention_ however_ at the higher airspeeds this condition is

more objectionable because of high angular accelerations away from the

desired trim and the possibility of high structural loads.

Test points taken from these same two flight conditions indicate

that the variation of lateral stick position with sideslip angle showed

a positive dihedral effect which pilots felt was adequate. These data

are plotted in figure 7 and show that as airspeed is increased, increased

dihedral effect is obtained.

Speed stability.- The speed-stability variation of longitudinal

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stick position with airspeed for each of several fixed wing angles and

i

constant power positions is shown in figure 8. The curves for airspeeds

below 40 knots have been replotted from reference 4. The steepness of

the slopes at the low-speed wing settings indicates that large pitching-

moment changes will be experienced with inadvertent changes in airspeed,

for example_ in gusty air and during longitudinal oscillations. Pilot's

comments indicated that flatter slopes would result in more favorable

flight characteristics.

Converting the longitudinal stick position from figure 8 into

pitching moment resulting from the combined tail fan and elevator control

ZkM decreases with increase in for-

(fig. 9) shows that speed stability

ward speed.

Maneuver Stability

At cruise speeds where significant changes in lift can be made by

changing the angle of attack, measurements of the buildup of the normal

acceleration are made in a wind-up turn, which is a constant-altitude

coordinated turn. At all speeds where appreciable normal acceleration

results from angle-of-attack change, the early normal-acceleration

response to a step input has been found to be of importance to the pilot.

At the lowest speeds, when a change in the attitude of the aircraft does

not produce appreciable lift change or change in flight path, only the

step inputs and not the wind-up turn maneuvers are made. In a step-input

maneuver, the buildup and peaking of the angular pitching velocity has

been found to be a primary parameter in the study of maneuver stability.

Reference 12 outlines acceptable angular-velocity and normal-

acceleration characteristics for the low-speed range for helicopters and

may be considered applicable for VTOL aircraft in general.

Wind-up turn.- Figure i0 presents results from a wlnd-up turn at

a constant velocity of lO0 knots and wing angle of 9 ° and shows that

increasing rearward longitudinal stick is required with increasing

normal acceleration_ as would be expected in normal airplane flight.

u

Step input.- Figure Ii gives time-histo_" plots of angular pitching

velocity 3 normal acceleration, and longitudinal stick position during

step-input maneuvers. At all test conditions the angular pitching

velocity is concave downward after 2 seconds &s required for helicopters

in reference 12 and is considered satisfacto_'. The configuration with

iw = 20 ° (fig. ll(a)) is the only configuration of those shown which

shows any discernible change in normal acceleration (initial positive

buildup limited by stall). At this particular flight configuration,

abrupt flow breakdown occurs over the wing causing a nose-down pitching

moment and a reduction in acceleration and angular pitching velocities

at the onset of stall. L

Variation of Power Required Wi_h Airspeed

The variation of wing angle of attack with airspeed is shown in

figure 12 and includes data where the fuselagc_ angle of attack was

varied up to i0 °. Figure 13 shows the power required for various level-

flight airspeeds of the aircraft before and after wing modification.

Data were obtained at aT _ 0 ° both before and after the aircraft wing

was modified, with no critical change except a slight decrease in the

power required as a result of unstalling the wing. Power measurements

presented at wing angles greater than 60 ° wer_ obtained near the ground

because of the lack of rotor horsepower avail_ble to hover out of ground

effect. All power readings indicate the totaJl, engine output and not

actual rotor horsepower.

Control

The control power, which is of primary il_ortance in the handling

qualities of an aircraft, is defined herein a:_ the moment on the air-

craft produced for a given control displaceme1_t. The control power in

the near-hovering configuration was considere([ by the pilots to be inad-

equate in yaw, marginal in pitch, and excessi_e in roll.

Roll control power.- Roll control power _n the hovering configura-

tion is about twice that of the cruise configuration as shown in fig-

ure 14. Pilots have objected to the excesslw_ roll control power in

the hovering configuration obtained with the ])resent linkage arrangement.

