NASA TN D-J389.
/V&J -Z b Y ¥ /
NASA TN D-1389
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
technical note D-1389
force-test investigation of the stability and control
CHARACTERISTICS OF A FOUR- PROPELLER TILT-WING
VTOL MODEL WITH A PROGRAMED FLAP
By William A. Newsom, Jr.
A wind-tunnel investigation has been ^^^g^^^of^model
tudinal and lateral stability £° w vTOL model. The model was
of a high-wing four- prope lie , , flaD which was programed to
equipped with a 55 -perce £<*“*£*£ 1 ^ ^ ^traded lor the 0 °
deflect as the wing rotated so th deflected downward for inter-
and 90° wing- incidence conditions and ^ performance and longi-
mediate angles of incidence to range. Three
tudinal trim characteristics m tested and the control effec-
tail aud the ailerons was also
It was found that, by
programing of the horizontal-tail incad^ chmge dur ing the tran-
eliminate the variation op to accomplish the same result
sition. It was not possib e, * fi ap effectiveness was augmented
with a partial-span flap, even horizontal- tai! deflec-
by the use of drooped conven “lerens^d^ used in conjunction
tion. It was also found a , satisfactory control, particularly
with the full-span flap did not Provide satisfy 7 ^^ ^ ^ hovering
because it was almost totally ine , . used dn conjunction with
= or by using the
rearward portion of the flap as an ai e
in the past, tests of various tilt-wing TOL ^tane^odels have
shown that such configurations ch^acteristica y transition
large nose -up pitching moment as the aircraiu
from hovering to forward flight. (See refs 1 and ? 'I -m, •
pitch trim with speed and wirL- 1 and 2. j This change in
of cente r- of - grav ity^pos i t ions Z ^ ?f n . severel y the range
transition successfully. Force tests ^of tilt S ?° SS1 ^ e to P erf °rm the
such as those of references 5 and h h tilt-wing- f lap combinations,
programing of flap deflection with the wing^Sf it ITl Wlt ^ proper
design a tilt-wing VTOL aircraft whioh k S * 11 ^ Possible to
trim change throughout the transit in r ^ ® ssentiall y no longitudinal
forward flight and h t IT fr ° m hoverin S to normal unstalled
performance characteristics! 1 C ° nflguratl °" also have favorable
trol characteristics^!" 3 !! moSl^ra'^ilt^n ° f 't“- S * abllit; > r 1111(1 oon -
landing high-vi ng transport o-r^n tilt-ving vertical take -off- and-
deflect as the ring U ? slottel flap Programed to
The flap programing was ar^ged so Saf ° ° f ° r f0rWard fli « ht
90 ° and 0° incidence oonditioS to • flap vas re tracted for the
ering and normal forward flight Lf Se%? “figuration for hov-
mediate angles of incidence fn nh+ ■ e flaP Was deflected for inter-
tudinal trim characteristics for thet av0: f^ dle Performance and longi-
re suits of the flight ^ests ofthf fli ght conditions. The
and the resnlts of^e^eleS S’ prSnt^hS 5 '
istics^in^transition ^nomaWr^Tti ° f ^ aerod y namic character-
aileron configurations. The contrSl^fectiveness ofth dl ?T erent flap ~
condition of stealy the
is an orthogonal^yste^with^he^'origin^t^the 1 ' 3 ^ 1 !^ ^^ 3 SySten1 ' " hi ‘
to the Z-axis, and the f-axis is perpeXSlar^S ^LfoTs^ry.
b wing span, ft
F L/ qS
pitching-moment coefficient, My/^Sc
yawing-moment coefficient, Mjy/qSb
side-force coefficient, Fy/qS
wing mean aerodynamic chord, ft
side force, lb
horizontal- tail incidence, positive when trailing edge is
wing incidence, deg
rolling moment, ft- lb
pitching moment , ft- lb
yawing moment, ft- lb
% ' ap
q free- stream dynamic pressure, lb/sq ft
S wing area, sq ft
V velocity, ft/sec
X,Y, Z coordinate axes
a angle of attack of fuselage, deg
0 angle of sideslip, deg
5 a'R deflection of right aileron, positive when trailing edge is
, mr m ° ded used ln the investigation was the model of a tilt-wing
VTOL transport used in the flight-test investigation of reference 5 .
igure i shows a three- view drawing and a photograph of the model and
able I presents the geometric characteristics. The model had four
3-blade propellers each of which was powered by an air motor. The pro-
pellers were not interconnected but the motors were all connected to a
common manifold and a valve was provided on each motor inlet by which
tne motor speeds could be synchronized, if necessary, before a test.
