REPORT N o. 700
REPORT No. 700
By H. À. SovLÉ
The National Advisory Committee for Aeronautics 1s
undertaking an investigation of the flying qualities of air-
planes. The work consists in the determination of the
significant qualities susceptible of measurement, the
development of the instruments required to make the
measurements, and the accumulation of data on the flying
qualities of existing airplanes, which data are to serve
as a basis for quantitative specifications for the flying
qualities of future designs.
A tentative schedule of measurable flying qualities has
been prepared and the instruments needed for their
measurement have been assembled. Two sets of instru-
ments have been used: One set consists of special N. A.C. A.
recording instruments and the other set consists of gen-
erally available commercial instruments including those
usually found on the instrument panels of large airplanes.
A trial of the schedule and the instruments has been
made using the Stinson SR-8E airplane.
The results showed that, although the original schedule
and instruments are basically satisfactory, some further
development is required to eliminate nonessential items
and to expedite flight testing. The report describes and
discusses the work done with thas airplane.
The National Advisory Committee for Aeronautics
is conducting an investigation of the flying qualities
of airplanes. The work in connection with this investi-
gation is divided into three phases that are proceeding
simultaneously. The first phase consists in determining
factors susceptible of measurement that can be used to
define quantitatively the flying qualities of an air-
plane. The second phase consists in the development
of instruments and test procedure for making the
required measurements. The third phase consists in
the accumulation of data on the flying qualities of
existing airplanes to serve as a basis for the establish-
ment of quantitative specifications for the flying quali-
ties of future designs. These data will also be useful
as a basis for any analysis that may be made regarding
the particular points which are shown to be in need of
The first work of the investigation was done by E. P.
Warner in an advisory capacity relative to the prepa-
ration of the specifications for the DC—4 airplane. He
consulted a considerable number of air-line pilots, en-
gineers connected with both the operating and the
manufacturing companies, and research men, including
members of the N. A. C. A. staff, in order to obtain
the general ideas of the industry as to the flying quail-
ties and the possible tests that could be made to deter-
mine them. The results of this survey, which were
used In the preparation of the specifications for the
flying qualities of the DC-4 airplane, were presented
to the Committee with a request for its cooperation.
The second phase of the work consisted in the consid-
eration of the results of the survey by members of the
N. A. C. A. staff and the preparation of a tentative
schedule of flying qualities and of tests by which they
could be determined. In the preparation of this
schedule, primary consideration was given to the defi-
nition of the flying qualities in terms of factors known
to be susceptible of measurement by existing N. À. C. À.
instruments or by instruments that could readily be
designed and developed. As the N. A. C. A. instru-
ments are not available to all agencies that might be
interested In measuring flying qualities, consideration
was also given to the possibility of making the meas-
urements with standard aircraft indicating instruments.
As a third step, it was necessary to male tests to
demonstrate that the items listed in the tentative
schedule could be measured with sufficient precision
and that they defined the flying qualities as intended.
The flying qualities of the Stinson SR-8E airplane were
determined in the fall of 1937. In addition to the
actual demonstration of the tests and the test procedure
of the tentative schedule, the possibility of using the
indicating instruments for obtaining the information
required was investigated. The results of the trials
with this airplane are reported herein.
The Investigation has since been extended to other
airplanes. A number of airplanes of various types, of
course, will have to be examined before it will be pos-
sible to prepare satisfactory quantitative specifications
for the flying qualities of airplanes.
The items for which measurement is required in order
to determine the flying qualities of an airplane and
the N. A. C. A. instruments with which these items
can be measured are given in table I. The table also
includes a list of standard aircraft instruments and
other indicating instruments that may be employed in
place of the N. A. C. A. instruments.
Items to be measured N. A. C. A. recording Indicating instruments
| Air speed Air-speed meter... Air-speed meter.
Time Timer Stop watch.
Control-force meters... Control-force meters.
Control-position meters. |
Force to operate the three
control surfaces.
Position of the three control
Angular motion of the air-
plane about the three air-
plane axes.
Control-position meters.
Artificial horizon, indi-
cating pitch and roll.
Directional gyro or rate-
of-turn indicator, indi-
cating yaw.
Normal ation Accelerometer....—_—.—.—. Accelerometer.
Altitude.....———-———e===--—— Altimeter. cee timeter.
Longitudinal acceleration.__| Accelerometer....__...__ Accelerometer.
All of the basic group of instruments were available at
the start of the trials of the Stinson airplane except
Aiteron force
Arras ts sti vl
r, To recording
goges я
measuring unit
measuring unit
FIGURE 1.—Arrangement of actuating eylinders of wheel-force recorders.
those needed for the measurement of the forces required
to operate the different control surfaces; the existing
N. À. C. À. control-force recorders were not adaptable
to the particular airplane.
For the measurement of the rudder forces, a brake-
pedal-force indicator, developed by the Bendix Com-
pany for automobile use, was obtained. It operated
hydraulically, the force being applied to a cylinder
interposed between the pilot’s right foot and the rudder
pedal and the reading being given by an indicating
pressure gage fitted with a maximum hand. Because
of the limited time available for preparation and
because the tests were so planned that they required
measurement of the maximum force applied and not of
the variation of force with time, it was decided to use
this instrument as an indicator and not to convert it
to a recorder. In the installation of the instrument in
the Stinson airplane, the actuating cylinder was
mounted on the right rudder pedal so that all maneuvers
requiring the measurement of rudder force had to be
made to the right.
The instrument developed for measuring the elevator
and the aileron forces was constructed according to
the same general principles as the rudder-force indicator
and involves the use of four actuating cylinders to
measure right aileron forces, left aileron forces, elevator
push, and elevator pull. The arrangement of these
cylinders on the control wheel is shown schematically
in figure 1. The instrument was used in conjunction
with four recording pressure gages.
Because the altimeter is used for only one measure-
ment, it was decided not to install the recording
instrument but to use a sensitive Kollsman indicating
altimeter. Where the tests specified the initial con-
ditions for a maneuver, indicating instruments were
used for attaining these conditions.
Of the alternate set of instruments, the air-speed
meter, the artificial horizon, the directional gyro, and
the altimeter normally appear on the instrument panel
of a transport airplane. The artificial horizon with the
standard dial, however, has no graduations to indicate
the angle of pitch. In order to make this instrument
suitable for quantitative measurements, the dial must
be modified by the addition of graduations of pitch
angles in degrees. The indicating accelerometer and
the stop watch, although not standard airplane equip-
ment, are readily obtainable. A rudder-force indicator,
as already noted, is also available. A wheel-force indi-
cator may be developed along the same lines as the
recorder previously mentioned.
