Natural Laminar Flow Experiments on Modem Airplane

NASA
Technical
Paper
225.6
June
!984
Natural
3
Flow
Experiments
on Modem
Airplane Surfaces
I"
Bruce
J. Holmes,
Clifford
and
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NASA
Technical
Paper
2256
1984
Natural
Laminar
Flow
Experiments
on Modern
Airplane
Surfaces
Bruce J. Holmes
Langley
Research
Hampton,
Virginta_
Clifford
Kentron
Hampton,
Long
Langley
Hampton,
Z-
,+
Center
J. Obara
International,
Inc.
Virginia
P. Yip
Research
Center
Virginia
NASA
Nahonal Aemnauhc:-,
and Spa(:+ _ Admmc-qtat_oc,
Scientific
Information
and Technical
Branch
t
................................................
-';, _ ..... _
.....%2._:o__.=.,
.............................................................
.+, ..;:2
Use
constitute
expressed
of
trademarks
an
or
official
implied,
or
names
of
endorsement
by
the
National
manufacturers
of-such
products
Aeronautics
in
this
or
and
report
does
not
manufacturers,
either
Space--Administration.
m_
®,
I
CONTENTS
INTRODUCTION
SYMBOLS
REVIEW
................................................................
AND
OF
AIRPLANE
ABBREVIATIONS
PAST
NATURAL
LAMINAR
DESCRIPTIONS
Airplanes
AND
FLOW
. .........
RESEARCH
CORRESPONDING
. .................
....................
EXPERIMENTS
......
. ......
. ........
...
. .......
. .......................
......................................................................
Rutan
VariEze
Rutan
Long-EZ
Rutan
Laser
Gates
Learjet
Cessna
24R
Bellanca
Beech
Testing
Other
Model
7
. .........
. ......................................
Racer
....................................................
9
28/29
Longhorn
9
Centurion
. ...............
9
. .......................
........................................................
........
of
of
boundary-layer
boundary-layer
procedures
10
................................................
detection
transition
transition
10
..................
.....................
.........
.... . ................................................
......
99
, ..................................................
.............
chemical
detection
. ..........................
. ......................................................
.....................................
gloves
testing
7
8
Procedures
Acoustic
5
...............
Skyrocket-II
Sublimating
3
8
Sierra
T-34C
I
................................................................
Biplane
P-210
Beech
RESULTS
..................
.....
10
11
11
.
/Transition
Effect
Flight
locations
of
.........................................................
transition
Experiments
on
canard
12
.........................................
12
.............................................................
13
Rutan
VariEze
................................................................
13
Rutan
Long-EZ
................................................................
14
Rutan
Laser
Gates
Learjet
Cessn_
Biplane
24R
Bellanca
Beech
gloves
............
Locations
Effects
of
Precipitation
Effects
of
Fixed
Propeller
Waviness
Sweep
Insect
Debris
CONCLUSIONS
APPENDIX
II
16
17
........................................................
17
19
........................
20
...................................
...........................................................
and
Cloud
Particles
Effects
...................................................
.....................
........................
......................................
WAVINESS
20
...................................
21
....................................................
Contamination
SURFACE
16
...........................................................
22
...........
............................
ON
RESEARCH
MODELS
J
• i|
±ii
. .............
...........................
23
24
24
....................................................
...........................................
-
15
...........................................
.......................................................
Transition
Slipstream
...............
Effects
Longhorn
.............................................................
skyrocket
Transition
....................................................
28/29
Centurion
Sierra
T-34C
DISCUSSION
Racer
Model
P-210
Beech
_4
fixed
25
.....
........
26
. .......
28
• imlm,,--T'--_--r'_-
+_''_m'_T
+
_
.
_+_
REFERENCES
TABLES
FIGURES
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_
SUMMARY
Flight
and
various
lifting
between
0.63
lifting
surface
selected
and
and
×
to
wind-tunnel
106
provide
waviness,
on
typically
for
heights
face
waviness.
were
,bserved.
trol
resulted
of
surface
Large
from
waviness
the
laminar
of
the
airplanes
exceeded
of
boundary
low-altitude
The
laminar
flow
tent
certain
as
results
exist
flow
a
taken
and
that
modern
the
by
on
this
with
indicate
that
airplane
the
and
is
rain
in
two
through
importance
modern
of
smooth
of
more-durable
than
coneffects
changes
regions
surfaces
surcriteria
flight
the
for
due
mea-
allowable
stability
significant
behavior
production
of
pitch-trim
procedure
peak
transition
Simulated
indicate
testing
pressure
for
at
were
roughness
contamination
and
and
transition
on
predicted
observed
observations
0.7,
tested
None
transition.
on
numbers
significant
tested.
nose-down
to
observed
effects
spanwise
was
0.1
calculated
performance
boundary-layer
p_actical
from
or
conducted
Reynolds
airplanes
The
empirically
forced
flight
a whole
free
surfaces
transition
These
standard
as
the
stick-fixed
effect
clouds.
tests
of
been
unit
from
The
discernible
flight-measured
caused
No
liquid-phase
frames.
in
63 ° .
measured
No
consistent
laminar
layer
tested.
fixed-transition
any
results
changes
loss
the
at
numbers
to
airframes.
of
on
Mach
conditions,
involved.
observed
have
airplanes
0°
production-type
downstream
Experimental
at
from
skin
experiments
several
ft -I,
angles
conditions
were
flow
of
106
stiff
modern
test
x
sweep
occurred
the
on
on
3.08
relatively
waviness
sured
and
smooth
locations
surface
ft -I
laminar
surfaces
leading-edge
locations
to
natural
nonlifting
air-
natural
and
persi§-
previously
expected.
INTRODUCTION
in
(NLF)
decades.past,
was-souqht
methods
of
wing
In
and
recent
current
at
two
milled
unit
than
20
tively
to
ate,
methods
skins,
the
lower
and
less
x
numbers
106 .
Therefore,
These
(wing
for
lower
loadings
modern
and
aspect
significant
that
sailplanes.
The
relative
to
lower
to
produce
the
1.5
x
106
ft -1
of
numbers
ratios
and
operate.
and
are
of
The
larger)
sur-
include
Reynolds
numbers
of
airplanes
Reynolds
surface
the
and
these
shorter
from
trend
at
which
crulse
numbers
quality
much
less
is
airfoil
the
com-
modern
unit
chord
have
aerodynamic
second
pro-
materials
Most
at
from
on
operations
techniques
NLF-compatible
result
performance
and
modern
early
therefore,
construction
skins.
chord
surfaces;
increased
production
These
However,
rela-
chord
higher
cruise
airplanes.
-
the
both
achievement
Reynolds
is
by
of
wavy
airframe
aluminum
airplanes
than
the
for
laminar'flow
range.
fabrication
modern
waviness.
bonded
range
for
airplane
First,
natural
and
rough,
airfoils
potential
and
speed
produced
in
NLF.
the
Reynolds
aircraft
itated
offer
business
altitudes
It
is
trends
to
high-performance
easy.
lengths
NLF
major
regions-of
airplane
maintenance
roughness
aluminum
extensive
6f.laminar-flow
achieved.
favorable
critical
of
increasinq
and
are
methods
without
favorable
of
years,
which
posites,
achievement
means
application
was
never
fabrication
faces
a
manufacture
the
successful
duction
aircraf%
developed
the
as
chord
most
power
smooth
NLF
has
been
a
achievement
Reynolds
numbers
airplanes,
complex
of
and
practical
laminar
(<4
by
x
the
reality
flow
106
use
on
for
typically)
of
one
sailplanes
composite
at
category
has
which
been
they
of
faciloper-
construction
shapes.
®
,
This
ments
report
recently
ber,
and
presents
sweep-angle
construction
These
of
environments.
those
to
meet
the
described
is,
the
Examine
2.
the
Observe
3.
that
NLF
the
the
favorable
issues
no
in
typical
conducted
on
for
and
flight
experi-
preparation
production-quality
by
NLF
gra-
operating
preflight
modification
requirements
pressure
achievability
_ecent
in
num-
airframe
of
these
difference
experiMach
practical
surfaces
were
received
waviness
in
flight
number,
modern,
distinguishes
is
experiments
which
Measure
experiments
5.
Document
6.
Observe
nation
7.
Investigate
Eiqht
large
locations
wings,
fuselage
and
propeller
filling
(minor
and
sanding
exceptions
are
were
conducted
numbers
and
with
the
following
Or
the
effect
the
equipped
with
additional
and
Based
on
a
airplane
results
appreciation
surfaces
transport
flight
were
flow
airfoil
clouds
at
aircraft.
testing
the
used
Model
on
of
VariEze
these
for
chord
The
(fixed
the
flight
28/29
and
The
vertical
and,
where
due
airplane
laminar
on
transition.
to
perfor-
flow.
sweep
an
on
NLF
fligh
t
(the
Rutan
implications
and
Longhorn);
a
of
the
other
flight.
two
VariEze
and
(the
Rutan
low-wing
high"
L0ng-EZ);
Biplane
airplanes
single-engine
general-
a Beech
conducted
flight
contami-
in
experiments:
airplane,
were
spanwise
airfoil
two
high-wing
eighth
experiments
T-34C,
to
was
provide
experiments.
The
wind-
airplane.
and
achievability
Reynolds
procedures,
surfaces
and
transition
on
on
in-the
II);
which
findings
the
laminar
predictions.
behavior,
leading-edge
Centurion).
gloves
on
surfaces),
empirical
configurations
Skyrocket
only
aerodynamic
tractor-propeller-confiqurati0n
P-210
support
used
the
new
on
_%rough
Learjet
Bellanca
(Cessna
to
laminar
slipstream
insect
types
biplane,
numbers
horizontal
airfoil
with
contamination
laminar-flow
data
propeller
of
(Gates
Mach
control.
pusher-propeller
experiments
of
rain)
flight
nature
jet
Sierra
airplane
blade
the
airplane
a business
and
transition
loss
of
fairinqs,
wing
canard,
24R
of
variety
wheel
practical
effects
of
the
attachment
line.
different
aviation
a
large
spinner
total
of
a
nose,
and
effect
the
of
on
simulated
stability
negative-stagger
Racer);
of
grit
Reynolds
Surfaces.
measured
effect
and
Observe
increasing
airframe
correlate
the
4.
aspect-ratio
of
transition
mance,
muter
for
address
and
Reynolds
of
airframe
1940's
recent
production
transition
designs,
to
wind-tunnel
effect
for
possible,
ern
and
or
and
stabilizers,
provides
factor
contour
(including
tunnel
smoothness
requirements
designed
surfaces
flight
flow
(Beech
meet
1930's
The
on
the
wind-tunnel
maximum
objectives:
I.
a
to
NLF
the
text).
Full-scale
specific
the
airfoil
in
which
several
determine
production-quality
tested.
that
over
were
on
of
to
significant
of
surfaces
surfaces,
results
NASA
fail
NLF
The
from
the
ranges
expeciments
maintainability
of
by
techniques
dients.
ments
the
conducted
numbers
of
further
wind-tunnel
and
experiments,
maintainabillty
representative
these
results
studies
are
of
to
also
this
of
business
further
paper
NLF
on
and
modcom-
airplane
discussed.
2
®
q
SYMBOLS
a
airfoil
avg
average
BL
butt
b
airplane
wing
span,
cD
airplane
drag
coeffleient
Cd -
eL
CL
mean-line
ABBREVIATIONS
drag
airplane
coefficient
trimmed
lift-curve
ft
slope,
lift
airplane
pitching-moment
Cp
pressure
coefficient,
c
local
coefficient
chord,
coefficient
(Pt
-
(referenced
to
c/4)
P)/q_
ft
aerodynamic
chord,
diameter,
fuselage
coefficient
deg "I
Cm
propeller
-
lift
section
_---FS
_
designation
c£
d
_
line
__section
mean
AND
r
in.
ft
station
h
indicated
hd
density
altitude,
J
advance
ratio,
LoE'
leading
edge
L.S.
lower
M
Mach
n
propeller
Ps
vapor
P
static
Pt
total
q
dynamic
R
free-stream
double-amplitude
wave
height,
in.
ft
V/nd
surface
number
rotational
pressure,
pressure,
pressure,
pressure,
unit
speed,
mm
rps
Hg
psf
psf
psf
Reynolds
number,
ft -I
3
®
i
Rc
chord
Reynolds
R0
attachment-line
eq.
1
number
I
boundary-layer
momentum
thickness
Reynolds
number
(see
(I) )
i
r
radius,
ft
S
lifting
surface
reference
s
surface
length,
in.
T
temperature,
t/c
wing
thickness
U.s.
upper
surface
/U e
free-stream
Vc
calibrated
or
velocity
true
airspeed
water
x/c
location
(x/h)
transition
Y
semispan
location,
z
vertical
dimension,
angle
airspeed,
layer
knots
or
mph
flow
field
and
indicator
errors
removed),
knots
knots
or
mph
local
chord
line
in
of
percent
location
attack,
boundary-layer
n
(local
boundary
mph
WL
e
in
airspeed,
indicated
6
ft
ratio
Vi
6
sq
°C
local-to-edge
V
area,
elevator
in
percent
length
ft
ft
deg
(relative
thickness,
deflection,
nondimensional
body
semispan
deg
to
longitudinal
reference
axis)
in.
(positive
trailing
edge
down)
position,
b/2
A
sweep
angl_,__deg
k
wavelength,
in.
Subscripts:
a
allowable
c
canard
L
lower
J.
• ,%
leading
edge
max
maximum
min
minimum
t
transition
U
upper
w
wing
free
location
stream
Notation:
--_
i.d.
inside
NLF
natural
laminar
oed.
outside
diameter
psf
,
diameter
pounds
....
flow
(force)
per
Square
REVIEW
The
use
on
achievement
for
airfoil
able
able
can
be
pressure
reduce
amounts
waviness.
of
protuberances
consideration
gradients
as
NLF
and
which
the
the
limit
and
gaps
by
conflicting
in
of
The
sition
swept
such
on
wings
at
transport
airplane
requirements
The
wings
for
will
be
contamination
kept
of
free,
(e.g.,
surface
of
is
understood
less
wing
rate
of
these
two-
and
about
for
or
ice),
Compared
the
from
conflicting
maintainability
is
of
of
with
sig-
distribution
pressure
vorti-
rapid
in
is not
presently
T-S
waves
affects
the
well
tran-
for
business,
commuter,
the
successful
design
gradient
of
design
instabilities.
with
NLF
critical
requires
_ounts
disturbances
phenomena
of
(crossflow
pressure
compatible
free-stream
wings
favorable
three-dimensiozLal
with
favor-
two-dimensional
w,rtices
to
on
influences
On
It
and
of
effect_
effects
pressure
less
growth
waviness
_hus,
the
above
interest
from
The
surface
disturbances
environment,
favor-
layers.
the
challenge
the
of
two-dimensional
These
other
crossflow
conditions
operating
of
boundary
of
runs
destabilizing
the
leading
edge.
crossflow
vortices
both
debris
sizes
of
meet
damage.
the
requirement
achieved
transition.
layer
between
technical
surface
an
insect
and
to
critical
conditions
The
of
in
turbulence),
NLF,
be
avoidance
maintenance
surfaces
free-stream
airplanes.
premature
boundary
laminar
growth
rapidly
falling
pressure
near
how
the
interaction
between
airline
to
its
is
layer.
of
to
flow
of
profiles.
three-dimensional
region
of
understood
or
effects
compromise
design
growth
layer.
boundary
velocity
long
growth
the
counteracting
the
the
the
lead
by
by- designing
limit
b_
laminar
govern
achieved
the
boundary
can
challenges
laminar
in
boundary-layer
waviness
steps
is
(_15 °)
which
principal
Natural
hand,
on
the
two
waves)
other
and
the
RESEARCH
today.
flow)
the
influences
are
angles
on
"protect"
surface
such
sweep,
in
sweep
stability
Similar
nificant
and
gradients
NLF
(T-S)
aggravated,
b0undary-layer
limited
small
LAMINAR-FLOW
airplanes
(accelerating
gradients
pressure
ces)
with
of
on
(Tollmien-Schlichting
waves
local
NATURAL
maintenance
gradients
disturbances
PAST
improvement
surfaces
pressure
T-S
and
performance
OF
foot
affecting
NLF
under
(e.g.,
the
the
that
of
the
surface
noise
and
achievability
wide
ranges
of
®
k
Reynolds
number_,
craft
generally
true,
Reynolds
ity
Math
configurations
meet
NLF
the
ating
drag
and
and
(2)
Past
research
airframe
airframe
of
waviness
and
(e.g.,
(refs.
faces,
Because
in
quality
little
laminar
flow
Heinkel
He.70
gloves
(see
matched
reported
successes
development
of
the
of
on
for
perception
that
surfaces.
Reynolds
This
number
fighters
on
make
laminar
the
In
which
NLF
would
be
was
2
×
106
applications
very
and
26)
of
of
that
the
period
surfaces
for
these
these
difficult
probably
have
to
the
to
surface
in
their
rough"
partly
even
the
War
such
from
exacerbate
achieve
World
attempted;
and
presented
been
by
twostrongly
conservative
stems
heightened
ft -I , for
is
may
may
guidance
was
provide
conservatism
conservatism
closely
initial
criteria
the
on
results.
allowable
Criteria
of
and
test
the
as
exten"
wings
typically
provided
disturbances
very
are
sur-
apparent
plywood
criteria
This
skin
(production)
was
well
of
at
experiments
were
determined
performance
as
exam-
excessive
techniques
wind-tunnel-model
tests
Close
been
widths
prepared
airfoil
that
metal
and
airfoil
gloves.
experiments
on the produc-
specially
sensitive
gap
production
these
were
5.)
unprepared
23,
was
the
have
flight
drag
exception
"stream
the
layer
the
surfaces.
past,
>
both
summary
where
to
riveting
general,
the
Rc
NLF
boundary
NLF
chapter
or
waviness
A
I,
certain
Development
perception
range,
early
oper-
Concerning
in
heights
surface
well.
tunnels
28).
high-speed
aircraft
1950)
achieved
from
21,
and
gloved
In
of
wind
under
(circa
step
smoothed
On
allowable
as
ref.
low-turbulence,
appendix).
in
and
heights.
manufacture
(ref.
the
production
unit
see
origins
problems'q
sible
and
for
practical-
surfaces
typical
shortcomings
single
locations
research
(also
for. the
development
13.
protuberance
27
guidance
reference
be
surfaces,
above,
the
18,
The
filled
prepared
wind-tunnel
reference
16,
the
criteria
NLF
is
as
the
conclusions
Previous
NLF
and/or
section
andsanded)
noted
11,
and
not
included
flow.
transition
of
in
air-
It
improves
production
research
(See
the
excessive
9,
predictions
three-dimensional
based
on
negative
time.
rivets).
location
laminar
in
I),
7,
and
protuberances
(filled
shortcomings
achieved
table
for
no
maintained
and
NLF.
concerning
achievement
could
reveals
experiments
(refs.
or
theoretical
The
ness
These
issues
practical
early
NLF
that
or dimpled
transition
I.
surfaces
in
of
of
for
surfaces
be
from
for
stringers,
specifically
prepared
of the
fabrication
resulted
sire
required
heights
table
positive
methods
and
press-countersunk
2 to 26)
in which
summarized
tion
ribs
excessive
profiles,
manner.
of
methods
NLF
Can
for
consensus
quality
of
(_)
flight
applications
critical
benefits
mixture
fabrication
between
joints,
a
production
those
the
twofold:
requirements
significant
surface
mass
ination
summary,
cost-effective
left
A
potential
maintenance
laminar-flow
a
conditions,
the
in
are
waviness
in
questions.
the
In
can
environments
these
ease
reduction
roughness
conditions,
_at
decreases.
for
meteorological
characterize
however,
number
of
numbers,
_%ich
respon-
on
modern
relatively
high
II high-performance
free-stream
conditions
imperfections
and
insect
contamination.
_i._J
Even
the
when
subject
ability.
Past
sition
mary
of
does
not
tical
have
•J k_
of
has
this
28).
