19740004604
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NASA
CR-132339
NASA-CR-
132339)
DESIGN
AND
FABRICATION
N7-12717
OF
AN
AEROELASTIC
FLAP
ELEMENT
FOR
A
SHORT
TAKEOFF
AND
LANDING
STOL)
AIRCEAFT
MODEL
Final
Boeinq
Commercial Airplane
Unclas
Co,
Seattle
-3-6
p
HC
14,0
CSCL
1C
G3 2
24431
DESIGN
AND
FABRICATION
OF
AN
AEROELASTIC
FLAP
ELEMENT
FOR
A
SHORT
TAKEOFF
AND
LANDING
(STOL)
AIRCRAFT
MODEL
Final
Report
By G. W. Belleman and R. R. June
Prepared
under
onitract NASI-11767
by
BOEING COMMERCIAL AIRPLANE
COMPANY
Seattle,
Washington
9$124
for
Langley
Research
Center,
Hampton,
Virginia
NATIONAL
AERONAUTICS
AND
SPACE
ADMINISTRATION
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1. Report
No. 2.
Government Accession No.
3. Recipient s
Catalog
No .
NASA CR-132339
4. Title
and
Subtitle
5.
Report
Date
DESIGN AND FABRICATION OF
AN
AEROELASTIC FLAP
December,
1973
ELEMENT
FOR A
SHORT
TAKEOFF AND
LANDING (STOL) 6. Performing Organization Code
AIRCRAFT
MODEL
7. Author s)
8. Performing
Organization Report
No.
G.
W.
Belleman and
R. R.
June
D6-41253
10. Work
Unit
No .
9. Performing
Organization
Name and Address
Boeing
Commercial Airplane Company
P.
O.
Box
3707
11
Contract or
Grant
No.
Seattle,
Washington
98124
NASl-
1l767
13.
Type
of
Report
and
Period
Covered
12.
Sponsoring
Agency
Name
and
Address
Contractor
report
National
Aeronautics and Space
Administration
Washington,
D.C.
20546
14
Sponsoring
Agency
Code
15. Supplementary Notes
The
work was monitored by Mr.
Marvin D. Rhodes of
the Structures
Division of NASA Langley
Research
Center.
16.
Abstract
A flap element typifying a
third
element in the
flap
system
of a short takeoff
and
landing aircraft
was designed,
fabricated, and
instrumented. It
was delivered
to NASA
for
flight-simulated
testing. The
flap element
was
aluminum
skin-stringer-rib
construction
with
adhesive
laminated skins.
17.
Key
Words (Suggested by
Author s) )
18. Distribution
Statement
Third
flap
element
STOL
Flight-simulated
tests
Unclassified-unlimited
19.
Security
Classif.
(of this
report)
20.
Security Classif.
(of
this
page)
21 No.
of Pages,
22. Price*
Unclassified
Unclassified
3 6,
*
,
For
sale
by the National
Technical
Information
Service,
Springfield,
Virginia
22151
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DESIGN
AND
FABRICATION
OF
AN
AEROELASTIC
FLAP
ELEMENT
FOR
A
SHORT
TAKEOFF
AND
LANDING
(STOL)
AIRCRAFT
MOI)EL
By
;.
.
Belleman
and
R.
R.
June
SUMMARY
A
flap
element
typifying
a
third
flap
element
of
a short
takeoff
almd
landing
(STOL)
airplane
was
designed
and
fabricated.
The
purpose
of
this
program
was
to
provide
NASA
with
a
representative
aircraft
part
which
could
be
evaluated
by
flight-simulated
tests.
The
tests
are
expected
to
provide
an
insight
to
the
loading
and
dynamic
response
of
an
airfoil
section
subjected
to
a
flight
airstream
and
jet
engine
exhaust
impingement.
INTRODUCTION
The
externally
blown
flap
concept
being
considered
in
preliminary
designs
of
a
short
takeoff
and
landing
aircraft
uses
special
high-lift
devices
to
decrease
runway
requirements.
One
element
which
is
critical
to
the
successful
operation
of
this
concept
is
the
trailing-edge
flap.
The
location
of
this
element
exposes
it
to
aerodynamic
loads,
elevated
temperatures,
jet
engine
exhaust
impingement.
boundary
layer
turbulence,
and
vortex
shedding.
The
object
of
this
program
is
to
design
and
fabricate
a
representative
third
flap
element
an d
provide
it
to
NASA
for Ilight-simulated
testing.