ITt however ' that the motors stayed in synchronization
S ° h ^ WaS ° nly necessai T to readjust the speed of a motor after
it had been disassembled for maintenance. The speed of the motors was
changed to vary the thrust of the propellers. The propeller blade angle
was 16 at 0.75 radius and the direction of rotation was as shown iii
1 igure 1.
, + ^\ Wing Was pivoted at the 65 -percent- chord station and could be
rotated between incidences of 0° and 90°. As the wing incidence changed,
he 55- percent- chord slotted flap was programed to deflect as shown in
igure 2. The model had an all-movable horizontal tail and conventional
rudder controls for forward flight. Two types of ailerons were used
unng e model tests. The original model configuration as shown in
1 igure 1 had a conventional aileron which was used in conjunction with
a partial-span single -slot ted flap. A second type of aileron was
installed on the model after, as a result of early tests, the slotted
flap had been extended to full span. (See fig. 1 .) A slot-lip aileron
v-as created by hinging the outer 30-percent span of each slot lip. A
iJ P fi^re r 3 SS SeCti ° n ° f the wing trough the slot- lip aileron is shown
The tests were made in the Langley full-scale tunnel with the model
T f T P ° rt StrUt near the l0Wer edge ° f the entrance cone and
about 5 feet above a groundboard. An electric strain- gage balance was
used to measure the forces and moments. All the tests were made at a
condition of zero forward acceleration by adjusting the tunnel speed
and model power for each wing incidence (fuselage a 0 ) until the
drag trim point was reached.
span invenSonS fu^Vroo^O 0 ,
span slotted flap with JO pe nercent-span slot-lip ailerons,
and a full- span slotted flap with 3 P (drag trimmed at fuselage
With the test condition set as “ “ “ 2^ stability and
a . 0°), the angle of attack was va * ® tail off and with the
control tests from -10° to 20 with from _ 10 o to 30°.
horizontal tail set at various angles of ^cid variation
For the full-span-flap configuration a testjas ma^ ^ ^ Qf side .
of rolling moment, yawing moment, , f , incidence from 0°
Slit) angles from -20° to 20° for each angle oi wing inc
^ 80^ Mleron control effectiveness was measured at 3 - 0
. ^-p n o i qO ana 20° were made at an
The tests at wing incidence , tive Reynolds number
airspeed of about 25 ta ? ots ,”“^_SeaTve(oc«y of aLrt 200,000. For
based on the wing chord an incidence, it was necessary to reduce
Z t^llirsSed S 25 knots to avoid exceeding the model motor
RESULTS AND DISCUSSION
The results of a force-test investigati^to ^mine^ae™-
dynamic characteristics of a model of a tilt wl g ^ ^ C enter-of-
gravity a positions a shown e in I1 figure v^^^m^uri^^hektests^described
r P d ri
approaches zero, the data for the rans f 51> 28 pounds. This
have been scaled up to the model y n 6 factor required
scaling of the data is accomplished ^determining ttefac ^ ^
to make the lift equal to the desired value (51^ multlplying
weight of the model dui ring ^ ^ « corresponding test velocities
all forces and moments by the iacror.
are scaled up by the square root of the factor.