Internal mechanical changes can be made to the: linkages to vary roll

control power, but have not been made during ",his series of tests.

From the hovering position to full wing-down ]_osition, roll control

per inch of stick goes from 1.08 radians/sec2 to approximately

0.6 radians/sec2 in.

in.

Maximum roll velocities induced by the pilot in hovering flight on

the test aircraft are greater than desirable. Figure 15 shows the

rolling velocity per inch of lateral stick deflection plotted against

airspeed Objectionable rates of 32 deg/sec • are noted in hovering while

in.

deg/sec

acceptable and desirable values of 10.9 are obtained at lO0 knots.

in.

This value corresponds to pb/2V = 0.068 at the high-speed condition

with full lateral stick deflection. These records were taken with roll

dampers turned off.

Yaw control.- The variation of yaw-fan static thrust with pedal

L

displacement is shown in figure 16. This nonlinear control charac-

teristic, particularly near zero fan thrust, is objectionable to pilots.

The average maximum yaw control power provided by the fan, experimentally

determined in hovering, is about 885 foot-pounds. This moment was

insufficient to handle the random yawing motions produced by the flow

characteristics near the ground. For a test vehicle of this type, little hazard results from the lack of yaw control in hovering; however,

for operational vehicles intended to perform precision maneuvers, the

control-power requirements of reference 12 are firmly recommended.

This subject is discussed more completely in reference 2.

Trim

Trim chan_e with airspeed.- The variation of longitudinal stick

position with trim airspeed is shown in figure 17. Two conditions were

explored: one at a fuselage angle of attack of 0 ° and the other at

varying fuselage attitudes. In trim unaccelerated flight, 30 percent

of the total longitudinal stick travel was used to maintain a constant

fuselage angle of attack of 0°. Pilots commented, however, that large

trim changes were not noticed at airspeeds below 35 knots. In normal

practice, the aircraft was usually flown through the speed range with

the fuselage attitude varying somewhat. The only trim change which was

objectionable under these conditions occurred during a rapid conversion

from a wing angle of approximately 20 ° to 9 ° and back. This change is

shown in figure 8.

Trim change with po_er.- A definite trim change with power is

experienced by the aircraft during transition. Test points taken at

two wing angles showing the variation of longitudinal stick position

for trim with power are given in figure 18. The curve representing trim

change with power at a wing angle of 40 ° is an example of the steepest

gradient encountered by this aircraft, but pilots indicate that it is

within tolerable limits although it is more desirable to have no trim

change. At a wing angle of 25 °, the trim longitudinal stick position

changes approximately 15 percent of the total travel from a rate of

climb of 1,400 feet per minute to a rate of descent of 13300 feet per

minute.

Ground Effect

Figure 19 showsthe behavior for the tilt-wlng configuration in a

near-hovering condition in and out of ground _ffect without automatic

stabilization. Figure 19(a) illustrates moderate aircraft and control

motions out of ground effect. For the aircraft in ground effect

L

(fig. 19(b)), aircraft and control motions are manytimes greater, with

erratic angular-velocity changes up to lO degrees per second and fre-

quent control motions of several inches. The effects of the ground

have been noted with the aircraft wheels as high as 19 feet above the

ground. The problem of erratic aircraft motion in ground effect may be

expected to arise in practice for a variety of designs, especially when

the aircraft is operating over rough ground, in gusty air, or when it

is not maintaining a level attitude of the wings. Buffeting is apt to

be encountered, also, in a variety of designs

Conversion

Data have been obtained for full convers_ions from hovering to for-

ward flight and back to hovering, and a typic_l resulting time-history

plot is presented in figure 20. In general, take-off, landing, and the

low-speed portion of the conversions have caulked the pilots most concern;

however, the converslonmaneuver presents no great difficulty to the

pilots.

Rate-of-Descent Limltat_ons

The most critical region of operation for the VZ-2 is encountered,

as a result of wing stall and separation, in ([ecelerating conversion

and/or descent. This wing stall and separation leads to buffeting and

erratic motions, with general difficulty in _ndling the aircraft. This

stall results in regions of the expected VTOL veloclty-rate-of-cllmb

envelope being completely unacceptable for no]_nal flight operations.