The
tests
will
allow
an
assessment
of
the
flap
response
to
the
flight
environment.
_I
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SYMBOLS
a
panel
length,
mm
(in.)
b
panel
width,
mm
(in.)
B
bar
width, mm
(in.)
C distance
from
neutral axis
to
outer
fiber, mm
(in.)
D
bar
thickness, mm
(in.)
ET Modulus
of elasticity,
tension,
GPa,
(psi)
EC
Modulus
of elasticity,
compression,
GPa
(psi)
Fe
stringer
stress,
MPa (psi)
Ftu
allowable
tensile
ultimate
stress,
MPa
(psi)
Fty
allowable
tensile
yield
stress,
MPa
(psi)
Fcy
allowable
compression
yield
stress,
MPa
(psi)
Fbru
allowable
bearing
ultimate
stress,
MPa
(psi)
G
modulus
of
rigidity,
GPa
(psi)
1
moment
of
inertia,
PM
4
(in.
4
)
J polar
moment of
inertia,p
M
4
(in.
4
)
K
2
torsional
shape
constant
L
bar
length,
mm
(in.)
M
bending
moment,
N-m
(in.-lbf)
t
panel
thickness,
mm
(in.)
2
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AT
temperature
differential,OK
(OF)
We
effective
skin
width, mm
(in.)
a
coefficient
of thermal expansion
6
deflection,
mm (in.)
e
expansion-contraction,
mm
(in.)
0
angular
rotation,
rad
DESIGN CONCEPT
The
overall
arrangement
and
shape
of the
third-element
flap
was provided
by
NASA.
This
arrangement
is shown
in figure
1. The
design
loads
and
airfoil
coordinates
were
also provided
by
NASA.
The
basic design
considerations
were
the
simultaneous
application
of a 36-m/sec
(70
kn)
airstream
and
exhaust
impingement
of a
General
Electric
TF34
turbofan engine.
The
design
life
is to
be
1000 hr,
an
environment
of 155
dB
noise
level (table
1)
and
temperatures
to 4220
K
(3000F)
(table 2).
It
was further specified
that
the
end attachment
sections
have
a bending
stiffness
of
approximately
3.18
mm/4.45
kN
(0.125
in./1000 lbf)
applied cantilever load and
3.29
rad/MN-m
(0.327
x
10-6 rad/in.-lbf)
torsional
stiffness.
This
was
accomplished
by
introducing
the
flap-end
reaction
loads
into a
solid
aluminum
bar
50.8
mm
x 203
mm
(2 in.
x 8
in.), approximately
670 mm
(26.4
in.)
long
(table 3).
These
end
attachment
sections
were
designed
but
not
fabricated.
They
will
be
fabricated
by NASA.
The end
bars
were
designed
with
a
nonstructural
airfoil-shaped
boilerplate
shroud
to
maintain
airfoil
continuity.
Schematics
of the
arrangement
are
shown
in
figures
2, 3, and
4.
The
entire
flap,
except
for
the
attachment
fitting,
was constructed
with
2024-T3
aluminum
alloy.
This alloy
was
selected
for the
following reasons:
(1)
the
flap
was basically
stiffness
designed;
(2) the
alloy
retains
95% of
its modulus
at
4220K
(3000F); (3)
it
retains
a
high
percentage
of its
tensile
ultimate
and
compression
yield
after
1000
hr
at
(300
0
F);
(4) it
is a
tough
alloy with
favorable
ductility,
and
(5)
it withstands
sonic
environments
better
than
alloys
such
as 7075-T6.
Material
properties
of 2024-T3
aluminum are
listed
in table
4.
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The
skins
were
laminated from
three
layers
of
0.64
mm
(0.025
in.) 2024-T3
sheet. The
removable
leading
edge
is
a solid
sheet
aluminum
1.8 mm
(0.071
in.).
The
spars
and
ribs
were
built-up,
riveted,
formed
aluminum,
as were
the
leading-edge
formers
and
trailing-edge
riblets.
The
end
ribs
were
machined
out
of aluminum
plate
with
a profile
mill. A
simple aluminum
sheet template
was
used
for the airfoil
profile.
The stringers
are
extruded
aluminum.
The
basic
structural
concept
is
a
conventional-type
skin-stringer
construction.
The
flap element
evolved
as
a
two-spar
box
construction
with
all
skin
effective
in
torsion
and
bending.
The
laminated
skins were
autoclave
bonded with Metlbond
329-7 adhesive.