(ice
identified
the
on
the
potential
laminar-flow-control
airplanes
ice-crystal
At
clouds.
lower
of
In
laminar
of
on
altitudes,
boundary
and
where
loss
layers
to
Reference
airframe
for
have
many
been
laminar
(refs.
flow
high-altitude
liquid-phase
cloud
remains
NLF
maintain-
physical
on
tran-
vibration,
28
is
a
sum-
vibration
important
no
transition
particles
of
on
the
that
boundary-layer
through
which
of
noise.
there
atmospheric
flight
concern
some
transition
significant
during
of
concludes
flight,
turbulence
effects
a
environments
understanding
literature
28).
for
achieved,
operating
boundary-layer
and
atmospheric
be
of
turbulence,
The
influence
Studies
our
crystals),
27
can
effect
exposure
work.
(refs.
of
quality
the
increased
from
past
significantly
observed
is
resulting
applications
and
surface
research
particles
much
effects
8,
proper
much
research
phenomena
atmospheric
4,
the
of
prac-
discernible
(refs.
27
and
2
to
28)
swept-wing
(stratospheric)
particles
exist,
-
little
research has been done to determine the influence of such cloud particles on
laminar flow of swept or unswept wings. Studies of the influence of noise on
boundary-layer transition have shown the potential for loss of laminar flow due to
turbine-englne and afterburner noise impingement on laminar surfaces (refs. 27
and 28). Limited evidence exists that engine/propeller noise on piston-driven
airplanes may slightly affect transition position on NLF surfaces (ref. I0). The
literature
is not conclusive on the operational seriousness of insect contamil%ation
and propeller slipstream disturbances to laminar flow.
....
AIRPLANEDESCRIPTIONS
ANDCORRESPONDING
EXPERIMENTS
Airplanes
Eightairplanes
were studied in these tests.
Seven of the airplanes utilized in
the flight experiments were selected because of smooth skin surface conditions existing on all or portions of the airframes. The eighth airplane utilized NLF gloves (as
opposed to a production-quality
wing surface).
The Rutan VariEze, Long-EZ, and Laser
Biplane Racer, and the Bellanca Skyrocket airplanes were constructed of composite
fiberglass
or
sandwich
Beech
24R
milled,
Sierra
or
of
five
T-34C
airplane
sections
transition
plane,
the
the
effect
of
fixed
by
and
(see
2
Transition
transition
is
a
flight
of
maximum
lift
of
the
Long-EZ.
rakes
Beech
on
with
on
the
left
support
and
bonded,
some
of
airplane
the
was
wing;
measuring
to
on
the
aerodynamics
canard
a
these
techniques
related
sur-
Beech
gloved
and
fo#
experimental
by
60-Foot
Tunnel
configuration
provided
and
by
canard;
either
airdata
and
on
(2)
artificial
and
chordwise
section
of
the
the
rough-
minimum
control
and
on
and
were
laminar-flow
data,
Rutan
Long-EZ
in
and
as
provided
Observation
the
affected
and
airplanes;
VariEze
and
instrumentation
airfoil
similar
included
performance
respectively.
behavior
sensors
the
fea"
airplane.
experiments
flight-test
measurements
drag
each
' visual
studied
extensive
airframe
for
Other
VariEze,
pressure
hot-film
a
transition
more
lift
as
unique
conditions
components.
Rutan
utilized
drawings,
test
included,
boundary-layer
stability
and
caused
and
airframe
Skyrocket,
measurements
Skyrocket,
winglet,
photographs,
airplanes
fixed
example,
30-
canard
Specifically
wing,
conducted,
various
on
Langley
rain.
airplanes
provided
the
descriptive
Bellanca
the
advanced
experiments
on
all
of
in
an
on
for
the
For
Skyrocket
provided
stream
wake
surveys
Boundary-layer
by
propeller
information
slipfor
the
T-34C.
Rutan
propeller,
The
of
others.
the
for
of
locations
of
effects
Some
conducted
The
experiments
effects
fixed-transition
on
listing
locations
the
29).
simulated
experiments
studies
was
ref.
construction,
transition
than
Centurian,
made
eighth
gloves
slipstream
aluminumhoneycomb
P-210
were
transition
propeller
or
structures
The
airfoil
boundary-layer
the
aluminum
appendix.)
characteristics
water-spray
Table
of
in
core
Cessna
measurements
(See
laminar-flow
develop
foam
28/29,
of
Waviness
investigation
VariEze
(I)
or
The
with
to
aerodynamic
following:
tures
Model
airplanes.
fitted
used
full-depth
constructed
skins.
these
wind-tunnel
study
ness
riveted
over
Leafier
-.....
A
to
were
measurements
results.
Gates
airplanes
of
were
skins
The
flush
faces
"--
carbon-fiber
structures.
VariEze.-Flight
two-place
airplane
tables
physical
3 and
difference
4.
between
The
and
airplane
wind-tunnel
type
characteristics
flight-test
the
full-scale
with
experiments
were
a high-aspect-ratio
and
airplane
wind-tunnel
design
is
coordinates
shown
model
in
and
conducted
canard.
figure
the
with
(See
are
presented
2.
The
flight
only
article
a pusherfig.
I .)
in
significant
was
the
®
I
installation
!
:
of
were
fiberglassto
sanded
conform
skins.
!,i
Both
_:_
,
an
airframes
visual
outboard
leadipg-edge
constructed
_e
using
wind-tunnel
of
and
of
the
flight
measurement
i,:_
winglets,
_:
experiments
included
observation
layer
transition
(using
acoustic
::
range
of the
flight
sublimation
technique
i!:
the
_
60-Foot
i_
an
Static-force
wind-tunnel
on
as
0.625
x
shown
the
canard
in
the
wind
of
the
canard
as
sprayed
rate
of
I gal/hr
canard
the
method
stability
and
to 148 knots.
unit
Reynolds
3.
The
and
canard
canard
force
and
included
surfaces,
of
ref.
30)
control.
_
and
of
The
wing,
flight
on boundary"
airspeed
transition
data
using
of 1.4
x 106 ft -I .
data
were
Collected
in the Langley
30-
was
data
and
filled
through
clouds
The
calibrated
Flight
number
mount
were
core
configuration
and
of the effect
of
flight
transition
detection).
figure
distribution
0.26,
0.53,
0.79,
by
water
spray
from
at
in
60
psi.
of
the
Long-EZ.-
isolated
were
The
the
from
collected
the
a
with
by
model
by
simultaneously
dynamic-presof attack
from
number
of
Flight
canard
to
with
spanwise
of
rain
airfoil-shaped
volume
about
such
6
ft
that
boom
downstream
mean
covered
simulated
located
and
diameter
water
stations
was
at
the
spray
ahead
located
on
total
flow
a
right
canard
enveloped
the
range.
experiments
similar
of
varied
four
effect
pointed
200-_m
span
was
from
The
Nozzles
about
boom
boom
recorded
0.95.
horizontal
4.
angle-of-attack
type
high-aspect'ratlo
.of
were
and
a
figure
droplets
height
airplane
data
=
diagrammed
throughout
Rutan
(using
and
model
_%is
Both
foam
of sideslip
from
-15 ° to 15 ° .
The
nominal
Tests
were
conducted
over
a range
of angle
10.5
psf
which
corresponds
to a unit Reynolds
water
The
propeller
a
_
tunnel
boom
canard
with
winglet,
transition
performance
balance,
pressure
at
the
semispan.
the
design
wind-tunnel
contours.
wing,
airplane.
full-depth
ft -I.
Chordwise
on
the
was
from
65
taken
at a
in
strain-gage
106
fixed
_llght-test
of
and
boundary-layer
flow
visualization
mounted
on an external
balance
system
-6 ° to 40 ° and
a range
with
model
force
data.
sure
of the tests
was
_!
of
on
airplane
tests
were
data
model
'runnel
internal
effect
canard
the
experiments
transition
i:_
:_'
on
structures
accurately
The
airfoil to surfaces
the airfo_l
on
determination
and
dr,:op
composite
were
the
also
conducted
VariEze.
different
The
wings
and
on
airplane
a
two-place,
pusher-
configuration
winglets
than
the
utilized
VariEze.
Two
t
different
Long-EZ
results.
The
to
L_I_I
_l_
airplanes
only
differences
aerodynamically
fair
the
faces.
Figure
and table
5 is
5 contains
a list
of
tograph
of
of
one
were
the
tested
in
main
these
Long-EZ
verify
and
the
was
The
the
trol
was
158
knots
dsring
i!_]
and
a
on
airplanes
wing,
of
determined.
at
it
was
foam
with
winglet,
fixed
1.51
assumed
Rutan
Laser
Biplane
tractor-propeller
this
canard,
on
indicated
altitudes
×
106
that
Racer.(figs.
of
ft -I.
the
_%e
and
size
shape
of
of
of
When
position
design
6.
table
fiberglass
airplane
included
fuselage
nose,
airplane
to
The
in table
given
in
with
airspeed
4700
were
size
tested.
core
transition
The
density
was
full-depth
conducted
the
effect
testing
purposes,
using
experiments
transition
tion,
built
repeatability
the
wheel
the
fairing
only
error
for
ft.
V i
was
used
sur-
designed,
6 is a pho-
coordinates
The
canard
4).
The
for
airfoil
composite
the
is
air-
skins.
visual
and
performance
range
7500
transition
rudder
the
geometry
of these
airplanes
as
geometric
characteristics.
Figure
NLF
airfoil
on the wing
and winglets
are
given
identical
to that
of the VariEze
(coordinates
frame
the
airplanes
wheels
a sketch
of
the detailed
two
to
observations
wheel
and
these
maximum
was
available
In
stability
tests
The
of
fairings.
was
unit
for
65
and
to
Reynolds
data
number
reduction
zero.
A single-place
biplane
with
7 and
8) was
tested
in flight.
addicon-
large
negative-stagger
Detailed
physical
_I
dinates are given in table 8• The composite airframe was built using full-depth foam
core
with fiberglass
on theareforward
and graphite
skinsairfoil
on thedesign
aft wing.
characteristics
of theskins
airplane
shown wing
in table
7• The wing
coor-
i,
:
Experiments
l!i!:
tions
side
_
165
conducted
with
on portions
of
of the propeller
knots
at
a
this
airplane
the
lower
slipstream,
(£orward)
The
density
altitude
of
10
included
determination
and
upper
(aft)
indicated
airspeed
000
ft.
The
of
wings
for
transition
both
these
corresponding
inside
tests
unit
loca-
and
was
out-
Reynolds
number_
i
_•i_
during
ii
these
Gates
ducted
wing
tests
Learjet
with
was
was
a
Model
in
determination-of
9.
numbers.
Mach
Cessna
in
figure
in
a
of
and
edge.
the
in
on
The
Beech
wing,
maximum
24R
Sierra.-
presented
in
The
propeller
uses
the
wing
The
rpm,
composite
ne,
nation
the
for
__!ii
outside
I
lation
,
The
,i I
wing
section
lift
and
scanning
_
A
maximum
The
Y
left
drag
The
these
to
for
in
incorporates
133
number
the
of
leading
transition
tests
I .48
was
×
and
139
106
the
four-seat,
12.
Geometric
NACA
on
propeller
details
airfoil.
portion
In
the
I .38
of
addi-
propeller.
was
was
co
low-
63-series
outboard
and
on
ft -1 •
with
the
left
The
figure
testing
the
the
roughness
near
visualization.
knots,
with
of
limited
stabilizer,
was
tail
during
The
A
surface
these
an
pre-
horizontal
percent.
rivets
bonded-aluminum-skin
was
are
region
horizontal
Conducted
the
surfaces.
observations.
reduce
testing
shown
shown
metal
observations
vertical
Experiments
distributions,
measurement
calibrated
the
dimpled
range
during
conducted
locations,
calculations,
propeller
done
airplane
the
9
f
i
operating
×
high-performance,
I0 u
at
ft-'.
single-
airplane
shown
in figure
13.
Geometric
details
are
II•Detailed
data
on an NACA
632-215
NLF airfoil
was
The
airframe
was built
of fiberglass,
aluminum-honeycomb
structure.
detailed
to
i
was
stringers•
chemical
on
tests
5
half-of
_ transition
Reynolds
testing
Mach
altitudes
airplane
and
are
subsonic
density
airfoil;
from
ribs,
chemical
made
at
business
64-series
c)n the
were
design
high
conventional
included
airplane
were
during
flush
airspeed
number
wing
unit
of
wing,
airfoil•
at
number
on
were
airplane
sublimating
for
11
experiments
e_eriments
transition
pressure
the
•
Reynolds
for
maximum
sandwich
of
Calibrated
observations
_u_-n_
the
unit
11.
engine,
retractable-gear
Bellanca
Skyrocket
presented
in table
12•
If!
of
visual
thickness
done
row
this
The
Clark
airspeed
and
obtain--_
a
selected
calibrated
figure
details
included
winglet
sublimating
The
leading-edge
physical
to 0.70
skins,
was
spanwise
with
Flight
table
transition
2700
in
a
retractable-gear
are
was
facilitate
portion
single-engine,
tion,
_
dark
spinner.
knots.
sanding
to
of
conducted
the
propeller
region
in
with
con-
9.)
airplane
on
NACA
fig.
Aircraft
retractable-gear
aluminum
and
illustrated
the
Experiments
locations
154
sanding
waviness
varying
riveted
skins
six-passenger
an
were
(See
0•55
measurements
single-engine,
airfoil
dark
and
was
Reynolds
incorporates
filling
painted
unit
this
wing
pressurized
this
of
with
tests
experiments
aluminum
material•
the
transition
wing
NACA
body-putty
was
The
of
The
on
maximum
for
constructed
which
filling
and
The
utilized
symmetric
was
amount
wing,
ft.
10.
filler
for-these
_light
airplane.
milled
conducted
locations
Centurion.-
was
table
uses
airframe
sanded
speed
business
Stiffened
range
characteristics
sented
tail
00
P-210
10
Physical
of
Higher
10-seat
Experiments
number
of 15 500 to
16
3.08
x 106 ft -I.
Longhorn.-
integrally
transition
The
ft -I.
turbojet,
made
table
106
28/29
of
modifications
presented
×
twin-engine,
constructed
contour
1,38
chordwise
and
system
slipstream.
description
airspeed
for
the
the
static
these
these
14
Skyrocket
effect
for
to velocity
measure
illustrates
experiments
tests
of
pressures
boundary-layer
was
utilized
Figure
of
with
including
was
the
is
176
include
propeller
analysis
profiles
airfoil
for
in
a
determi-
slipstream•
of
A
section
insideprofiles
and
wake
instrumentation
contained
knots
visual
instal-
reference
maximum
31
unit
•
Reynolds numberof-1.90 × 106 ft _I. During the observations of propeller slips£ream
effects on the laminar boundary layer, the propeller
was
operating
at
1800
rpm.
Beech
T-34C
turbine-engine
surfaces
to
utilizing
were
gloves.-
Support
mounted
inside
these
was
hot
were
films
phase
maximum
of
and
layer
166
used
to
All
these
knots
detection
transition
the
over
a
range
experiments
Reynolds
Testing
Sublimating
method
for
surface
with
stream
in
of
heat
rates
in
rate_
for
given
tempera£ure,
the
rates;
(ref.
15.
if
faster
velocities
less
of
sati§factory
in
is
5
an
I
tO
removed
hand
250
8
10
20
To
particles
by
gently
testing.
In
minimizes
technique,
to
by
30
formation
which
occasionally
the
addition,
the
several
shown
in
chemical
the photographs
in
roughness
particles.
protect
condition,
the
pit
for
flight
not
necessary
the
testing.
to
of
As
this
"bag"
can
be
However,
the
surface
using
in
0°C
flat
typical
This
The
the
soft
figure
allows
from
relatively
manner.
with
a
any
a
slow
Even
or
or
can
prior
brushing
chemical
patterns
wedges
reaching
running
atmospheric
be
rubber-gloved
the
transition
to
large
the
brushing
cord
spray-
cheesecloth
learning
sublimating
at
psi
thickness
particles
tuDbulent
rip
25
solution
coating
vinyl-
prior
a
conven-
at
unusually
the
chemical
contain
with
chemical
without
diffusing
paper
to
for
free-stream
operated
brush
with
Prior
the
and
1,1,1-trichl0roethane
coating,
coating
frequently
30°C
a
Sublima-
times
spraying"
bristle
For
different
the
nozzle
from
are
temperatures.
produces
wedges
to
relative-
sublimation
"dry
of
faster
subsonic-flight
33).
acenaphthene.
rate
than
is
produce
with
is
rate
result,
this
to
fan
conducted
with
sublimation
have
free-
stress
temperatures
predicted,
particles.
chemicals
covered
of
in
a
turbulence
relative
figure
solvent
chemical
report
a
test
to
shear
rate
the
the
local
pressures
(ref.
utilizes
a
of
Typical-sublimation
A
with
a
with
sublimation
exposure
areas
free-stream
the
be
turbulent
these
described.
sublimating
surface
the
liquid-
airspeeds
The
testing
5 minutes
adhere
experiments-were
in
To
manner
the
coatings
_2 ,'ZI
the
of
to
The
surface
addition,
through
coating
atmospheric
from
surface.
of
rubbing
occurrence
can
suifiable
volume.
ft 2 of
avoid
brushing
of
method
A
The
for
in
equipment.
coatings.
In
the
The
ft _I.
higher
determine
variety
60
in
fluorene
chemicals.
application
I solution
per
to
shown
from
of
During
vapor
at
temperatures
are
spray-paint
to
_m.
at
knots
chemical,
quart
a
reacting
flight
coating
uniform
chemical
to
than
compressed-air
mixed
ing
in
used
temperatures
slower
range
of
blade.
rpm.
flight
involves
layer.
Chemicals
be
2000
of
films
behavior
propeller
10 -6
the
higher
and
chemicals
operating
or
to
well-suided
under
various
due
boundary
these
can
conducted
hot
high-frequency
transition.-
rapidly
acenaphthene,
figure
x
more
pressure-
Of
1.5
solid.
is
Chemicals
chemicals
the
thus,
of
produces
32).
The
various
indication
A
turbulent
the
a
chemical
sublimates
vapor
.over
of
the
to
transition
difference
diphenyl,
transition
tional
the
chemical
The
flow
boundary-layer
volatile
film
characteristics
figure
selection
p,.
the
naphthalene,
shown
of
This
within
sublimation
pressure
film
flow.
transfer
of
were
detectors.
- single-
smooth
Procedures
boundary-layer
chemical
laminar
to
tion
thin
the
vapor
detedtion
indicating
very
proportional
include
ii,
a
airflow,
areas
and
chemical
visually
experiments
of
laminar
number
with
time-dependent
pass
conducted
low-wing,
fabricated
(feathered)
on
were
the
each
150
effect
two-place,
slipstream;
of
from
from
the
unit
propelle_
observation
disturbances
a
this
gloves
hot-film
pgrmitted
for
on
Transition
outside
determine
on
conducted
with
operated
clouds.
flow.
sensors
boundary
propeller
experiments
were
surface-mountid,
both
of
l_minar
flight
airplane
laminar
glue-on,
response
The
training
Caused
the
to
chemicals,
the
by
test
cockit
is--
temperatures
I0
®
as
high
time
as
for
30°C,
developed
at
recorded
has
shown
test
the
no
NLF
effects
on
transition
L
of
Further
Acoustic
detection
described
in
section
permitted
documentation
this
method
to
the
listening
to
stainless
steel
ness
of-0.015
ble
tubing
tubing
the
The
and
within
about
heard
4000
to
was
tube
sound
turns
could
response
local
A
accomplished
in
a
each
validity
be
"calibrated"
when
Airspeed
tests.
airplane
utilized
airspeed
by
installing
the
figure
18.