The limits of the flight envelope have been d_fined by pilot opinion

and are presented in figure 21.

In an attempt to alleviate the flow sepaJBtion, a modification

to the leading edge was made. The modification consisted of additional

thickness near the leading edge and an increased leading-edge radius

and resulted in approximately 6 ° of leading-edge droop. This modifica-

tion is illustrated in figures l(b) and l(c).

Figure 21 shows the successive shifts in the rate-of-descent bound-

aries for the addition of partial-span and full-span droop. It is not

to be implied from one success with this device that a thorough under-

standing of this flow-separation problem has been attained. The leading-

edge droop, as such, was not expected to be nearly so effective, and the

changed relative position of the wing and propeller axis may have had a

material effect on the results.

Basic wing.- Figure 21(a) presents the rate-of-descent limitations

for the basic wing condition. In level flight, as the wing is raised

and the airspeed is decreased from 75 knots to approximately 60 knots

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(i w = 20o), abrupt flow breakdown occurs over the wing and causes wing

i

dropping or aircraft roll-off accompanied by buffeting, yaw disturbances,

and a nose-down trim change (requiring rearward-stick displacement to

offset).

In a constant-descent condition, as the wing angle is increased

from 30 ° to 50 ° and the aircraft airspeed drops from 55 knots to 35 knots,

limiting stall conditions are not as severe as those found at lower wing

angles because of the decreased dynamic pressure; but considerable direc-

tional and longitudinal disturbances and a reduction in control effec-

tiveness make flying in the shaded areas equally undesirable. At wing

angles greater than 50 ° , the limiting feature seems to arise chiefly

from the general lack of control effectiveness about the directional

and longitudinal axes of the aircraft.

Basic win_ with droop.- The first improvement, which is shown in

figure l(b), was accomplished through the addition of a drooped leading

edge over the outboard portion of the wing. This modification increased

the rate-of-descent capability at an airspeed of approximately 60 knots

by approximately 600 feet per minute (fig. 21(b)). The second improve-

ment, which increased the maximum rate of controllable descent at this

same airspeed by approximately another 500 feet per minute (fig. 21(c)),

was due to the further addition of inboard drooped leading edges. Other

modifications, such as the inboard wing fences illustrated in figure i

and aileron droop, indicated no significant improvement over that shown

by the full-span drooped leading edge.

The leading-edge droop not only lowered the "unacceptable" bound-

aries, but it also made flying in the "poor" areas much easier.

It should be pointed out that in a decelerating condition in level

flight, the boundary for the equivalent descent rate indicates the region

of unacceptable flight characteristics; that is, the boundaries of fig-

ure 21 are effectively raised by deceleration. Therefore, in level

flight the conversion rate can also be limited by wing stall and

separation.

lO

CONCLUSIONS

Pertinent results of the flight-test investigation of the VZ-2

tilt-wing VTOL/STOL aircraft indicate the following conclusions:

i. Pitch and roll pulse inputs initiated _n oscillation which

expandedat such a rapid rate as to appear as _. divergence on the first

swing through the trim position.

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2. The aircraft shows increasing positive speed stability with

decreasing airspeed, a condition which can cau_;elarge variations in the

pitching momentwith inadvertent changes in airspeed.

3. Hovering control power of the aircraft is considered by the

pilot to be inadequate in yaw, marginal in pitch, and excessive in roll.

4. Ground interference causes erratic airc:raft motions which, with-

out the use of automatic stabilization, limit operation when the air-

craft wheels are within 19 feet of the ground.

5. Wing stall and separation leading to buffeting, erratic motions,

and general difficulty in handling the alrcraf_, result in the desired

VTOLvelocity-rate-of-cllmb envelope having regions completely unac-

ceptable for normal flight operations. The a@lition of a full-span

leading-edge droop decreased the regions that were unacceptable for

normal flight and thereby permitted an additional 1,100 feet per minute

descent capability at an airspeed of approxima_ely 60 knots.