Manufacturer data
on
this material
indicates
a 90%
strength
retention
after
aging
1000
hr
at 450
0
K (350*F).
All
metal-to-metal
contact
surfaces
were
coated
with
Products
Research
and
Chemical
Corporation
PR
1750-A-4.
This
material
is
a polysulphide
composition
that
is
used as
a faying
surface
sealant.
Boeing
data
indicates
structures
assembled
this
way
are better
in sonic
environments
to the
extent
of about
1 dB.
The
total
gross
mass
of
the completed
part
was
45.5
kg
(96 lb).
The
net
mass minus
instrumentation
was 38.6
kg
(85 lb).
The
bottom
skin
of
the
flap
element
was
instrumented
with
22
axial
strain
gages
and
tw o
rosettes.
Eight
accelerometers
were positioned
inside
the
flap.
An
analysis
of the element
was
conducted
to
assess
the
structural
adequacy
of the
part.
The
aerodynamic
loadings
on
the
part
were
provided
by
NASA
and
result
in
the
loads
shown
in
figures
5,
6,
and
7. Figures
8
to
show
the
physical
properties
of
the
flap
element.
Figures
12 and
13
show
the
Boeing
design
curves
for
sonic
fatigue
life
that
were
used
to check
part
details.
The
fabrication
sequence
was
straightforward.
The
skins
were
laminated
and
roll
formed
to
contour.
The
ribs
and
spars
were
assembled
in
a simple
jig
tool. The
upper
skin
was
riveted
on
and
the
instrumented
bottom
skin
was
blind
fastened
with
bulbed
Cherrylock
rivets.
Lastly,
the
removable
leading
edge
was
installed
with
bolts
into
nutplates.
Figures
14
to
24
show
the
subassembly
and
some
details
of the
construction.
4
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ANALYSIS
The
design
loads
were
provided by
NASA.
The flap
is supported
at each
end by pinned
connections and is therefore
analyzed as a simply supported
beam (tables 5
and 6). Various
members were
checked for their sonic fatigue
capability
(table
I).
using Boeing-derived
test
data.
Boeing
Commercial Airplane Company
P.O. Box
3707
Seattle, Washington
98124, December
14.
1973
5
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TABLE
1.-SONIC
CHECK
GEOMETRY
Rib
spacing,
a
=
10
in.
Stringer
spacing,
b
=
8 in.
(maximum)
Skin
gage,
t
=
0.075
in.
Spar
gage,
t
=
0.063
in.
CALCULATIONS
Skin:
b
8
-
-
=
107
t
0.075
=
1.25
b
then,
from
figures
12 and
13,
162
- 1.7
=
160.3
dB
Spar:
b
4
.. . 64
t 0.063
a
=
2.5
b
then,
from
figures
12 and
13,
163
4.7 =
158.3
dB
-
Trailing-edge
skin:
b
11
t 0.063
a
-=
1
b
then,
from
figures
12
and
13,
156
0
156 dB
PRECEDTING
PAGE
BLANK
NOT
FILMED
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TABLE
2.-THERMAL
EXPANSION
CAL CULA
TION
Assume
third-element
flap is
fixed
in expansion-contraction
length
by steel
support
structure.
Also
assume,
as worst
case,
that
third
element reaches
3000F uniform
temperature;
support structure
remains
at
ambient
75*F.
Then, maximum
thermal
expansion
stress is
e
= cATE
=
12x
10 -
6
)(300
75)(10.1
x 106)
=
27,300 Ibf/in.
2
Maximum
bending stress is
a
e
= 7218
Ibf/in.
2
The
total
stress is then
27,300
+ 7218
= 34,518
Ibf/in.
2
(237.97
MPa)
Since
yield
stress
is
35,000
Ibf/in
2
(241.29 MPa),
we have
excessive
thermal
stress.
The
expansion
relief
required
to alleviate
thermal stress
is
E= xATL
=
(12 x
10
6
)(225)(80)
=
0.216 in.
(5.49 mm)
This
relief
is
maximum
due to
conservative
assumptions
NOTE:
Calculations
were
performed
in customary
units
and
answers
converted
to
SI units.
8
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TABLE
3.-END
FITTING
STIFFNESS
Where
desired
end
fitting
stiffnesses
are
Torsion:
8 =
0.372
x 10-6
rad/in.-lb
Bending:
6
=
0.125
in./1000
lb
end
load,
the
length
of a
solid
aluminum
rectangle
with assumed
cross-sectional
dimensions
of
2
by
8
in.
is
calculated
from
Strength
of
Materials
by
F. R.
Shanley,
McGraw-Hill
Book
Company,
Inc.,
p.