The
three
elevator
outside
thickby
The
of
the
flexi-
flexible
listening.
than
of
To
the
turbulent
background
of
engine
during
calibrated
deflections
shown
the
in
fig-
locations
elevator
were
Indicated
a
density
its
airspeed
9_e
and
of
deflection
recorded
both
VariEze
free
about
2.0
fashion,
particular
acoustic
flow.
was
pace-airplane
of
being
Inthis
boundary-layer
for
were
signal
factors
air-
altitude
location
tube.
for
conducted
fixed
at
load
indicated
by
flight.
acoustic
selected
turbulent
testing
both
are
transition
spanwise
the
normal
VariEze,
were
in
each
fashion
and
error)
for
at
the
absolute
the
test
conducted
selected,
pulling
laminar
the
determined
was
tubes
was
flights
calibrated
for
for
o.d.
connected
cabin.
sound
recorded
Testing
an
For
and
tubes
used
transition.
manually
(position
calibration
was
of
forward
in
effects
an
then
also
flow.
0.060-in.
attain
plug
during
aft
means
between
procedures,-
tubes
and
tube
transition
passing
laminar
were
airplane
ear
visually
the
by
were
quieter
the
were
As _ each
force
a
at
knots.
speed,
to
tubes
acoustic
technique
components
necessaryattenuation
pressure
i50
time.
for
an
on
tubes
the
technique
This
the
Shape
to
exhibits
chosen
data
to
i.d.
provide
noise.
forward
to
testing
chase
deflections
75
At
at
flow-field
(ref.
34).
transition
shown
ft.
one
surface
were
chord
from
checked
Other
for
the
with
airstream
transition
varied
banked
each
of
in.
layer
oval
temperatures
total-pressure
shows
pressure
an
to
of
chemical
31.
acoustic
clouds
surface
16
pressure
provided
and
pdsitions
acoustic
cabin
defenders
noise,
surface
0.060
boundary
_5-percent
was
listened
and
the
employed
to
order
tests.
through
at
reference
the
flight
Figure
These
In
durability
the
testing
in
fact
films
beneficial
the
coatings,
after
are
locations,
flight
tests
flattened
o.d.
hours
method
VariEze
of
The
in.
Ear
These
The
speed
layer.
in
the
VariEze.
17.)
[
locations
17.
the
fig.
laminar
on
present
end
48
the
thick
transition.-
effect
(See
0.080
propeller
ure
in
boundary
layer.
noise,
on
termina%
ear,
boundary
the
one
to
of
With
that
This
hot
additional
is
and
implies
using
testing
and
chemical
6-series
transition.
An
flight
run.
transition
used
the
for
with
up
use
NACA
pattern
observed
of
effects
of
glove.
or
be
layer
determination
t2te T-34C
been
lasted
was
surface
in.
of
Was
human
of
the
movement
boundary-layer
of
used
wing
forward
of
first-order
ample
chemical
can
thin
variety
of
transition
has
boundary-layer
the
the
absence
on
test
on-the
of
of
on
permits
the
locations
described;
wind-tunnel
transition
measure
The
the
details
redundant
taped
either
after
indication
no
chemicals
for
pattern
-The
example,
affecting
transition
manner
simultaneous
sublimating
acenaphthene
20°C.
date.
causes
conducting
without
pattern
to
roughness
by
chemical
near
tested
for
without
the
first-order
confirmed
the
Hence,
the
airfoils
of
condition.
acenaphthene,
landing
in
chemical-coatlng
and
for
and
Applied
the
feature
reaction
approach,
ground.
was
with
slower
climb,
the
on
modern
the
takeoff,
calibrated
technique
fixed-
to
and
measure
transition.
pointer
visually
free-
elevator
This
and
from
was
markings
the
chase
airplane.
11
r
Airplane
using
all
a
the
used
level-flight
calibrated
flight
to
geometric
clinometer
tests,
calculate
angle
during
measured
of
attack
testing
pressure
in
both
altitude
and
......
_
was
recorded
the
VariEze
outside
11
,
onboard
and
air
manually
Long-EZ.
During
temperature
were
densityaltitude.
RESULTS
L
Table
test
2
is
a
conditions
summary
in
of
both
measured
the
and
wind-tunnel
Wind-Tunnel
Transition
conducted
locations.-
at
a
test
Boundary-layer
transition
demarcation
where
cal
formed
the
by
chemical
coating
is
the
has
The
rapidly
results
in
the
coating
indicate
the
that
(x/c) t = 55 percent
and
on the
region
of
laminar
flow
was
also
transition
angle
attack
line
of
"frosty"
of
various
boundary
transition
was
winglet
by
the
lift.
by
the
the
formed
the
t = 65
pattern
line
darker
as
exposing
on
at
(x/c)
chemical
19
and
is
were
cruise
figure
layer,
obtained
tests
for
area)
demarcation
turbulent
wing
and
indicated
detection
of
photographs
(white
This
in
for
Experiments
1.5 ° , the
indicated
chemical
locations
experiments.
chemical
_ =
sublimated.
sublimates
surface.
of
transition
flight
VariEze
Sublimating
condition
predicted
and
the
chemi-
the
canard
of
area
wing
at
percent.
on the
A limited
highly
swept
strake.
Measured
wave
on
2.0
surface
the
in.
wing
The
allowable
waviness
has
an
amplitude
maximum
of
(h
=
data
indicated
this
0.036
presented
wave
wave
in.
height
is
only
for
in
the
of
0.009
appendix
one-fourth
k =
2
in.)
show
in.,
Of
for
a
and
the
that
a
the
of
empirically
single
largest
wavelength
wave
determined
at
the
test
conditions.
Effect
rain
or
with
a-large
determine
dinal
of
amount
and
of
loss
configuration
of
lift
a
result
by
The
data
flow;
not
when
possess
by
balance
during
about
results
indicates
30
the
percent
as
water-spray
that
with
the
artificial
indicate
that
moments
energy
shown
in
tests
are
effect
of
that
figure
shown
spray
that
the
22.
in
Data
figure
was
is,
obtained
similar
the
slope
of
the
loss
on
canard
on
the
pressure
from
the
effect
of
the
lift
this
recov-
slope
the
to
is
for
leading
lift-curve
Comparison
the
lift
designed
induced
during
canard
23.
to
was
from
attached
The
the
on
function
in
is
con-
23.
of
related
airfoil
turbulent
remain
to
transition
a
loss
separation
of
the
distribution
this
canard
roughness,
water
roughness;
to
is
carbo-
chord
20
as
pressure
becomes
effects
60
simulate
fixing
curve
The
No.
5-percent
figures
to
The
of
to
in
longitu-
characteristics
chordwise
layer
artificial
in
pitching
trailing-edge
boundary
wing
interest
flow.
strip
at
and
of
flight
conditions
stick-free
laminar
surfaces
indicates
For
is
or
presented
separation.
sufficient
with
the
21)
it
narrow
pitching-moment
of
however,
fixed
20
airplane
(fig.
the
of
35,
flow.
pitching-moment
figure
the
loss
canard
are
and
the
in
trailing-edge
airfoil
of
a
I/8-in.
the
reference
stick-fixed
lower
study
examination
extensive
and
on
lift
decreased
An
is
in
laminar
surfaces,
the
a
upper
on
of
airfoil
there
this
discussed
loss
in
water
of
reduction
transition
transition
the
transition
does
For
of
artifical
NLF
and
when
flow
shown.
the
changes
spraying
laminar
The
on
in
by_applying
Results
canard.
reduced
ing
or
span
boundary-layer
particular
ery.
wing
of
the
of
attached
edge
full
attack.
with
flow
studied
significantly
of
canard
the
As
result
significant
were
are
canard
canard.-
can
laminar
are
conditions.
effects
on
types
characteristics
on
canard
taminated
of
of
there
transition
grit
angle
transition
cloud
aerodynamic
rundum
the
fixed
certaih
if
fixed
the
of
in
of
canard
these
of
curve
fixis
is
'
reduced. It should be noted that only half of the canard span was enveloped in the
water spray; ?herefore, the results from a fully enveloped canard would be in closer
agreement with the fixed-transition
canard data Shownin figure 23. These data indicate that a nose-downpitch-trim change (with stick fixed) would result •from flight
through rain or from artificial
transition
(grit)
in this
airplane.
Tests
wing
were
with
laminar
the
flow
on
of
the
shorter
of
the
movement
pressure
also
conducted
canard
the
wing
moment
of
angle
of
has
arm
less
from
on
determine
already
of
this
the
fixed
effect
the
transftion
distribution
degrees
to
transition
on
wing
wing
airfoil
is
Rutan
6,
VariEze.-
error
may
(discussed
the
side
is
expected
in
this
the
upper
Since
the
surface
at
the canard
identical,
Long-EZ-apply
region.
airfoil
the
calculated
in
in
in
an_
at
55
and
and
because
is,
above
canard
as
well.
using
a
the
amplitude
on
in
region
are
listed
the
few
has
in
on
indicator
of
the
of
slope
located
the
2
on
canard.
for
wing.
on
the
shown
0.012
maximum
this
in
AI
the
Transicanard
in.
the
section
figure
and
in
in.
the
fo:
AI.
fig.
A2)
the
allowable
wave
is 0.020
existing
on
exceeded
is
table
in
(table
wing-as
appendix)
small
(x/c) t = 10 percent.
the VariEze
and
Long-EZ
discussed
the largest
the waviness
some
lift-curve
edge
25(b))
strake
at
for both
the
not
On
an
constant
and
R = 1.40
x 106 ft -I.
60 percent
behind
the out-
(fig.
appendix
show
_ At
airplane
elsewhere
dial
in
the
trailing
transition
24
error.
the
percent
the
figure
Therefore,
for
0.35,
x/c) t =
percent
and
on the
conditions
presented
laminar
below
tCL(=
55
in
transition
port
a
t =
of
measured
total
loss
minimum
lift
trim
laminar
caused
for
the
wind-tunnel
experiments
by
show
laminar
amplitude
for
airplane
empirically
edge
canard.
with
(fixed
a
at
and
the
transition)
10-knot
to
The
magnitude
determined
the
isolated
lift,
drag,
these
canard
and
decrease
in
speed
the
edge,
maximum
was
canard,
flow
moment,
effects
airplane
was
presented
a
in
(corre-
in elevator
trim
lift-curve
slope
wing,
separation
and
perfor-
26.
The data
any airspeed,
20-percent
The
changes
changes
in
on
leading
of
a
decrease
cruise).
by
large
transition
canard
affected
on
are
presented
in figure
deflections
required
at
in•
CD
near
were
caused
leading-edge
fixed
flow
(corresponding
and
increase
trim
speed
trailing
deflections
speed
coefficient),
transition
the
of
trim
characteristics
in the
trim
elevator
lift
With
near
Of
because
That
flow
position
fixed
25
3 °,
at the
location
of
conditions.
Thus,
a 23-percent
and minimum
maximum
induced
gravity
laminar
strength.
of
figure
percent,
operating
data
t_e
of
trimmed
winglet.
moment
attack.
the
loss
value.
effects
to
for
On
pressure
value-(equation
single
wave
at
the test
increase
sponding
deflections
(x/c)
double'wave
and
longitudinal
a large
increase
maximum
of
presented
circulation
effects
in
and
at
data
behind
shown
VariEze
was
indicated
surfaces
The
of
angle
favorable
transition
static
strake
waviness
maximumallowable
and
the
illustrations
to
waviness
The
for a
2 in.
7-knot
bound
of
I ft
(x/c) t = 55
airfoils
and
surface
maximum
mance
show
about
free
by
The
occurred
the
Surface
Measured
h
k a=
because
drooped
winglet
the
the
dominated
locations-are
are
pitching
center
with
on
However,
Experiments
Winglet
at
V c = 135 knots,
_ =
on
the wing
(fig.
25(a))occurred
on
aircraft
calibration
versus
section).
leading-edge
tion
airspeed
fixed
fuselage
Transition
board
of
be
of
wing
and
Transition
The
effect
position
effect
the
transition
chord.
attack.
Flight
insignificant
fixing
configuration
airfoil
not
of
5-percent
the
to
the
effect
at
elevator
determined
previously
and
was
trim
during
in
13
®
figures
20
transition
is
to
23.
on
all
about
The
reduction
lifting
30
in
surfaces
percent
and
total
is
the
airplane
shown
in
reduction
in
lift-curve
figure
The
airplane
due
to
reduction
is
CL_
To
slope
27.
fixed
in
about
13
canard
percent.
CL_
analyze
the
stick-free
pitch
changes
due
to
loss
of
laminar
flow,
wind-tunnel-
I',/
L
measured
elevator
These
data
thus,
for
showed
the
deflections
stick-fixed
of
the
0.20
pressure
The
no
visible
laminar
observation
change
was
while
ficiently
as
large
discussed
at
2
is
the
153
two
a
= ---I .5 °,
locations
made
of
between
both
the
(x/c) t =
beginning
32 to 34 percent.
at the
leading
juncture
between
vortex
impinged
figure
were
unbrushed
the
on
calsed
by
chemical
Winglet
the dark
in
the
transition
leading-edge
boundary
chord.
chord)
and
near
caused
juncture,
the
the
lines
black
transition
local
in
moved
junc%ure
of
two
on
and
remaining
transition
interference
figure
the
which
on
tip,
to
each;
there
loss
pitch-trim
been
trim
sufchange,
the
effects
on
On
the
nearer
have
been
are
and
wing
to
As
in
was
locations
upper
by
(fig.
the
surface.
vortex
on
percent
the
of
_c_.al
wing-winglet
highlighted
the
by
winglet,
Onthe
which
as shown
in
an elevator
_igure
28(e),
deflection
was
6
of
(x/c) t =
1.8 °.
55
the
figure
shows
aft-facing
step
the
of
in
28(b))
forward
are
side
the
strake,
the
transition.
(in
At
tip
in
surface
The
small
well
28(c).
(inboard)
canard
airfoil
aft
was
in width
o_ the
seen
strake
farther
figure
wing
the
of the
wing
leading-edge
step
was
resulting
distinc28(a))
wedges
wing
transition
caused
where
in
fairing
previously
and
(fig.
the
listed
wheel
no
(fig.
28(c)).
presented
a
this
suction
the
and
ft -I .
wing
turbulent
outboard
step
seen
28
nose,
forms
wing
at
upper
the
surfaces.
transition,
occurred
with
of
This
nose-down
conditions
main
main
e
14
the
No
Clouds.
identical,
adhered
root,
the
wedges
forward
appears
I minute
× 106
test
the
inboard
portion
coating
indicated
winglet
28(d).
1 .42
the
of
the
using
this
Boundary-Layer
turbulent
flow
about
I ft
the
location
just
outboard
strake
winglet
a
figure
essentially
of
at
the
=
the
Most
Transition
in
(x/c) t = 32 to 35 percent
paint
stripe
which
physically
slightly
transition
were
particles
Near
flow,
one
at
of
concentration
fuselage
R
Since
wing
chemical
the
the
surface,
Canard
transition
wing.
layer.
However,
and
Transition
15 percent.
On
of the chemical
was
0.16,
small
region
was
observed
edge
coating.
was
(x/c) t = 10 to
complete
sublimation
=
shown
canard,
tested.
results.
outboard
the
CL
airplanes
A
are
winglet,
were
laminar
by
windshield.
and
at
tempera-
located
longitudinal
size
observed
aft
of
these
to
occurred.
strake,
airplanes
of
was
and
than
or
pronounced
were
transition
through
particle
loss
have
any
less
elevator
similar
Ambient
Detection
wing
flight
of
cloud
locations
wing
knots,
absence
the
the
moments;
conducted
ft.
chemicals
was
on
during
were
port
of
"Acoustic
cloud
significant
would
wing,
Long-EZ
transition
tion
the
Had
' Long-EZ.--Transition
for
Vi =
noted,
cause
previously,
Rutan
table
by
sublimating
be
flow
tests
2000
hinge
produce
would
laminar
existence
on
not
total-pressure
The
encounters
the
on
altitude,
section
cloud
would
These
transition.
edge
behavior
clouds
surface
by
the
from
rain
stick-free
17.)
acoustically
clouds.
to
the
mist
detected
the
in
of
of
reinforced
in
The
fig.
natural
leading
through
density
determined
duration
was
one
and
the
technique.
68°F.
(See
described
deposit
flow
was
0.35.
previously
Transition."
was
knots,
and
fixed
at
liquid-phase
and
=
fixed
changes
through
with
flight
130
_
was
transition
detection
altitude
technique
compared
stick-free
flight
and
port
of
moment
test
maneuvering
were
transition
airspeed,
at
=
effect
tested,
acoustic
calibrated
x/c
model
effects
the
ture
no
moments
due
to hinge
behavioro
The
using
hinge
percent.
This
,_
V
¥
On
tance
the
of
flow
16
represents
boundary
the
at
fuselage
about
11
layer
about
the
tersunk
same
tube
the
laminar
the
the
k
=
in.
surface
fairing
The
3.0
The
nose,
the
and
in
trim
knots,
discussed
in
upper
(aft)
on
propeller
wake.
chemical
On
the
the
chemical
in
the
propeller
the
of
decreases
thereby
the
the
laminar
shear
leaving
the
the
of
laminar
laminar
leading
edge
step,
The
hatch-cove_
the
no
no
of
occurred
0.25-in.
couno.d.
observable
pit0t
effects
on
on
R
(table
wing
the
1.42
decrease
increase
of-fixed
in
in
transition
figure
slope
is
was
multiple
2 in.,
has
maximum
about
7
manifested
with
speed
transition
The
in
a slight
shown
in
the
were
aerodynam-
fixed
percent.
coef-
by
airfoil
by
mini-
lift
As
trim
canard
are
in
transition
cruise.
caused
winglets,
increase
trimmed
in
on
wings,
characteristics
minimum
slope
30
a
surfaces
on
fixed
C
an_
lift-curve
with
k =
airfoil
length
A3).show
in.
for
11-knot
with
deflections
fig.
ft -I,
trim
an
reduced
percent
Total
and
h a
transition
longitudinal
was
33
value.
experienced
27-percent
=
0.006
106
airplane
(fixed
and
AI
was
×
allowable
flow
airplane
lift-curve
=
t
surface.
amplitude
for
on
(x/_)
side
the
allowable
in.
laminar
in
the
appendix
maximum
airplane
on
previously
reduction
in
speed.
Transition
lower
locations
(hereinafter
=-165
knots,
forward
and
turbulent
which
the
existing
a
on
at
CL
aft
wedges
adhered
=
to
0.13, _ and
wings
seen
to
are
referred
was
in
the
R
(x/c)
the
wing
as
=
1.38
x
t =-61
figure
for
surface
figure
forward)
106
without
and
and
the
ft -1.
percent
both
31
wing
outside
wings
the
were
brushing
caused
during
coating.
portion
of
wake
flow
aft
of
film
could
in
the
the
aft
be
propeller
could
wing
propeller
at
(x/c)
propeller
by
sheet
the
chemical
the
a
in
61
in
the
the
A
on
propeller
of
layer.
slow
the
the
turbulent
outside
feature
film
the
laminar
that
dissimilar
wake
layer.
boundary
to
chemical
location
loss
slipstream
similar
thin
boundary
in
propeller
was
a
significantly
observed
the
percent.
was
transient
turbulent
to
film
t =
transition
vortex
thicken
in
wake
wake
transition
caused
sufficiently
immersed
the
observed
of
stresses
thin
aft
transition
inside
thin
the
the
pattern
showing
wake)
of
06020
effects
chemical
this
in
maximum
Racer.-
patterns
propeller
impingement
loss
of
with
at
had
_at
Transition
in.)
nose
of
the
surface
occurred
amplitude
24-percent
cruise
V i
wake
of
the
percent
trim
the
The
the
propeller
existence
at
inboard
31(b)),
of
at
dis-
extent
shows
nose.
existence
28(g))
52
is
presented
at
both
of
the
a
canard
particles
application
(fig.
as
for
t =
elevator
total
Biplane
2
wing
Transition
by
in
in
Laser
tip
configuration
the
to
damping
table
0.035
The
of
to
for
in
surfaces
Rutan
the
=
performance
This
speed
change
for
(h
(fig.
of
airplane
significant
short-period
listed
loss
29.
reduction
lifting
as
step
determined
total
changes
the
The
in.)
step
the
waviness
corresponding
large
causedby
ics.
the
corresponding
VariEze,
0.035
the
longitudinal
This
figure
from
presented
value
Thus,
on
Maximum
=
a
in.
