FULL-SCALE WIND-TUNNEL TESTS OF

NASA
TECHNICAL
NASA TN D-6573
NOTE
!
Z
Z
FULL-SCALE
WIND-TUNNEL
TESTS
OF A SMALL UNPOWERED
AIRCRAFT
WITH
A T-TAIL
by Paul
Ames
T.
Soderman
Research
and
Thomas
JET
N.
Aiken
Center
and
U.S.
Army
Air
Mobility
Calif.
R&D
Laboratory
Moffett
Field,
94035
NATIONAL
AERONAUTICS AND SPACE ADMINISTRATION
•
WASHINGTON,
D. C. •
NOVEMBER 1971
.L
I.
i
_ 2. Government
Report No.
NASA
TN
Accession No.
I
D-6573
3. Recipient's Catalog No.
4. Title and Subtitle
FULL-SCALE
JET
5. Report
WIND-TUNNEL
AIRCRAFT
WITH
TFSTS
OF
A
SMALL
UNPOWERED
A T-TAIL
Organization
Code
Performing Organization
T. Soderman
and
Thomas
10. Work
Research
Moffett
Center,
Field,
Calif.,
Report No.
A-3135
N. Aiken
9. Performing Organization Name and Address
Ames
1971
6. Performing
7. Author(sl
Paul
Date
November
Unit No.
126-13-01-43-00-21
NASA
11. Contract
94035
or Grant
No.
13. Type of Report and Period Covered
12. Sponsoring Agency Name and Address
National
Aeronautics
Washington,
and
Space
Technical
Note
Administration
14. Sponsoring
D. C., 20546
Agency Code
15, Supptementary Notes
16. Abstract
The aerodynamic
characteristics
of a full-scale
execfitive
type jet transport
aircraft
with
the Ames
40- by 80-Foot
(12.2by 24.4-meter)
Wind Tunnel
(subsonic).
Static
longitudinal
control
characteristics
were determined
at angles of attack
from -2 ° to +42 ° .
The
leadingnacelles,
plane
mean
aircraft
wing
had
and trailing-edge
wing tip tanks
location
were
aerodynamic
13 ° of
sweep
and
high-lift
devices.
and empannage_
calculated.
The
an
aspect
ratio
of
5.02.
The
aircraft
The basic conf'tguration
was tested
Hinge-moment
data
were obtained
data
were
obtained
at
Reynolds
numbers
was tested
between
at angles
Hinge-moment
data
and
of attack
showed
flight-test
data
no
30 ° and
regions
that
40 ° (depending
would
result
on e.g.
in
adverse
18, Distribution
stall
aircraft
General
with
various
and
8.7X106
based
on
occurred
ofT-tail
location
effects
and
near the angle
configurations.
wing
flap setting).
on
stick
force.
Comparisons
Unclassified
Statement
Unlimited
aviation
19. Security
wing
of attack
for
A stable
trim
are presented.
! 17. Key Words (Suggested by Author(s))
Deep
T-tail
off
in
and
such components
as engine
angles in the horizontal-tail
of4.1×106
point
data
were investigated
lateral
stability
chord.
stability
through
initial stall. Severe tail buffet
aircraft
had pronounced
pitch-up,
characteristic
wind-tunnel
power
with and without
and downwash
The model had static longitudinal
maximum
lift. Above
initial stall the
was possible
a T-tail
and
Classif. (of this report)
Unclassified
/ 20. Security Classif. (of this page)
l
Unclassified
21. No. of Pages
Technical
99
Information
Service, Springfield,
Virginia
Price"
$3.00
ii
For sale by the National
22.
22151
of
NOTATION
b
wing span,
10.40 in (34.1
c
wing chord
measured
mean aerodynamic
ft)
parallel
S_ /"
,,o b/2
chord,
CD
drag coefficient
(wind
CI
rolling-moment
coefficient
to the plane of symmetry,
c2 dy, 2.14 m (7.04
ft)
axes), drag
qooS
about
stability
axis,
rolling moment
OCl
%
m (ft)
qooSb
_"7" lateral
stability
ac__2/
aileron
parameter,
effectiveness
CI6 a
aCSa
CL
lift coefficient
(wind
per deg
parameter,
axes),
per deg
lift
qooS
OCL
CLfi
_8"-_ flap effectiveness
Cm
pitching-moment
parameter,
coefficient
Cmo_
,CJ_n,
• longitudinal
_)o_
Cn
yawing-moment
coefficient
yawing moment
qooSb
per deg
d
_ (stability
about
stability
parameter,
about
pitching
axes),
moment
qooSd
per dog
moment
center
shown
in figure 2(a) (stability
axes),
aCn
Cnb3
a#
directional
stability
parameter,
per deg
effectiveness
parameter,
per deg
OCn
rudder
Clt81.
OSr
Cy
side-force
it
horizontal-tail
q
dynamic
pressure,
R
Reynolds
number,
coefficient
(wind axes),
incidence
angle,
N/m 2 (lb/sq
side force
q_S
deg
ft)
Vood
P
S
wing area, 21.50 m 2 (231.77
V_
free-stream
velocity,
m/sec
ft 2)
(ft/sec)
211
Y
spanwise
distance
O_
angle of attack
#
angle of sideslip,
_Sa
trailing-edge
8e
elevator
8f
trailing-edge
8r
rudder
deflection
angle, deg; positive
8s
spoiler
deflection
angle, deg
r7
of fuselage,
m (ft)
deg
- nose to left
deflection
angle, deg; positive
angle, deg; positive
flap deflection
downwash
to tile plane of symmetry,
deg; positive
aileron
deflection
average
perpendicular
- left aileron
trailing
edge down
- trailing edge down
measured
from wing chord
- trailing
at the tail location
line, deg
edge left
with respect
to free stream,
deg
Y
wing semispan
station,
At/4
sweep angle of quarter-chord
v
free-stream
leading-edge
kinematic
contours
line,
viscosity,
defined
13°
na2/sec (ft 2/sec)
on figure 2(d)
Subscripts
L
left
max
maximum
R
right
t
tail
u
uncorrected
A
change
free stream
Hinge Moments
Positive
hinge moments
tend to move the control
surface in the direction
deflection.
The average chord aft of the hinge line was used for the reference length.
iv
of positive
Aileron
hinge moment
qSada
where
Cha =
Sa
da
= 0.544 m z (5.85 ft 2)
=0.38 m (1.24 ft)
=
hinge moment
qSrdr
where
Chr
Sr
dr
= 0.609 m 2 (6.56 ft 2)
= 0.46 m (1.51 ft)
Che =
hinge moment
qSede
where
Se
=0.635
de
=0.29
hinge moment
qShdh
where
Sh
eh
= 5.02 m 2 (54.0
=l.17m(3.83ft)
Rudder
Elevator
Horizontal
ft 2)
ft)
stabilizer
ChhNote:
m 2 (6.83
m (0.96
Se is the area of the right or left elevator;
ft 2)
S h is the total area of the horizontal
stabilizer.