Longitudinal Stability and Control
are presented in figures 5 and 6. ai -*- erons n °t drooped)
pitchSg^OTt^Sh of e attaJk 8 at he Varl f lon of lif1 e drag, and
tlons as well as for the *™ rsl horizontal- tail defies-
and trim characteristics from these data n condition. The stability
the tail-on case, the stability parameter Sy^s^asuref q P*
possiWe. ta in m ° ment WhSneVer
the effect of tail incidence on My was^irt^ll^ pitching m °ment,
average slope was used At , a virtuald y negligible and some
model was neutrally stable tSl off°l incldence from 90° to 60° the
tail did not increase tL 3 «ii1 , “e ca^eT^ir, ' ^
At lower angles of wing incidence tL mn ,! I f th 1 dynamic pressure,
off but, because of the higher dynamic- -n 6Came unstable with the tail
zontal tail made the model stabl^T The P £; ssure ; the addition of a hori-
of wing incidence from 80° to 4o°' the alS ° Sh ° WS that > at ^S^es
which cannot be trimmed with the horizontal tail & ft° Se-up P itchin g moment
curves shows that for the worst condition t 6 0 of ^ ? ° f the
require an upward force of about 2 percent of th d i ^ 61 V ° Uld
auxiliary control device at the tail which seems to^ a've^^SSf leant
slotted flap^and^HjO- percent-span con^ 8 ^ 1 ? 11 haV±ng a Partial- span
drooped 200 are pre in?ed in fW 6 s ? IT adallei ’ on v ^h the ailerons
variation of lift, dra g ^ ? ?' Flgure 7 P^sents the
several horizontal- tail deflections IndTo h^^ 311816 ° f &ttack for
summary plot showing stability and trim v, 1 " °P z ° ntal tail off, and a
this basic data is presented as fi OT1TV , A Cha £f Cter * Stics extr acted from
to determine quantitatively the effect of polnt of . thes e tests was
viatmg the trim problem at hieh sn^ip r • dr00ped aile rons in alle-
tesfs of reference 5 had Si ” Ce the fllght
effective in eliminating the pitching ^ thlS P roce dure- was not
data from the summary figure (fig- ft? f+f Problem. Comparison of the
figure 6 shows thatTven^Su^ tJ ^ ° f the desponding
at wing incidences above and below th ° Se ~V P pi ^ cllln S moments measured
vere reduced, the pitching LLTnear X M8r 6 °°
bad as for the previous configuration whirh h a n cidence was still as
The longitudinal stability and\he w* ? ? ad undro °P ed ailerons,
little changed h y the "»
The longitudinal data for the full- span- flap configuration are
presented in figures 9 and 10. Figure 9 presents the variation of lift,
drag* and pitching moment with angle of attack for several horizonta
tail deflections, and for horizontal tail off, for each angle of wing
incidence. The longitudinal stability and trim characteristics measured
from these basic data are summarized in figure 10. These data show that,
when the model was fitted with the full- span slotted flap, the pitching
moment in the critical range near i w = 50 ° or 60 was reduced to the
point where it was almost trimmed by use of the horizontal tail even
though at such low speeds the tail has little effectiveness. In fact,
analysis of the data of figure 9(d) indicates that the pitching moment
in this most critical condition could probably have been trimmed by the
use of a higher tail incidence of about 35°- For this full- span- flap
configuration the longitudinal stability was little changed from that o
the other configurations.
Lateral Stability and Control
Figure 11 shows the effectiveness of the aileron throughout the
wing incidence range for the partial- span- flap configuration with
undrooped conventional ailerons. Qualitatively, these data show the .
results that would be expected; that is, that the ailerons produce prin-
cipally yawing moments in hovering flight and rolling moments in forward
flight. One point that bears further study is the magnitudes of the
yawing moments produced in hovering flight. Inspection of figure 11
shows that at i w = 80 °, ±20° deflection of both ailerons would give a
yawing moment of ±10 foot-pounds. This value would correspond to a^
force of approximately ±2.5 pounds at the tail jet which was approximately
the force used in the flight tests of reference 5 for hovering in still
air. A better indication of the suitability of these ailerons for yaw
control in hovering might be obtained by a comparison of the control
power of these ailerons with that required in the handling requirements
for these airplanes. In this case if the model is a dynamically scaled
model in the range from l/5 to l/lO scale, the yaw control power of the
ailerons would be only about one-fourth to one-third of that indicated
as being required by reference 6.
Figure 12 shows the effectiveness of the aileron for the partial-
span-flap drooped-aileron configuration. These data show the expected
result in that, as the ailerons are deflected downward from the 20
drooped position, they tend to lose their effectiveness in producing
rolling moments at the low angles of incidence or yawing moments at the
high angles of incidence.
The effectiveness of the slot- lip aileron used with the full- span
flap is shown by the data of figure 13- There seem to be two important
points to note. First, the slot- lip aileron would not be usable as a
yaw control in hovering since it does not produce any yawing moment as
indicated by the data for the near hovering condition of i w = 80°.
Second, in the normal forward- flight conditions as represented by the
■'■w — 0 and 10 tests the slot-lip ailerons produce only about one-
third of the rolling moment of the conventional ailerons. (See fig. 11 . )
It would seem that a more satisfactory system for providing aileron con-
trol would be to actuate the entire full- span flap as an aileron or to
actuate the rearward portion of the flap. The effectiveness of such a
full- span aileron is shown in reference 7.