The
in.
double-wave
appendix)
figure
speed,
ficient.
data
in
of
canard)
presented
(x/_)
ft.
empirically
effects
from
fairing
at
2.75
calculated
ft.
exceeded
all
and
was
equation
=
in.
(h
at
18
14
transition.
wheel
indicated
c
ml_
the
waviness
and
11
on
surface
(see
of
the
about
layer.
maximum
2
not
immediate
0.50
step
length
aft-facing
the
wave
surface
occurred
of
length.
forward-facing
the
The
that
length
fuselage
behind
boundary
upper
wheel
the
that
about
Transition
on
of
a
transition
surface
location
caused
protruding
28(f)),
a
of
a
at
shows
screw
at
percent
hatch
also
(fig.
or
survived
removable
figure
nose
in.,
aft
remaining
wing.
(and
not
flow
due
The
outsile
to
Such
a
This
thickening
sublimation
pressure
the
transient
process,
recovery
15
region
of
quently
the
for
the
Gates
and
are
airfoil.
Skyrocket
Lear_et
listed
of
higher
are
and
2
for
3.08
×
conducive
T-34C
Transition
wing
and
106 ft -I.
to rapid
cruise
slipstreams
this
at
M
=
altitude
of the
resulting
in
locations
winglet
This
test
sublimation
The
values;
propeller
are
made
subse-
airplanes.
Longhorn.the
conditions.
typical
of
number
the
shown
hd
=
was
chosen
chemicals.
Reynolds
sense,
are
0.7,
of
in
figure
500
it,
to provide
It isnot
was
results
16
about
a
repre-
400
these
32 ....
percent
experiments
conservative.
Transition
figure,
which
The
ated
the
largest
k
=
maximum
2.0
wave
rows
t
were
during
on
=
40
terminate
and
the
of
the
to
45
in
the
attributed
percent.
In
natural
to
the
transition
large
chemical
par-
application.
winglet
(fig.
wedges
were
observed
well
as from
surface
edge
is
wave
On
step
near
tion
observed
fig.
32(b).)
NO
the
The
Re
at
the
are
listed
stabilizer
surface
32(b))
was
emanating
from
irregularities
skin
flush-countersunk
portion
to
locations
on
the
chemical
at
the
suction
structural
of
of
junc-
(inner)
screwheads
initi-
V
=
139
the
to
of
Vc,
knots
the
the
using
the
measured
height
0.002
it,
and
of
the
wing
this
sweep
an
the
lami-
aft-facing
transi-
step.
A.
of
le
Reynolds
thickness
the
maximum
in
premature
to
in.
the
in
the-wing
bg_attributed
momentum
_
6.58
determined
on
leading-edge
h
=
equation
existing
height,
cad
was
c
empirically
the
attachment-line
Transition
locations
wing
and
154
upper
knots,
Observations
angle
with
by
allowable
winglet
to
region
(See
17 ° was
number
74.
Centurion.-
0.32 c.
i
of
was
for
the
laminar
condition
waviness
winglet,
the
due
value
2
Thus,
by
the
the
test
determined
in.
exceeded
contamination
table
with
2.0
in
the
as
exceeded
condition
wing
For
=
span
edge
P-210
in
k
on
that
the
height,
not
lower
maximum
for
0,36
for
leading
test
Cessna
wave
was
the
spanwise
observed.
in.
height
on
appendix.)
single
0.008
region.
measured
(See
allowable
allowable
=
wedges
transition
leading
(x/c)
which
surface
turbulent
surface
as
chordwise
in.
appendix,
nar
and
was
seen
turbulent
wing
natural
Many
the
to
32(a))
are
transition.
The
with
the
the
winglet
Spanwise
the
of
rearward
(fig.
wedges
to
55 percent.
adhering
between
side.
wing
Most
adhered
most
(x/c) t =
particles
the
turbulent
noted.
ticles
ture
on
several
location
CL
=
observations
28/29
table
cruise
_an
II
Model
in
C L = 0.12,
and
R
static
temperature
sentative
Further
on
attack_are
R
the
given
CL
are
1.34
x
variation
in
shown
in
lower_-surfaces
the
106
of
to
the
ft "I
1.48
x
and
106
ft -I,
and
transition
table:
(X/C)t,
percen£
1.34
×
106
5
.28
1.43
×
106
29
154
.26
1.48
×
106
44
0.35
33
horizontal_
surface
149
1 39
the
upper
following
R,
figure
and
7:
Figure
of
this
33(a)
figure,
shows
transition
(x/c)
of
t =
29
(x/c)
percent
t
=
44
at
percent
V c
=
at
149
Vc
knots.
=
154
On
knots
the
is
lower
part
faintly
®
the
wing
to
the
dark
painted
reduced
skin-surface
skin)
reduced
the
I
i
ment
of
free
transition
in the white
region.
little
significant
difference
in transition
It is
locations
i
sanded)
and
production
wing
surfaces.
(See
thick
skin
was
sufficient
at the
unprepared
fig.
A4.)
The
stiffness
surface
location
tested
rimental
waviness
of
ure
initiated
;_
were
!
Figure
Vc
=
149
step
_i
at
c
joint
=
at
c
=
=
12
rate
was
ft.
Beech
Vc
For
133
chemical
eral
a
single
knots,
R
25
J
of
=
75
the
Reynolds
wing
on
about
the
the
fig-
coating.
• at
aft-facing
horizontal-tail
free
at
this
was
22
_J
measure-
transition
length
in.
radius
no
in
to
8 in.
(forward)
and
34(e),
the
blade
and
was
2.89
of
at
a
35
unit
×
laminar
on
coating.
Figure
by_paint
surface
on
the
was
by
Reynolds
on
wake
34(b)
boundary
leading
less
edge,
(aft)
propeller
blade
unit
the
local
layer.
or
crit-
1.88
of
station
the-
_as
on
the
the aft
face.
radial
location
was
operating
Reynolds
Mach
in
x
about
faces
was
at
number
flight
106
at
number
are
transition•locations
of
sev-
than
Transition
lower
surface.
At these
flight
probe
shows
imperfections.
are
propellers
the
wing
(x/c) t = 42 percent.
the aft-facing
pressure
number
the
much
of the area
wedges
caused
by
laminar
stripes
the
ft -I , and
illustrates
inboard
of
The
to 46 percent
on
the
Math
number
was
0.31.
at
the
the
listed
transition
local
flow
are
Natural
(x/c) t = 80 percent
was
about
6.5
in.;
106
and
either
for
and
the
on
34
and
indi-
is
stabilizer
respectively.
length.
conditions,
figure
paint
aft
exceeded
percent,
over
of turbulent
effect
of these
not
criterion
vertical
0.30.
and
face
and
locations
Figure
Rc
had
suction
of
(x/c) t = 45
was
0.22
and
shown
ft
_I, thek =maximum
2 in.,
show
the
were
and
remains
6
34(d)
these
II.-
are
CL =
heights
the
forward
measurement
surfaces
number
percent
length
=the1.48
× 106
appendix
conditions,
chemical
insect
observations
31.
Skyrocket
lower
0.035-in.
the
in
propeller-spinner
the wing,
free
transition
(fig.
34(c))
was
triggered
step
At
blade
_J
40
spinner
(x/c) t = 45
by convergence
stripes
of
tail
percent
0.84.
seen
chemical
the
chord
the
criteria
tested.
and
unbrushed
paint
inR
same
propeller,
ft -I,
roughness
joint
on
these
and
occurred
at
coefficient
chord
by
in• figures
Additional
reference
Bellanca
and
the
locations
shown
or
in
caused
the
38 percent
chord
at
percent
was
0.46.
cussed
in
_
shows
propeller
locations
surface,
106
lower
surface
the
vertical
are
between
x
dark
skin
lap
chord.
the
figure
34(a)
was
was
obliterated
stuck
measured,
the
on
(x/c) t =
The
local
1.38
wedges
Transition
the
=
shown
in
transition
sloping
not
propeller
_i_
the
the surface
waviness
wing
surface
regions
wing
particles
rpm
wave-under
upper
0.0020-in.
10-percent
tion
lift
33(d)
of
_*
t =
shows
local
of
the 0.020-into preclude
det-
_i
Transition
spanwise
upper
33(c)
The
by
there
exists
(filled
and
wedges
unbrushed
(x/c)
case
The
dark
measure-
rpm.
the
ical.
On
Transition
_L,
Figure
sierra.-
turbulent
Though
50
Figure
percent.
length
for
2
upper
surface
tested,
free
2700
ft.
The
24R
table
was
27
the
area.
the hotter
successful
that
prepared
turbulent
of
this
double-wave surface
amplitude
0.010 presented
in.
for
Measured
waviness was data
4.83
The
location.
in
to
the
noteworthy
on
the
in
transition
t
=
the
adhering
initiated
(x/c)
3.67
Most
surface
that
1900
h a = 0.020
in.
Thus,
prepared
or production
in
loads.
was
in.
the
white
area
(relative
rate
sufficiently
for
particles
lower
of
was
st
rotation
cated
_!
shows
location
of
flight
chemical
Transition
skin
location
location
the
by
33(b)
transition
ment
under
knots.
a
temperatures
on
chemical
sublimation
adjacent
il
!
on
ft -I .
dis-
the
Transi-
Airplane
trimmed
conditions,
9.7
x
106
and
17
/
k!
k
K
,
9.0
X
106
caused
by
the
at
the
coating.
marked
outboard
large
chemical
In
with
an
figure
patch
induced
wedges
in
cles
brushing
the
slipstream.
were
at
the
were
_
the
joint.
the
This
Skyrocket
varies
since
testing
the
at
the
Skyrocket
the
predicted
From
the
was
typical
36)
an
example
downstream
an
also
of
Figure
37,
from
illustrate
wind-tunnel
same
that
a
single
from
the
and
is
of
is
shown
of
minimum
measured
in
method
and
h
where
=
0.002.
criterion
on
of
the
root.
the
the
in
However,
allowable
waviness
existing
on
NLF.
location
measured
the
mid-
attachment
2 in.)
speeds,
Thus,
for
boundary-layer
to
high
larger.
allowable
the
tip
=
actual
appeared
empirical
(k
the
station
about
the
wave
wing
altitudes
of
Both
i_cation
measurements
of
transition
figure
with
predicted
pressure
range
transition
section
as
on
drag
36.
the
(pressure
relative
Predicted
Granville
transition
on
to
tran-
transi-
locations
peak)
the
During
above
and-to
the
(See
dard
time.
along
the
debris
in
Figure
span
of
lift
both
occur
upper
and
contamination
tests,
38
flow
ref.
a
flight
depicts
right
wedges
was
the
wing,
for
coefficients.
At
(C_ _ 0.3).
the Skyrocket
(ref.
or
on
31).
No
Based
gained
on
25
significant
on
the
flightvalues
of
only
by
speed-power
percent
in
effect
high-angle-of-attack
increased
flight
conducted
and
36)
and
higher
of
handling
about
fixed
qualities
4 percent
was
conducted
at-less
to collect
a samp!e
of insect
caused
transition
(supercr±tical)
of
heights
this
<ref.
predicted
with
31.)
2.2-hr
region
polars,
low-turbulence
performance
analytically
lift
drag
with
was
apparently
responsible
for
increased
drag
on template-measured
Skyrocket
airfoil
coordiappears
as an 80-percent
increase
in wing-
coefficient
in
airfoil
comparisons
airfoil
between
lower
slope
V c = 178 knots
insect
strikes
Tidewater
the
at
laminar
details
This
the
polars
of
Skyrocket
(subcritigal).
predicted
exists
lift-curve
ground
level
at
determine
which
warmweather
and
and
cruise
lift
coefficients
and
fixed
transition,
maximum
transition.
flight-measured
transition
air
leakage
were
based
transition
result
Skyrocket
observed;
fixed
the
37)
agreement
section
profile
drag
for
measurements
with
natural
cruise
presents
fixed
(ref.
Excellent
airfoil
31,
effects
, lower
to upper
surface
in flight.
The
predict4ons
nates.
The
effect
of fixed
18
low
integral
reference
the
airfoil.
measured
not
for
height
probe
were
and
and
the
presented
leading-edge
wings
test,
of
twist
surZaces.
which
was
flight
wake
bonded
near
are
wave
inboard
the
Skyrocket
the
distribution
shown.
the
the
the
of
between
Skyrocket
parti-
summary
measurements
thickness
indicated
at
conditions
than
pressure
using
at
0.015
at
less
31,
is
and
the
large§t
occurred
on
cruise
was
for
the
a
effects
Deviations
excess
Particleis
surface-waviness
(of
data
wave-height
0.017
in.
surface
of
and
35(e)
the
of
are
roughness
chemical
loosening
shows
were
insects
artificial
any
Figure
It
35(a)
v
application
by
mechanically
Skyrocket.
The
wave
conducted
chordwise
well
lower
heights
between
the
0.117
At).
conditions
reference
(ref.
lower
wave
wing
criterion
the
allowable
more
tion
table
of
on
by
_Rmlmlllll_lNiiillnp
fiqure
during
caused
of
flight.
accuracy
as
were
caused
semispan.
waviness
particular
typical
27,
sition
and
free-stream
reference
waviness
A5
edge
in.
More
Using
Detailed
(fig.
wing
to
in
surface
absence
from
prior
contour
large
the
resulted
the
stati6ns
as
measured.
leading
0.015
wing
contours
appendix
near
h
acrose
were
Note
seen
the
which
wedges
coating
Airfoil
several
theoretical
chord)
chemical
to
wedges
grit).
this
wedges
adhered
unmarked
80
pattern;
locations
turbulent
turbulent
The
No.
this
propeller
made
35(b),
of
the
transition
and
The
_ich
asterisk.
(I/4-in-square
by
station.
particles
"
in
Virginia
and
figure
flight.
late
March
between
positions
35(b)
shows
after
1430
of
and
the
the
several
1630
insects
lower
than
500
debris
and
ft
patterns
which
did
weeks
eastern
of
stan-
collected
surface
insect
As
of
illustrated
height
protrude
out
near
stagnation
the
cause
of
the
record
a
the
line
surface
rather
long
this
and
large
duration
in
25
of
the
the
the
propeller
the
indicated
Skyrocket
by
experiments,
slipstream.
the
chemical
detailed
Figures
pattern,
moved
of
was
the
where
lack
they
might
bances
outside
apparent
on
the
×
estimated
the
the
3°
Very
which
did
rapid
forwaL'd
occurred
pa_t
=
of
did
washout,
x/c
not
response
which
Wing
betwc;;n
transition
106
mean
boundary-layer
the
and
the
not
the
0 and
on
to
(x/c)
similar
the
the
show
the
that
upper
0.002
surface
tip
is
slipstream.
along
the
chord
inside.
interesting
for
wake
from
vortices
explanation
propeller
were
transition,
t = 36 percent
increment.
An
propeller
propeller
earlier
measurements
35(c)
possible
in
in
the
the
the
transi-
forward
effect
These
than
on
the
On
of
_arge
an
distur-
smaller
distur-
slipstream.
boundary-laYer
(See
1.715
One
transition
to
the
the-pr0peller
Skyrocket.
effect
wing.
environment
amplify
Time-averaged
=
any
chemical-indicated
bances
R
of
impinged
disturbance
outside
on
strikes
forward
detail
of
relatively
For
35(b)
a
increased
t]_e
were
insects
inward.
recorded
approximately
(x/c) t = 42 percent
outside
the slipstream
lower
surface,
transition
moved
f0rward
by
motion
protrude
especially
collected
supercritical
were
and
pattern.
varied
insects
conditions.
in
tion
ones
flight
-
the
figure,
remains
insect
chemical
edge
of
the
subcritical
insect
supercritical
leading
percent
In
turbulence
that
wedge
on
about
transition.
boundary-layer
transition
test
airfoil
unlikely
During
made
as
to
it
only
caused
point,
The
- make
38,
and
the
chemicals
stagnation
at
from
transition.
airfoil
i
I
in figure
supercritical
profiles
slipstream
fig.
39.)
ft -I ,
M
unit
with
These
=
free
and
measurements
0.31,
Reynolds
were-measured
both
and
n
number
=
was
were
1800
by
fixed
made
rpm.
1.778
x
rakes
at
s/c
Inside
106
inside
transition
ft -I
=
the
and ....
on
the
28.7
percent,
slipstream,
(using
the
propeller
momentum
theory).
With
the
free
thickened
The
which
was
profile
ness
the
an
to
create
seen
near
the
in
the
of
this
fluctuations
i!ii
A
the
boundary
this
as
time-averaged
the
the
conducted
Thus,
The
It
is
and
and
resulting
is
=
thick-
in
front
turbulent
that
profile
which
thickened
fixed
apparent
in
position
shape
was
has
turbulent
this
the
transition
thickness_(6
using
illustrated
the
voltage
propeller
sensor
on
section
environment
boundary-layer
the
boundary'layer
preceding
are
The
more
chordwise
verify
boundary-layer
hot
films
behavior
in
figure
fluctuations
heat-transfer
at
To
symbols.
Skyrocket
slipstream
layer.
a
patterns.
slipstream.
boundary-layer
propeller
behavior,
at
outside
profile
the
effect
measurements
is
increased
in
thick-
0.28
for
the
in.
case).
in
therefore
slip[tream
appearance
layer
the
appearing
position,
solid
in
boundary
slipstream,
sense.
propeller
on
Since
of
the
normal
at
laminar
changed,
chemical
the
discussed
laminar
has
in
the
thin
Inside-the
profile
turbulent
were
the
in.
39
turbulent
experiment
(and
in
the
high-speed
traces
cyclic
is
gloves.-
oscilloscope
laminar
figure
actual
the
in
understanding
are
outside
which
shows
0.06
sublimating
profile
experiments
of
by
turbulent
in
T-34C
the
one
slipstream
phenomena
results
shown
and
_
positioned
slipstream
symbols
frequency,
and
was
39
6
turbulent
inside
shape
Beech
the
as
a
a
propeller
in.,
rake
propeller
to
where
0.24
actual
are
the
solid
not
rakes
of
_
laminar
profiles
ness
6
inside
was
of
figure
slipstream
to
shape.
of
transition,
propeller
fluctuations)
blade-passing
the
miniglove
measurements
were
are
on
in
time
time
the
dependent
T-34C
this
40.
the
due
a
to
leading
is
since
of
high
better
The
signals
local
boundary
freque_Lcy,
at-the
gain
hot-film
Occur
in
to
and
and
environment.
The
which
in-the
averaged,
shown
velocity
layer.
seen
edge
in
the
records
an
fl_2 '."
19
®
apparent
small
disturbance
in
velocity
reaches
amplitude
and
slipstream
the
of
data
the
in
show
are
hot-film
the
(the
transition
the
canopy
cated
boundary
to
ary
layer
layer
the
and
boundary-layer
very
quickly
laminar
near
at
all
in
•
Although
general
behavior
During
or
glove
laminar
40-percent-chord
when
clouds,
chordwise
Upon
laminar
same
grown
propeller
oscilloscope.
air).
edge
the
signals.
the
this
rpm.
_%e
clear
locations.
the
rpm,
2000
When
progressively
outside
windscreen
to
in
has
turbulent
onboard
the
leading
hot-film
to
an
flight
glove
reverted
and
of
on
frequency.
it
glove
propeller
mist
conditions
40,percent-chord
NLF
the
laminar
for
the
on
using
of
remained
on
sensors,
conditions
observed
as
fourth
10w
at
deposit
location
windscreen
turbulent
edge
no
of
relatively
were
which
blade-passing
and
sensors
observed
signals
for
Occurred,
same
a
was
propeller
third,
magnitudes
for
layer
clouds
the
The
relative
presented
The
inside
at
second,
duration.
the
boundary
rise
the
accumulated
hot
stations
exiting
station
mist
the
films
from
the
flight
the
cloud,
on
indileading
the
bound-
state.