FULL-SCALE WIND-TUNNEL TESTS
UNPOWERED
JET AIRCRAFT
Paul T. Soderman
WITH
and Thomas
Ames Research
and
OF
A SMALL
A T-TAlL
N. Aiken
Center
U.S. Army Air Mobility R&D Laboratory
Moffett Field, California 94035
SUMMARY
The
aerodynamic
characteristics
T-tail were investigated
in the
Static longitudinal
and lateral
attack from -2 ° to +42 ° .
of a full-scale
executive
type jet
transport
aircraft
with a
Ames 40- by 80-Foot
(12.2- by 24.4-m) Wind Tunnel (subsonic).
stability
and control
characteristics
were determined
at angles of
The aircraft wing had 13 ° of sweep and an aspect ratio of 5.02. The aircraft was tested power
off with various wing leading- and trailing-edge
high-lift devices. The basic configuration
was tested
with
and without
such
components
as engine
nacelles,
wing tip tanks,
and empennage.
Hinge-moment
data were obtained and downwash
angles in the horizontal-tail
plane location were
calculated.
The data were obtained
at Reynolds numbers of 4.1XI06
and 8.7X106
based on wing
mean aerodynamic
chord.
The model had static longitudinal
stability through initial stall. Severe tail buffet occurred near
the angle of attack
for maximum
lift. Above initial stall the aircraft had pronounced
pitch-up,
characteristics
of T-tail configurations.
A stable trim point was possible at angles of attack between
30 ° and 40 ° (depending
on c.g. location and flap setting).
Hinge-moment
data showed no regions that would result in adverse
Comparisons
of wind-tunnel
data and flight-test data are presented.
effects
on stick
force.
INTRODUCTION
Most small aircraft,
including
executive
jet transports,
are designed
with a minimum
of
wind-tunnel
data. Furthermore,
flight tests are likely to be qualitative
rather than quantitative.
As a
result, the designer has little opportunity
to verify his design predictions.
Therefore,
to aid designers
longitudinal
and lateral
stability
executive
jet aircraft.
The deep
unfavorable
characteristics
the
present
investigation
and control
stall testing
at high angles of attack
was conducted
characteristics
was conducted
because
to determine
the static
through
deep stall of a full-scale
to see if the aircraft
exhibited
of its T-tail. Some of these problems
and
relatedresearchcan be found in references1 through4. Unfortunately,it cannotbe determined
from wind-tunneltestsof unpoweredaircraft whetherthe poweredaircraft canbecomelockedin
deepstall.
AIRCRAFT AND APPARATUS
In figuresl(a) and (b) the model is shownmountedin the Ames40- by 80-Foot(12.2-by
24.4-m) Wind Tunnel. Pertinent dimensionsof the basic model configurationsare given in
figures2(a)and(b).
Wing
Thewinghada quarterchordsweepof 13°, an aspect ratio of 5.02, a taper ratio of 0.507, and
a dihedral angle of 2.5 °. The airfoil section was an NACA 64A
109 modified by increased camber
and chord at the leading edge (fig. 2(d)) which was minimum
at the root and maximum
at the
wing-tip
tank junction.
High Lift Devices
Flap details- The basic wing had a single slotted, extendable
(Fowler)
flap (fig. 2(c)) located
from the edge of the fuselage at 7.1 percent to 61.2 percent r/. Maximum flap angle was 40 ° at the
lower Reynolds
number
and 38 ° at the higher Reynolds
number
because of air load effects. A
center section
flap that extended
under the fuselage was tested (fig. l(b)). There were no gaps
between the sides of the center section flap and the main flaps.
Leading-edge
contoursThe drooped
leading edges of the basic wing were removed, part way
through the test, and replaced by various leading-edge
contours (fig. 2(d)). The dimensions
of the
leading edges varied linearly from root to tip.
Wing plan form modificationIn an attempt
to delay the stalling of the wing tip region,
were placed first on the tops and then on the sides of the tip tanks (fig. 2(e)).
Lateral
Aileronsdecrease stick
such that
The
force.
fence,
Controls
ailerons
(fig. 2(b)) had relatively
blunt
As the ailerons were moved, the balance
leading edges and balance
tabs to
tabs moved in the opposite direction
e_tab = -(5/6)6a
where C_tab is the tab angle relative
wing chord.
2
to the aileron
chord
and 6a is the aileron
angle relative
to the
Spoilers- The chords of the spoilers were 10.25 percent of the wing chord at midspoiler
and
were located from 22.2 percent
to 49.4 percent semispan (see fig. 2(b)). Spoiler angles ranged from
0° to 42 °. In addition to the basic wing spoilers, dummy spoilers were tested outboard
(fig. 2(f)).
Tail
The geometry
of the horizontal
and vertical tails is described in figure 2(g). Pitch control was
provided
by an all-movable
tail that had an available deflection
range of 0.4 ° to -7.0 ° and by a
32 percent chord elevator with balance horn. The elevator angle was variable from 15 ° to -15 °. The
rudder (25 percent chord) had a deflection
range of 30 ° to-30 ° and had a trim tab that was locked
at 0 °. The horizontal
stabilizer was used for aircraft trim.
Nacelles
Engine nacelle detail and location are shown in figure 2(h). A constant-area
circular duct was
installed in each nacelle to allow mass flow conditions
of 4.81 kg/sec (10.6 lb/sec) of air at standard
conditions,
similar to that of the jet engines for idle airflow at a Mach number of 0.2. Static and
total pressures
were measured
with rakes at the aft ends of the ducts to determine
the actual
dynamic
pressure of the nacelle
data). The nacelles were removed
flow and the internal naceIle drag (which
from the pylons during a part of the test.
was removed
from
the
Tip Tanks
tanks
Wing tip tank detail and location
on unless stated otherwise.
are shown
TESTING
AND
in figure 2(b). All data are presented
with the tip
PROCEDURE
Forces and moments
were measured
for the model through an angle-of-attack
range from -2 °
to 42 °. Pitching-moment
data were computed
about a moment center location at 25 percent & The
center-of-gravity
range for this aircraft is 16 percent C to 31.5 percent _, Tests were conducted
at
Reynolds
numbers
of 4.1×106
and 8.7X10
6 based
on a mean aerodynamic
chord of 2.14 m
(7.04 ft) and speeds of 27.8 m/see (54.2 knots) and 59.0 m/sec (115.0 knots), respectively.
These
speeds correspond
to dynamic pressures of 478.8 N/m s (q = 10 psf) and 2156 N/m 2 (q = 45 psf).