The data of figure 14 show the variation of rolling moment, yawing
moment, and side force with sideslip angle, and these data are summarized
m figure 15 in terms of the directional stability parameter M^p and
the effective dihedral parameter M Xp . The plots of rolling moment and
yawing moment in figure 14 are, as in other lateral data figures, to
some degree erratic. It is believed that, in general, the erratic data
are due in part to random gusts in the tunnel and wing stalling which
can cause large changes in some of the moments. For example, a rolling
moment of about 1.5 foot-pounds, which is representative of the scatter
in the data, can be produced by a difference in lift of 1 pound, which
is only 2 percent of the total lift, distributed over one semispan. In
this connection, a tuft survey showed that there was a severe stall over
the wing center section which at times, possibly due to wing asymmetry,
extended over the inboard portion of the right wing. The plots of
rolling-moment variation with sideslip are extremely unsymmetrical, but
seem to show, in general, the trends indicated by the slopes M Xp pre-
sented in figure 15.
The directional stability data show, in general, that the model
was unstable in the low- speed portion of the transition range and that
it was stable at higher speeds where the wing incidence was less than 30°.
Actually, the directional instability shown is very small; for example,
at i w = 50° which was the worst condition, M^ = 0. 3 to 0.2 ft-lb/deg.
This value is small compared with the 10 foot-pounds which was available
from the tail- jet reaction yaw control used on the model.
SUMMARY OF RESULTS
The following results were obtained from the investigation of the
static stability and control characteristics of a four- propeller tilt-
wing VT 0 L model having a single slotted flap programed to deflect as
the wing rotates.
1. For the full- span- flap configuration, the variation of trim
pitching moment throughout the transition range was small for the tail-
off condition with the particular flap programing built into the model.
2. With the horizontal tail fixed at low angles of incidence, the
model experienced large nose-up pitching moments during the transition
because of the download on the tail induced by the downwash from the
3* By properly programing "the horizontal- ia.il incidence to vary
with wing incidence, it would be possible to reduce the pitching-moment
variation through the transition range to zero or to a very low level.
4. It was not possible with a partial- span flap (even when augmented
by drooped ailerons and horizontal-tail deflection) to eliminate the
nose-up pitching moments encountered in the transition range.
5. The tests show that the slot- lip aileron tested in conjunction
with the full-span flap did not provide satisfactory control. Specifi-
cally, it was alm ost totally ineffective as a yaw control for the hovering
condition and was about one-third as effective for roll control in for-
ward flight as were the conventional ailerons tested.
6. Tests of the conventional ailerons used with the partial-span
flap and previous tests of the effectiveness of a full-span aileron
indicated that a more effective control could be obtained by actuating
the f ull -sp an flap itself as an aileron or by using the rearward portion
of the flap as an aileron.
Langley Research Center,
National Aeronautics and Space Administration,
Langley Station, Hampton, Va., June 7, 19°2.
1. Tosti, Louis P. : Flight Investigation of Stability and Control
Characteristics of a l/8-Seale Model of a Tilt-Wing Vertical-
Take-Off- And- Landing Airplane. NASA TN D-45, i960.
2. Tosti, Louis P. : Flight Investigation of the Stability and Control
Characteristics of a l/4-Scale Model of a Tilt- Wing Vertical- Take-
Off- And- Landing Aircraft. NASA MEMO 11-4- 58L, 1959.
5- Newsom, William A., Jr.: Effect of Propeller Location and Flap
Deflection on the Aerodynamic Characteristics of a Wing- Propeller
Combination for Angles of Attack From 0° to 80°. NACA TN 3917,
4. Kuhn, Richard E., and Hayes, William C., Jr.: Wind-Tunnel Investi-
gation of Longitudinal Aerodynamic Characteristics of Three
Propeller-Driven VTOL Configurations in the Transition Speed Range,
Including Effects of Ground Proximity. NASA TN D-55, i960.
5. Newsom, William A., Jr.: Flight Investigation of the Longitudinal
Stability and Control Characteristics of a Four- Propeller Tilt-
Wing VTOL Model With a Programed Flap. NASA TN D-I390, 1962.
6. Tapscott, Robert J. : Helicopters and VTOL Aircraft - Criteria for
Control and Response Characteristics in Hovering and Low- Speed
Flight. Aero /Space Eng., vol. 19, no. 6, June i960, pp. 38-41.