DISCUSSION
When
on
viewed
modern
ity
as
a
whole,
the
production-quality
and
maintainability
speeds
up
mental
results
to
of
about
M
and
results
of
airframes
=
NLF
0.7.
their
for
chord
The
these
provide
In
studying
of
background
airframe
discussion
of
location
with
numbers
of
of
the
using
sition
ods
the
attack
used
at
for
trimmed
the
In
the
ble,
i'
sure
(e.g.!
fig.
analyses
reported
N
and
23).
sition
where
38,
The
are
and
of
the
39).
surfaces
pressure
on
transition
forward
disturbances
other
of
comparisons
transition
Typically,
occurred
to
36,
wind-tunnel
38,
and
did
not
the
about
achievabil-
30
x
106
these
or
and.
experi-
For
analysis
the
2.
effects
for
minimal,
using
transition
the
point
of
of
the
measured
locations
minimum
of
attack
airplane
this
procedure
locations.
of
was
possi-
minimum
and
with
pres-
predicted
little
transition
two-dimensional
of
value
of
of
similar
comparative
14 to 20,
22,
measured
surfaces
meth-
angle
angle
distributions
point
tran-
transition.
local
the
transition
the
waves.
were
empirical
tested,
with
other
4, 7 to 12,
experiments
be
of
of
evi-
T-S
measured
predict
flight
pressure
consistent
refs.
3,
predicted
measured
downstream
the
measured
of
of
locations
made,
wings
downstream
Generally,
should
location
the
present
to
measure
were
using
and
occurred
for
direct
estimated
predicted
where
in table
a
Reynolds
provides
with
(or
transition
the
These
factors
moderate-aspect-ratio
of
locations
comparison
influ-
effects
amplification
experiments.
observations
was
the
For
pressure
transition
for
the
noise,
measured
airfoil.
shape
provide
determine
first-order
the
minimum
normal
39
to
engine/propeller
Any
the
than
and
transition
predictions
as
comparing
of
flight
parameters
where
interest
boundary-layer
36).
This
observation
is
in the
literature
(e.g.,
summarized
experiments
for
summarizes
waviness.
minimum
experiments,
transition
three-dimensional
stream
these
the
and
of
things
by
comparisons
present
F
roughness
was
apparent
experiments
transition
up
follows
become
boundary-layer
coefficient.
it
Such
reference
the
locations
wind-tunnel
Locations
including
predictions
of
from
meaningful
all
of
method
flight
the
lift
produces
k.L"
the
integral
Because
of
tests,
empirical
locations
use
location
existence
Two-dimensional
made
surface
present
numbers
which
locations_
disturbances
the
the
transition
and
such
and
appreciation
implications.
disturbances
vibration,
bypasses)
dence
the
new
Reynolds
Transition
ence
flight
a
reported
pressure;
tran-
sweep
occurred
analysis
in
in
down-
(refs.
table
fact,
2
36,
for
for
the
2O
®
+*
Skyrocket,
transition
adverse
pressure
higher
Reynolds
nated
by
_n
amplification
by
occurred
30
_
extent
inlet
to
of
inlet
noise
on
by
an
environment
ents,
upper
relatively
large
surface
be
chord
Under
certain
either
stability
cloud
par£icles
laminar
which,
in
edge.
Cloud
wakes
flux
from
limited
data
The
total
of
the
mist
mist
loss
near
the
of
occurred
layer
early
became
measurements
mist
deposit
loss
of
ing
temperatures),
heavy
on
effect
laminar-flow
turbulent.
with
a
(at
where
the
even
at
trip
rain
layer.
At
sufficient
on
the
canard
characteris-
fixed
the
number,
provide
NLF.
aerodynamic
on
leading
Reynolds
on
of
surface
turbulent
experiments
transition
drops
loss
the
of
impinging
the
cause
near
particle
particles
affected
free-stream
airfoil
shedding
present
of
be
of
can
boundary
cloud
the
deposit
6.5
Hawcon
Hawcon
by
artificial
airfoil
42-percent
<
Rc
<
is
×
flight
to
move
a
deposit
on
wing
is
at
on
drag
possible
suggest
low
when
a
the
I ), wake-rake
clouds
section
results
clouds
the
It
that
clouds,
table
through
in
!06).
These
through
showed
through
(see
flight
increase
8.5
13)
flight
flights
from
a
(ref.
during
roughness.
during
mist
the
surface
mist
supercritical
flow
On
During
a
showed
_ing
oan
flux
on
by. the
The
water
gradi-
edge.
experiments
made
the
sufficient
with
of
condi-
flight-
pressure,
wing
by
elements
that
test
the
pressure
minimum
Precipitation
Comparison
engine
noise
flows.
or
laminar
and
the
the
in
about
Particles
flow
occur.
of
This
some
favorable
boundary-layer
and
spray-and
the
flight
the
a
demonstrated
on
the
laminar
can
f+low.
water
that
measurements
creates
for
tests
leading
were
Hawcon
deposit
a
the
time)
precipitation
laminar
shows
as
by
this
engine
of
under
laminar-flow
layer.
laminar
unit
flow
of
a
root)
turbojet
layers
of
Cloud
surface
act
traverse
laminar
of
in
deposit
The
of
effects
23)
of
of
chord.
aided
of
point
by
gradient
transition
wing
the
chord
result
the
roughness
size,
per
wind-tunnel
a
Results
wing.
the
the
canard
to
boundary
drag
on
operation
loss.of
they
area
perhaps
domi-
dominated
influence
boundary
and
boundary
and
cause
as
unit
loss
(fig°
transition
|
per
caused
roughness
can
particles
VariEze
surfaces
quantity
the
laminar'flow
laminar
of
two-dimensional
three-dimensional
particles
the
or
the
creating
(particles
the
Ks
of
in
of
lack
not
Transition
the
the
at
free-stream
be
35-percent
70-percent
Precipitation
the
onto
sufficient
partial
tics
by
of
conditions,
through
flow
as
numbers
a
laminar
downstream
Reynolds
precipitation
about
since
at
proximity
which
at
t_at
expected
Effects
by
shock
the
is
flight-measured
(near
was
in
flight
pressure
separation).
number
indicates
layer,
suggest
sufficient
Can
of
This
to
adverse
where
in
turbulent
appears
wing,
pressure
spite
boundary
data
possess
transition
in
that
process
or
the
separation
Ks
transition
(laminar
Reynolds
minimum
laminar
surface,
in
28/29
0hord
surfaces.
laminar
These
a
of
occurred
upper
the
at
point
the
transition
layer
Model
of
comparison
instabilities
free-shear
Lear_et
this
acoustic,
Rather,
T-S
chord
flow
wing
flown.
li
I
laminar
the
attenuation
tions
the
predicted
as
layer.
Gates
40-percent
The
!ii
__
in
location
of
thought,
such
boundary
the
predicted
previously
two-dlmensional
on
at
106 .
L
_+,
the
of
occurred
_lan
_%e
implication
disturbances
instabilities
type.
at
The
numbers
background
disturbances
or
occurred
gradient.
that
the
altitudes
the
due
to
the
mechanism
(above
freez"
occurs.
&
The
VariEze
flight
experiments
liquid-phase
clouds
on
ous
on
effects
research
9rincipally
_.he X-21
•;-hrough
with
flight
ice-crystal
the
ice
laminar
of
crystals
demonstrated
the
when
deposit
cloud
(ref.
+
In
40)
the
no
mist
particles
occurring
experiments
clouds+
flow
at
high
laminar
present
on
effects
NLF
(refs.
altitudes
flow
flight
was
of
occurs
40
(in
lost
as
experiments,
flight
through
on
the
wing.
to
42)
has
the
a
Previdealt
stratosphere).
result
when
no
,_ >
of
mist
In
flight
deposit
21
........
"
-
®
_T"_T
occurred
on
cloud
the
laminar
particles
critical
the
in
spherical
calculated
Since
no
critical
loss
However,
would
loss
if
of
the
on
laminar
Reynolds
number
to
size
flow
cloud
flight
in
would
For
of
the
mance.
These
a%d
Skyrocket
"the
could
erosion,
the
or
to
ice
changes
in
to
the
drag
area
which
with
less
profile
slze
curve
25
been
was
a
layer
near
on
had
no
is
flow
a
_m.
than
condition.
of
587
knots
to cause
laminar
particles
the
a
do
,lot
leading
flight),
a
of
debris,
edge
perforthe
loss
flow.
and
the
coefficient
should
(see
figs.
able
to
27
due
and
no
30).
significant
on
VariEze
separation
Long-EZ,.
laminar
flow
leading-edge
Whatever-the
without
cause,
laminar
of
flow
are
area,
not
effect
Long-EZ
the
occur.
and
lift
on
on
lift-curve
Long-EZ
fixed
As
flow
loss
wing
lift-curve
separation
on
the
values
of
is,
in
such
loss
the
such
of
laminar
also
repro-
trip
the
transition
was
canard
leading
the
edge.
can
that
airfoil
that
be
This
designed
pitch-down
with
highly
selected
which
flow.
induced
For
no
with
during
loaded
do
not
the
separation
and
31).
reductions
reductions
these
by
be
to
fashion
airfoils
surfaces
was
flow
This
predominantly
the
a
lift-
canard
fixed).
fixed
is
fixed
total
This
VariEze,
of
on
where
roughness
effect
should
slope,
induced
wetted
transition
(stick
configurations
large
the
fixed
and
effect
from
(ref.
experienced
transition;
with
canard
slope
These
Airplanes
airplanes,
30).
Canard
tail
measured
29.)
airplane
larger
airplane
That
upon
of
change
the
airfoils
and
with
artificial
existed
On
were
separated.
27
conditions
NLF
and
the
canard,
figs.
this
airfoils.
canard),
or
the
designed
flow
of
or
and
tested,
was
26
NLF.
experiments
Although
NLF
proportion
pitch-trim
using
transition
figs.
turbulent
(see
Long-EZ
laminar
fixed
(See
on
percent
boundary-layer
the
and
to
data
the
edge.
transition
VariEze
provided
both
and
typical
effect
to
13
to
wetted
due
airplanes)
if
separation
measurable
due
configuration-related.
both
(i.e.,
fixed
lift
22..
88
aerodynamic
and
laminar
with
large
total
wind-tunnel
turbulent
rain
II,
the
near
on
insect
either
Configurations
than
for
not
surfaces
Both
fixed
section,
airplanes.
became
to
to
leading
separated
through
Skyrocket
the
for
_,maller
airspeed
WheL'c
tunnel
in
drag
II
longitudinal
the
canard
under
experience
losses
relatively
flow
7
VariEze
the
same
feature
separation
For
in
rather
the
layer
trimming
(wind
qualities
benefits
nose-down
the
two
airfoil-related
flight
is
flight
flow
Additionally,
then
also
from
observed
the
the
significant
observed
no
was
laminar
preceding
cruise
relative
reduced
boundary
the
smaller
slope.
on
design
altitudes
of
the
produce
and
airplanes
flight
boundary
condition
an
the
diameter),
considerably
_m,
by
41)
T
Transition
VariEze
in
Skyrocket
laminar
produced
was
and
particle
VariEze
20
transition
handling
from
surface
three
first
(which
the
been
low
loss
the
in
and
resul£
induced
in
percent
experience
slope
duced
are
40
altitude
and
temperature)
an insensitivity
of the
Fixed
effects.
and
lift-curve
result,
of
discussed
could
lifting
separahion
at
a
flight
stream
had
at
tested,
included
As
Long-EZ,
drag,
transition
VariEze
expected
complete
accretion
changes
These
a
atmospheric
of
had
airplane
of
performance
VariEze,
large
free
unaffected
(refs.
on
_'_'
understand.
Increases
on
the
clouds
airplanes
If.
due
important
the
effects
airplanes
occur
(based
400
was
-
surface.
several
determine
of
the VariEze
test
results
illustrahe
througll
layer
criterion
the
}_
Effects
to
boundary
Hall's
for
particle
required
(at
flow.
These
the
].aminar
Using
particles
average
layer
deposit
t|%e
stream.
particle
aloud
have
been
of
laminar
boundary
free
particle
liquid-phase
tills,
surface,
tile
_ 7] I _
fixed
range
large
in
maximum
from
20
changes
transition
trimmed
to
are
on
the
27
percent
attributparticular
--.
canard
airfoil
transition
coeffielent
cient
incorporated
does
not
occurred.
actually
The
result
tests
increased
loss
as
a
fact,
as
laminar
standard
two
as
the
changes
of
those
significant
In
significant
of
on
induce
which
procedure
in
Skyrocket,
where
reduction
31t
in
maximum
fixed
maximum
lift
lift
coeffi-
transition.
performance
the
any
the
no
reference
fixed
occur
for
in
Of
indicate
On
separation,
discussed
result
flow
airplanes.
flow
or
importance
airplane
handling
of
with
qualities
as
fixed-transitfon
surfaces
smooth
the
flight
enough
to
support
NLF.
Propeller
Past
observations
of
transition-(refs.
5,
6,
research
by
Young
reported
effect
Of
the
behind
the
measured
Where
thickness,
tion
near
by
the
these
that
indicated
by
propeller
his
flight
the
Three
T-34C)
the
slipstream
on
in
the
slipstream
figs.
35(b)
slightly
and
hot
airplane
the
of
increase
due
to
These
experiments
the
question
Analysis
laminar
that
flow
the
of
total
recent
in
Wenzinger
the
observations
in
cyclic
of
(ref.
flight
the
nature
45)
suggest
of
Such
(i.e.,
that
be
in
previous
incorrect,
time-average-measured
66-series
latter
at
flight
14,
and
of
to
transition.
Biplane
boundary
Racer,
layer
40.
the
On
aft
effect
the
since
pattern
on
using
the
laminar
laminar
behavior
benefits
and
31
indicates
about
some
boundary-layer
of
on
empennages).
significantly
conclusions
the
conducted
time-dependent
is
the
<see
slipstream
nacelles,
and
in
wing
of
chemical
experiments
reference
as
slipstream
drag-reduction
slipstreams
NACA
measurements
by
gave
measurements
apparent
cyclic
wings,
presented
an
changes
7,
45)
moderate-effects
These
propeller
laminar-flow
in propeller
fl0w.-
on
indicated
The
slipstreams
slipstreams'may
mistakenlyc.depended
as
(ref.
for
portion
any
the
detrimental
propeller
experiment,
slipstreams.
possibility
data
if
methods,
a
about
as
rake
figures
little
layer
the
drag
the
Skyrocket,
the
of
the laminar
inboard
transition
be
showed
profile
in
at
Wenzinger
not
in
the
similir
Concerns
airplanes.
Skyrocket
boundary
laminar
airfoils
loss
of
laminar
propeller
the
propeller
propeller
flow
lamitransi-
paragraphs.
and
might
large
slipstream.
illustrate
in
of
in
laminar
40)
showed
During
inside
(fig.
immersed
slipstream
edge.
transition-
reported
mounted
boundary-layer
on
edge
calculated
using
experiments
illustrated
the
leading
judge
thus
_7)
P-51
by
pattern
showed
a
and
experiments
(the
and measurements
35(c))
in
and
to
Young
section
detailed
that
wing
the
following
laminar
indicated
chemical
31(b)).
behavior
surfaces
drag
that
as
forward
films
boundary-layer
raises
fig.
16
tunnel
P-47
on
configurations
the
propeller
(see
moved
T-34C
the
reiM
used
propeller
slipstream
extensive
on
to
the
wake-probe-measured
locations
the
(refs.
the
exceeded
Hood,
a
leading
in
propeller
the
first
Racer,
wing
Wenzinger's
on
the
the
Zalovcik
the
of the present
flight
included
observations
Biplane
surface
by
of
was
airplanes.
discussed
to
occurred.
for
The
indicates
experiments,_boundary-layer
probe,
have
boundary-layer
conclusions.
44)
transition
flight
tests
of
are
(ref.
thickness
to
on
varying
Hood
move
different
front
reported
transition
immersed
in
experiments
were
determine
two
Hood.
Zalovcik
experiments
Rutan
and
and
survey
wind-tunnel
slipstream
produced
Young's
assumed
on
in
slipstream
airfoil.
during
edge
effect
Young
6)
effectively
of
Effects
propeller
45)
boundary-layer
was
reported
the
the
to
5 and
case
conclusions
Experiments
43
total-pressure
position
of
evidence
NLF
a
results
20-percent-chord
of
the
of
and
to
measured
leading
similar
validity
In
transition
the
reported
was
propeller.
location.
effect
17,
(refs.
slipstream
thickness,
nar
the
16,
Slipstream
that
less
the
the
loss
the
than
of
early
..thickness
or
23
shape
as
that
an
the
indication
section
slipstreams
may
airfoils
may
wing-mounted
of
drag
not
transition.
increase
be
as
large
provide
drag
tractor
engines.
The
implication
associated
as
with
_lat
reduction
for
the
fixed
benefits,
of
the
transition
present
observations
changes
in
leading-edge
even
on
transition.
multiengine
is
propeller
Thus,
NLF
configurations
with
Waviness
No
premature
attributed
to
smooth
wave
and
contour
As
a
the
waviness
the
King
Cobra
experiments
_ie
produced
Re. =
fact
which
numbers.
17
that
of
the
106 •
A
a
compatible
the
is
moderate
represented
waviness
level
of
The
two
of
the
tion).
no
Obvious
these
flight
winglets
in
ination.
Crossflow
streamwise
and
A
is
tion
criterion
no
spanwise
line.
On
the
figure
42
show
both
24
airplanes;
swept
that
R@
the
to
those
waviness
the
some
in
special
metal
shown.
methods,
relatively
that
no
waviness
fabrication
to
flight
Cobra
the
21)
for
measurements
with
King
achieve
modern
point
in
(ref.
waviness
high
significant
favorable
of
Reynolds
amount
pressure
Effects
for
A
no
from
at
be
in
flow
was
recognized
the
data
variEze
46
the
the
the
centers
the
and
the
on
of
chemical
spanwise
contamina-
swept
wings
leading-edge
existence
The
and
contam-
closely
spaced
coating.
contamination
L0ng-EZ.
affect
turbulent
spanwise
criterion
contamina-
as
rle
I + (t/c)
occurs
(I)
for
R 0
<
100.
spanwise
contamination
for
any
source
freely
propagates
did
adversely
leading-edge
on
sublimating
and
and
reference
(or
observed
by
can
instability
discussion
in
which
crossflow
line
this
flight
the
_R
VariEze
R o
phenomena
are
transition
the
42
sin
be
special
test
This
Skyrocket
surfaces
instability
can
contamination
tions,
there
may
lent
contamination
the
laminar-flow
experiments,
A
Where
the
at-medium
attachment
crossflow
summarized
0.404
no
Cobra
41.
measured
than
sanding
flow
surfaces
leading-edge
figure
is
=
swept
betweeen
in
R8
on
preceding
comparison
presented
received
King
figure
of
on
, with
illustrate
instability
streaks
be
production-quality
drag
waviness
wing-geometry-related
layers
contamination
- Since
could
perfectly
strength.
significant
boundary
were
which
1950
in
profile
of
Sweep
laminar
not
comparison
level
on
the
surface
laminar
results
which
tested
modern
from
areshown
filling
acceptable
experiments
composites.
achievability
with
the
surfaces
tests
lower
the
on
composite
extensive
of
surfaces
qualitative
modern
illustrates
waviness
gradients
×
the
-Conversely,
surface
or
minimum
provides
waviness
tested
Skyrocket
the
required
comparison
surface
the
any
the
occurred
metal
comparison,
preparation
surface,
surfaces
in
though
results
either
from
at
confirms
contour
in
observed
even
These
historical
the
of
free,
achieved
and
was
waviness
preparation;
smoothness
This
transition
-surface
(A =
not
root
27 °)
and
the
exceed
100.
was
for
51
R o
Long-EZ
The
the
same
VariEze
For
various
< 240.
spanwise
(A
=
was
and
roughness
_or
along
23 ° ) wings,
true
36
for
for
condi-
R 0 > 240,
turbuthe attachment
the
data
thewinglets
the
Long-EZ.
in.
on
On
the swept strakes of both the VariEze (A
exceeded
I00.
leading
have
edges
been
tions,
the
show
surface
_
of
240,
no
between
swept
region
On
high
from
A
the
64 °
not
flow
on
was
A
Learjet
148
=
wing
(A
=
spanwise
turbulent.
heads
and
a
Even
uncertainty
step
which
region.