Tests were conducted
with
elevator, rudder, aileron, spoiler,
The maximum
angle of attack
limitations.
Most data, tail on, at
of4.1×106
.
the basic configuration
_ at several tail incidences
with variable
and flap settings. Data were also obtained with landing gear down.
at R=8.7×106
was 16 ° (tail on) because
of tail buffet
load
angles of attack higher than 16 ° were taken at a Reynolds number
_Basic configuration refers to the airplane as shown in figure l(a) with engine nacelles, tip tanks, and
empennage on model. Control surfaces were at zero angle unless stated otherwise.
DATA ACQUISITIONAND REDUCTION
Forcesandmomentsweremeasured
on thewind-tunnelsix-component
balance.Torquetubes
in the elevatorsandrudderweregagedto providehinge-moment
data.
All datawerecorrectedfor strut tares,
Nacelle
internal
AC D = 0.0005
effects were
flow drag
cos a
was
was calculated
subtracted
from
Ac_ = 0.506
model
drag.
Corrections
CLu 2
ACm
CLu (tail on runs only)
= 0.0171
measured
reading,
added
OF
wind-tunnel
wall
MEASUREMENT
were accurate within the following
and reducing the data.
Angle of attack
Angle of sideslip
Free-stream
dynamic pressure
Control surface settings
Force
for
CLu
AC D = 0.0088
ACCURACY
The various quantities
limits involved in calibrating,
nacelle internal flow drag, and wind-tunnel
wall effects.
from pressure measurements
in the nacelle ducts, and
or moment
N (+5 lb)
limits,
which
include
error
_+0.2°
+0.5 °
-+0.5 percent
-+0.5°
Coefficients
R = 8.7×10
Lift
+22.4
Drag
Side force
-+13.4 N (+3 lb)
_+13.4 N (+3 lb)
_+.0003
+.0003
Pitching moment
Yawing moment
Rolling moment
+271 J (+200
ft-lb)
-+136 J (_+100 ft-lb)
--.475 J (+350 ft-lb)
-+.0027
+.0003
_+.0010
at
6
+0.0005
RESULTS
Table 1 is the index to the figures. The longitudinal
data are presented in figures 3 to 18 and
the lateral data in figures 19 to 30. Downwash
and hinge-moment
data are given in figures 31 and
32, respectively.
4
DISCUSSION
LongitudinalCharacteristics
Flap effectiveness-
The longitudinal
characteristics
of the basic airplane
at R = 8.7×106
with
three flap settings are shown in figure 3(a). The flap effectiveness
parameter,
CL6, was 0.015/deg
for the 20 ° flap setting and 0.013/deg
for the 38 ° flap setting. A theoretical
flap effectiveness
estimate was made using the simplified lifting-surface
theory of reference
5, which gave the value of
CL6 as 0.022/deg,
almost 60 percent higher than measured. This discrepancy
was probably due to a
nonoptimum
gap setting for the single-slotted,
Fowler type flaps. A comparison
of small-scale with
full-scale wind-tunnel
data to be discussed in a later section shows that small-scale flap effectiveness
is closer to the theoretical
value. This suggests that the flap gap choice was based on small-scale
data and not corrected properly for full-scale Reynolds number effects.
test
Maximum
lift- Figure 3(a) shows the basic stall characteristics
of the aircraft at R = 8.7× 106 .
Because of severe buffet on the tail as it penetrated
the wing wake, the tail was guy-wired as shown
in figure l(a). 2 In addition,
some of the data were taken at a reduced
Reynolds
number
of
4.1×106.
The tail buffet
acted as a strong stall warning.
Figure 3(b) shows the longitudinal
characteristics
at R = 4.1X 106 . Increasing
the Reynolds number from 4.1 × 106 to 8.7× 106 caused
an increase
in maximum
lift coefficient
of 0.19 (flaps down) and 0.20 (flaps up) as shown in
figure 4. The high Reynolds
number condition
is closer to actual flight conditions.
Observation
of
tufts on the left wing indicated that a region of separated flow developed near the wing leading edge
tip tank junction
at 8 ° angle of attack (this did not happen with tip tanks off). As angle of attack
was increased the region of separated flow spread aft and inboard. Near CLmax the wing root began
to stall. Both regions grew with angle of attack until most of the wing stalled and lift dropped.
Static stabilityA study of the variation of the stick-fixed
pitching-moment
coefficient
with
angle of attack (fig. 5(a)) shows that the airplane was stable through maximum lift (even for aft c.g.
limit of 31.5 percent e). Above maximum lift, the classic deep stall situation
occurred that will be
discussed
later. The data presented
for c.g. at 25 percent
d give Cma = -0.0186/deg.
At stall the
aircraft
experienced
a slight
nose
down pitching
moment.
The
stick-free
static
stability
characteristics,
determined
from
hinge-moment
and pitching-moment
curves, are shown
in
figure 5(b) (data are shown for c.g. at 25 percent d). Freeing the elevator reduced the stability, but
the aircraft did not become unstable.
reduced from -0.0185 to -0.005/deg.
For the aft c.g. case (31.5
percent
d), 6 f = 0°, o_= 0 °, Cmo_ was
Deep stall- As illustrated
in figure 6, the airplane
was unstable
above maximum
lift
(stick-fixed)
with the center of gravity at the quarter chord, and maximum nose-down
trim until an
angle of attack of 32 ° was reached at which point static stability was again attained. Furthermore,
the pitching moments
became zero or slightly positive above a = 28 ° flaps down. Thus it may be
possible (at Iow Reynolds
number)
for the airplane to reach a region of positive pitching moment
and pitch up to e_= 32 °, a trim condition
(power off) if the pilot does not take corrective action.
However,
aircraft rolloff may preclude
this possibility,
as will be discussed in a later section. As
shown by the axes superimposed
on figure 6(b), at forward
c.g. the pitching moments
do not
become positive, but at the aft c.g. the aircraft would reach the positive pitching-moment
region
2The wires had very little effect on the data.
sooner and could pitch up to trim at
Figure 6(c) shows that while the effect
moment only 0.06 at c_= 32 ° .
o_= 39 ° (flaps up or down) while completely
stalled.
of sideslip was beneficial, 8 ° of sideslip changed the pitching
With the flaps up, elevator
effectiveness
was maintained
at all angles of attack
but
pitching-moment
increment
due to full elevator deflection
at angles of attack greater than 24 ° is
approximately
one fourth that at angles of attack below stall (see fig. 7(a)). Therefore,
recovery
from deep stall (flaps up, c.g. at 25 percent 6) would be possible using the elevator, but the time it
takes to rotate the nose down may be long. With the c.g. in the aft location there is insufficient
elevator effectiveness
to recover from deep stall. With the flaps full down (fig. 7(b)) there was an
almost complete
loss of elevator control power above o_= 24 °. Since the data (flaps up and down)
were taken with the horizontal
stabilizer leading edge full up, any movement of that control surface
would only make the pitching moments more positive.