In the trial of the indicating instruments on the
Stinson airplane, difficulty was expected in connection
with the use of the artificial horizon, the directional
gyro, and the rate-of-turn indicator. The test program
required the measurement of small angular changes,
particularly in pitching and yawing, and the measure-
ment of these small values simultaneously at a particu-
lar time. A preliminary investigation showed that it
was practically impossible, where the motions are rapid,
to obtain these readings visually. For the actual tests,
therefore, these instruments were especially grouped
on the pilot’s instrument board along with the stop
watch and were photographed with a small motion-
picture camera. As no difficulty was anticipated in
the use of the wheel-force indicator or in its develop-
ment from the wheel-force recorder, this instrument
was not employed in the tests. For similar reasons,
a control-position indicator was not used but, in order
to demonstrate that one could be readily made, an
instrument was developed and installed in the airplane.
(a) Front.
FIGURE 2.—Three-quarter views of Stinson airplane.
The Stinson airplane used in the investigation 15 а
five-place, externally braced, high-wing, cabin mono-
plane. It is equipped with partial-span, pneumatically
operated, balanced flaps and has a single engine fitted
with a two-position controllable propeller. General
views are given in figure 2. The specifications of the
airplane, as given in the manufacturer’s handbook,
are included in the following list:
Name and type_-.------_--_- Stinson Model SR-8E 5PCLM
Engine. coo emcee Wright R-760-E2
Rating_-_-----------0000 0000000 320 horsepower at 2,150 rpm
Propeller_._-_---------- Hamilton controllable (two positions)
Fuel capacity oo ovo coe oom 82 gallons
Oil capacity -...----------=--=-=-===--- nm 5 gallons
Gross weight . ooo cco ee 3,800 pounds
Empty weight oo ee eee 2,417 pounds
Useful load....--.—--——-———e-—-—= eee 1,383 pounds
Allowable center-of-gravity limits... 22.1 percent to 33.1
percent of mean aero-
dynamic chord.
Power loading... --.------------ 11.88 pounds per horsepower
Wing loading... ----------—----- 14.70 pounds per square foot
Length over-all. _ meen 28 feet, Y inch
Wing span. cco ooo 41 feet, 10% inches
Height (tail down) ee 8 feet, 6 inches
Wing chord (tapered)- -------------—=---- 96 inches maximum
Dihedral._...--..----—-—-—--=-==-- mn aaa 000 = 2°
Incidence. o-oo eee eee eee eo CC - A m0 0 00 2°
Wing area (gross, including flaps and ailerons). 285.5 square feet
Aileron area (both). -...------------=------- 22.02 square feet
Flap area (both) cocoa 23.82 square feet
Stabilizer area co eee 25.06 square feet
Elevator area (including balance) _______ 19.77 square feet
Tin ares ann 10.8 square feet
Rudder area (including balance). _.._ 14.38 square feet
The measured angular deflection of the control surfaces
Flap deflection . — ooo oo eee 38°
Aileron deflection. oo ooo mms 36° up, 23° down
Rudder deflection. coo nen + 22°
3-27° (stabilizer tail heavy)
0° to nose down 9°
Elevator deflection ooo -
Stabilizer deflection.._----—--—-.. ----—-—-—----=
The relations between control-wheel position and aileron
and elevator position are given in figures 3 and 4.
As flown in the tests, the airplane weighed 3,655 pounds
and the center of gravity was located at 25.4 percent of
the mean aerodynamic chord and 7.2 inches above the
thrust axis of the airplane.
Left aileron--t- ight aileron
Aileron deflection
Degrees down
180 120 60 180
Degrees left
O .. 60 120
Whee/ position
Degrees right
FIGURE 3.—Relation between aileron and control-wheel positions.
The tentative schedule of flying qualities and tests
was prepared to include multiengine airplanes equipped
with trimming tabs on all three sets of control surfaces.
Complete trial was, of course, not possible with the
Stinson airplane, which has only one engine and was not
fitted with aileron and rudder trimming tabs. For
reference, however, the complete tentative schedule
has been included in the following presentation of the
trials made with the Stinson. The requirements listed
are believed to be essential to a completely satisfactory
Because of the preliminary nature of the trials, vari-
ations of propeller pitch and of airplane weight and
center-of-gravity position were not tried. All tests
were made with the propeller in the high-pitch setting.
The welght and the center-of-gravity location for the
airplane as flown have already been noted. Two engine
conditions were investigated: For the first condition,
Trailing edge up
So, So
Elevator angle relative to thrust axis, deg
Trailing edge down
2 14 16
6 8 0 2
Confro/-wheel position, in. from dosh
FIGURE 4.—Relation between elevator and control-wheel positions.
the throttle was closed and, for the second, the throttle
was set at the position that gave level flight at cruising
speed with the flap up. These two throttle settings
were used in both the flap-up and the flap-down tests.
Requirement.—With the elevator free, the airplane
shall be dynamically stable in pitch at all speeds
throughout the speed range. The period of the longi-
tudinal oscillation shall never be less than 40 seconds
and the damping shall be sufficient to reduce the
amplitude of the oscillation to one-fifth the original
amplitude in four cycles.
(Although the numerical limits have been suggested
in this and the following requirements, they are, for
the most part, considered to be quantitatively unre-
liable owing to the present lack of data concerning what
constitutes satisfactory flying qualities. It is appreci-
ated that, in the final forms for the requirements,
airplanes may have to be classified as to purpose and
have different numerical values assigned for each
Procedure.— Trim the airplane for the desired speed
and then push the stick forward by an amount sufficient
М —— Se Te
to increase the steady speed by 5 to 10 miles per hour.
Release the stick and record the variation of air speed
with time during the ensuing oscillations as the airplane
returns to a steady state at the original trimming speed.
Results and discussion.—As the procedure for the
measurement of the period and the damping of the
longitudinal oscillation had been employed before, no
development was required for the present investigation.
With regard to measurements of the damping, however,
it may be necessary for precision to vary the speed by
more than the 5 or 10 miles per hour suggested in the
procedure. The amount depends on the speed range
of the airplane being tested and the sensitivity of the
recording or the indicating instrument used.
The method of making the measurements by photo-
graphing the indicating instruments did not work out
as well as expected but 1t may be further improved.
Some difficulty experienced in obtaining sufficient light
for photography was overcome by using the fastest
commercial film and a relatively fast (f:2) lens. Even
with this arrangement, however, no clear pictures were
obtained of the stop watch. The stop watch used was
unsatisfactory because of the slenderness of the hand,
the fineness of the dial markings, and because the face
was white with black markings rather than black with
white markings. A more suitable watch could probably
have been obtained but not within the time available
for the tests. Lighting conditions are always likely to
be critical where natural illumination is depended upon.