Langley Research Center, National Aeronautics and Space Administra;ion, Langley Air Force Base, Va., Septemb,_r 25, 1961.

ii

REFERENCES

i. Salmirs, Seymour, and Tapscott, Robert J.: The Effects of Various

Combinations of Damping and Control Power on Helicopter Handling

Qualities During Both Instrument and Visual Flight. NASA TN D-58,

1959-

2. Tapscott 3 Robert J.: Criteria for Control and Response Characteris-

tics of Helicopters and VTOL Aircraft in Hovering and Low-Speed

Flight. Paper No. 60-51, Inst. Aero. Sci., Jan. 1960.

L

3- Reeder, John P.: Handling Qualities Experience With Several VTOL

Research Aircraft. NASA TN D-735, 1961.

4. Thongs, Lovic P., llI: A Flight Study of the Conversion Maneuver of

a Tilt-Wing VTOL Aircraft. NASA TN D-153, 1959.

5. Segner, D. R.: Navy Evaluation of the Vertol Model V-76C Tilt Wing

Aircraft. Report No. i, Flight Test Div., U.S. Naval Air Test

Center (Patuxent River, Md.), Oct. 31, 1960.

6. Pegg, Robert J.: Damage Incurred on a Tilt-Wing Multipropeller

VTOL/STOL Aircraft Operating Over a Level_ Gravel-Covered Surface.

NASA TN D-535, 1960.

7. Tosti, Louis P.: Aerodynamic Characteristics of a i/4-Scale Model

of a Tilt-Wing VTOL Aircraft at High Angles of Wing Incidence.

NASA TN D-390, 1960.

8. Newsom, William A., Jr., and Tosti, Louis P.: Force-Test Investigation

of the Stability and Control Characteristics of a i/4-Scale Model

of a Tilt-Wing Vertical-Take-Off-and-Landing Aircraft. NASA

MEMO II-3-58L, 1959.

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

Characteristics of a 1/4-Scale Model of a Tilt-Wing Vertical-Take-

Off-and-Landing Aircraft. NASA MEMO 11-4-58L, 1959.

i0. Tosti, Louis P.: Rapid-Transition Tests of a i/4-Scale Model of the

VZ-2 Tilt-Wing Aircraft. NASA TN D-946, 1961.

ii. Ward_ John F.: Structural-Loads Survey on Two Tilt-Wing VTOL Con-

figurations. NASA TN D-729, 1961.

12. Anon.: Helicopter Flying and Ground Handling Qualities_ General

Specification for Milltary Specification MIL-H-8501A, Jan. ll, 1961.

TABLEI .- PHYSICAL CHARACTERISTICS OFTHEVZ-2 AIRCRAFT

Rotors :

Diameter, ft ......................... 9.5

Blade twist (linear_ root to tip), deg ............ 19.2

Airfoil section ........... NACA 0009 with 0.5-in. cusp

Solidity (_) ........................ 0.218

L

Distance between propeller axes, ft ............. 14.67

i

Operational speed, rpm .................... 1,416

Differential pitch, deg ................... _+2

Wing:

Span (excluding tips), ft .................. 24.88

Chord, ft .......................... 4.75

Airfoil section ..................... NACA 4415

Taper ratio ......................... i

Pivot, percent chord ..................... 37.6

Ailerons -

Chord, ft ......................... 1.25

Tilt range (referenced to upper longeron), deg ...... 9 to 85

Vertical tail:

Height, ft .......................... 5.43

Approximate mean geometric chord, ft ............. 5.90

Basic airfoil section .................. NACA 0012

Rudder -

Chord, in ......................... 21.5

Span, in .......................... 58.0

Horizontal tail:

Span (less tips), ft ..................... 9.90

Chord, ft .......................... 3.00

Airfoil section ..................... NACA 0012

Length (distance from wing pivot to leading

edge of tail), ft .................. 10.475

Hinge point (distance from leading edge), in ......... 8.3

TABLE I.- PHYSICAL CHARACTERISTICS OF THE VZ-2 AIRCRAFT - Concluded

Control fans:

Diameter (both fans), ft .................. 2.00

Moment arm abQut wing pivot (both fans), ft ........ 12.35

Rotor speed, rpm ...................... 5,850

Fuselage length, ft ..................... 26.4

Engine ......................... Lycoming T53

Weight as flown with ejection seat, ib ............ 3,500

L

Center of gravity (for 9 ° wing incidence), percent M.A.C. 33.5

i

Center of gravity (for 85 ° wing incidence), ft forward of

pivot point (measured along longitudinal axis) ....... 0.135

Moments of inertia:

Aircraft weight = 3,432 ib -

IX , slug_ft 2 ....................... i, 634

Iy, slug-ft 2 ...................... 2,937

IZ/slug-ft 2 ....................... 3,988

Aircraft weight = 3,204 ib -

IX , slug_ft 2 ....................... i, 560

Iy, slug-ft 2 ....................... 2,899

IZ, slug-ft 2 ....................... 3,985

Total control travels:

Lateral stick, in ................. • •

Longitudinal stick, in ................... ii_

¢J ® o _ !,_

o _

I1) .,-4 % -,-4 _ o 0 _ _ ,,,-4 _m_m _ o % __--Q-- oJ

.; a

¢.1 r/l _ , 0.1 II!

/ Leed_n_-e_ve droop L'% (b) Airfoil section showing wing fence and droop leading edge.

wl.g moclf Ica_1on Upper Lo.er ordtnete ordinate O. 7125 i.75 -2.28 I._25 2.37_ -Z.79 2._5 3._72 -3.15 1. _.l_ -3.z75 5.70 -_'. o9 7 • 125 -- - Z. 92 8.55 5.2_3 -2.715 -- -_. 5 9.075 -8._ 12.825 I_.25 6.92g .?. 26_ Leadln@(-edl_e L.E. redluij2._?8 In.

Center at (2.691 tn...0.2_ In. ) droop ill l (c) Wing planform showing location of fences and leading-edge droops.

Figure i.- Concluded.

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t-J .-.1 4; .,-I f-t _H .,-t cH fJ ._t u) Q) I C_J Q_ t_ f f

2N

Noise up .2 ol

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u o

¢

t_ -.2 ,--I i= NOSe down -.3 5 o Full Rearward cf J o 'o _ .4 o 4, o

o_

Full Forward 5 0 1 2 3 _ 5 Time, see

(a) Longitudinal pulse input in the hovering configuration in ground

effect. 5-knot headwind; iw = 80°; P = 630 hp.

Figure 3.- Longitudinal dynamic-stability characteristics in the

hovering and transition flight configurations measured with pitch

dampers off.

,D ,.el u'A 0,1 I1 p.., II

j, I

k.n ,_ 0

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_ o

w

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r_ ,-4 I 0J n_ '_' 0

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o

I o Iaa_a% IP.%0 5O

_ t %u-ooagd =uoI%Isod >I.oT_ IT_ulpn_3Hucq

0 0 ._ 0 Z Z Nose up • 15 ,..-i c

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gh Nose d own--. 15 Full rearward 50 o

O----_. , C_-------O O .... _

-N p 4_IJ 4J .H Full forward 50 , , , , , L . ,

o I _ 3 _ 5 6 ?

Time, sec (c) Longitudinal pulse input at forward velocity of 62 knots.

iw = 20°; P = 490 hp.

Figure 3-- Continued.

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.!5 .IO

P

.05 i --- C, (_ ---- -.05 gull J -- --T _..... 'f:' < _, C> Full Forward o 1 2 3 I, 5 Time, _ec (d) Longitudinal pulse input at forwa_t velocity of i00 knots.

iw = 9°; P = 600 hp.

Figure 3.- Conclud '-_d.

I 0 ,,r-4 C)

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(3", iO o r-'l +_ C) ,43 o o Ul

o _

4._ -0 II ",4 0 0,-'1 •,-t r-'l / /

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-0 r-'t /

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0,'-t

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0 0 -0

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!

i ..... _1___1 I I t__ o Od oe_ _uoT_eiZToso Ieu;on_Z_uoT ,_o poT_e_ w .6 Right h q I i i .2 o o point of recovery initiation _n --,2 -.LL Left Full Right

J_

b o i u point of recovery Inltimtlon a a • ..........