Figures 8 and 9 illustrate the effects
empennage,
respectively,
on the longitudinal
of horizontal
characteristics.
stabilizer
incidence
and
removal
of the
Effect of wing tip tanks, engine nacelles, and landing gear- Figure I0 shows that the wing tip
tanks caused an increase in lift coefficient
and lift curve slope primarily because of the increased
wing area and aspect ratio (reference
area was not changed).
The drag change was small tip to
CLmax.
The addition
of the tip tanks
The engine nacelles caused
was probably due to interference
made the aircraft
slightly
more stable
in pitch.
a decrease in lift, especially with flaps down (fig. 1 1). This decrease
with flow around the wing that redtlced wing lift since the nacelles
did not develop negative lift or reduce tail lift. This explanation
is substantiated
by the increase in
nose-down
pitching moment with the nacelles on the aircraft. If the nacelles had developed negative
lift or if the tail lift had been reduced, the pitching-moment
change would have been nose up. The
fact that the wing tips were probably
not affected
by the nacelles accounts
for the nose-down
pitching-moment
change (i.e., the lift loss was inboard).
The landing
gear effect on the longitudinal
characteristics
is small (fig. 12).
High-lift devices- The effects of four wing leading edges are given in figures 13(a) and (b), flaps
up and down. For the flap down case the leading edge 14, which had the greatest droop, increased
maximum lift beyond the value achieved by l 3 , the basic configuration
leading edge.
In an attempt
to improve
the CLmax
of the airplane,
fences
were placed
on the tops and,
later, sides of the tip tanks to alleviate flow separation
at the junction
of the tip tank and wing.
Fences on the sides of the tip tanks caused an increase in lift due to the increased wing area and
aspect ratio (fig. 14). In no case was the flow separation
alleviated near the tip.
The center body flap (fig. l(b)) caused a very small reduction in lift and drag of the model and
a very slight change in pitching moment (fig. 15). The reason for the reduction
in lift and drag is
unknown.
Drooping
the ailerons
13.7 ° increased
drooped ailerons reduce roll control, outboard
lateral control section.
6
maximum
lift coefficient
spoilers were tested. These
by 0.1 (fig. 16). Since
will be discussed in the
EfJ_,ct of spoilersRuns were made with various right and left spoiler deflections
(see
figs. 17(a)
and
(b)).
The
deflection
of one or both
spoilers
42 ° caused
a nose-down
pitching-moment
change probably
because
of an induced
increase of tail angle of attack. This
supposition
checks
with figure 17(c) that shows very little change in pitching moment
with
outboard
spoiler deflection.
It was expected
that the flow field of the tail would not be affected
greatly by deflection
deflection.
of the outboard
spoilers.
The
drag was increased
80 percent
with full spoiler
Comparison
of wind-tunnel
and flight-test
data- A comparison
of Ames 40- by 80-Foot (12.2by 24.4-m) Wind Tunnel data, Wichita State University 7-by 10-Foot (2.1-by
3.l-m)Wind
Tunnel
data and Lear Jet Flight-test data is made in reference 6. Two figures from that paper are presented
in this report as figures 18(a) and (b). Results show good agreement
between full-scale wind-tunnel
and flight-test
data.
Reynolds
number
effects account
for most of the difference
between
small-scale and full-scale results.
Lateral
The
lateral
characteristics
and Directional
of
the
airplane
Stability
are
and Control
shown
in figures
19 to 23, and
lateral
and
directional
control
effectiveness
in figures 24 to 29. Stability derivatives Cn/3 and C//3 are plotted
versus angle of attack in figure 30. These data show that the airplane had positive effective dihedral
(-CI/3) over the normal operating
range and was directionally
stable statically
(positive Cn/3). With
the tail removed (fig. 22) the nonzero rolling moment and side force at/3 = 0 ° were probably due to
flap misalinement.
The flaps had been removed and reinstalled
on the model prior to these runs.
The data in figure 23 show that as the model stalled with flaps up, it tried to roll left (left wind
down) and with flaps down, it tried to roll right (right wing down). The change in roll direction at
stall was probably
caused by asymmetric
deflection
of the flaps. The rolling moment,
flaps down,
was greater than that produced
by full opposite aileron deflection.
This severe rolloff in stall would
complicate
recovery, but it might prevent a deep stall condition.
Control effectivenessAileron
in stall (fig. 24). The airplane had
Figures 24(b)
(d) show the control
due
attack.to model
Rudder
misalinement
deflection
lateral effects of rudder
between-15
° --.</3_< 15 °.
The control
versus left spoiler
outboard
spoilers
roll power was fairly constant below stall but decreased rapidly
slight favorable yaw due to aileron above 6 ° angle of attack.
power due to one aileron. The nonzero side force was probably
in the test section.
affected
Figure 25 is a summary
the longitudinal
deflection.
The
rudder
very
little.
of holding
Figure
angle of
26 shows
the airplane
the
in sideslip
power of the basic spoiler is shown in figure 28(a) as plots of Cy, Cn, and C l
angle (right spoiler full down). Figure 28(b) shows the effectiveness
of dummy
S_, $2, and $3. These spoilers were more effective
than ailerons
or inboard
spoilers for lateral control. The lateral characteristics
are shown in figure 29. Comparison
with the results
that the landing
characteristics
was capable
plot of C16 a versus
gear had a small effect
of the airplane with the landing gear extended
in figure 19(a) (landing gear retracted)
indicates
on Cy vs./3 but only a slight effect
on Cn and C l vs./3.
!
\
!
Downwash
at the Horizontal
An average downwash
angle at the horizontal
for several tail incidence
angles. The intersection
points where tail lift is zero; and for a symmetrical
stabilizer was estimated
from curves of Cm vs. ot
of the tail-on curves with the tail-off curve are
horizontal
stabilizer
e=o_+i
Figure 31 shows
number cases.
the results
of the above
Tail
t
calculation,
which were identical
for both Reynolds
Hinge Moments
Typical curves of hinge-moment
coefficient
C h versus angle of attack and C h versus control
position
are presented
in figures 32(a)-(h)
for aileron, elevator, rudder, and horizontal
stabilizer.
The data were obtained at R = 8.7X 106 to approximate
actual flight conditions.
These results show
no control force reversal for any of the controls within the normal operating range.