For this reason, an additional set of instruments,
especially mounted and artificially illuminated for
photography, seems to be desirable if the airplane size
The results of the longitudinal-stability measure-
ments are given in figure 5. Figure 5 (a) presents the
period of oscillation as a function of speed for the four
test conditions. From the figure, the actual periods
are noted to be less than the suggested minimum value
for all speeds up to 140 miles per hour with either power
off or power on. No difference occurs between the
flap-up and the flap-down periods for equal speeds. In
connection with the disagreement between the actual
and the suggested periods, a comparison has been made
between the data for the Stinson airplane and the data
given In reference 1 for several other airplanes. In
this reference, it was observed that the designer has
little control over the period of oscillation of a con-
ventional airplane. An empirical formula for the
period for the power-off condition, P=0.262V, where
P is the period in seconds and Y is the velocity in
miles per hour, was given. The straight line given by
this equation has been plotted in figure 5 (a). The
agreement between it and the curves for the Stinson
airplane indicates that the specification of a constant
period is illogical.
The damping characteristics of the longitudinal os-
cillations are given in figures 5 (b) and 5 (c). Figure
5 (b) presents the measured damping coefficient; and
figure 5 (c), the computed number of cycles required
for an oscillation to damp to one-fifth amplitude. In
figure 5 (c), it should be noted that the actual plot is’
made of the inverse of the number of oscillations. For
most of the speed range for all conditions, the damping
is less than the suggested value. For three conditions,
the airplane is unstable for portions of the speed range.
The importance of the damping specification is ques-
tionable in the light of present knowledge. The Stinson
tests and also those reported in reference 1 show that
an airplane which, from the pilot’s viewpoint, has satis-
factory flying qualities may still be dynamically un-
stable. It 1s not at all certain that airplanes can be
made longitudinally stable under all conditions of flight
without adversely affecting some other qualities. Until
this point is settled, however, the requirement seems
desirable that the airplane be stable in the range
wherein the airplane may be flown with the elevator
free. The amount of damping is considered of no
importance at the present time.
Requirement.—The range of the elevator control
shall be sufficient to meet the following conditions:
a. With every setting of the trimming device, it shall
be possible to maintain steady flight at any speed from
the design probable diving speed to the minimum speed
for any power condition, flap up.
b. With every setting of the trimming device, it shall
be possible to maintain steady flight at any speed from
the placarded to the minimum, flap down. *
c. With the conventional type of landing gear, it
shall be possible to make three-point landings and to
hold the tail down while braking enough to give a de-
celeration of 0.39 during the landing run down to a
speed of 30 miles per hour.
d. In the take-off run, it shall be possible to raise the
tail off the ground by the time a speed of 30 miles per
hour is attained.
e. If a tricycle type of landing gear is used, it shall
be possible to raise the nose wheel off the ground in a
take-off run by the time a speed of 30 miles per hour is
attained. (As has been noted, this and other require-
ments not relating to the Stinson airplane are included
for reference.)
Procedure for items a and b.—Measure the elevator
angles at different speeds with different tab or stabilizer
settings and different throttle positions.
Procedure for item c.—Merely demonstrate the abil-
ity to make three-point landings. For the braking
tests, run the airplane along the ground at a speed of
approximately 50 miles per hour. Close the throttle
and apply brakes to the maximum extent for which the
pilot can maintain contact between the tail wheel and
the ground. Record the air speed and the longitudinal
acceleration as the airplane decelerates to less than 30
miles’ per hour.
Procedure for item d.—Apply full throttle while hold-
ing the airplane with the brakes. Release brakes and
attempt to raise the tail as soon as possible. Record
speed at which the tail leaves the ground.
Results and discussion—No difficulty was encoun-
tered in carrying out the procedure and obtaining the
information desired.
The results of the measurements of the elevator posi-
tion for steady flight for the Stinson airplane for the
o flop up, power of
+ "e " - O
X =» down power off
О ” ч " on
Period, sec
70 80 100. 10
Air speed, mph
(a) Period of oscillation.
/20 130
© o————— Flap up, power off
5 Dem sn n и on
E x—— * down power off
© C— - — " " “ on
70 80 90 100 ¡10
Air speed, mph
20 130 140
(b) Damping coefficient.
OX Flop up, power off
=D ————— re de —— n a“ “ or?
“3 2 — —— " down, power off
OT — я “ On
00 71
Ss X
38. = ©)
со 70 8 90 100 no 120 130 140
Air speed, mph
(e) Number of cycles to damp to one-fifth amplitude.
FIGURE 5.—Longitudinal-stability characteristics.
four test conditions are presented in figure 6. Al-
though complete measurements were not made for all
stabilizer settings 8, because of high control forces,
figure 6 indicates that the range of elevator control
should be ample for any desired speed from the mini-
mum up to the high speed in level flight with the flap
up. Because of the preliminary nature of the tests, it
was considered unwarranted to carry the measurements
to the design probable diving speed. With the flap
down, the airplane could be flown steadily at any speed
from the minimum to the placarded of 125 miles per
The data from the landing tests showed that the air-
plane could be landed smoothly in a three-point attitude
and that the elevator was just sufficient to hold the tail
wheel in contact with the ground while braking vigor-
ously enough to give a deceleration of 0.39 at 30 miles
per hour. The elevator was sufficiently powerful to
lift the tail wheel from the ground in a take-off run at
Troiling edge up
Elevotor angle relative to thrust axis, deg
Trailing edge down Troiling edge up Trailing edge down
60 80 100 120 140
(a) Flap up, power off.
(c) Flap down, power off.
Air speed, mph
amount the elevator deflected relative to the control
column per unit applied force was determined by
static tests. The constant thus obtained and the
recorded elevator-control forces were used to compute
the errors in the individual control-position readings.
The actual elevator angles as a function of speed are
shown in figure 6. Because the pilot's opinions of
ô, , deg
Nose heavy, O
=—— o q чай аи EEE
-4.8 ——-——-
7? —— ——
Tail heavy, -96 ——--—
- (d)
100 120 140 160
60 80
(b) Flap up, power on.
(d) Flap down, power on.
FIGURE 6.—Elevator angles corrected for cable stretch.
a speed so low that precise measurements could not be
Requirement—The curve of equilibrium elevator
angle against speed for every setting of the trimming
tab shall be smooth and shall everywhere within the
permissible speed range have a negative slope.
Procedure.—The information needed for this re-
quirement is obtained from the measurements of ele-
vator angle previously discussed under the section
Range of Elevator Control.
Results and discussion.—The control-position re-
corder used for the measurement of elevator angles
was, for convenience, connected to the control column
in the cockpit rather than to the elevator. The
results had to be corrected for the stretching of the
elevator cables. In order to make the correction, the
stability are influenced by control-column movement,
which differs from the elevator movement owing to
the cable stretch, figure 7 has also been included pre-
senting the variation of control-wheel position, or
apparent elevator angle, with speed.