509,
as
follows:
K
2
ML
0-
BD
3
G
where
D/B
=
0.25
and
K
2
=
3.6.
Then,
substituting
desired
torsion,
0
=
0.372
x
10
. 6
=
3.6) 1) L)
8) 2)3 4
x
106)
Solving
for
length,
L =
26.4
in. (670.6
mm)-
length
of
a 2-
by
8-in.
(50.8-
by
203-
mm)
aluminum
bar
required
for torsional
stiffness
Checking
bending
deflection
for this
length
PL
3
3El
(1000)(26.4)
3
(3)(10.1
x
106)
(8)(2)3
12
=
0.114
in.
(2.9
mm)
vs desired
0.125
in.
(3.18
mm)
NOTE:
Calculations
were
performed
in
customary
units
and
answers
converted
to
SI units.
9
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TABLE
4.-MA
TERIAL
PROPER
TIES
2024-T3
aluminum
aged 1000
hr
at
4220 K
(3000
F) and tested
at 4220
K
(3000 F)
Fts
=
333.71
MPa
48,400
psi)
Fty
=
279.24
MPa
40,500
psi)
Fcy
=
241.32 MPa
35,000
psi)
Fbru
=
558.48
MPa
81,000
psi)
Fbry
=
434.37
MPa
63,000
psi)
E
t
= 69.64
GPa
10.1 x
106
psi)
Ec
= 71.02
GPa
10.3
x
106 psi)
G
=
26.48
GPa
(3.84 x 10
6
psi)
*Reference:
Boeing
Design
Manual,
D-5000,
volume
84A2
10
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TABLE
5.-MAXIMUM
BENDING
STRESS
IN
SKINS
GEOMETRY
t t
135.7
kN
3050
Ibf)
137.9
kN
3100
Ibf)
FLAP
ELEMENT
PROPERTIES
Mmax
=
75,465
in.-Ibf
at station
117
(see
fig.
7)
I
=
23
in.
4
(see
fig.
10)
Cmax
=
2.6
(tension
side)
cmax
=
2.2
(compression
side)
CALCULATIONS
Mc
75,465)
(2.6)
23
= 8530
Ibf/in.
2
(58.81
MPa)(tension)
75,465) 2.2)
23
=
7218
Ibf/in.
2
(49.77
MPa)(compression)
NOTE:
Calculations
were
performed
in customary
units
and
answers
converted
to
SI units
11
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TABLE
6.-EFFECTIVE
SKIN
WIDTH
AND
STRINGER
SPACING
GEOMETRY
t
SWe
0.
9 1
We
CALCULATIONS
W
e
= 0 85t
E
e
fc
Where
f =
stringer
stress,*
and
assuming
that
maximum
compressive
stress
equals
stringer stress,
then
from
table
5)
fc
=
7218
Ibf/in.
2
and
We =
0.85) 0.075)
1
e
7218
= 2.41
in.
Total
effective
skin
width
=
2
We
+ 0.9
=
(2)(2.41)
+
0.9
=
5.72
in.
(145.3
mm)
For
all
skin
to
be
effective,
Stringer
spacing
5.72
in.
(145.3
mm)
Actual
stringer
spacing
=
5 in.
(127
mm)
Therefore,
all
skin
between
stringers
is effective.
Note:
Calculations
were
performed
in customary
units
and
answers
were
converted
to
SI units.
*Reference
Boeing
Stress
Manual,
D6-22695,
p.
11-11
12
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Leading
edge
3.70
m 145.69
in.)
3.81
railing
edge
38
150.
9
in.)
1.326
rad
FIGURE
1.-DIMENSIONAL
ENVELOPE
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End
fitting
2.108
m 83.0
in.)
End
fitting
Test
part
Aluminum
bar
Aluminum
bar
FIGURE
2.-GEOMETRICAL
CONCEPT
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Expansion
link
End-fitting
shroud
Test
part
Aluminum
bar
End
fitting
Support
box
Aluminum
bar
FIGURE
3.-A TTA CHMENT SCHEMA
TIC
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Front
spar
Leading
edge
Rear
spar
Trailing-edge
riblet
Stringer
Rib
chord
RIB CROSS
SECTION
FIGURE
4. SPANWISE
A RRANGEMENT
16
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160
140
25
-
120
o 0
S100
E
10
20
?