CONCLUSIONS
A full-scale
wind-tunnel
investigation
was made of a small jet aircraft
determine
the longitudinal
and lateral stability and control characteristics
through
off. The following conclusions
were drawn from the results of the investigation:
I. The airplane
the full
unstable.
c.g.
range.
had stick-fixed
With
the
stick
static
free,
longitudinal
stability
stability
was reduced
with a T-tail to
deep stall, power
at angles
of attack
but
aircraft
the
up to stall for
did not become
2. Before stall the tail experienced
severe buffet as it penetrated
the wing wake, and, in stall,
the airplane tended to roll right wing down or left wing down depending
on flap angle. The tail
buffet
acted as a strong stall warning,
that might prevent deep stall entry during actual flight
conditions.
However, the rolling moment in stall, flaps down, was greater than that produced
by full
opposite aileron deflection.
3. Above stall, the airplane was unstable in pitch, and the pitching moments could become
positive, depending
on c.g. A trim condition
in deep stall (a = 39 °) with a large reduction
in elevator
control
was possible. With the c.g. in the aft position,
elevator control and horizontal
stabilizer
control were insufficient
for recovery from deep stall trim.
4. The airplane
Ames Research
was directionally
stable,
below
Center
National Aeronautics
and Space Administration
Moffett Field, Calif., 94035, July 6, 1971
8
stall, and had positive
effective
dihedral.
REFERENCES
1. Aoyagi, Kiyoshi; and Tolhurst, William H., Jr.: Large-Scale Wind-Tunnel
Engine Nacelles and High Tail. NASA TN D 3797, 1967.
Tests of a Subsonic
2. Ray, Edward J.; and Taylor, Robert T.: Effect of Configuration
Variables on the Subsonic
Characteristics
of a High-Tail Transport Configuration.
NASA TM X-1165,
1965.
3. Trubshaw,
E. B.: Low Speed Handling
no. 667, pp. 695 704, July 1966.
4. Thomas, H. H. B. M.: A Study
R.A.E. TM Aero 953, 1966.
With Special
Reference
of the Longitudinal
Behavior
5. DeYoung, John: Theoretical Symmetric Span Loading
at Subsonic Speeds. NASA Rep. 1071,1952.
Transport
Longitudinal
to the Super Stall. J. Roy. Aeronaut.
of an Aircraft
Near-Stall
Due to Flap Deflection
and Post-Stall
for Wings of Arbitrary
With Aft
Stability
Soc., vol. 70,
Conditions.
Plan Form
6. Neal, Ronald D.: Correlation of Small-Scale and Full-Scale Wind-Tunnel Data With Flight Test Data on the Lear
Jet Model 23. Paper 700237 presented at SAE National Business Aircraft Meeting (Wichita, Kansas), March
1970.
TABLE
1.-INDEX
TO
FIGURES
The model in the wind tunnel
....................................
Geometric
details of the model
...................................
Longitudinal
characteristics
with flap deflections
.........................
Reynolds number effect on longitudinal
characteristics
.....................
Variation of pitching-moment
coefficient
with angle of attack
.................
Longitudinal
characteristics
through deep stall
..........................
Elevator effectiveness
.........................................
Figure
!
2
3
4
5
6
7
Longitudinal
characteristics
with horizontal
tail deflection
...................
Longitudinal
characteristics
with empennage
removed
......................
Longitudinal
characteristics
with tip tanks removed
.......................
Longitudinal
characteristics
with engine nacelles removed
....................
Longitudinal
characteristics
with landing gear down
.......................
Effect of four leading edges on longitudinal
characteristics
...................
Effect of fences on longitudinal
and lateral characteristics
....................
Effect of center section flap on longitudinal
characteristics
...................
Effect of drooped ailerons on longitudinal
characteristics
....................
Longitudinal
characteristics
with spoiler deflection
........................
Comparison
of wind-tunnel
ahd flight-test data ..........................
Lateral characteristics
versus sideslip angle/3 ............................
Lateral characteristics
versus angle of attack c_ ..........................
Lateral characteristics
versus lift coefficient
............................
8
9
10
11
12
13
14
15
16
17
18
19
20
2I
Lateral characteristics
Lateral characteristics
Lateral characteristics
Aileron effectiveness
Lateral characteristics
Rudder effectiveness
versus/3, empennage
off
..........................
versus or, empennage
off ..........................
with aileron deflection
...........................
.........................................
with rudder deflection
...........................
.........................................
22
23
24
25
26
27
Lateral
Lateral
with spoiler deflection
with landing gear down
28
29
characteristics
characteristics
...........................
..........................
Lateral stability derivatives Cn8 and Clt_ ..............................
Downwash at tail ...........................................
30
31
Control
32
10
surface
hinge moments
...................................
E-_D
cq
.0
0
0
LT_
oo
o
<
o
o
E
i
P_
ll
(b) Center
section
flap, nose and tip booms
Figure
3_2
1.- Concluded.
on model.
Wing
Aspect ratio
5.02
Horz.
tail
4.0
Taper ratio
0.507
Area, sq m
(2],1.77)
(54.00)
I3
25
Sweep, deg
(25%
21,51
0.469
5.02.
Vert.
tail
-3.495
34.48
34.48)
40
Airfoil section
64A-10964A008 64A010
All dimensionsin meters (feet)
i)5.86 __
2.47
(8. _o)
4.47
(14.6 7)
Moment
C/4
Line
2.5 ° Dihedral
2.51
(8.25)
(a) General
Figure
arrangement
2.- Geometric
details
10.40
(54.10)
of model.
of the
model.
13
4.40
(14.45)
(4.57)
I .39
/
1
=
i
0.431
(I
/
/
Aileron
hinge line
78% c
0.55
(I.75)
-_
l
(4.91)
13°
10.76
0
35)
5
0.24
i
(o.
/
/
73% c
/
/
Flap leading
edge
J25 '
Moment
center
i
__
Aircraft
(0.93)(I
2.75
(9.02)
(b) Basic wing detail.
Figure
3_4
t
2.- Continued.
.39)
0.37
(I.2 I)
0.268c
Flap angle,
0
2O
deg
Gap,cm(in)
0
1.57 (0.62)
13.97(5.5)
38
2.46(0.97)
18.55 (7.3)
4O
2.46(0.97)
18.80(7.4)
(c) Trailing-edge
Figure
flap.
2.- Continued.
17
Zl
1,z(STD NACA
64A-109)
X
_f
Wing and tip tank junction, cm (in)
(blunt)
/ _/-J
-'/
/6% Chordline
of STD NACA
J
×,:,o._,
<.,.,>
,,-,.,,,o.,>-/tc--_.
"' x,
t 4
/
configuration)
_-
_
--_×_
"
;
-
_
Hinge point
of _4
14
junction,
cm (in)
×1 =8.64(3.4)
R l =2.28(0.9)
X2=13.71 (5.4)
Rz = 1.78(0.7)
(STD NACA
k
drooped
64A-109)
30 °
Chordline
of STD NACA
#
Wing and fuselage
1,1
linch
Scale:
I
I cm
(d)
Leading-edge
Figure
16
modifications.