The curves of figure 6 indicate that, for most of the
test conditions and stabilizer settings, the slopes are
negative as required. With flaps up and with both
power on and power off for certain stabilizer settings,
a small range of speeds apparently exists in which the
slopes of the curves are slightly positive. “The reason
for the change of sign of the slopes of the curves for
the particular conditions is unknown. The measure-
ments and the corrections were possibly in error by an
amount sufficient to produce the positive slopes.
Figure 7 indicates that, because of the extension of
the control cables, the airplane with the tail-heavy
stabilizer settings gave the pilot an impression of
greater stability than did the nose-heavy stabilizer
settings, although figure 6 shows very little change
in the actual elevator curves with stabilizer setting.
With the nose-heavy stabilizer setting for the two
flap-up conditions, the airplane appeared unstable to
the pilot. The comparison between the two sets of
curves indicates the need for exercising considerable
care In obtaining and interpreting data relating to
static longitudinal stability. It is of interest to note
Apparent elevator ongle relative to thrust oxis, deg
Nose heavy
5 = 0° 1
Trailing edge up
Control-wheel position,in from dash
© ©
Trailing edge down
80 100 120
Air speed, mph
1/40 160
(a) Flap up, power off.
t elevator angle relotive fo thrust oxis, deg
Troiling edge down
Trailing edge
Control-wheel position, in. From dosh
Toil heavy
60 100 120
Air speed, mp A
140 /60
(0) Flap down, power off.
that neither half of the elevator travel nor half of the
wheel movement was required during the tests.
Requirement.—With every setting of the trimming
tabs or the stabilizer, it shall be possible to fly the
airplane from the minimum to the maximum permis-
sible speed with a change of elevator-control force no
greater than 100 pounds.
Apparent elevator angle relative to thrust axis, deg
Trailing edge up
Control wheel position, in from dosh
Trailing edge down
Tail heavy
60 80
700 120
Air speed, mph
140 760
(b) Flap up, power on.
e 70 thrust axis, deg
Trailing edge up
Control-whee/ position infrom dash
% 0
Ss Nose heavy
CO =0°
UL 6
O 5
60 80
100 120
Air speed, mp h
(d) Flap down, power on.
FIGURE 7.—~Apparent elevator angles.
Procedure.—Measure the elevator-control force at
difrerent speeds with different tab or stabilizer settings
and different throttle positions. (These measurements
may be made simultaneously with the measurements
of elevator angle.)
Results and discussion.—The control-force recorder
Elevator force, /b
50 100 120
Air speed, mp A
140 760
(a) Flap up, power off.
Nose heavy
Ё —
Elevator force, Ib
60 80 700
Air speed, mph
(c) Flap down, power off.
ment would give erroneous readings. The trouble was
eliminated by arranging for the operation of the in-
struments by the observer, thus permitting the pilot
the use of both hands to apply symmetrical loads.
The results of the measurements of the elevator forces
are given in figure 8, which shows that, for the Stinson
Nose heavy
Elevator force, Ib
Toil heavy
“ ,-9.6°
60 50 1/00 120
Air speed, mp h
(b) Flap up, power on.
Nose heavy
Elevotor force, Ib
60 80 {00
Air speed, mph
(d) Flap down, power on.
FIGURE 8.—Elevator forces.
used the first time for these measurements behaved
satisfactorily. A slight difficulty was experienced at
the start of the investigation because, owing to the
arrangement of the electrical system, 1b was necessary
for the pilot to apply the control force with one hand in
order to have the other free to operate the Instrument
switch. With this unsymmetrical application of the
load, the actuating pistons would bind and the instru-
airplane, the range of the elevator-control forces
depends on the stabilizer setting. For the nose-heavy
settings, the force variation was small.
As the airplane was trimmed toward the tail-heavy
condition, the range of the elevator forces rapidly in-
creased. This increased variation of force with tail-
heavy stabilizer settings is accounted for, in part, by a
spring system interconnected with the elevator and
trimming control so that, when the stabilizer is set tail
heavy, a spring comes into action and increases the
force required to pull back on the control column.
With the flap down, the suggested range of force was
met for all stabilizer settings. With the flap up,
although the tests did not cover the complete flight
range, it appears that the force variation of less than
100 pounds would not be met with the tail-heavy
stabilizer settings, probably because of the spring
just discussed.
Requirement.—The curves of stick force required
for steady flight plotted agamst the speed of flight for
all tab or stabilizer settings shall be smooth without dis-
continuities or sudden changes of curvature. The
slopes shall be everywhere negative throughout the
specified speed range and shall nowhere be less than
one-fourth pound per mile per hour.
Procedure.—The data relating to this requirement
were obtained from the measurements of elevator-
control force made in the section Range of Elevator-
Control Force.
Results and discussion.—Figure 8 shows that the
same nose-heavy stabilizer position that gave an un-
desirable variation of wheel position (figs. 7 (a) and (b))
also gives an undesirable variation of elevator force.
It will be noted, however, that with this stabilizer
setting no balance speed exists. The control column
must be pulled back throughout the complete speed
range, On the basis that the stabilizer range should
be insufficient for balance at speeds above the maximum
level-flight speed, the results for this stabilizer setting
may be neglected in the present discussion. It might
be desirable to limit the stabilizer to a setting that would
permit trimming at the maximum speed with flap up
and power on. In order to make such a decision, how-
ever, tests would have to be made with the most tail-
heavy positions of center of gravity. With the stab-
ilizer settings that gave balance between the minimum
and the maximum speeds, the variation of the elevator-
control force with speed was satisfactory.
Requirement.—The force required on the stick to
overcome, without change of tab or stabilizer setting
or speed, the effect of any change in the engine operat-
ing condition from full power to fully throttled shall
not exceed 100 pounds.
Procedure.—The data relating to this requirement
are obtained from the measurements of elevator-
control force made in the section Range of Elevator-
Control Force.
Results and discussion.—The effect of opening the
throttle is to make the airplane trim more tail heavy.
The change of force required to maintain a given speed
when the throttle is varied from the closed to the
power-on condition varies from 0 to 40 pounds, de-
pending on the speed and on the settings of the flap
and the stabilizer.
a. It shall be possible to trim the airplane at a low
enough speed so that no greater force than the 30-
pound pull would be required in performing a three-
point landing.
b. It shall be possible to trim the airplane at its
maximum level-flight speed.
Procedure for item a.—Measure the maximum ele-
vator force in a landing with the stabilizer or the trim-
ming tab set full tail heavy.
Procedure for item b.—The data relating to this re-
quirement were obtained from the measurement of
elevator-control force made in the section Range of
Elevator-Control Force.