RA
surface
spanwise
keep
R8
<
This
size
in
100-for
an
the
R@
varies
tion..
AS
R@
fact,
no
for
the
example,
spanwise
on
jet
1.9
×
106
ft-1.
potential
relatively
large
at
The
effect
in
laminar-flow
tics,
of
NLF
on
seriousness
characteristics
porous,
and
protection.
The
reference
debris
55,
For
a
is
Skyrocket,
NLF
as
well
in
GIII
of
of
the
64 °
fact
that
the
winglet
the
extremely
which
form
leading
condition
unit
J
I
of
R_
spanwlse
Reynolds
as
were
screw
where
the
values
ensuring
no
large
of
edge.
1.5
number
in.
and
relatively
not
airplane
45
000
ft
large
be
a
on
still
(chosen
and
at
it
appears
M
con-
for
=
its
large
0.85,
contaminacriterion,
40 at
the tip
for
Reynolds
number
of
that
need
lifting
serious
the tip,
precluding
spanwise
the
spanwise
contamination
for
certain
not
be
a
concern
is
an
important
for
Contamination
wings
as
by
in
as
the
needed,
may
wetted
discussed
insect
only
in
well
percent
as
of
insect
active
serve
48
of
of
edges
56
and
contamination
the
of
such
to
to
In
practice,
on
airplane
protection
both
insect
systems
protect
con-
with
characteris-
54).
dependent
insect
purposes
features
leading
of
be
methods
the
airplanes
population
(refs.
will-likely
references
debris
25
debris
operation
literature
performance
of
insect
the
contamination
edges
ability
representative
Bellanca
If
leading
contamination
the
debris
mission.
the
Debris
on
detail
ice-protection
and
the
may
contamination
considerations,
some
insect
fluid-exuding
leading-edge
the
spite
the
cruise
as
certain
observations,
spanwise
design
These
In
of
of
in
test
cruise
altitude
68 at
below
these
contamination
discussed
lead-
surfaces.
airfoil
wings.
are
on
Aerospace
an
Insect
sideration
the
from
47) varies
from
64 at
the root
to
of 35.000
ft, and
at a cruise
unit
applications,
lifting
general,
wing
r6ot-to
of operations
- Based
the
be
high-altitude
Gulfstream
the DC-10
winglet
(ref.
at a cruise
altitude
important
in
at
M = 6.82,
about
of
106 ft-1),
tip,
thus
could
the
contamination.
class),
from
80 at
the
final
example
in
of
Learjet
radius
that,
the
at
of
to
upper
winglet,
where
R 0
varied
could
not be ascertained
roughness
regions
present
At
spite
100
portions
some
aforementioned
leading'edge
contamination
business
a
at
in
were
the
excessive
the
42(b).)
exceed
cruise
(R = 0.87
x
root
and
40 at
the
implies
spanwise
As
typical
winglet
40 °-, the
on
by
transition
in
_
64 ° ) where
contamination
On the
Learjet
the tests,
it
present
caused
contamination
observation
surfaces,
cern.
In
swept
was
on
chord
=
pattern.
fig.
not
(A
Calcula
relaminarization
l-percent
strake
region
(See
did
been
the
might
strakes.
present
chemical
51 ° swept
the
have
about
Re
near
acceleration
for
leading-edge
this
on
(A = 51°),
observed
necessary
inboard
the
the
was
caused
106 ft -I, at
to 80 at the
contamination.
on
by
17°),
contamination
This
if
R = 3.08
×
would
drop
the
within
51 o , the
flow
observed
may
short
onto
for
runs
rapid
conditions
strakes
very
recorded
and
to
of
acceleration
the
propagate
127
laminar
46,
Long-EZ
flow
=
and
the Long-EZ
flow
were
still
by
unit
Reynolds
number
during
the
t_st.
151 at
the root
to 75 at the
tipduring
whether
a
short
necessary
laminar
from
the
and
= 61 °)
laminar
of
Relaminarization
reference
Long-EZ,
did
varied
for
VariEze
the
regions
strakes.
of
_e
the
On
break
Re
bo_%
method
that
edge*
R8
of
small
responsible
by
occur
-ing
However,
are
against
such
and
as
ice
discussed
in
insect
57.
pattern
insects
accumulated
caused
transition
in
flight
at
sea
25
®
_,
_
_
_
a,_
level.
Analysis
shows
thicker
boundary
layer,
of
the
insects
numbers
can
of
be
would
to
sample
try
and
ity
of
nation
can
nation
levels"
year,
transition
of
It
is
collected
airfoil
at
a
data
of
to
degrade
geometry,
and
here
even
serve
to
and
9
with
a
percent
though
large
few
of
illustrate
the
of
Of
of
them
are
sufficient
of
time
geome-
presented
day,
in
contami-
seridus
of
cer-
sensitiv-
insect
occurrence
place,
a
airfoil
varying
effects
although
combinations
mission
ft
about
combination
contamination
that
000
relatively
Examples
performance,
many
Thus,
edge,
particular
insect
25
only
altitudes.
presented
recognize
airplane
for
35_b)).
contamination.
to
to
of
number,
leading
cruise
this
insect
geometries
wing
high
altitude
Reynolds
(fig.
on
insensitivity
is-infrequent
cruise
unit
contimination
important
seriously
typical
lower
transition
cause
airfoil
54.
a
caused
conditions
different
reference
more
by
be
level
operating
a
might
insect
inherent
at
caused
have
insects
expected
The
tain
that
contami-
time
of
profile.
CONCLUSIONS
Flight
ducted
and
on
Reynolds
significant
cant
were
and
the-investigation.
1.
Taken
durable
on
exist
c_rtain
business
the
results
_hat
where
comparisons
could
occurred
downstream
of
calculated
of
ment
the
disturbances
favorable
tively
2.
of
indicate
any
of
airplane
of
with
the
same
in
evidence
is
are
for
measured
and
in
were
stability
of
lift
slope
as
fixed-transition
and
using
increases
as
tests
smooth
aerodynamic
surfaces.
airfoil
aerodynamic
changes
the
small
to
design
flight
don-
environ-
that
occur
13
as
in
typical
even
at
27
rela-
from
to
large
as
percent,
and
observations
flight-test
water
the
as
as
These
standard
fixing
drag
large
percent.
a
resulting
roughness
cruise
as
Heavy
as
_ontrol
artificial
coefficient
large
signifi-
persistent
flows.
flight
made
trimmed
find-
observed
that
this
modern
previously
for
Sufficiently
stability
that
more
than
provided
from
smooth,
locations
locations
two-dimensional
performance
is
surfaces
The
free
significant
suggest
transition
layer
typical
most
behavior
pressure
enough
on
maximum
the
Thus,
of
investigations
airplane
made.
were
lift-curve
importance
No
the
discernible
surfaces
In
criteria
all
effects
tested.
the-allowable
4.
tion
airplane
these
The
structures.
conditions,
the
con-
chord
airplanes.
aluminum
skin
to
boundary-layer
boundary
Measurements
or
relate
minimum
provide
transport
stiff
been
at
spray
to
procedure
simulate
transition
near
the
edge.
3.
than
the
causes
leading
any
flow
in
be
numbers
effects
lamina@
decreases
in
laminar
Reynolds
transition.
decrea§es
rain
the
of
tested.
gradients
chord
-percent,
for
to
Significant
loss
trigger
shapes
pressure
large
total
24
surface
commuter
have
airplanes
representative
p@oduction
expected,
tours
were
this
experiments
several
composite
c6nclusions
and
the
of
relatively
and
practical
(NLF)
and
either
provide
following
a whole,
NLF
to
flow
surfaces
using
waviness,
The
Of
as
of
selected
roughness
of
laminar
nonlifting
constructed
airframes.
regions
and
were
tested
production
and
representative
tested
surfaces
natural
lifting
numbers
airplanes
ings
wind-tunnel
various
maximui
cases
and
the
on
transition
Measured
wave
tested,
observed
heights
the
due
surface
agreement
laminar-flow
to
wave
determined
between
results
surface
waviness
amplitudes
by
an
the
was
were
empirical
empiricel
consisten_
were
observed
generally
on
smaller
criterion.
spanwise
with
contaminaprevious
_esearch.
26
®
5.
throuqh
The
effect
of
low-altitude,
windscreen
(or
fliqht
throuqh
liquid-phase
winq),
laminar
clouds
_c]ouds.
flow
is
on
With
unaffected
transition
no
for
mist
was
deposit
.@ubsonic
observed
for
oce_,_rinq
flight
at
fliqht
on
the
low
altitudes.
Lanqley
Research
National
Aeronautics
Hampton,
VA
May
3,
Center
and
Space
Administration
23665
1984
i_
!
."
L.
27
®
APPENDIX
i<
1
11
SURFACE
nar
The-accurate
measurement
boundary-layer
research
ence
of
local
waves
on
pressure
critical
callv
a
amplitudes
related
=
9
000c
k
where
h
is
the
c
multiple
waves,
is
dial
2
in.
indicator
The
over
at
from
the
is
placed
over
are
sured
on
arises
were
pretation
chord
without
the
ground
the
an
added
source
meaning
have
lami-
the
The
been
by
pres-
empiri
the
t
equation
error
thin
that
was
be
the
the
exists
a wave.
the
actual
leg
for
the
calculations
This
1/32
gauge
the
readings.
legs
During
the
minimized
by
Swept
will
rest
measurements
care
in
at
the
is
placed
the
type
recorded
around
1/4-in.
the
intervals)
curvature
was
of
to
provide
of
measurement
fact
that
certain
waviness
not
the
artificial
the
a
(e.g.,
addition
Additionally,
to
tapered
distorted
If
the
wings
gauge
at
a different
on
the
can
each
of
wave
also
is_skewed
of
and
the
more
measured.
on the
affect
will
airplanes-discussed
alignment
mea-
the
with
slightly
level
that
difficulty
as
than
the
surface
being
and
the
I/4-in.
in£ervals
or
device
waviness
structures
in
successively
yields
streamwise
leading
were
distance
as
chordwise
chord
representative
the
loads.
amplitudes
10 -_ in.,
in.
this
is
deflected
deflection
smaller
of 1 x
is
With
flight
(NLF)
base.
with
skins),
is
indicator
surface
plots
surface.
dial
permit
flow
was
(for
two
wing
laminar
tape
versus
indicator
is
The
dial
average
airfoil
this
leg
to
deflections
the
arise
under
a
gauge
plotted
For
used.
Beginning
Foremost
the
metal
using
during
single
simply
natural
originally
and
in
sweep.
stability.
selected
early
running
dial
A
for
transparent
then
the
waviness
legs.
apart
was
was
tape,
chosen
on
surface
measured.
which
larger
and
is one-half
of
co
on
the-center
within
from
waviness.
loads
probably
data
the wavelength
leading-edge
wave.
fixed
in.
method
was
of
flight
0.6
is
wing
single
three
nine-point
were
a
This
between
calculate
measured,
deflection.
of
in
turbulence.
27
k
the
measuring
design
because
stressed
to
being
data,
points
through
accurate
A
is
marking,
reading
edge.
shortcomings
fact
passes
or
line
both
physical
changes
to
reference
inches,
measurements
was
marked
2-in.-length
to
cycles
and with
both
The dial
resolution
wing
waviness
difference
several
lightly
from
legs
in
for
The
transition
in
of
are
with
refere-ce
raw
The
used
is
with
base
the
the
procedure
those
important
wings.
macroscopic
A
with
gauge
in
gauge
for
which
data
were
value
center.
waviness
Nine
over
legs,
_le
and
base
making
leading
waviness.
measured
wave
height
used
waviness
The
the
known.
There
is
transition
trigger
single
the
solid
waviness
which
interval.
smoothing
a
can
inches,
AI)
a
at
for
on
in
paired
intervals
plotted
and
modern
line
accurately
for
create
trigqer
which
one-third
convenience
from
actual
is
on
1/4-in.
each
Can
turn
wave
chord
(fig.
For
surface
was
h/k
procedure
the
edge,
wing
for.which_this
follows.
in
double-amplitude
the
mounted
of
research,
surface
can
waviness
laminar-flow
cos
is
leg
comparison
of
number
indicator
investigation
spaced
surface
MODELS
kRcl.5
inches,
The
RESEARCH
airfoil
wavelengths
Revnolds
ON
production
airfoil
which
and
to
of
and
laminar
gradient
WAVINESS
interfrom
herein,
dial
the
produce
indicator
this
APPENDIX
base. Because of these shortcomings in the dial indicator
surement, the data are defined as "indicated" waviness.
The indicated
figures A2 through
number of waves at
wavelengths of the
each
spanwise
counted;
region
of
quality.
waves
smaller
location.
waves
fell
in
the
chord
were
included
The
maximum
comparison
the
waviness data measuredon the airplanes tested are presented in
A7. Table AI is a summaryof the waviness data in terms of the
each location and the chordwise position, double amplitude, and
largest wave in * _e laminar
region
and over
the total
chord
at
measurement
most
between
existing
than
methodof waviness mea-
this
in
allowable
the
in
allowable
maximum
the
for
Only
waves
category.
the
multiple
measured
laminar
premature
table
wave
and
region
that
Waves
of
were-2
which
as
an
heights
in.
indication
are
also
maximum
allowable
all
one
but
or
occurred
of
in
the
were
turbulent
of-overall
given
wave
the
shorter
test
in
surface
table
heights
airplanes
AI.
shows
A
that
were
transition.
29
l
APPENDIX
TABLE
AI o- SUMMARY
OF
INDICATED
Largest
Airplane
Surface
WAVINESS
wave
MEASURED
measured
wing
Right
Long-Ez
Right
Right
Right
AIRPLANES
Largest
wave
in laminar
measured
region
Positioni"I
k, in.
s/c
Right
TEST
I]
Position,
VariEze
in
flight
ON
0.25
,40
.55
.85
.95
O. 736
.309
.578
.704
win@let
h/k
k, in.
s/c
0.0035
.0060
.0030
.0075
.0030
0.194
.535
2.0
2.0
2.0
2.0
2.0
O. 55
0.678
2.0
O.0O7O
0,465
wing
O. 55
.85
0.189
.208
2.0
2.0
0.0030
.0020
.189
•208
wingle£
0.55
0.270
3.0
0.0020
canard
0.45
o 356
1.75
,¢ing
O. 25
.55
.75
0.333
.433
.511
winglet
0.25
.80
3.0
h/k
0.0030
0.0100
.0060
.0020
.0015
.0020
.0100
.0105
.0115
.0120
3.5
0.0017
0.0125
2.0
2.0
0.0030
.0020
0.0100
o0110
0.270
3.0
0.0020
0.0115
0.0046
0.356
1.75
0;0046
0.0135
2.0
2.5
2.0
0.00_5
.0024
.0030
0.333
.433
.511
2.0
2.5
2.0
0.004[
.0024
.0030
O.0180
.0193
.020_
0.223
.533
2.0
• 3.25
0.0045
.0018
0.223
.533
2.0
3.25
0.0045
.0018
0.0215
.0248
0.2?8
.095
.125
3.5
2.0
3.5
0.0020
.0050
.0011
0.228
• 095
,1 25
3.5
2.0
3.5
0.0020
.0050
.0011
" 0.0100
.0100
.0100
0.312
.065
2.0
2.0
0_0015
.0075
0,312
•065
2.0
2.0
0.0015
.0075
0.0078
.0078
2.02.0
0.0045
.0070
0,065
,065
2.0
2.0
0.0045
.0070
.309
.316
.477
.347
2.0
3.0
4.02.0
i
VariEze -in
£unne I
Right
Right
Cessna
P-210
0.S.
u.s.
L.S.
Bellanca
Skyrocket
- pr0ductiSn
- filled and
- filled
and
Inboard
wake
probe
II
Outboard
wake
probe
sanded
sanded
Upper
Lower
Upper
Lower
0.065
.065
i
Gates
Learjet
Right
wing
Model
28/29
Longhorn
Right
3O
winglet
0.48
.72
0.430
.450
2.0
3.0
0.0020
.0030
0.10
0.10
.63
0.17
.69
2.0
2.0
0.0135
.0010
0,17
O.0079
.0079
',.
2.0
0.0010
0.0040
.0041
2.0
0.0135
0.0050
.0070
®
APPENDIX
ORIGINAl,.
OF POOR
PAGE IB
QUALITY
L-81-9530
Figure
AI .-
Airfoil
surface
waviness
gauge
w_th
2-in.
base.
'2
APPENDIX
J'AGE Ig.
-3
8O
I I"I
I,
I -.
n =0.25
X
c = 36.0
I I I
I
80
-in.]_
I
/%
£'F "\
r,_
X
I
[l
II
_O
"I
U
__
0
_0
\
20
reading,
l
c,:
in.Z
31.6
Data
\,
Relative
I I I
n : 0.40
gauge-
A
in.
I
Y
A
;'
"7
r_
tO
_
.
......
Nine,point
avg"
4
%
"/
I
I
i
V
0
aO
I
i
.J....
I
j
c =
8O
I
I
;l
/h tl g
I
n : 0.55
_--
27.7 in
I
I
I
^,-L
I/ _ n
_0
_./
v
t _
X
t
0
8
4
B
B
Distance
(a)
Figure
A2.-
Indicated
Upper
waviness
tO
18
J4
along
surface
data
VariEze
tB
18
80
surface
of
on
right
88
24
from
L.E.,
88
aO 82
84
36
38
40
in.
wing.
flight-test
airplane.
88
airfoil
surface
of
APPENDIX
ORIGINAL
PAGE 18
OF POOR
QUALITY
30 x 10 -3
•
1 I
y,
I 1
_
n = 0.75
t
c = 22.5
-in.
2O r
L_
I0
v
I
I
0
3O
1
-/1
|
I
c = 19.9
A
^ /I
2O
Relative
I
!
q : 0.85
in,--
]]I]
ll-ll
gauge
Data
reading,
in.
\\t
:
_./
iO
J
V;
V
.....
Nine-point avg
0
3O
n = 0.95
.I
I
c
= 17.3
'
2Q
-in.
,_.
j
':
Aw
/
_O
A
t} '¸
0
, ....
•,
0
2
4
6
B
iO 12 14 IB IB 20 22 24 2B 2B 30 32 34 S8 3B 40
Distance
(a)
Figure
along
surface
from
L.E.,
in,
--
Concluded.
A2.-
continued.
1
• .,._ _,.- ..,--.. _,,_ ._/_,, ..........
.;., .._............... . ..................
__:..... :,.._
,_
.-wT
7-
APPENDIX
ORIGINAL
PAGE 19
OF POOR
QUAUW
-3
BO
n = 0.25
50
c = 16.8 in. -
%----
4O
l,,
"\
_
',,:_:./
BO
.i
,,
....
:k
20
-,,
j
'_
_
,,
u ,,.;
I0
'
0
BO
I
,n = 0.55
BO
,,/
\,
40
Relative gauge
reading, in.
I
\''_"
BO.
c = 12.9 in.
%.
I
";'_1'
I....
Data
.....
L
20
Nine-point
avg
I0
.I
o
BO
:1
I
....
I
....
I;_',r
I
I
n : 0.80
-
IV
\"
,,--
,
•
,
c : 9.6
in."
I
H
all
....
,,,
211 ..
t0
k'
0
F'
&.
0
2
4
B
B
t0
t2
i4
from
L,E,,
tB
_',
Distance
(b) Upper
b
surf&ce
Figure
'
34
_
along
A2.-
surface
of right winglet.
Continued.
in,
i8
20
APPENDIX
ORIGINAL
PAGE
19
OF POOR
QUALITY
50 x 10-3
I
.....
q = 0.45
4O
C, :
13.0
in
.
3O
2O
\_/
iO
o
50
4O
'_n
Oat • "-=!
:r--,
"
......
Nine-poi
nt ._---_
aug
..... ,_"
3O
'= 0.65
cj=
'
]
13.0
I
_,
Relative gauge
in.
....
--
reading, in. _0
io
0
50
'
....
n = 0.85
i
4O
-- c = 13..0 in.
I
3O
./
20-
iO
--
0
/
%
"f'
.\ t
''
\
0
2
B
4
Distance along surface from
(C)
Upper
surface
Figure
of
right
8
L.E.,
iO
in.
canard.
