Document
N 62 13850
NASA TN D-1365
COPY
TECHNICAL NOTE
D-1365
LONGITUDINAL STABILITY AND CONTROL OF A TILT-WING VTOL AIRCRAFT MODEL WITH RIGID AND FLAPPING PROPELLER BLADES By Louis P. Tosti Langley Research Center Langley Station, Hampton, Va.
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON July 1962 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION TECHNICAL NOTE D-136_ LONGITUDINAL STABILITY AND CONTROL OF A TILT-WING VTOL AIRCRAFT MODEL WITH RIGID AND FLAPPING PROPELLER BLADES By Louis P. Tosti SUMMARY An experimental investigation has been conducted to compare the effects of rigid and flapping propellers on the dynamic longitudinal- stability and control characteristics of a tilt-wing vertical-take-off- and-landing aircraft model in hovering and in the transition range from hovering to normal forward flight. The model had two interconnected three-blade propellers rotating in opposite directions with the thrust axes parallel to the wing chord. The investigation included both static force tests and flight tests with the model center of gravity at various locations.
In hovering flight, the pitching motions of the rigid-propeller configuration could be more easily controlled than those of the flapping- propeller configuration for similar conditions. In transition flight the model experienced large nose-up pitching moments that had to be trimmed out at low transition speeds, and these pitching moments were somewhat lower for the rigid-propeller configuration than for the flapping- propeller configuration. Consequently, with the same amount of pitch control available, the configuration with the rigid propellers could be controlled with a center-of-gravity location about 2-percent wing chord more rearward for the rigid-propeller configuration than for the flapping- propeller configuration.
INTRODUCTION The relative merit of rigid and flapping propellers for use on tilt- wing VIOL aircraft is a subject of interest to aircraft designers from consideration of aerodynamic and structural problems. In an attempt to provide some aerodynamic information on this subject, the present inves- tigation has been made with both rigid and flapping propellers on a tilt- wing VTOL aircraft model and has included dynamic longitudinal stability and control tests in both hovering and transition flight. The two sets
of propellers used in these tests were identical except that one had
conventional (rigidly mounted) blades, whereas the other had the blades
mounted with flapping hinges. The flight tests included hovering well
above the ground and slow constant-altitude transitions from hovering
to forward flight. The force tests consisted of longitudinal stability
and control tests in the transltlon-speed range.
SYMBOLS
!
drag, Ib F D lift, ib FL horizontal-tail incidence, positive trailing edge down, deg it lw wing incidence, deg My pitching moment, ft-lb V scaled-up aircraft velocity, knots angle of attack of fuselage, deg APPARATUS AND TESTS Model The i/4-scale model of the VZ-2 (Vertol 76) tilt-wing VTOL aircraft used in the tests of references 1 to 4 was also used in this investiga- tion. A photograph of the model is shown in figure 1 and a three-vlew sketch showing some of the more important dimensions is shown in figure 2.
Tables I and II list the geometric and mass characteristics of the model.
The wing was pivoted at the 57-percent-chord station and could be rotated between incidences of 4 ° and 86 ° during flight. The model was powered by a 6-horsepower electric motor which drove the two propellers through shafting and right-angle gear boxes. The speed of the motor was changed to vary the thrust of the propellers. The blade-form curves of the propellers are shown in figure 3.
The closeup photographs of the three-blade propeller hubs, which are seen in figure 4, show details of both the rigid and flapping pro- peller hubs. The flapping propeller provided for blade-flapping freedom
and variable blade pitch, whereas the rigid propeller hub was variable
only in blade pitch. In order to change from the flapping to the rigid
configuration or vice versa, the complete propeller assemblies (including
the hubs) were interchanged on the propeller shafts. The flapping pro-
peller was designed for pure flapping motion with a flapping hinge off-
set 5/8 inch from the axis of rotation and with no drag hlnges_ no
feathering, and no cyclic pitch. The three blades for any given hub
the
(at
were collectively controllable in pitch ±l_ U from a 12° setting
0.75 blade radius) for roll control in hovering and low-speed flight with the pitch of the right and left propellers being varied 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 full-scale aircraft. The model also had an all-movable horizontal tail and conventional aileron and rudder controls for forward flight.
The controls were deflected by fllcker-type (full on or off) pneu- matic actuators which were operated remotely by the pilots by means of solenoid-operated valves. The control actuators were equipped with integrating-type trimmers which trimmed the controls a small amount each time a control was applied. With actuators of this type, a model becomes accurately trimmed after flying a short time in a given flight condition.
Test Equipment and Setup The static force tests were conducted in the Langley full-scale tunnel under the same conditions as the tests presented in reference 1.
The model was strut mounted, and the forces and moments were measured by using an internal strain-gage balance.