Results and discussion.—The range of the stabilizer
is greater than required with the airplane as loaded for
the tests. No force measurements were made in the
actual landings but, with flap down and power off, the
landing condition, the airplane could be stalled with a
pull of the order of 25 pounds with the stabilizer in
the tail-heavy position. As previously noted, the air-
plane may be trimmed with the stabilizer at speeds
above the maximum.
a. As an indication of the effectiveness of the eleva-
tor for maneuvering the airplane, it shall be possible
to obtain an acceleration of 0.8 of the design applied
load factor at any speed with the elevator alone and
with the airplane originally trimmed for cruising speed,
when a force of not more than 200 pounds and not less
than 60 pounds is applied to the control wheel.
b. At low speeds down to 10 miles per hour above
the minimum where the theoretical maximum accelera-
tion approaches 1, 1t shall be possible to change the
attitude of the airplane in space with respect to its
transverse axis in either direction by 5° in 1% seconds
by use of the elevator alone.
Procedure for item a.—Trim the airplane for cruising
speed. Increase speed to the design probable diving
speed. Make pull-up to 0.8 of the applied load factor,
using an indicating accelerometer for reference. Record
acceleration, speed, and stick force. Repeat at various
speeds throughout the speed range.
Procedure for item b.—Trim the airplane for a speed
20 miles above the minimum. Apply full-up elevator
and record angular velocity and air speed. Repeat,
applying full-down elevator. Repeat at various speeds
to the minimum.
Results and discussion.—The tests to determine the
effectiveness of the elevator for maneuvering were
made in the manner outlined with no difficulty as far
as the procedure and the operation of the recording
instruments were concerned. No measurements were
Lood factor
> 1
Elevator à
100 120 140 160 180
Air speed, mph
made at speeds greater than the maximum for level
flight. In the trial of the indicating Instruments, the
camera had to be rigidly mounted. This rigid mount-
ing may occasion a little difficulty in small airplanes,
although it was readily accomplished with the Stinson.
An inspection of the photographs made showed that the
indicating accelerometer and the artificial horizon with
the modified dial were suitable for the measurements.
The photographs were not evaluated because of the
previously mentioned trouble with the photographs
of the stop watch.
The results of the measurements are given in figures
9 and 10 and in the following table. Figure 9 shows
the forces required to obtain an acceleration of approxi-
mately 0.8 of the applied load factor at various speeds
when the airplane is trimmed for cruising speed. The
results show that, within the range of the tests, the
forces fell well below the upper limit of 200 pounds
and they all were, in fact, below the minimum limit.
The requirements need some revision. The force
probably should be defined as a function of speed and
the values of 200 and 60 pounds made to apply at some
definite speed.
FIGURE 10.—Variation of attitude with time following abrupt
elevator deflections at low speeds. Flap up. power off.
Deviation from initial attitude, deg
A ©
Obtained in flight
+ Power on
factors >
FIGURE 9.-—Elevator forces required tu obtain normal ac-
celerations at various speeds.
” orf
200 ECO
Figure 10 shows the manner in which the attitude
of the airplane can be changed in pitch at low speed
with the elevator alone. The figure has been prepared
for only the flap-up power-off condition. The lengths
of time to rotate the airplane through angles of 5° and
10° in pitch given by figure 10 are listed in table IL.
Similar data are also given for the other three test
5° 10°
РР Speed
- Test condition (mph) Pull- Push- Push
UP | down Pull-up down
(sec) (sec) (sec) (sec)
Flap up, power off. _______. 71 0. 51 0. 44 0.81 0.72
77 .46 .46 . .71
81 .39 .46 59 ‚ 18
86 .38 .46 83 | -.----
Flap up, poweron.________ 66 . 56 ‚46 | ea... ‚ 19
70 „49 .48 69 .71
77 43 .39 58 57
40 .39 72 54
86 .45 ‚ 45 ‚ 67 ‚ 63
Flap down, power off... 60 ‚ 50 ‚ 51 ‚ 82 ‚ 75
65 ‚51 „47 79 71
70 ‚ 48 „50 ‚ 66 ‚72
Flap down, poweron._...__ 60 ‚ 46 ‚ 49 ‚ 66 ‚ 78
65 42 ‚ 54 62 ‚ 84
70 43 ‚ 38 60 58
B-——— 71 mph
O-———77 *
x——- 61
‚2 3 . ‚6 ‚7
Time, sec
The values given in table IL do not represent ultimate
conditions. The pilot made the maneuvers only as
abrupt as he thought necessary to meet the requirement.
Even then, the values of time obtained for a 5% change
of attitude were in the order of one-third of the 1%
seconds specified.
a. The airplane shall be laterally stable for the same
conditions for which longitudinal stability 1s required.
The period of the lateral oscillations shall not be less
than 20 seconds and the damping shall be sufficient to
reduce the amplitude of the oscillation to one-half
the original amplitude in two cycles.
b. Between the minimum trimming spéed and the
minimum speed, the airplane shall show no negative
dihedral effect nor any autorotative tendencies.
ce. From the minimum speed to the limit of the
elevator control, there shall be no sudden development
of marked autorotative tendencies nor any sudden
change in lateral stability characteristics.
Procedure for item a.—Trim the airplane about all
three axes for flight at the desired speed. Start a
lateral oscillation by rolling the airplane with the
ailerons, by yawing the airplane with the rudder, or by
placing the airplane in a sideslip through the combined
use of the rudder and ailerons. Free the controls and
record the ensuing angular motion in yawing and pitch-
ing and the air speed. Because of unsymmetrical
rigging and power effects, these tests should be made to
both sides.
Procedure for items b and c.—Obtain information by
direct observation of low-speed and stalled-flight
characteristics. A sample copy of the data sheet used
for the Stinson airplane is given in table IIT.
Flap up, power off:
1. Can normal turns with 15° bank be made within 5 miles
per hour of stall?
2. As speed is decreased to stall, will airplane bank normally
with rudder movements?
3. Do ailerons lose effectiveness progressively?
4. Is there any tendency for airplane to fall off and spin
before minimum speed is reached?
5. Is stall progressive or sudden and violent?
6. Beyond the stall, what is relative effectiveness of rudder
and ailerons?
7. With stick hard back, can airplane be flown reasonably
steadily for 10 seconds?
8. Is tail buffeting violent?
Results and discussion.—It will be noted in require-
ment a that no differentiation is made between the
different types of control-induced disturbances. This
procedure is in accordance with the theory which pre-
dicts that the ultimate motion is independent of the
type of disturbance. Because of the usual procedure
in the past of differentiating between the lateral and
the spiral motion, some uncertainty was felt regarding
this conclusion from the theory. With the Stinson,
therefore, the three types of disturbances given under
the procedure were tested. The requirement also pre-
supposed that the airplane could be laterally balanced
at any speed through the use of aileron and rudder
tabs. As the Stinson airplane was not fitted with
these additional controls, three series of aileron maneu-
vers were made. In the first series, all controls were
freed after the airplane attained a bank angle of 15°;
in the second series, the ailerons were returned to neu-
tral and held there during the ensuing motion; and, in
the third series, the ailerons were freed but the rudder
was held neutral.