A2.-COncluded.
35
APPENDIX
OfllO,
IN/U- pmee,IS
OF mOORQUALn'Y
ii
;:7"
--..i
iI
m
2
4
6
Distance
8
along
t0
surface
12
from
t4
L.E
16
,
IB
2O
in.
:,
(a) Upper
Fiqure
A3.-
Indicated
waviness
surface
data
on
of
right
airfoil
wing.
surface
of
Long-EZ
airplane,
36
r,
•_
$
®
t
APPENDIX
_NAL,
PAGE |9
OF POOR QUALITY
_00
10'3
I
I
90
_q = 0.85
BO
i,,
c = 25.3
in.
70
i
Data
BO
%%%
%
-,,_,
5O
Nine-point
avg
40
30
'%'1
",, k..
20
_0
Relative
gauge
0
7, ,
_4
reading,
in.
lO0
I
90
I<
70
I
I.
60
2:
I
,I
50
"
k
%
""
c = 218 in.---_-
I
t
I
i
I
I
I
,
I
I
I
40
I
I
f_
L
I
_ = 0.95
I
\1
BO
m,
I
I
I
-
=
i
t
I
30
20
I
I
]0
0
0
I
I
I
I
2
I
I
I'
I
I
I
i
I
4
6
Distance
8
iO -"
along
Surface
i2
from
t4
L.E.,
i6
i8
20
in.
(&) Concluded.
i.,
Figure
A3.-
Continued.
[
r.-
k_
37
.!
':i
APPENDIX
ORIGINAL PAeE II;
OF POOR QUALITY
BO x 10-.3
h,-
70
n = 0.25
BO
c = 23.5 in.
m
5O
Data
4O
Nine-point'avg
=
=
1
3O
2O
I0
0
BO
r
m,-_.
Relative gauge 70
reading, in.
BO
n : 0,55
c = 18.5 in
BO
4O
3O
2O
I0
"vi",
0
2
4
Distance
(b)
Upper
surface
Figure
38
along
A3.-
of
B
surface
right
from
8
L.E.,
10
in.
winglet.
Continued.
®
I
]
APPENDIX
IlAg£ 18
_F
POOI_ _UALITY"
60x 10' 3
]
7o
: 0.45
c : 13.0 in..
,o I
50
40
3O
\
, \
20
,¢'.f
_"_-_, _
re"
_ _,/
Data
-
.-=
iO
Relative
reading,
gauge
ifl.
Nine,point
0
BO
avg -
|
7O
q : 0.55
BO
c = 13.0
in,_
5O
! \
4O
3O
2O
_0
0
2
-
4
B
B
_0
Distance along surface from L.E., in.
(c)
Upper
Figure
L
surface
A3.-
of
right
Canard.
Concluded.
39
®
L
APPENDIX
01#_
OV_11"v
70 x 10-3
I
B0
I
c -
l
l
50
l
58.0
in.
i
I.
\;,_.=.\
4o
u.S. - production
quality
LI
3O
_
/ _
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w
Z'h
_
20
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5O
Data
Relative gauge 40
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sanded
qo
f
•
20
10
0
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I I I
1
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50
c : 58.0
in.
\
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40
30
20
tO
0
0
2
4
6
B
Distance
Fiqure
A4.-
Indicated
.
•
waviness
.
.....
iO
12
along
data
on
14
16
surface
airfoi.l
IB
from
20
22
L.E.,
surface
24
26
2B
in.
of
Cessna
P-210
airplane.
®;
J,
"_-"
A[_PEHi)IX
,31,.
-
_lmwJlllmlliJlllmll_
ORIQINAL
PAGE' I@
OF
QUALITY
POOR
"
x 10-3
Relative
gau3e
reading,
in.
20
v.
10
,,I
0
2
4
6
8
i0
12
14_16
Distance
(a)
x
5o
Inboard
..
18
20
22
24
wake-probe
28
30
32
34
36
38"
station.
i0-3
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l
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26
along surface from L.E., in.
......
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Upper surface
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30
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"
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I
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I
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50
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Lower surface
It
40
_,_
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v
I:
30
20
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i
-:
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Distance
(b)
Figure-A5.-
Indicated
Outboard
waviness
data
14
16
along
18
surface
wake-probe
on
20
22
frem
24
L.E,,
28
30
32
34
36
•
38
in.
station.
airfoil.surface
airplane.
26
of
Bellanca
Skyrocket
II
'
APPENDIX
ORIGIN_P_EIg
I
i
•
I
I
I
I
I
j
'
'
i
I
i
,-.
,%
A
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gauge
in.
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50
n : 0.40
_____
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I
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c : 31.6.-in._Data
i
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Relative
reading,
I
Nine-point
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I
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70
i
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l
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30
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i
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[
]
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50
Figure
avg _
I
I
waviness
t2
along
surface
data
VariEze
i
surface
of
right
I
i
from
26 28 30
L.E.,
32-34
36 3B
in.
wing.
on wind-tunnel
airfoil
surface
of
airplane.
42
®
APPENDIX
ORIGINAL
PAGE I_
OF POOR
QUALITY
!L
il
n = 0.75
c i: 22.5
l
ii
''
---
! in.-i_
I
i
i
!
I
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Relative
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gauge
--
Data
.....
Nine-point
in.
J
avg
I
I
I
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c :
I
'
19.9
|
i
I
4
6
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t2
14
i6
:[8 20
22
24
26
28
in.-
'
I
I
I
_
.
I
'i
I
I
i
30
32
34
36
38
Distance along surface from L.E., in.
(a)
Figure
Concluded.
A6.-
Continued.
43
APPE.DZX
ORIGINAL PAGE IS
OF POOR OUALITY
-3
6O
I
I
I
q = 0.25
50
-
c = 16.8
40
in _-
k k
./
30
20
iO
•
\
0
B0
k
. •
,
I
I
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50
_
= 12.9 in._
f
40
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Data
Relative gauge 30
reading, in.
Nine-point
avg
20
f
L
• •
-¸
10
/
0
60
I
I
n = 0.80
Z
50
" , c-- 9 o6 in- -40
%%
30
2o
10
0
2
4 -
6
Distance
(b)
Upper
Surface
Figure
A6.-
B
along
of
iO
surface
right
Continued.
winqlet.
t2
from
i4
L.E.,
iB
in.
iB
APPENDIX
ORIGINAL PAGE 18
OF POOR QUALITY
70 x i0"3
,w
-
1
B0
n= 0.45
50
c = 13.0
-_
in.-
40
3O
<
20
I
_,,m ,r"
i0
0
70
I
I
I
I
60
Data
50
Nine-point avg _-
Relative gauge
reading, in. 40
_
n = 0.65
-_
c = 13.0 in.-
i\
30
20
-- _--._
/,
iO
0
7O
m
I
60
50 "
c = 13.0 in._
,,,, ,
40
_i
n = 0.85
_-
30
20
./\
t0
0
0
2
Distance
(c)
4
along
Upper-surface
Figure
6
surface
of
A6.-
right
B
from
L.E.,
iO
i2
in.
canard.
Concluded.
L
45
_
i
APPENDIX
70
k
I
I
' I
•
"
i
BO
q = 0.72
l
ic,.=.,63.3in.:..,
50
L!I
4O
30
%
20
,
i i_'
"
i _i
I"
iO
t
0
Relative gauge iO0
reading, in. BO
BO
BO
50
L
40
30
20
C
iO
0
o
2 4 6 B t0 12 :_4 IB 1B 20 22 24 26 2B 30 32 34 3B 3R 40 42 44 4B
Distance
(a)
Figure
A7.-
Indicated
Upper
waviness
Model
along
surface
data
28/29
surface
of
on
right
airfoil
from
L.E.,
in.
wing.
surface
of
Gates
Learjet
airplane,
46
®
i_
APPENDIX
ORIGINAL
PAGE
OF
QUALITY
POOR
IS
Bo x 10-3
-:q
GO
:
0.63
c :
40
16.8
in._...__
'I
3O
r
.
,
• .
,
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t
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0
70
I
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[
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,
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.
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m
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'
""
m
Data
_m
Relative gauge
reading, in. 50
i(
_n
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.
_
= 26.5 in.
i
4o
m
A
L_
\
\ _ -L
20
L.fi
%
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%
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_
10
v
']
O
,,,
i
,i
0
2
4
6
8
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_.2
14
16
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20
Distance along surface from L.E., in.
(b) Upper
surface
Figure
of
right
winglet.
A7.-Concluded.
47
.....
, ....
IT
if
.........
i,_
............................
®
REFERENCES
I.
Loftin,
to
2.
Laurence
Stuper,
J.:
NACA
3.
TM
Jones,
5.
A.
Young,
V.;
in
A.
D.;
stream
6.
on
Young,
A.
7.
Young,
Goett,
Haslam,
pp.
and
the
Matching
of
Size
Layers
on
an
Airplane
Wing
in
Free
Flight.
E.:
D.
the
Boundary
G.:
Note
R.
E.:
Layer
Flight
Drag.
Flow.
Morris,
Boundary
A.
Profile
D.
Layer
on
Layer.
J.
Aeronaut.
Sci.,
81-94.
J.
to
Morris_
and
D.;
Serby,
Finish
Harry
J. E.;
on
J.;
Wing
and
&
Experiments
R.
&
on
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M.
No.
Further
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693,
Bi@knell,
M.
No.
on
No.
on
1800,
Tests
1957,
Note
Rep.
Wetmore,
on
the
British
Flight
B.A.
Boundary
A.R.C.,
Effect
of
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Tests
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Layer
British
Slip-
1939.
on
the
British
Effect
of
R.A.E.,
D.
No.
&
M
in
E.:
Flight
2258,
Comparison
Flight
and
in
of
the
Profile
NACA
Rep.
Tests
British
of
the
on
the
A.R.C.,
Effect
Profile-Drag
Full-Scale
of
1939.
and
Wind
Boundary-
Tunnel.
1939.
at
High
J.
W.;
Determination
Reynolds
Number_.
Zalovcik,
Boundary-Layer
Flow
Morris,
R.
Joseph:
Obtained
Joseph:
Flight
and
Drag.
Bicknell,
Measurements
TN
the
and
Relation
on
A.
NACA
10.
Evolution
1939.
Layer
9.
1938,
and
D.;
Surface
8.
Boundary
Experiments
Jan.
Boundary
Slipstream
Sept.
Flight
3,
Transition
1938.
of
Aircraft:
1980.
1934.
no.
Stephens,
Subsonic
RP-1060,
Investiqation
751,
5,
Jr.:
NASA
Melvill:
vol.
4.
K.,
Performance.
Airfoil
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A.;
and
Platt,
Characteristics
at
High
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and
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667,
Robert
Profile
Numbers.
of
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A
Drag
NASA
WR
an
Airplane
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in
1939.
of
Flight
the
L-532,
Investigation
NACA
1941.
35-215
of
Laminar-
(Formerly
NACA
MR.)
11.
Zalovcik,
John
A.:
Fighter-Type
NACA
12.
of
13.
ACR,
Serby,
J.
14%
Serby,
14.
Tani,
48
E.;
is
-
Morgan,
25%
and
On
The
Investigation
North
B.;
No.
the
and
Wings.
Morgan,
A.:
Obtained
M.
Thick
Delayed.
John
as
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American
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XP-15
Flight
(Air
on
an
Experimental
Corps
Serial
No.
41-38).
Tests
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Profile
1942.
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Itiro:
Zalovcik,
foils
_I
E.;
Drag.
Layer
15.
Nov.
and
J.
Wing
A
Airplane
M.
B.A.
B.:
of
TM
Profile-Drag
in
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&
M.
E.
No_
Note
1360,
Design
NACA
Cooper,
R.
on
British
Airfoils
1351,
R.:
Flight
1826,
British
the
Progress
R.A.E.,
in
Dec.
Which
the
the
A.R.C.,
of
Drag
1937.
Flight
Experiments
on
1936.
Transition
of
the
Boundary
1952._
Coefficients
NACA
WR
L-139,
of
Conventional
1944.
(Formerly
and
Low-Drag
NACA
ACR
AirL4E31.)
ij;
!
16. Zalovaik, John A.; and Skoog, Richard B.: Flight Investiqation of B0undary-Layer
Transition and Profile Drag of an Experimental Low-DragWing Installed on a
Fighter-Type Airplane. _.NACA
WRL-94, 1945. (Former]] NACAACRLSC08a.)
17.
Zalovcik,
John
A.:
acteristics
of
(Formerly
18o
ACR
Zalovcik,
Profile
With
19.
20.
A.;
Drag
of
H.;
Davies,
Gray,
W.
&
Montoya,
Banner,
NASA
29.
Bushnell,
Yip,
D.;
On
Dennis
M.;
Flow
RP-I035,
Long
the
Systems
eds.,
as
and
and
Coy,
II
1945.
High
Speeds
ACR
L6B21.)
Z.3687
Fitted
R.A.E.,
Bramwell,
A.
R.:
of
"Low-Drag"
of
Low
of
Covered
With
Sept.
1946.
Flight
Tests
Design.
R.
&
on
M.
2375,
the
L.;
G.;
AF
Maintenance
Coefficients
Aug.
FZ.440
Layer
Moderate
R.A.E.,
J.
Flight
1945.
To
Investigate
Transition
on
Reynolds
Number
Sept.
1951.
R.
Aeronaut.
of
Laminar-Flow
1981,
and
pp.
Petty,
Soc.,
Laminar
and
a
and
vol.
55,
Wings.
Trujillo,
Advanced
Bianca:
Aerodynamics
-
11-20.
Gilbert,
Flight.
Data
David;
R_sults.
33(657)-13930),
819
Air
Drag
A.R.C.,
"King
Cobra"
of Boundary
Christopher,
Glove
CP-2208,
Tuttle,
Profile
British
1952.
Design
in
British
Research.
Full-Scale
AD
Paul
General
International
and
Char-
L-86,
Jr.:
NACA
RM
Flow
Control
Boundary-Layer-
H58E28,
Northrop
1958.
Demonstration
Corp.,
June
Pro-
1967.
317.)
Marie
H.:
Using
Pressure
Survey
and
Bibliography
Gradient
and
on
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Attainment
Volume
I.
1979.
P,;
Aug.
Hurricane
in Flight
at
193,
British
on
John
in
Control
Canard-Configured
of
NASA
Aircraft
DTIC
NACA
Wings
No.
Flight
Louis
(Contract
from
Laminar
WR
Surfaces
(Formerly
2153,
and
at
Production
Achievement
Wing
Aero
Flow
McTigue,
LFC
on
F.;
the
Note
Steers,
Resear@h,
1946.
Special
A.R.C.,
Laminar
NOR-67-136
(Available
NASA
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Investigation
Having
Aero.
& M.
of
H.:
Measurements
Report
gram.
Profile-Drag
NACA
1946.
Of
No.
British
Natural
Richard
Final
of
Davies,
2485,
Lawrence
TACT
and
Airplane.
Flight
Tests
on
of the Position
Aspects
and
No.
R.
Section
Memo.
325-361.
L-98,
Smith,
for
Aerofoil.
Tech.
L.:
With
Requirements
Some
No.
Transition
28.
J.;
Fitted
pp.
E.;
M.
Boundary-Layer
P-47D
Measurements
Sept.
Handel:
1951,
a
Airpl_ne
WR
A.R.C.,
Drag"
Selected
27.
Flight
Smith,
F.; and
Higton,
D. J.:
Britland,
C, _ M.:
Determination
F-111
26.
Fred
Rep.
D.
Z.3687
Practical
"Low
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2546,
of
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Drag
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R.
25.
a
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H.:
II,
June
24.
of
R.
Specially-Prepared
Mach
Number.
23.
Daum,
NACA
R.
a
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Wing
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the
21.
22.
John
Drag"
of
Sections
LSH11a.)
Paint.
Plascott,
on
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Camouflage
Plascott,
"Low
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the
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of
Proceedings
2,
Sciences
B.
as
Laschka
a
Full-Scale
of-the
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AIAA
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ICAS-82-6.8.2.)
49
®i
_i _
30.
Braslow,
Albert
Cri£ioal
tion
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at
Holmes,
Mach
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37.
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Dan
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1980.
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Manley
J.; and
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1982.
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45.
S.
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Mead,
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Laminar
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1983.
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Airspeed
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Solids
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Natural
From
the
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1958.
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f.._. Boundary-Layer
4363,
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to
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Of
830717,
R.
Summary
NACA
On
an
TN
Gregorek,
Reference
Viscid/Inviscid
R.:
Through
41.
Von
E._
of
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Transition
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R.
NACA
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O.:
Hulti-Component
of
tation
40.
H.:
Richard;
ysis
A.
Low-Speed
Data.
Eppler,
5.
Paper
Airplane.
A.;
Ira
Airfoil
SAE
Program
Dimensional
Particles
1950.
Bruce
Stevens,
II.
Chemical
J.:
Experimental
Simplified
to
Flight
Boundary-Layer
Bruce
C.:
Roughness
Clifford
Particular
Research-Support
35.
0
Transition).
cations
Holmes,
Eugene
J.:
Ormerod,
Boundary-Layer
34.
From
Rick
(With
Main-Smith,
Knox,
Distributed
Sk?_ocket
Airstream
33.
and
of
Numb_rs
Bruce
and
L.;
Height
ACR,
J.:
of
Novo
a
NACA
Wind-Tunnel
High-Speed
F_ward:
on the
Effects
N.A.C.A.
of Propellers
27-212
Airfoil.
and of
NACA
Vibration
WR L-784,
ACR,)
Investigation
Pur§uit
Airplane
of
With
Several
Factors
Air-Cooled
Affecting
Radial
the
Engine.
1841.
5O
®
46.
Beasley,
J.
A.:
Transition
47,
Gilkey,
NASA
48.
Gliek,
Bull,
49.
R.
D.:
P.
Freeman,
J.
14,
Atkins,
P.
52.
The
A.:
Studies
D.:
Lachmann,
53.
Volume
54.
55.
57.
Aspects
Tests
of
and
British
Prediction
A.R.C.,
Winglets
S.:
on
a
of
_976.
DC-10
Wing.
M.
David
a
M.;
and
Insects
to
to
by
by
300
the
Aerial
Feet.
Air.
Tech.
Currents
J.
May
Anim.
- The
Ecol.,
Insects.
J.
Note
17,
on
Aircraft
Wings.
in
Relation
to
Laminar
Flow
1960.
Boundary
Pergamon
J.
Flight
1952.
Contamination
W.:
VTH-LR-326,
Insects.
Contamination
R.A.E.,
A.R.C.,
ed.,
Selen,
Airfoil.
Layer
Press,
1961,
On
Design
Dep.
Co.:
B.,
William
Porous
NASA
the
Aerospace
Jr.;
and
Its
CP-2036,
I,
1978,
and
G.;
Leading
CR-165444,
Evaluation
Transport
Contamination
Part
in
1951.
Insect
Insect
Schweikhard,
Commercial
John
Mites
pp.
of
Eng.,
and
Flow
Control,
682-747.
Some
Airfoils
Delft
for
Univ.
of
1981.
L.;
Lockheed-Georgia
of
Level
Oct.
of
Due
Lachmann,
Glycol-Exuding
Aviation
and
1939.
Contamination
British
Roughness
V.
Spiders,
May
Ground
British
of
484,
Apr.
Peterson,
Layer
3787,
Distribution
From
F_ge
2164,
No.
G.
the
Air
Measurements
Aero.
Technol.,
Subsonic
Boundary
No.
Insects,
(Melbourne),
Application.
eral
_klnnel
of
Leading
C.P.
2,
of
in
V.:
L.
Kohlman,
M.
Agriculture,
the
Labs.
Sailplane
Tests
56.
Wing
W.
Boermans,
&
128-154.
Brief
G.
Coleman,
pp.
No.
Aircraft.
Wind
Dep.
of
Res.
Note
Laminar
R.
Distribution
u.S.
B.:
Tech.
and
673,
1945,
Johnson,
the
Wing.
197g.
Population
vol.
of
Sheared
Design
A.:
No.
Aeronaut.