2.- Continued.
I_
1.61
-
(5.28)
---f0.48
0.18(0.58)
F
Area =(5.81 ff 2)
(I.58)
/
l
Side view
1.61
(5.28)
0.54m
z
Area = 5.81 ft 2
0.1
0.48
(I.58)
Top view
(e) Fence
located
Figure
on side and top of tip tank.
2.- Continued.
17
I
!
A
70 % chord
Symbol
dimension,(in)
Si
(2.25)
$2
(5.50)13.98
S3
(6.75)17.15
(f) Outboard
Figure
18
"A"
.-7
spoiler.
2.- Continued.
5.71
cm
0.45(I
/
--_
0.75
1.67
(5.46)
(0.83)
(4.33)
0.74
(2.42)
/
i.32
/
1.04
(3.42)
1
P
1.43
(4.6 9)
Fuseloge
L
Ventral
fin
r
2.80
vI
(9.17)
CL Aircraft
5_
mom
Hinge
pin
,
1.52
.JIL(5.OO)
o.18
(o.58)
-I
0'.52
(I. 70)
I
J
2.24
(7.35)
0.25 (o.75)L
(g) Horizontal
Figure
and vertical
2.- Continued.
-I
stabilizer.
3-9
(3'61
.I I
4-)
it
0.76
0.63
(2.50)(2.08)i
O.
--(1_
Cross-section
line
ot engine inlet
o-- 0.66
(2_.17) 3
0.60__
I
9-( I, 98)-"
]
Lo.3eJ
Cross-section ot wing leading
edge-fuselage
(h) Fuselage-wing
cross section
Figure
2O
junction
at two locations.
2.- Concluded.
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R = 8.7X
10 6 ,
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48
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oa
0
oa
I•
.J
(D
0
_i•
I
_J --EL_ _L
(xl
i°
J
0
\
PJ
oo_
ffl
©D<
_
°_..q
0
R
ff
_k
×
i
II
II
•
CO
_D
,q
oJ
to
0
_.1
(.P
.52
o
_.
oq.
o
I
!
!
E
0
o
0
_D
o
El
0
--
"-d
0
o
r,.)
!
g
_
M
t_
P_
II
0
o
o
0
II
_c_
E
•
I
017
-4D
I
_D
o
O0
0
_.
m
_.
o
i°
--I
0
53
2.8
I
I
0
I
Flight
1.4
I
test
r
r
Flaps
up
1.2
Ames 40 x 80,
R = 8.6 x I06
Wichita St. Univ. 7xlO
R = 1.4XlO 6
1.0
1.6
F
.8
i
CL
//__Sf:
40 °
CL
1.2
.6
I
/
//
.8
.4:
0
-4
_'
8
12
16
2O
I
0
St. Univ. --
7xlO
I
I
.12
.16
1
.08
.04
1.4
20 °
!
1.2
I
.24
/
full
down
1.2
1
I.O
/
I
I
I
,8
.8-
/,
......
CL
CL
.6
O_
.4.
I
0
Flight
Ames
Wichita
7xlO
.04
.6
--1
.4
o
I
,2
i
.08
[
.12
I
.16
test
40x
Flight
Ames
80
St. Univ.--
I
test
40 x 80
Wichita
.2
St. Univ.--
7xlO
L
.20
.24
0
I
.04
1
.08
I
.12
I
.16
CD
CD
(a) Lift and drag characteristics.
Figure
_4
I
.20
I
Flaps
i
,o
0
test x 80
40
Co
I
Flaps
23
Wichita
a , deg
1.4
Model
flight
Ames
.2
4
0
I
0
.4
/
i
ff
18.- Comparison
of wind-tunnel
and flight-test
data (taken
from ref. 6).
.20
.24
0.2
I
Tail - on
Flops
up
I
Cm
o _..__
_._
_
-0.2
0.2
Tail-off
Cm
0
w
Ames 40X80,
R = 8.6 X 06
0.2
Flops
Toil - on
C
m
Wichita
full down
St. Univ.
7X O,
R = 1.4XIO
_
-0.2
6
-
,lib.
-0.4
0
Tail-off
Cm
-0.2
T
- 0.41
0
,
0.2
0.4
0.6
0.8
1.0
CL
(b) Pitching-moment
Figure
characteristics.
18.- Concluded.
1.2
1.4
1.6
1.8
.24
.2O
[
.16
.12
Cy
.O8
.O4
0
8f, deg
0
-.04
0
038
.O4
0
C
n
-.04
-.08.04
-.0_204
-16
-12
-8
-4
0
4
.8, deg
(a) or= 0°, R = 8.7X106,
Figure
76
19.- Lateral
characteristics
flaps 0 ° and 38 °
of basic configuration
in sideslip.
8
.24
.20
a, deg
0
0
D 8.4
O 14.6
.16
Cy
\
.12
.08
.04
0
0
C n
-.04
.08
C_
.04
0
-20
-16
-12
-8
-4
0
4
8
/3, deg
(b) $f=
0 °, R = 8.7X106,or
= 0 °, 8.4 ° , and 14.6 ° .
Figure 19.- Continued.
77
.12
a, deg
0
12
.08!
Cy
.O4
0
24
-.04
.04
-.08
a,deg
0
Cn
-.04
.04
a,deg
24, :32, I1_
C_
0
-16
- 2
-8
-4
0
_, deg
(c)6f=0
.58
°,R=4.lX106,e=0°,12°,16°,24
Figure
19.- Concluded.
° 32 °
4
,12
.08
Cy
0 (i)--_--_-_
G.--___.
-.04
_3
.O2 -.08
0
Cn
-.O2
_C_
-.04
0
.O6
,8,deg
0
[] -4
0
-8
.O4
C_
.O2
E}-----ID-'OI
-'020
4
8
12
16
20
24
28
52
:56
e, deg
(a) 6f=0
Figure
20.- Lateral
characteristics
°, R=4.1X106,/3=0
of basic configuration
°,-4 ° , _8°
versus angle of attack.
79
.O2
.'3
0
.:>--
Cy
<_
1
-.02
-.04
8f
, deg
0
0
[]
2O
<>
38
.O2
C
n
0
-.02
.O4
i
.O2
C_
,.,-I0
-.02
0
z
8
2
16
a,
20
deg
(b) R = 8.7× 106, f3= 0 °, 6f= 0 ° , 20 ° , 38 °
60
Figure
20.- Concluded.
24
28
q
d
J
o
f
o"
o_
0
0
11
q
0
!
X
II
0
o
II
o
0"'
I-7
0
[]
0
I"
o
GO
o.