The test procedure given for item a was satisfac-
torily employed for the tests. Visual observations had
to be made of the spiral motion. The recorded yawing
velocity was used as the basic parameter for the com-
putations of the characteristics of the lateral oscilla-
tions. With regard to the indicating instruments, the
movement of the indicator on the directional gyro was
found to be too small for the angular displacements
involved in the oscillation to permit obtaining quanti-
tative data. It is believed that, if the dial of the rate-of-
turn indicator were graduated for angular velocity, this
instrument could be used for the measurements. Com-
putations of the period of oscillation made from the
data for this instrument gave fair agreement with the
results obtained from the recording instruments.
The tests showed that the Stinson airplane was
laterally unstable with the flap up below speeds of 140
miles per hour. With the flap down, it was unstable
at the highest test speed, 110 miles per hour. This
instability took the form of a spiral divergence, as was
determined from observations and not from the instru-
ment records. In connection with the spiral charac-
teristics of the airplane, it was noted that the ultimate
motion of the airplane depended upon the amount of
bank obtained before the controls were freed. For
both aileron rolls and sideslips, the initial angle of bank
was greater than 10°. For these maneuvers, the con-
sequent spiral tightened up somewhat when the controls
were freed before an ultimate steady condition was
attained. In the abrupt rudder kicks, the first motion
was primarily yawing. Only a small amount of roll
occurred in the first second. If the rudder was freed
before the airplane had time to bank appreciably, the
spiral motion was imperceptible so that, when the
lateral oscillation damped out, the airplane remained
in sensibly straight flight. When the rudder deflection
was held long enough for the airplane to attain an
angle of bank equivalent to that obtained in the aileron
maneuvers, the airplane behaved approximately the
same as in the aileron tests. The results of the measure-
ments of the lateral oscillation are given in figure 11.
Figure 11 (a) shows the period of oscillation; figure 11
(b), the damping coefficient; and figure 11 (c), the
number of cycles to damp to one-half amplitude. It
Aleron moneuver ~
Rudder a -
Sideslip " /
o4—— Flop up power on and o
Vn down, power off
o—--—— * 7 ” on
Period, sec
{20 1130 1
80 90 100 110
Air speed, m ph
(a) Period of oscillation.
Air speed,m ph
70 80 90 /00 110 120 130 140
ileron maneuver x
+ Rudder a —
E Sideslip 7
0 -~4
Q 4A——— Flop up, power off
Q > + E e o 7
Q down, power off
(b) [= ea a Ir ” on
(b) Damping coefficient,
Flap up, power off
Q o po “ " « on
E « down, power of.
o ~ ” “ on
> QQ
70 80 90 100 140 /20 130 140
Air speed, mph
(e) Number of cycles to damp to one-half amplitude.
FiGURE 11.—Lateral-stability characteristics.
will be noted from these figures that the characteristics
of the lateral oscillation were independent of the type
of maneuver used to set up the disturbance. In all
cases and at all speeds, the period is below the suggested
minimum of 20 seconds. The damping, except for a
15-mile-per-hour range with flap down and power off,
is greater than suggested.
From the results of this investigation, the conclusion
is drawn that the complete information on lateral-
stability characteristics can be obtained from any of
the three maneuvers so that in the future the perform-
ance of the other two will be unnecessary. It is
suggested that the aileron maneuvers, which are the
least violent of the three, be used for the purpose.
No difference was noted in the lateral-stability
characteristics when the ailerons were held neutral
instead of being free. Presumably the control friction
was so small that the ailerons returned to neutral of
their own accord. With the rudder held neutral, the
increase of the effective fin area was indicated by slight
changes in the speed at which the airplane became
stable and in the period of the lateral oscillation.
The observations made for items b and c showed that
the airplane had no negative dihedral effect at any
point in the flying range to the limit of the elevator
travel. Autorotative tendencies were noted at and
beyond the stall in straight flight. Stall could be in-
duced 3 to 4 miles per hour above the minimum speed
by abrupt full use of the ailerons. The stall was pro-
gressive and not particularly violent.
O 6
SS -
cx O
+ O
0-Q с © 0
особе Level flight flop up ©
OS SS ” ” " down x
Se a 80 90 100 Mo 120
Air speed, mph
FIGURE 12.—Dihedral effect; recovery from a 15° bank through the action of roll due
to sideslip.
a. The dihedral effect shall be sufficient that, when
the ailerons are freed immediately after putting the
airplane in a 15° bank and using the rudder to avoid a
change of heading, the angle of bank shall be reduced to
2° within 15 seconds and with a loss of altitude of not
over 300 feet.
b. The rolling acceleration accompanying abrupt
rudder displacement shall be less than one-half the
-| yawing acceleration.
Procedure for item a.—Place the airplane in a steady
sideslip with one wing down 15°. Release the ailerons
and record the rolling motion of the airplane and the
altitude lost in recovery to straight flight.
Procedure for item b.—In the steady-flight condition,
apply full rudder abruptly and record the resultant
yawing and rolling motion.
Results and discussion —The tests regarding the
minimum dihedral effect, item a, were made with no
particular difficulty except that, at low speeds, the
rudder was insufficient to prevent the airplane from
yawing in a 15% bank. Inspection of the photographs
of the indicating instruments showed that the artificial
horizon could probably be used for the measurements,
but no evaluation was made because of the poor photo-
graphs of the stop watch. The data obtained from
the recording instruments are presented in figure 12.
The figure shows that the airplane has an appreciably
greater dihedral eftect than required, the maximum
time recorded for the recovery being only 5% seconds
as against the 15 seconds specified. The loss of altitude
during the recovery for the power-on condition was
negligible. For the power-off condition, no apparent
increase occurred in the vertical velocity during recovery.
The tests to determine the maximum dihedral effect,
item b, were unsatisfactory because of the violence of
the maneuver. The measurements were made in con-
nection with the rudder tests and the results will be
discussed later in the section dealing with the rudder
a. At a speed of 70 miles per hour with the flaps
down and 80 miles per hour with the flaps up, 1t shall
be possible to bank the airplane 15° in 2% seconds with
the ailerons alone and, at 120 miles per hour or higher,
the same angle of bank shall be obtained in 2 seconds.
b. At a speed 2 miles per hour above the stall with
the flaps down, it shall be possible to bank the airplane
10° in 2 seconds with the ailerons alone.
c. The aileron effectiveness shall be proportional to
the aileron deflection.
Procedure.—At the specified speeds, apply full
aileron control and record rolling velocity and aileron
position. Repeat with %, %, %, and % aileron deflection.
Results and discussion.—The procedure and the re-
cording instruments had been used in previous aileron
tests and, as expected, were satisfactory for the purposes
of the present investigation. The artificial horizon in
conjunction with a suitable stop watch appeared to be a
satisfactory indicating instrument for obtaining the
desired information.