7- Newsom, William A., Jr.: Experimental Investigation of the Lateral
Trim of a Wing- Propeller Combination at Angles of Attack up to 90°
With All Propellers Turning in the Same Direction. NACA TN 4190
1958 • y }
TABLE I.- GEOMETRIC CHARACTERISTICS OF MODEL
Length, in 84.8
Diameter (maximum), in 10.4
Area, sq in 1,002.25
Aspect ratio 9
Mean aerodynamic chord, in 10.77
Airfoil section NACA 65-210
Tip chord, in 7*9
Root chord, in 13 . 2
Span, in 95
Taper ratio 0.6
Sweepback of 0.65 chord 0
Dihedral angle, deg 0
Pivot station, percent chord 65
Flap chord, percent wing chord 35
Aileron, conventional (each):
Chord, percent wing chord 35
Span, percent wing semispan 30
Aileron, slot- lip (each):
Chord, in 0.75
Span, percent wing semispan 30
Area (total to center line), sq in 269
Aspect ratio 1*97
Airfoil section NACA 0009
Tip chord, in 5*4
Root chord (at center line), in 18.0
Span, in 23.0
Taper ratio 0.3
Sweepback (leading edge), deg 25
Rudder (hinge line perpendicular to fuselage center line):
Tip chord, in 2.5
Root chord, in 4.05
Span, in 14.03
Area, sq in 24l.9
Aspect ratio 5«8l
Airfoil section NACA 0009
Tip chord, in 4.60
Root chord, in 8.3
Span, in 37 • 5
Taper ratio 0.55
Sweepback (leading edge), deg 7*3
Mean aerodynamic chord, in 6.62
Propellers (three blades each):
Diameter, in 20
Chord, in 2.5
Wing with full - span flap
(a) Sketch of model. All dimensions are in inches.
Figure 1.- Model used in tests.
Figure 1.- Concluded.
Figure 2.- Variation of model flap angle with wing incidence.
Distance of center of gravity
from wing pivot
Distance of center of gravity
forward of wing pivot
(horizonta I) , percent c
Figure 4.- Variation of model center of gravity with wing incidence.
(a) i w = 80°; V = 1-9 feet per second.
Figure 5.- Longitudinal stability and control characteristics of the
partial- span- flap configuration with undrooped conventional aileron
Figure 5.- Continued.
(d.) i w = 50 °; V = 27.6 feet per second.
Figure 5.- Continued.
(e) i w = 40°; V = 35*1 feet per second-
Figure 5-- Continued.
(f) i v = 30°; V = 4l.9 feet per second.
Figure 5 *" Continued.
(g) i w = 20°; V = 55-3 feet per second
Figure 5-- Continued.
(h) i w = 10 °.
Figure 5 .- Continued
't> de 9
Figure 6.- Variation of longitudinal stability and trim wn-h • . J
(a) i w = 80°; V =1.2 feet per second.
Figure 7 .- Longitudinal stability and control characteristics of the
partial- span-flap configuration with drooped conventional aileron.
i w = 70°; V = 7-6 feet per second.
Figure 7*- Continued.
Figure J.- Continued.
(f) i w - 30°i V = 40.0 feet per second.
Figure 7>- Continued.
Figure J.- Continued
Figure 7*- Continued
Figure 8.- Variation of longitudinal stability and trim with wing inci
dence for the partial- span-flap configuration with drooped conven-
tional aileron. Data from figure 7.
(a) i w = 80°; V = 1.2 feet per second.
Figure 9 .- Longitudinal stability and control characteristics of the
full-span-flap configuration with slot-lip aileron.
Figure 9*- Continued.
-10 0 10 20
(d.) i w = 50°j V = 23.8 feet per second.
Figure 9 .- Continued,
-10 0 10 20
( e )
40° ; V =31.2 feet per second.
Figure 9»- Continued.
Figure 9*- Continued
(g) i w = 20°; V = 52.6 feet per second.
Figure 9 .- Continued.
(i) iv = 0°.
1 w i ^9
Figure 10.- Variation of longitudinal stability and trim with wing inci
dence for the full- span- flap configuration with slot-lip aileron.
Data from figure 9 .
Figure 11.- Concluded.
(a) Transition range.
Figure 12.- Aileron effectiveness. Partial-span flap; drooped conven-
Figure 12.- Concluded.
0 10 20 30 40 50
Slot-lip aileron deflection, deg
(a) Transition range.
Figure 13.- Effectiveness of the slot-lip aileron used with the full-span
(b) Normal forward flight.
Figure 13.- Concluded.
O t ‘ ; r ; ; ] ; ; t [ ! ;rri 1 1 1 . ; i i . ■ — u ■ l-m
-20 -10 0 10 20
(a) Transition range.
Figure 14.- Lateral stability characteristics of the model. Full-span-
Figure 15.- Variation of directional stability and effective dihedral
with wing incidence for the full -span- flap configuration. Data
from figure 14 .
NASA-Langley, 1962 L“ 3003