I
I
OD
<]
12)
P_
o
c,l
o
o._
\
E
D
o.
i
O.
n
0
TO
OI
q
,_J
61
,O8
O4
Cy
O
mm
-.04
0
[]
0
8
0
0
I
If-
0
Z_
0
8
58
58
|
l
.O4
Crl
0
-.04
O8
.O4
C1
I
0
-0420
-I(
-12
-8
-z
0
4
8
,8, deg
Figure
62
22.- Lateral
characteristics
of the
model in sideslip
R = 8.7X 106 .
with
the empennage
removed;
.O4
.O2
-.04
.04 -.06
,O2
Cn
0
-.02
.O6
Tail
.O4
• off
[] on
• off
8f,
deg
0
40
40
.O2
__
Oon
O
Cl
-.02
--'040
4
8
12
16
20
24
28
32
36
4O
a, deg
Figure
23.- Lateral
characteristics
versus
angle of attack,
R = 4.1X 106.
tail on and
off;
8f = 0 ° and
40 °,
63
.O4
.15,+15
.O2
Cy
0
.O2 -.02
Cn
0
-.02
,deg
+10,0
+5,0
0
0,0
C_
-15,+15
0
4
8
12
a, deg
(a) 6f = 0°, R = 4.1X106
Figure 24.- Aileron
64
effectiveness.
16
20
24
.O2
_o
i
I
0
i
4X
Cy
-.02
-.04
.02
Cn
0
-.02
.O6
.04
8aR, deg
+15 _---
.O2
+10
Z
C_
0
_a
-.02
0
4
8
12
16
2O
24
a, deg
(b)Sf=40
Figure
°, R=4.|X|06
24.- Continued.
6_
.O5
s
.02
$
g
Cy
.OJ [
t_
0
.02
.01
Cn
0
- .01
- .02
.O3
.O2
>
.01
C_
0
a, deg
-.0
- ,02
-15
-I0
-5
0
BaR, deg
(c) 5f= 0 °, R = 8.7X10 e
Figure 24.- Continued.
66
5
©
0
[]
8.4
0
14.6
I0
15
Cy
-.02
.02
Cn
r
0
.,.._...._...,.--.--_
---------K
-.02
.O4
a,deg
.O2
[]
8.7
0
12.9
0
.5
C?.
0
-.02
D
5
-I0
-5
0
5
I0
15
BaR, deg
(d) 6f = 38 ° , R = 8.7X106
Figure
24.- Concluded.
67
.004
I
8f, deg
0
38
0
[]
I
.002
\
\
b
0
-.002
0
5
I0
15
20
e, deg
Figure
25.- C16 a for right aileron
versus angle of attack
(measured
at 6aR = 0°); R = 8.7X 106 .
.O8
.O4
0
Cy
O, deg
0
-%
[
0
[]
8.4
O
16.4
- .08
0
On
- .02
-.04
.O2
C_
- .02
-.04
-3O
-20
-I0
0
St,
(a) 6f= 0°,#
Figure
26.- Effect
of rudder
I0
3O
deg
= 0°, R = 4.1X10
deflection
20
6
on directional
characteristics.
69
,16
.12
.08
,,,,
-
Cy
.04
0
I
- .O4<
0
14.6
-.08
.O4
0
C n
- .02
- .04
-.06
_>
.04
<
-.----41
.02[
C_
0
T
<)
-,02
-30
-20
0
-10
I0
_r,deg
(b) 8f=
0°,/3 = -4 °, R = 8.7X 10 6
Figure
70
26.- Continued.
20
30
.20
.16
of
t
.12
f
Cy
f
J
f
.08
.04
a, deg
O
O,
[]
0
0
8.4
14.6
-.04
.02
C n
-.04
-.06
-.08
.06
.04
C_
.02
_
0
- 50
A
- 20
- I0
0
'
I0
20
3(
B r,deg
(c)6f= 0°,#=-8 °,R = 8.7X106
Figure
26.- Continued.
71
.02
-. 04
•
,
-.06
-.08
-.08
_ f
f
"--"
- ,04
f
-.02[
0
-30
-20
-I0
0
I0
8r, deg
(d) 6f=
0°,/3 =-12 °, R = 8.7XI06
Figure
72
26.- Continued.
2O
30
,32
....X )
,28
/
,24
tf
.20
J_
f_
_.
.08
_.._ _-_-
--
.04 i
0
[]
a,deg
0
8.4
0
14.5
0
-.02
- .04
C n
- .06
-.08
-.10
.O8
>
<_-.
.06 <
C_
.04
.02
3--
0
-3O
_
-2O
_,__........_..._
-IO
...-----'---"
'-
0
I0
2O
3O
Br , deg
(e) 6f=
0°,/3 = -16 °, R = 8.7X l06
Figure 26.- Continued.
73
.08
_4
.O4
Cy
_, deg
©
.3
[-1
8.7
O
{6.6
C n
-.02
Q
-.04
.O4
.O2
C_.
-.02
-30
-20
-I0
0
8r, deg
(f) _ f = 40 ° ,/3 = 0° , R = 4. l X 106
Figure
7/,
26.- Continued.
I0
20
30
.12
._t
.08
J
:, deg
0
[]
.3
8.7
12.8
-.OE
.04
.02
0
Cn
___.
_.
-.02
-.04
-.06
.O6
.04
........
,
_......._>-- -
....
_
C7.
..---()
.02
0 _
-:50
-20
-I0
0
I0
20
50
_r, deg
(g) 6f = 38°,/3 = -4° , R = 8.7X106
Figure
26.- Continued.
75
• 16
J
_I
.12
J
1f
_
J
a,deg
o
0
.3
[]
8.7
0
t2.8
-.04
0
C n
- .02
-.04
-.06
.O6
.04_
C_
_
.02[-]
__"-
--'-
"-
_)-_-
O3o
- 20
-I0
0
Br,deg
(h) 6f = 38 °,/3 = -8 °, R = 8.7X 106
Figure
76
26.- Continued.
I0
2O
3O
.20
.16
_-
.._ /
.12
_
Cy
.08
a, deg
0
.04
<
.3
I-i
8.7
0
12.8
0
.O2
- .02
Cn
-.04
-.06
-.08
-30
-20
-I0
0
I0
20
30
8r,deg
(i) 6f= 38 ° ,/3 =-12 ° , R = 8.7X106
Figure 26.- Continued.
77
.28
i
.24
l
/
i
I
fl
.20;
!