The transltlon-flight investigation was also conducted in the Langley full-scale tunnel with the test setup illustrated in figure 5.
The electric power and compressed air are supplied to the model through a slack overhead line which also acts as a safety cable to prevent the model from crashing if it should go out of control. A more complete description of the test technique used in making free-flight model tests is given in reference 2. The hovering- and pltching-oscillation tests were made with an almost identical setup in a large building that pro- vided protection from inclement weather and the random effects of out- side air currents.
Flight Tests
Flight tests were madewith both the rigid- and flapping-propeller
configurations in hovering and transition flight. The detailed flight-
test results for the flapping-proReller configuration have been presented
previously in reference 2. The results that are most pertinent _Ith
regard to the comparison of the flapping- and rigid-propeller configura- tions are repeated herein for comparison with the test results obtained with the rigld-propeller configuration. Since the comparison of the flapping- and rigid-propeller configurations relies to a considerable extent on the pilot's qualitative observations, it is important to note that the flight tests for the two configurations were made within a week of each other under comparable conditions.
Hovering-flight tests were made to determine the stability of the uncontrolled pitching motions and the ease with which the pilot could control these motions. The hovering flights were made with a pitch Jet force of ±3.6 percent of the model weight. Transition-flight tests representing slow constant-altitude transitions were made with both the flapping- and the rigid-propeller configurations in order to study the stability and control characteristics of the model and to determine the effects of center-of-gravity position.
The center of gravity of the model for the hoverlng-flight tests was located 1-percent chord forward of and 15.8-percent chord below the wing pivot. For the transition flight tests the model was ballasted to several different center-of-gravity conditions. Actually, for the transi- tion condition, the center of gravity of the model moved downward approxi- mately 5-percent chord and forward approximately 5-percent chord as the wing rotated from 86 ° to 4° incidence.
Force Tests Force tests were made to measure the longitudlnal-stability and con- trol characteristics of the flapping- and rigid-propeller configurations in the transition condition. The data for the flapping-propeller con- figuration have been previously presented in reference 1 scaled up to represent the full-scale VZ-2 aircraft at a weight of 5, 159 pounds. Some of these data are repeated in this paper for comparison with data for the rigid-propeller configuration which are scaled up to the same condition.
The tests were made with power settings (propeller rotational speed) which, with the fuselage at zero angle of attack, gave zero forward acceleration at wing incidences of 20 °, 40 °, 60 °, and 80 °. The angle of attack was varied from -15 ° to 20 ° with the horizontal tail either off or set at a 15 ° angle of incidence. Since conventional aerodynamic coefficients approach infinity and lose their significance as the airspeed
approaches zero, the data have been scaled up to a full-scale weight
of 3, 139 pounds, and the center-of-gravity locations of the full-scale
aircraft are listed in table III.
RESULTS ANDDISCUSSION
Hovering
Whenthe model was first flown with the rigid propellers, after
having been flown previously with the flapping propellers, the pitch
pilot was immediately impressed by the fact that the model was steadier
in pitch and that the pitching motions could be controlled more easily
for normal hovering flight for the samecenter-of-gravlty position and
with the sameamount of pitch control available. Whenthe pitch pilot
subsequently allowed the unstable controls-fixed pitching oscillation
to build up and then tried to stop the oscillation and regain control
of the model, the difference was also very apparent. With the rigid
propellers installed, the pitch pilot was able to stop the oscillation
consistently and to regain control of the model fairly easily; whereas,
with the flapping-propeller configuration, he was unable to stop the
oscillation, as pointed out in reference 2. This difference in steadi-
ness and ease of control probably resulted from the fact that the model
seemedto have more damping in pitch and a less unstable pitching oscil-
lation with the rigid blades than with the flapping blades.
The present investlgation was intended only as a study of longitu-
dinal stability and control, and no attempt was madeto determine the
characteristics of the model in roll and yaw; however, there was no
effect of the change in propeller configuration on the rolling or yawing
behavior of the model that was sufficiently pronounced to be noted by
the roll and yaw pilots, who were only concerned with controlling their
phases of the model motion as smoothly as possible to facilitate the
study of the pitching motions.
Transition
The model developed a large nose-up pitching momentas it started
through transition as pointed out in reference 2. This samecharacter-
istic was observed in flight tests of the full-scale VZ-2 aircraft as
indicated in references 9 to 7 which might be of interest in connection
with the present investigation. The nose-up momentduring the low-
speed part of the transition, together with the ability of the pitch
control to trim it out, establishes the rearward end of the allowable
center-of-gravity range. Force and flight tests were madeto establish
any difference in the allowable rearward center of gravity between the rigid- and flapplng-blade configurations.