The results of the measurements are given in figures
13 to 15. Figure 13 gives the time for the airplane to
roll 10° at two speeds with the flap up and at two speeds
with the flap down. No reason being apparent for
changing angles between items a and b of the require-
ments, figure 13 was prepared on the basis of a single
arbitrarily chosen angle. The speeds for the tests were
not in strict accordance with those of the requirements.
Figure 13 shows that Stinson ailerons will meet the
conditions suggested in item a of the requirements.
Tests were not made at a speed of 2 miles per hour
above the stall with the flaps down because observa-
tions had shown that the airplane could be stalled at
that speed by abrupt use of the ailerons.
Figures 14 and 15 present the data on aileron power
in terms of the maximum rolling acceleration and the
maximum rolling velocity. These figures show that,
although the aileron effectiveness is not in direct pro-
portion to the deflection, the variation with deflection
is smooth and progressive.
a. The force required to attain the aileron reactions
listed under the section Aileron Power shall not exceed
50 pounds applied tangentially at the rim of the wheel.
3 Air speed
Lo o 65 miles per hour
© $ 2 A 122 " г „
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(a) SS
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0 20 40 60 80 100 So
Percent of full aileron movement =
(a) Flap up, power off.
= Air speed
Lo miles per hour
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29 Со
(b) OS
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20 60 80 /00 ©
Fercent of Full aileron movement
(b) Flap down, power off,
FIGURE 13.—Time to roll 10° with ailerons.
b. The aileron force shall be approximately propor-
tional to the aileron deflection.
Procedure.—The information relating to these re-
quirements may be obtained by supplementing the
procedure listed in the section Aileron Power with
measurements of aileron forces.
Results and discussion.—The data have been plotted
in figures 13 to 15. For the highest test speed, the
figures show that 40 pounds were required for full
deflection of the ailerons. At all speeds, the aileron
forces varied somewhat in proportion to the aileron
deflection. Although the variation was not strictly
linear, the curve was smooth and showed no abrupt
changes in slope.
Requirement.—At speeds beyond 10 percent more
than the minimum speed, the ailerons shall not produce
a yawing acceleration greater than one-tenth the accel-
eration in roll. At speeds below 10 percent more than
the minimum, the acceleration in yaw shall be less than
one-fifth the acceleration in roll.
Procedure.—The information relating to this require-
ment may be obtained by supplementing the procedure
5 8
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0 0
0 20 40 60 80 100
Percent of full wheel movement
(a) Air speed, 65 miles per hour.
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Percent of Full wheel movement
(b) Air speed, 122 miles per hour.
FIGURE 14.—Aileron characteristics. Flap up, power off.
listed in the section Aileron Power with measurements
of angular motion of yaw.
Results and discussion.—The information obtained
has been plotted in figures 13 to 15. For no condition
was the yawing acceleration caused by the aileron as
great as one-tenth of the rolling acceleration. Simi-
larly, the airplane did not yaw 1° in the time taken to
obtain a bank of 10°.
a. It shall be possible by the use of the aileron tabs to
balance the airplane against the dissymmetry in loading
mum rolling velocity, radian pers
Rolling velocity--
q and yawing acceleration,
radians per second per second
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Percent of full wheel mo ment
(a) Air speed, 65 miles per hour.
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(b) Vowing lerat =
20 40 60 80
Percent of full wheel movement
(b) Air speed, 80 miles per hour.
FIGURE 15.—Aileron characteristics. Flap down, power off.
corresponding to full gas tanks on one side of the center
of gravity and empty gas tanks on the other.
b. It shall be possible by the use of the aileron tabs to
compensate for any rolling tendency accompanying
steady flight with asymmetrical power conditions.
a. It shall be possible during steady flight at 70
miles per hour with the flaps down and 80 miles per hour
with the flaps up to produce a change of heading of 15°
in 3 seconds by the use of the rudder alone. At and
beyond 120 miles per hour, the same change shall be
possible in 2 seconds.
b. At 2 miles per hour above the stall with the flaps
down, it shall be possible to make flat turns up to a
change of heading of 10° in 2 seconds.
c. At 20 miles per hour above the minimum speed, as
well as at any higher speed, it shall be possible to hold
a straight course with the wings laterally level with
both engines on either side cut out and with those on
the other side operating at full rated power.
d. With any three engines operating or with one out-
board and the opposite inboard one cut out, it shall be
> speed
28 х————^ >
9 120 mp h OD
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VQ ©
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0 100
20 40 60
Percent of full rudder movement
FIGURE 16.—Time to yaw 10° with rudder. Flap up, power off.
possible to hold a straight course on the ground down to
50 miles per hour with the flaps either up or down.
(Items c and d are not applicable to the Stinson.)
e. The rudder effectiveness shall be proportional to
the rudder deflection.
Procedure for items a and b.—Trim the airplane
laterally and longitudinally at the desired speed. Apply
abrupt full rudder control. Record rudder position and
rolling and yawing velocities. “Repeat with %, %, %, and
% rudder deflection. This procedure is similar to that
for item b for Dihedral Effect.
Results and discussion.—Insofar as the recording in-
struments were concerned, these tests were made satis-
factorily. 'The actual maneuvers, however, were very
violent and the requirements should preferably be re-
written to eliminate the need for fully deflecting the
rudder. Because of the violence of the maneuvers, the
tests were not made with the flaps down. The results
of the tests for the flap-up condition are given in figures
16 and 17. The manner of presentation is similar to
that for the aileron data.
The data of the figures, although not conclusive be-
cause they do not include the flap-down condition, in-
dicate that the rudder is considerably more powerful
than necessary to meet the requirements as written.
The specifications, in addition to requiring more violent
maneuvers than are considered necessary, are believed
not to define the maximum rudder power that may be
desired. It is therefore suggested that the primary
requirement for rudder power be based on the maximum
angle of bank which can be held in the steady sideslip.
The present requirement may be treated in a subsidiary
manner similar to the requirement regarding the ability
to change altitude with the elevator.
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De Yawing accelerat
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0 20 40 60 80 100
Percent of full rudder movement
(a) Air speed, 70 miles per hour.
9 8
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3 Rolling
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2 80 $
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8 &
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o 80 100
20 40 60
Percent of full rudder movement
(b) Air speed, 120 miles per hour.
FIGURE 17.—Rudder characteristics. Flap up, power off.
Figure 17 shows that the rolling acceleration resulting
from an abrupt rudder kick is of the order of one-half
the yawing acceleration. This fact indicates that the
effective dihedral is at the upper limit permitted by the
requirements for dihedral. If the rudder specification
is changed, the specification for the maximum dihedral
should also be revised. It is suggested that the di-
hedral be defined in terms of the amount of aileron
required to hold the specified angle of bank in a steady
a. It shall be possible to obtain the rudder reactions
without applying a force greater than 180 pounds to the
rudder pedals.
b. The rudder force shall be proportional to the
rudder deflection.