...A :]
it
f
f
I
f
1t
Cy
.i6
I
t-
0
.08
.o4
>'7
0
-.02
- .04
C n
-.06
-.08
-.10
.08
<
.06
CZ
.04 I
.02(
0
-30
8r,deg
(j) 6f = 38 ° ,/7 =-16 ° , R = 8.7XI06
Figure
78
i
f
26.- Concluded.
a,deg
0
f
0
,_ deg
0
I
I
Sf, deg
0
Rnx I0 -_'-
-.0005
_f,deg
__
40
Rn x 10-6
4.1
4.1
I
I
-.0010
2
,,,O
_:,.-_.._..._ J
-.0015
Cnsr
0
-.0005
I
I
m
_f, deg
__
:58 __
Rn x 10-6
deg
_f,
0
__
Rn x 10 -6
8.7
8.7
-.0010
_.0015'_""
0
7-------o--'-o
5
I0
15
20
0
u,deg
5
I0
15
2O
(:z,deg
(a)/3 = 0 ° and -4 °
Figure
27.- Cn6r versus
angle of attack
for flaps up and down;
R = 4.1X l 06 and 8.7X 106 .
79
0
/3,deg
-8
I
I
_f, deg
Sf, deg
38
Rnx 10 -6
0
- .0005
Rn x I0 -6
8.7
8.7
- .0010
• I
f
.----.0--
"--'-0"-
"-o
-.0015
Cn_r
0
"
/9_deg
-12
I
Sf, deg
0
-.0005
I
_f, deg
__
38
-Rn x I0 -6
Rn x 10 .6
8.7
8.7
I
-.0010
-.0015
_.._.._ .--...o-.-..._
_
0
5
I0
:2...__......-.-.o
_
15
2O
a,deg
5
I0
a,deg
(b)/3 = -8 ° and - 12 °
Figure 27.- Concluded.
80
0
15
20
o0
0
_D
X
p..,
eJ
/-
_
II
;.2¸
""
0
I
/
I
0
(k.l
..9
"_
¢)
I
I
I
!
1
/
i
I
T
I
--
('-,1
g
.,.._
%
0
0
0
q
!
¢.)
c
¢.)
0
I"
J
81
cO
0
o
0
II
0
X
_6
lJ
o
i
o6
¢-q
o
L_
o
0
o
8
I
0
0
o
0
I"
I
C
82
0
q
q
o.
o
!
_f
0
I"
.16
.12
a, deg
O8
Cy
[]
2
0
ZX
6
0
0
-2
.O4
0
-.04
.O4
C n
0
-.04
.O4
C_
-.04
6
-12
- 8
- 4
0
4
8
2
#, deg
Figure
29.-
Lateral
characteristics
with
landing
gear
down;
R = 8.7×
106 , 8 f = 38 °,
83
.003
.002
Rn xlO -6
0
CnB
8.7
r-I 4.1
.001
0
o
-.001
-.002
_
c_e
- .003
-.004
0
4
8
12
16
a _ deg
Figure
84
30.- Stability
derivatives
Cnt 3 and C//3 versus angle of attack;
6f = 0 °.
\\ ?-
-
\
X
N
c-
X
0
II
0
O0
-o
0
m:
m=
o
l
i
f_
c
0
QL_
Q_ Q.m
O
O
LL
oJ
LL
L_
_f
li
_dd
0
Sap ' •
' alSUO qso_uMoo
85
.t2
/
8f, deg
0
[]
.10
0
2O
/
58
.O8
Ch o
.06
.04
.02
.,,d
0
0
4
8
12
16
a,deo
(a) Right aileron
hinge-moment
coefficient
versus angle of attack
for three
flap angles, 8a R = 0 °.
l
Figure 32.- Hinge-moment
coefficients;
R = 8.7X ! 06 .
,
K_
I:
!
L
I
86
_=
m
i
=
.10
Cha
I
0
a, deg
0
[]
O
8.4
14.6
8f,deg
0
.05
[
(,'_
0
.15
0
F!
O
e, deg
.4
8.7
2.9
8f, deg
38
.10
Cho
.05
]
0
- .05
15
10
5
0
-5
-I0
-15
8a,deg
(b)
Right
aileron
hinge-moment
coefficient
Figure
versus aileron
(bottom).
32.- Continued.
angle for $f=
0 ° (top)
and 6i"= 38 °
87
0
OJ
II
0
0
!
o_
.e
0
o
E
0
E
_b
o
o
c_
8..
o
0,,I
0
o.
I
I
I
g
88
i_r) l
o
.10
.05
_'_
8f ,deg
0
0
0
,',,
d _)g
0
14.6
Che , left
\
l
- .05
N
-.I0
.15
)
8f,deg
38
Che , right
-t0
-5
0
5
I0
15
Be , deg
(d) Elevator
hinge-moment
coefficient
versus elevator
Figure
angle for 6f = 0 ° (top)
and 5f = 38 ° (bottom).
32.- Continued.
89
.3O
0
[]
,,, deg
0
B
•20,
.I0
L
0
Ch r
-:10
-. 20
-.50
-30
-20
-I0
0
I0
20
8r, deg
(e) Rudder hinge-moment
coefficient
Figure
90
versus rudder angle at two angles of attack; 6f = 0°.
32.-
Continued.
30
o
O
II
tO
c_
¢.D
"3
a)
-0
°_""
.,=-i
o
_.
I".I
O
O
O<>
.d
c_
,===_
el)
o
O
"_
i
c,i
a0
e,D
,,,=_
[.r.,
o
/
O
,b
o
I
CD
• ,,.-.t
..o
tO
o
eq
o
.-.;
o
o
OI
Od
o
O
I
¢..
NASA-Langley,
19'71
--
1
A-3135
,.el
95
=
l-
oJ
O
o
I
("4
O_
o
°_._
o
x_
-
&
OOOD
OJl_O
..=
q..
_o ODO
O
O
N
r
oJ
O
o
O
N
O
I
r
(.)
92
o.
I
I
.2O
,B, deg
.I
8f,
deg
0
-8
0
[]
-8
58
0
A
-16
58
-16
0
0
]hr
-.10
\
\\\
-.20
\
1
-'3930
-20
-I0
0
I0
20
30
8 r , deg
(0 Rudder
hinge-moment
coettlcient
versus
rudder
Figure
angle
32.-
at two
sideslip
angles
and
two
flap
angles;
c_ = 0 °.
Continued.
91
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SPECIAL PUBLICATIONS:
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information
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on technology
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interest in commercial and other non-aerospace
applications. Publications
include Tech Briefs,
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Scientific and
under a NASA
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to existing knowledge.
Technology
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Details on the availability of these publications may be obtained from:
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If Undeliverable
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published in a foreign language considered
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knowledge.
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SPACE
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States shaft be
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thereo[."
TECHNICAL
technical information
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complete, and a lasting contribution
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POSTMASTER
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o[
edge o[ phenomena
in the atmosphere and space. The
shall provide/or
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o[ inJormation concerning
its acth,ities and the results
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