Force tests.- The basic force-test data in figure 6 are presented for both the flapping- and rlgid-propeller configurations in the transi- tion range for angles of wing incidence from 80 o to 20 °. A cross plot of the pitching-moment data of figure 6 at 0 ° fuselage angle of attack against forward velocity is presented in figure 7 to illustrate more clearly the trim problems at low speeds in the transition-flight range.
The trim problem due to the large nose-up pitching moments at low speeds has been discussed in reference 4 for the flapping-propeller con- figuration and would apply similarly to the rigid-propeller configura- tion. That is, there is a tendency with both the rigid- and flapping- propeller configurations for the nose-up pitching moments on the model to build up to a maximum at the start of transition and then to decrease with buildup of forward speed. The nose-up pitching moments reach a higher maximum for the flapping-propeller configuration than for the rigid-propeller configuration, and tend to remain higher with increasing forward speed. It can thus be seen from the force tests that for a given margin of pitch control, a pilot could control a rigid-propeller aircraft with a center-of-gravity location slightly more rearward than for a flapping-propeller aircraft. In fact, calculations based on the fairing of the curves shown in figure 7 indicate that the model could have been trimmed with the center of gravity about 2-percent wing chord farther rearward for the rigid-propeller configuration than for the flapping-propeller configurations.
Flight tests.- The flight tests also showed that the model exper- ienced a large nose-up pitching moment at the lower speeds during the transition and that this nose-up moment was larger for the flapping- propeller configuration than for the rigid-propeller configuration.
This fact was evident to the pilot in a qualitative way in that he had to apply considerable nose-down trim at low transition speeds and the amount of trim required for a given center-of-gravity position was greater for the flapplng-propeller configuration. It was also evident in a more qualitative way in that with the same amount of pitch-control moment available, the rigid-propeller configuration could be flown with a center of gravity 3-percent wing chord farther rearward for the rigid- propeller configuration than for the flapping-propeller configuration.
This value of 3 percent should be considered only approximate, however, since the center of gravity was moved in 3-percent increments.
Inasmuch as the forward end of the center-of-gravity range, which was established by the ability of the pilot to trim the model in hovering flight as shown in reference 2, would not be changed by the use of rigid propellers, the effect of the propellers on the rearward end of the
center-of-gravlty range would indicate that the model would have an
allowable center-of-gravity range about 3-percent wing chord larger with
rigid propellers than with flapping propellers.
Here again, it seemsworthwhile to point out that the differences
in propeller configuration (flapping or rigid) had no effects on the
rolling and yawing behavior of the model that were great enough to be
noticed by the roll and yaw pilots, who were primarily concerned with
controlling their phases of the model motion as steadily as possible
in order to facilitate the study of the characteristics of the model in
pitch.
CONCLUSIONS The results of this investigation of the longitudinal stability and control can be summarized as follows: 1. The pilot found that the model could be controlled more easily in hovering flight with the rlgid-propeller configuration than with the flapplng-propeller configuration under similar conditions, apparently because of greater damping in pitch and a less unstable pitching oscil- lation for the rlgld-propeller configuration.
2. In transition flight, the model experienced large nose-up pitching moments that had to be trimmed out at low transition speeds; these pitching moments were somewhat lower for the rigid-propeller con- figuration than for the flapplng-propeller configuration. ConsequentlY, with the same amount of pitch control available_ the pitch pilot could control the rigld-propeller configuration with a center-of-gravity loca- tion about 2-percent wing chord more rearward for the rigid-propeller configuration than for the flapplng-propeller configuration.
Langley Research Center, National Aeronautics and Space Administration, Langley Station, Hampton, Va., April 12, 1962.
REFERENCES
i. Newsom,William A., Jr., and Tosti, Louis P.: Force-Test Investiga-
tion 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-5-58L, 1959.
2. 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-Landlng Aircraft. NASA MEMO 11-4-58L, 1959.
3. Tosti, Louis P. : Aerodynamic Characteristics of a i/4-Scale Model
of a Tilt-Wing VTOLAircraft at High Angles of Wing Incidence.
NASATN D-390, 1960.
4. Tosti, Louis P. : Rapid-Transition Tests of a i/4-Scale Model of the
VZ-2 Tilt-Wing Aircraft. NASATN D-946, 1961.
5. Thomas_Lovic P., III: A Flight Study of the Conversion Maneuver
of a Tilt-Wing VTOLAircraft. NASA TN D-155, 1959.
6. Reeder, John P. : Handling Qualities Experience With Several VTOL
Research Aircraft. NASATN D-735, 1961.
7. Pegg, Robert J. : Summary of Flight-Test Results of the VZ-2 Tilt-
Wing Aircraft. NASATN D-989, 1962.