Procedure.—The data may be obtained by supple-
menting the tests of the previous section by measure-
ments of rudder force.
Results and discussion.—The data for rudder force
have been plotted in figures 17 and 18. The maximum
force recorded was 120 pounds. At both test speeds,
the rudder force varied almost linearly with rudder
a. Down to 10 percent in excess of the minimum
speed, it shall be possible, by adjustment of the trim-
ming tabs, to fly straight with any three engines operat-
ing or with one inboard and the opposite outboard
engines cut out with no force on the rudder pedals.
b. Down to 20 miles per hour above the minimum
speed, flaps down, it shall be possible with the trim-
ming tabs to reduce to 30 pounds the force on the
rudder pedals required for straight flight with both
engines on either side cut out and with those on the
opposite side operating at full rated power.
a. It shall be possible to enter a 45° banked turn at
140 miles per hour in 5 seconds without having the
rudder force exceed 100 pounds or the aileron force
exceed 75 pounds. The same limitations on forces
shall apply to a 30° banked turn at 200 miles per hour
entered in 4 seconds.
b. It shall be possible to make normal banked turns
up to a 15° bank at speeds within 5 miles per hour of
the minimum with the flaps either up or down. It shall
be possible with flaps either up or down to fly the air-
plane steadily for at least 10 seconds up to the Limit
of the elevator control if the elevator control is sufficient
to stall the airplane.
c. It shall be possible, at speeds beyond 10 percent
more than the minimum, to maintain a steady sideslip
with an angle of bank of 20°.
Procedure for item a.—At the specified speeds, make
the required turns using the artificial horizon for
reference and measure the maximum rudder and aileron
Procedure for item b.—Obtain the desired informa-
tion by direct observation.
Procedure for item c.—At a series of speeds beyond
10 percent more than the minimum, place the airplane
at the maximum steady sideslip that can be maintained.
Obtain the angle of bank from the artificial horizon.
Record the speed and the rudder and the aileron posi-
tions and the rudder and the aileron forces.
Rudd i e
80 £2 8
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Air speed, mph
(8) Flap up, power off.
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Airspeed, mph
(b) Flap up, power on.
FIGURE 18.—Sideslipping characteristics.
Results and discussion.—The results of the measure-
ment of the forces in normal turns are given in table IV.
In no case do the forces approach those specified.
Maxi- Maximum forces (1b)
Fla Speed | mum АНУ Allerons
р (mph) | bank (sec) Eleva- Rud-
(deg) tor Right Left der
Up. 81 45 6 16 7 4 [eee
Do__.| 137 50 5 38 10 10 40
Down....| 78 60 6 31 9 11 35
Do...| 122 60 6 35 10 | 8 40
From direct observation, it was found possible to
make banked turns with a 15° angle of bank down to a
speed of 5 miles per hour more than the minimum for
all conditions. With power off, the airplane handled
normally down to the stalling speeds. With power on
at about 3 miles per hour above the stalling speed, an
uncontrollable lateral oscillation of constant amplitude
was encountered. At the stall with power on, this os-
cillation became unstable and the airplane consistently
fell off to the left. With power off, the airplane could
be controlled indefinitely beyond the stall to the limit
of the elevator control. The flight was not very steady
however, because any movement disturbed equilibrium
The results of the sideslip tests are given in figure 18,
which is indicative of the type of data that will be
obtained if the rudder and the dihedral specifications
are changed. It will be noted that, in the sideslip,
approximately 20 percent of the full aileron control
was used and practically the entire rudder deflection
was required. With power off, only at the highest speed
could the specified angle of bank of 20° be maintained.
The results of the tests of the Stinson airplane have
indicated that the tentative schedule of flying qualities
and flight tests is generally satisfactory. The follow-
ing revisions are suggested:
(2) The specification for the period of the longi-
tudinal oscillation should be changed so as to de-
fine the period on the basis of the empirical equa-
P=0.262 Y
where P is the period in seconds and V is the air
speed in miles per hour. 'The period is permitted
to deviate from the value given by this equation
with the power off +5 seconds and with the power
on +10 seconds.
(b) The specification relating to the amount of
damping should be eliminated. If this omission
seems inadvisable, the amount of damping should
be based on the number of cycles for the oscillation
to decrease to one-half amplitude instead of to one-
fifth amplitude.
(c) The specification for the maximum amount
of effective dihedral should be based on the amount
of aileron control required to obtain a specified
amount of bank in the steady sideslip.
(d) The rudder specification should be based on
the maximum angle of bank that can be held in the
steady sideslip with full rudder control. The
present specification relating to the change of
heading in 1 second should be made subsidiary.
The recording instruments used for the measurements
functioned satisfactorily. The control-force recorder,
however, is cumbersome and difficult to install because
flying time.
of the four high-pressure rubber tubes required to
connect the recorder to the indicator. In order to im-
prove the situation regarding this instrument, the deci-
sion has been made to convert it to an indicator by in-
stalling four indicating pressure gages directly on the
wheel. Thus, in future investigations, a recording air-
speed meter, an accelerometer, two turnmeters, and a
control-position recorder will be used in conjunction
with a wheel-force indicator, a rudder-force indicator,
and an indicating altimeter. The further use of the
complete set of indicating instruments by the Commit-
tee 1s considered inadvisable.
The possibility of using only indicating instruments
has been demonstrated and refinements have been in-
dicated. The refinements should include an improve-
ment to the arrangements for photographing the in-
struments, probably through the use of an extra set
especially mounted and artificially illuminated, and
should include modification and the calibration of the
rate-of-turn indicator.
The tests of the Stinson airplane involved approxi-
mately 20 hours of flying time. About one-fourth of
this time may be attributed to repeat flights resulting
from improper operation of the control-force recorder,
which was developed and tried for the first time during
these tests. On the basis of the experience obtained,
it 1s estimated that similar tests with one center-of-
gravity position could be repeated on another airplane
with a total flying time of approximately 10 hours.
The over-all time would depend on the weather condi-
tions but, with ordinary weather, it is estimated that
tests could be completed in 2 weeks. This estimate refers
to tests with one propeller pitch setting and one center-
of-gravity location. A change of propeller pitch is
expected to have little effect on the flying qualities and
check tests at different pitch settings would add little
The center-of-gravity changes permitted
on most large airplanes are so great as to require a
complete repetition of the test program; the total time
required would therefore depend on the number of
center-of-gravity positions tested. Probably only the
rearmost and the foremost positions would have to be
tested, in which case 1t appears that not more than 4
weeks would be required.
LanGLEY FIELD, Va., March 29, 1940.
1. Soulé, Hartley A.: Flight Measurements of the Dynamic
Longitudinal Stability of Several Airplanes and a Correla-
tion of the Measurements with Pilots” Observations of
Handling Characteristics. Rep. No. 578, №. А. С. А.,
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