I
I
I
1
.5
0
.5
1
Sta
Distance
from
engine
C
,
m
Sta
154.8
82.8
I
I
I
I
I
I
I
I
40
30
20
10
0 10
20
30
40
Distance
from engine L
, n.
FIGURE
5.-AERODYNAMIC
LOADING
ON
THIRD-ELEMENT
STOL
FLAP
REF
NASA
A0332
17
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15
3000
10
2000
-
1000
-
5
z
0
0
-1000
-5
-2000
-10
-3000
-15
I
I
I
0
.5
1
1.5
2.0
Distance
from
left
end,
m
I
I
I
I
I
I
I
I
I
0
10
20 30
40
50
60
70
80
Distance from
left end, in.
FIGURE
6.-SHEAR
DISTRIBUTION
THIRD-ELEMENT
STOL
FLAP
18
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100-
10-
80-
75,465
in.-lbf
maximum
-.
8
60
E
0
z 6
0
40
w
0
4
4
20-
2
0
0
0
.5
1
1.5
2.0
Distance
from
left
end,
m
0
10
20
30
40
50
60
70
80
Distance
from
left
end,
in.
FIGURE
7.-MOMENT
DISTRIBUTION,
THIRD-ELEMENT
STOL
FLAP
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6
4000
3000
3Test
part
CN
N
E
.S
E
4
C
3u
2000-
2
1000
1
0-
0II
4.0
3.5
3.0
2.5
2
Spanwise
station,
m
160 150
140 130
120
110 100 90
80 70
Spanwise
station,
in.
FIGURE 8.-AIRFOIL
SKIN AREA,
1.91-mm 0.75-IN.) SKINS
20
-
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.08
120
-
100 -
.06
ev
CE
Test
part
.
80 -
E
60 04
o
60
0
40-
.02
20
0
I
I
I
I
4.0
3.5
3.0
2.5
2
Spanwise station, m
I
I
I
I
I
I
I
I
I
160
150
140
130
120
110
100
90
80
70
Spanwise station,in.
FIGURE
9.-AIRFOIL
CROSS-SECTIONAL
A
REA
21
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60
25
50
sio
olar
20
of
inertia
40
15
E
30-
10
Moment
f
inertia
20
5
10
0 0I
I
I
I
2.0
2.5
3.0
3.5
4.0
Spanwise
station,
m
80
90
100
110
120
130
140
150
160
Spanwise
station,
in.
FIGURE
10.-TORSIONAL
AND
BENDING
MOMENT
OF
INERTIA
22
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.28
-
7
.24
6
.20
5
E
._
.16
E
4
o
.2
S.12-
3
.08
2
.04
1
0
0
I
I
0
.5
1 0
1 5
2.0
Distance
from
left
end,
m
I
I
I
I
I
I
I
I
0
10
20
30
40
50
60
70
80
Distance
from
left end,
in.
FIGURE
11. FLAP
DEFLECTION
23
-
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172
168
164
160
=
156
-
0
05 152
U-
148
144
140
60
80
100
120
140
160
180
200
b t
FIGURE
12.-SONIC
FA
TIGUE
DESIGN
SKIN/STRINGER
STRUCTURE
1000-HR
LIFE
AT
SOUND
24
-
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6
5
4
2-
1-
a/b
FIGURE
13. LEVEL
CORRECTION
FACTOR
FOR
RECTANGULAR
PANELS
IN SONIC FATIGUE
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FIGURE
14. COMPL
ETED
FLAP SSEMBL
Y
-
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FIGURE
15. FLAP
FINAL
ASSEMBLY
-
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FIGURE
16. FLAP
SUBASSEMBL
Y LOWER
SURFACE
S
5
5
-
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FIGURE 17. FLAP
SUBASSEMBL
Y TOP SURFACE
-
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FIGURE
18. FLAP
SUBASSEMBLY
AND
INSTRUMENTED
LOWER
SKIN
S
S
S
-
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FIGURE 19. INSTRUMENTATION
LA
YOUT
-
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FIGURE
20. INSTRUMENTATION
WIRE
ROUTING
THROUGH
FORW RD
SP R
-
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FIG
URE
21. A
CCEL
EROMETER
INSTAL
LA
TION
-
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F/G
URE
22. FLAP A TTACHMENT
FITTING
-
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F/IGURE
23.
TYP/CAL
INSPAR
RIB
ARRANGEMENT
5
-
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FIGURE
24. TRAILING EDGE
RIBLET
ATTACHED
TO RE R
SP R AND
TOP SKIN