@ TABLE I SCALED-UP GEOMETRIC CHARACTERISTICS OF THE MODEL Propellers (3 blades each rotor): Diameter, ft ......................... 9.33 Solidity ........................... O. 239 Chord, ft .......................... 1.0 Flapping hinge offset, ft .................. 0.208 Wing: Airfoil section ..................... NACA 4415 Aspect ratio ......................... 5.42 Chord, ft .......................... 4.75 Taper ratio ......................... 1.0 Area, sq ft ........................ I18.2 Span, ft ........................... 24.88 Ailerons (each) - Chord, ft ......................... 1.22 Spanj ft .......................... 5.83 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 00]2 Aspect ratio ......................... 3.10 Chord, ft .......................... 3.0 Center-section chord, ft ................... 4.21 Area (including center body), sq ft ............. 33.2 Span, ft ........................... 9.90 l0 TABLE II COMPARISON OF MASS CHARACTERISTICS OF MODEL (SCALED-UP) AND FULL- SCALE AIBCRA_V_ Model Full-scale aircraft (scaled-up) Gross take-off weight (including one pilot and research instrumentation), lb. ..................... 3,553 3,290 Rolling moment of inertia, IX, slug-ft 2 i, 811 3,280 (hovering configuration) ..... "--A Pitching moment of inertia, Iy, slug-ft = (hovering configuration) ........... 5,890 2,851 Yawing moment of inertia, IZ, slug-ft 2 (hovering configuration) ........... 5,330 3,779
/
/
o a_ .el r_ -r'l ,-4 ,._ o m co ro I (1) © i cd © .el ...... I I Rhode St. Genese 35 airfoil sections modified by varying blade thickness ratio I 20.4 °. 14.5°_ _0.124 _- O. 294 • 05 .S • t ! LLLAI [ I 1_ 4 i I I L_kl_l I_II
_lti/ttltt_
,, llllt
I .I._I ,Itll .6 .2 II/IL lllli II ll.L h i ;t I/II1111111 t I II i b !_ lltll_llltil ! ] i±_] ill!
-t., P • 5 m .i D I I 1 llJ i i ::f 2:tltTllfllllltH-_tft .4 0 .2 .4 .6 .8 1.0 r R Figure 3.- Blade-form curves. In this figure, symbols are as follows: D, diameter (D = 28 in.); R, radius; r, station radius; b, section chord (b = 5.0 in.); h, section thickness; p, geometric pitch (p = 2_r tan _); _, section blade angle.
TQII irlcIdel_ce 3 it,deg Propellers off flopping o 15 flapping B_ off rfgld O 15 rigid A v,_ )- I000 r :.. : < , q.J E E Cn C - I000 J;3 IOO0 _6 LL c_ -I000
I
a r _--i,. L_ i - -= 3000 ..3 2000 I00O 0 -15 -I0 -5 0 5 I0 15 20 Angle ofattack, a, deg (a) iw = 800; V = 8.9 knots.
Figure 6.- Effects of flapping and rigid propellers on the longitudinal stability and control characteristics in transition flight.
Toll incidence, it,deg Propellers off flopping H 15 flapping off rtqld <> Z_ _d-- 15 rIi:Jld 30OO .... -- ----_r ¸ T T. . ] Z ...... V--' _ _.._ _[ >.- O_ _ r ..---[ .....
E I IO00 _n e- e-.- I I I , , , n- O I r--! .... I .L Lm I000 .c5 LI_
i • i
( > -I000 400O ,.0 3000 ._l LL -6- 2OO0 ....I 15 20 5" -I 0 -5 0 5 I0 Angle of attack, a, deg (b) iw = 60o; V = 29.8 knots, Figure 6.- Continued.
Toll incidence, i_lfe g Propellers flapping i-t 15 flapping off rigid 15 rigid 2OOO ..1:3 >.- I000 --( ¢..
E o E _n -I000 n- I000 .o _e, ,S C3 I -I000 4OOO J_ p.-,, 3OOO '4-- -I I000
-J5 -Io
-5 0 5 I0 15 20 Angle ofoltock, a, deg (c) iw = 400; V = 43.0 knots.
Figure 6.- Continued.
I000 J:} +t N,- >- -IOOO 4- r- E o -2OOO E \ _liP -i, t-
I 1I
OT_ -3000 _& J I000 LL o; o
I
- I000 5OOO 400O a_ __l LL
.,4r..-"
m __I I To_I _ncldence, i_deg Propellers J, off tlopp g
#/
15 flapp ] i000 off rigid <> 15 rigid
I I
-15 -I 0 -5 0 5 I0 15 20 Angle of attack,, a, deg (d) iw = 200; V = 68.5 knots.
Figure 6.- Concluded.
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