5O Full 1 2 3 4 Left Time, see (a) Lateral pulse input while hovering into a _-knot headwind.

iw = Bo°_ P = 630 hp.

Figure 5-- Lateral dynamic stability characteristics in the hovering

a_d cruiae CO_f_tieae meesureclwi_ roll dampe_s off.

.2 Rlght .I O --.I I T t I t i i --.2 Le ft 5O Full Right o o o

_o

Fu 11 _ .J ...... J .... L ....... _ Left 5O I 2 3 n

5 6 7

T_m@, sec (b) Lateral pulse input at forward velocltyof 99 knots.

iw = 9o; P = 530 hp.

Figure 5.- Concluded.

0 a.l._ I b_ ,,-I b0 RD O",O o o c_ _ off', 0_ o oJ e_ I::I o oo 0 _ o__ .t-I _ v-t -,-4 o

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,--t Nose down --.I Rearward 50 i i i i i i o

C_C'_"-" _ -t;O OC

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Full Fozwsrd 50 1 2 3 _ 5 6 Time, sec (a) V = 62 knots; P = 450 hp; iw = 20 ° .

Figure ll.- Time-history records of normal acceleration and angular

pitching velocity resulting from a longitudinal control input at

three different flight conditions.

3O

2.0

_J

1.0 z .2 Noee up .1 No_e _owrl Full 5o K_ o o o .p H Full i i rJ Fo_-w_rd 50 I 2 3 Time, seo (b) V = 40 knots; P = _2_ hp; iw = 40 ° .

Figure ii.- Contlnued_

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NOSe down -, I 5o

o

_t

.o o

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Full i L L m Forward 50 1 2 3 _ 5 T_me, 8ec (c) V = 27 knots; P = 550 hP; i w = 50 °.

Figure ii.- Concluded.

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e-4 id i0 o t9_ ,--I f,.._ o ,--I o ID ,-H 4_ ,'t q-t o ,'-t J-I I ,"-I I O o o o _j mid _UOT_OeIJep _OTqe Ie_e_eI jo _ouT _ed _TOOIeA IIO_ Zpeg_

120 -

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,=

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_0

GO

I r i l I I I I

0 5O

Full left

percent Full right

Pedal position,

of

total travel

Figure 16.- Variation of yaw-fan static thrust with pedal displacement.

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o

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%

a; hD f_

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2.., le^_a% I_O_ jo %uooaod 'uoTaTsod _oI_m l_uTpnql gu°q 3o eI_US g_Iomn_ .,-4 Nole up 20 © .J / ,_ zo

} °

Ncee down -10 . Noee up _4 gr, i i i i i i i , _ _°2 gose dovn Full Re s r',qr.d 50 F.

o_ _ o ,-4 _o Full i I i i i i i i i ._ i i Forvar_ 50 io 2a _o ho 50 60 7o 80 90 _oo ii0 Time_ see Figure 20.- Time history of typical level-flight conversion performed by the tilt-wing VTOL aircraft.

_0 _0 o ) _0 o Llft-off Touchdown

I

I @

o

4-O0 z/ (3 © // "\ 2OO H i i i i i i i i I0 20 30 hO 50 60 70 _0 gO I00 II0 Time, sec Figure 20.- Concluded.

8OO _00 .80O Not explo:.ed (a) Aircraft with basic x-ing.

._oo c Aoeeptable .8o0 ,_°i _ ' ..... , , ,__. , , .16oo (b) Aircraft with basic wing with outboard droop.

_i_ Poor -&oo _Unscceptable-Sangeroua(_eavy b_ffettlns) l UnaccepCa_1,-=an_eroua (w_n¢ _oppln=_ .soo Acoep_able .12oo Not explored i i i -16oo 10 _0 30 hO 50 6o o AI _*epee_, Mnots (c) Aircraft with basic wing with full-span droop.

Figure 21.- Bate-of-descent boundaries.