51.
a
CR-3119_
Inse_t
50.
Calou]dtion
on
of
Fisher,
David
CTOL
Albright,
Ice
Alan
Protection
E.:
Icing
System
on
Tunnel
a
Gen-
1981.
Laminar
Aircraft.
Alleviation.
pp.
and
Edge
Flow
NASA
F._
Control
CR-159253,
Flight
Transport
System
Concepts
for
1980.
Investigation
Technology
of
-
1978,
Insect
NASA
357-373.
51
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ORIGINAL
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URIGINAL
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ORIGINAL PAGE lg
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56
7L I
_7
TABLE
Gross
weight,
3.-
PHYSICAL
CHARACTERISTICS
OF
VARIEZE
ib .....
eeeoeeooeeeeQeeeeaI0eeoQe,eQeeQeeeeleleQQ,_iQe_eie_o_etiQ_._eo
1050
Wing:
Area,
ft 2
Span,
in ....................................................................
Aspect
ratio
Tap_r
Root
(main
chord
aerodynamic
Root
chord
Twist
wing)
(washout),
Incidence
. .......
....
.................
9.28
.................................
0.44
LS(I)-0417
(Modified)
35.75
16.
in ..................................................
31
in .....................................................
88
in ......................................................
deg
deg
53.6
267.6
in ...................................................
chord,
(strake),
Dihedral,
............
. ..... ..................
in ...................................................
wing),
(strake),
chord
......
,.. ....
.....................................................
wing),
(main
_ ...................
.........
(main
(main
chord
.....
wing)
section
Mean
Tip
.......
........
ratio
Airfoil
Tip
....
..................
35.75
* .....................................
3.0
...............................................................
at
root,
deg
.........................
Sweep
at
leading
edge
(main
wing),
Sweep
at
leading
edge
(strake),
deg
deg
-4.0
. ...........................
.
1.2
......................................
.........
27
................................
61
Canard:
Area,
ft 2
Span,
in .......................................
Aspect
.....
ratio
Taper
Airfoil
section
deg
deg
Incidence
at
at
winglet
chord,
141.6
.
11.32
1"O
11)
in ..............................................
.....
13
, .............................
0
................................................................
root,
leading
deg
edge,
.......................
deg
0
................................
O
...................................................
0
(upper):
Length,
Root
Tip
12.3
.............
isee'_ble'4;'[[[[j[[[2[[[[[2[2[[[[[[[[jfjj[j2j[[22j2j[J'GU25_5(
....................................
Dihedral,
Sweep
.....................
. ...............
....
aerodynamic
Twist,
. .......
................................................................
ratio
Mean
. ................................
in ................................
chord,
in.
chord,
aerodynamic
Area
(projected
Taper
ratio
Sweep
at
Airfoil
ft 2
.......
vertically
, ........
14.5
......................................
projec£ed
geometry)
3.35
........................
2.6
..................................................................
....
at
angle,
on
7
in ..........................................
vertically),
(based
leading
deg
Incidence
Cant
chord,
ratio
Twist,
36
20
in ................................................................
Mean
Aspect
. ..................................
..............................................................
edge,
deg
....................
0.35
, ...............................
...................................................
root,
deg
deg
...............
. ....
...................
section
29
• ...........
0
. ..................................
0
...........................................
5
.......................................................
See
table
2
Powerplant:
Manufacturer
Model
....
Takeoff
and
Revolutions
Propeller
maximum
per
(fixed
Manufacturer
..................
Teledyne
Continental
.....
....................
, ................
continuous
minute,
maximum
power,
hp
Motors
Corp.
0-200A
, ..................................
................
......
..........
100
. ..........
.
2750
pitch):
.............
Number
of
Pitch,
in ........
Diameter,
..............
............................
blades
..................
Ted
Hendrickson,
...................................
in .........
....
, ..........
, ..........................................
, .....................
...................
Snohomish,
, .....
......
Washington
,..,
.....
. ....
.............
2
..
70
..
56
57
®
OI_IGINAL
_F POOR
TABLE
AIRFOIL
4.-
DESIGN
AND
COORDINATES
CANARD
(a)
(x/c)
U
0 00000
00500
01000
02600
03000
04000
05000
O600O
07000
08600
09000
10OOO
15000
20000
25000
30600
35000
40000
45000
50000
55000
60600
65000
70000
75000
80000
85000
90600
95000
1 00000
(z/c)
OF
VariEze
u
0 00000
02090
030ZO
03910
04750
05310
05870
06280
06620
06980
.07320
,07630
.08770
.09690
10340
10700
11000
11060
11030
10950
10590
10060
09100
08000
06980
05730
04190
03040
01.680
00000
PAGE-Ig
QUALrT'V
_R
VARIEZE
wing
WING
AND
at
AND
WINGLET
OF
VARIEZE
LONG-EZ
BL32
(x/c)
L
0.00000
,00500
.01000
.020£0
.03000
.04000
.05000
.OGO00
.07000
.08060
,09000
,!0000
,15000
,20000
.25000
.30060
,35000
,40000
45000
50000
55000
60060
65000
70000
76000
80000
85000
90060
35000
I 00000
(z/c)
L
0.00o00
- 01120
- 01760
62370
02790
0321.0
03600
- 03910
04230
- 64476
-.04610
--.{14030
-.05680
-.06150
-.0G540
-.0679C
-.06955
-.06899
-.06676
- 06310
- 05720
- 05'1.16
- 04190
- 03320
- 02680
- 01930
- 01120
- 60596
- 00170
00000
58
®
TABLE
(b)
(x/c)
U
0 00000
00250
00500
01060
015OO
02000
03000
04000
06000
08060
10000
12500.
15000
175O0
20000
,22500
25000
27500
30000
32500
35000
37500
40000
4250045000
47500
50000
52500
55000
57500
GO000
62500
65000
.67560
70000
72500
75000
77500
80000
8256085000
87500
.90000
,92500
,95000
.97560
1.00000
(z/c)
4.-
VariEze
U
0 00000
01150
01550
02150
02600
02900
03500
04000
04950
65706.
06500
07300
08050
08700
09250
69806
10300
10750
1.1150
11500
11700
11906
12000
12050
12000
11900
11700
11406
11000
10650
10200
09700
09150
08556
08000
07350
06700
06000
05250
04556
03850
03100
.02400
.01650
,00850
.60006
,00900
ORI_NAL
PAGE
19
OF POOR
QUALITY
'iI
Continued
winqlet
root
(x/c)L
(z/c)L
0 00000
00250
00500
61006
01500
02000
03000
04000
06000
08006
10000
12500
15000
17500
20000
Z_501,
25000
27500
30000
32500
35000
37506
40000
42500
45000
47500
50000
52506
55000
57500
60000
62500
65000
67506
70000
72500
75000
7.7500
.80000
.82501;
.85000
.87500
,90000
.92500
,95000
.97506
1.00000
0
-
00000
00700
01i-50
0170rl
02050
02300
02700
02950
03200
03400
03500
03650
03750
03800
03800
03800
03750
03700
03G50
03600
03550
{ )._:,500
"%
-.03450
- 03300
•- 03150
03050
02950
- 02800
- 02650
02450
02200
-.02000
-.01750
-.01500
--.01300
-,01100
- 00950
- 00750
- 00550
- 06400
- 00300
- 00200
- 00250
- 00350
- oo55n
- 06800
- 0 Ir)O0
59
®,
/
_IOINJU,, pAOli Ill
_I p_OR OUAUTY
TABLE
(C)
4.-
Var_.Eze
Continued
winglet
tip
i
(x/c)
U
0.00000
00430
00710
01430
02860
04290
05710
08570
11430
17140
-2660
• ,"8_,70
,34290
.40000
.45710
.51z,30
,571 40
,62860
.68570
74290
80000
85710
91430
97140
1 00000
(z/c)u
00000
01570
02140
02860
03710
04430
05000
06140
07140
085_0
09860
10710
11290
11930
11930
11930
.11000
.10140
.09140
.07710
.05860
.04060
,01860
,00430
.01430
(x/c)
L
0 00000
00430
00710
01430
02860
04290
05710
08570
11430
17140
22860
28570
34290
40000
45710
51430
57140
62860
68570
74290
80000
85710
91430
97140
1 O0000
(z/c)
L
0.00000
-,01570
-.02140
- 0300(;
.- 03710
- 04140
- 04290
- 04570
- 04860
- 65000
- 05290
..................
=..05140
- 04860
- 04570
- 04290
- 63570
02860
- 02290
- 01860
- 01.570
- 01430
-,O129O
-,01140
-,01290
-.01430
60
®
TABLE
(d)
(x/c)
U
0,00000
.00130
.00492
.01057
.01777
.02676
.03907
05088
07473
10035
12444
15030
17745
20049
22661
25016
27677
30006
32718
34970
37"502
.39902
.42566
.44662
.47449
49954
52330
54555
57413
59865
62444
64946
69951
74980
79934
84963
89993
.94869
1.00000
VariEze
(z/c)
ORIGINAL
PAGE IS
OF POOR
QUALITY
4,-
and
U
0 00000
00583
01387
62288
03-154
04068
05166
06099
07725
09211
10402
11494
.12488
13280
14103
14713
15771
16248
16539
16757
1691_
17.041
17077
16972
16686
16275
15758
15045
14332
13530
12720
10997
09171
07345
05496
03638 .....
01843
00025
Concluded
Long-EZ
(x/c)
canard
L
0.00000
.00076
.00203
60406
00814
01707
02678
03827
05079
07482
10063
12747
1.5074
17784
0
,0034
22668
25019
27550
30056
.32512
.35043
.:,7606
39978
42536
45041
50003
55066
59977
65015
69951
74937
79976
84937
89949
94477
1.00000
(z/c)
L
0 00000
- 00273
00438
- 06651
- 00955
01412
01749
02030
- 02256
- 02558
- 02752
- 02822
o
- 0_946
-.03013
-.03046
• U
OUl_,.
- .03143
- 03178
- 03145
- 03076
- 03017
- 02990
02925
•- 02815
- 02700
-. O2472
- 02253
02057
- 01850
01639
- 01417
- 01183
-.00924
-.00617
-.00312
-.00025
" :
-
'
TABLE
Gross
weight
]b
•
,
_
•
e
•
.'go
5.-
qe
00
o
PHYSICAL
.o
•
•
•
•
,
CIIARACTERISTICS
e
...,..go
OF
eooeoooo
o.
LONG-EZ
ooe.ooe(Jooeoe
1325
o.oo.o.oeoo
Wing:
Area,
ft 2
in .
Span ,
Aspect
Root
(main
section
chord
.... .....
wing)
(main
(main
chord
Mean
Root
Tip
..........
ratio
Airfoil
.... .....
........
. .... ..,,......,,...%.._..........
.....
(main
....
..........................
i .........
wing)
wing),
81.99
313 • 2
"
eeeoeeee0eeo0ooeeeeoeo.oeeieoeeeeoeeeeoeo.ee.ee.eooe..eeoeo..eo0o..
ratio
Taper
Tip
.,..............
.... .. .....
....
8.3
, .....................................
0,48
...-........................................
See
table
in ..................................................
wing),
in ...................................................
20
aerodynamic
chord,
in .........
.... . ....................................
chord
(outboard
strake),
in ............................................
chord
Twist
(outboard
(washout),
4
41.4
strake),
deg
37.6
76
in ......................................
• ......
...................................................
41.4
BL157:-2.7
BLI06.25:-0.46
BL55.5:-0.6
Dihedral,
deg
Incidence
at
.. .... .......
root,
deg
............o.........,...........oo...............
Sweep
at
leading
edge
(main
Sweep
at
leading
edge
(outboard
Sweep
at
leading
edge
(inb0ar_
Canard:
Area,
Span,
Airfoil
J ...............
...
(see
aerodynamic
deg
Incidence
at
51
64
chord,
2)
12.8
..........................
....
......
• .......................
141.6
.,.
10,88
• ................................
1.0
.........................................
GU25-5(11)8
in ...................................................
13
• ..............................
0
................................................................
root,
leading
deg
edge,
...............
deq
......
0
. ..........................
.. .........
. .... , .........................
. ...........
..
0.6
..
0
(upper):
Length,
_oot
in ........................................
chord,
in ................
chord,
aerodynamic
Area
(projected
ratio
Taper
ratio
Sweep
at
Twist,
chord,
edge,
...........,
at
root,
Cant
angle,
deg
Airfoil
section
Powerplant:
Manufacturer
.deq
deg
...0.
49
..
27._
(fixed
.. ............................................
projected
geometry)
. .... . ...........
. .....
.,.....
... .......
......
. ......
. .........
• .....
.....
•...........
• ........
.,,..,,.,
. ......
. .......
. .... • .......................
continuous
minute,
6.57
........................
..................................
........
.......................................................
maximum
per
11
20.5
2.54
. .........
.....
0.40
...................................................
......
..............
and
Manufacturer
Number
of
ft 2
vertically
........
..............
Revolutions
Propeller
on
.....................
deg
Takeoff
• .....................
. ................................
in ...................................................
vertically),
(based
leading
Incidence
Model
. ......
in .............................................................
Mean
Aspect
maximum
power,.hp
28
.......
......
............
•...........
.... .... .......
.. .........
• .....
.......................
...........................
Avco
0
•
0.5
.....
See
table
Lycoming
_ ...............
. .............
• ................
0
4
Corp.
-0.235
118
2800
pitch):
..... i .......................
blades
-... .........
.........
.. Ted
Hendrickson,
Snohomish,
.... ..........,._...,....i............o..
Pitch,
in ......
. .................................
Diameter,
in.
....._ ..... .. .... ,......
.......
62
.................................
.....................................
deg
at
table
23
..................................
.....................................
Dihedral,
Tip
deg
deg
................................
section
Twist,
Winglet
strake),
strake),
..
..........
0
.......................................
...................................................................
ratio
Sweep
deg
in.
Aspect
Mean
wing),
ft 2
Taper-ratio
0
.......................................................
Washington
-..... , ...........
• ...........
... .... ...............,...,.......
2
70
58
TABLE
6.-
AIRFOIL
DESIGN
(a)
•_/e)u
t
o.o00GG
00296
00688
01236
01937
02789
03787
04928
06206
.07618
09157
10817
12591
14478
16479
18594
20823
23166
25623
.28190
.30860
.33627
,36482
39414
42415
45472
48574
51.708
54861
58019
51167
64290
67373
70399
73353
76219
78980
,81622
,84128
.85484
88676
90691
92517
94141
95562
96797
97859
98742
93418
99850
I 00000
[i_3R
COORDINATES
WING
AND
(x/c)L
0,00581
01331
02129
02958
03807
04665
05521
06366
07191
07987
08744
09454
10105
10683
11180
11591
11911
12138
12274
12325
12294
12186
12005
11758
11449
II085
1067"1
10214
09722
09200
08655
08094
07524
06950
06378
05812
0,00002
(I0126
00507
01154
02015
03077
0_4334
05789
07441
•09284
11309
13505
15860
18361
20997
23753
26617
29574
32610
35711
38861
42045
45249
48460
51674
54884
58083
61264
64419
f_7547
70000
72000
74000
76860
79400
.82200
.85O00
.88000
,89800
.92800
• 952[} 0
.98000
1.00000
04720
04199
03699
03221
02765
•02331
.01915
,01 c
.01095
.00715
,00396
,00167
.00089
,O(iOlO
OF
(z/e) L
-0.110103
- 00675
01165
- 61641
- 02102
02538
- 02937
- 03294
- 03613
-.03898
- 04152
- 04378
.- 04575
- O4745
- 04887
- 05006
- 05084
- 05139
- O5162
- 05153
- 05110
- 65031
- 04910
- 04739
- 04516
- 04239
- 03912
- 03537
- 03116
- 02648
-- 02400
- 02200
- 02000
- Cl80O
- 01600
- 01400
- 01200
- 01000
- 00800
- 6060O
- 00400
- 00200
- 00010
ORIGINAL
-
LONG-EZ
Winq
(z/e)u
EOE
OJ_-,8
WINGLET
PAGE
I@
OF POOR QUAL-tTY
i
/
ORIGIINAL PAGE IS
OF POOR QUALITY
TABLE
6,-
Continued
(b) Winglet
(x/e)u
0 0000O
00000
00196
00644
01175
02265
03607
05621
07858
10653
12891
15435
18314
21977
.25415
.29944
,33970
.38331
.41798
,45991
49122
53120
56251
60639
69525
68272
72102
77748
82528
87588
91838
86366
00000
64
root
(z/c)u
(x/OIL
(z/C)L
0,00000
0000900846
01516
02010
02727
03461
.04392
,05205
.06062
,06538
,07007
.07295
,07499
,07656
.07731
,07625
.07377
,07154
,06848
,06576
06_o5
0 00000
00000
00140
00280
00866
02012
03018
04696
07128
09840
14005
16884
.20266
,23342
,27451
,31812
34943
38354
42016
46154
50767
55129
'60217
64858
,69862
.73608
,79060
.84288
,89124
,95052
,99972
o ooooo
05974
05562
05116
04599
04060
03310
02686
.01996
.01406
,00774
,00296
- 00009
-. 00783
- 61116
- 01798
- 02324
- 02537
- 02906
- 03141
- 63289
-- 03528
- 03737
- 03961
- 04081
- 04211
- 04301
- 04276
- 04154
- 04021
-,03892
-.03753
-.03580
-,03289
,03005
-,02713
-,02466
-,02080
-.01721
-,01397
-.00946
-,00292
ORIGINAL
OF POOR
TABLE
6.-. Concluded
(c).-Winqlet
¸
tip
¢xlc)u
(z/c.).u
(×/C)L
(zlC)L
O, 00000
,00000
,00026
.00212
.00503
,00793
01215.
01665
02244
02S52
03617
0 00000
(i 0018
0fi283
00809
01365
018.14
02315
02754
03245
03695
04213
05134
06052
06838
07551
08187
08619
08940
09211
(19431
09609
,09815
,09818
09663
09418
0£132
08802
EO
084_,_
08048
,07621
,07147
,06575
,05958
05382
04719
04109
03436
02869
0-2194
01408
00695
00185
0 00000
00000
00053
00210
00394
00841
01369
0210702923
03977
07508
11172
14940
18656
22346
2601O
29489
33311
37185
40744
44302
48057
,_. _'_1124
Lr
._,_821
59037
52938
66549
70371
13930
78200
82734
86768
,91196
,94366
,97620
,99947
0 (10000
- 00018
--00370
- 00727
- 00951
- 01251.
- 01495
- 017%7
-01968
- 02189
- 02793
- 03275
-,03597
-- 03770
- 03885
- 6395?
- 04012
- 04061
- 04052
- 03976
- 03876
- 6374:';
-. 03527
05180
06966
08944
11106
1321-6
15114
16828
18647
20492
22311
26654
30034
33566
,37308
41103
44872
r,
48077
52225
55967
59550
63346
67035
,70513
74518
7-8129
82003
85191
88959
9,:,o,35
97392
1 00000
PAGE IS
QUALITY
""
%'13"%_
J,,:
L i
-
- 02998
19
t,...710
- 02459
- 6218_,
- 01933
- 01632
- 01332
- 01072
-00779
-, 00559
- ,00323
-.00184
®
TABLE
7.-
+PHYSICAL
CHARACTERISTICS
OF
BIPLANE
RACER
!+ •
Gross
Wing
+
weight,
................................................................
i200
(forward):
Area,
f+t2
Span,
in ....................................................................
Aspect
Airfoil
section
wing)
(main
(main
chord
Mean
(main
Anhedral,
chord,
deg.,
deg
Incidence
Sweep
0.68
...........................................
See
table
root,
leading
6
in ...................................................
37
in ....................................................
26
in ...................................................
32
.......................................................
0
................................................................
at
at
wing)
wing),
(washout),
6.68
....................................................
wing),
aerodynamic
Twist
Wing
(main
chord
47.6
213.6
................................................................
ratio
Root
Tip
..................................................................
ratio
Taper
_;+
Ib
deg
edge,
6.5
.......................................................
0
deg
6
...................................................
(aft):
Area,
ft 2
Span,
in ...................................................................
Aspect
...................................................................
ratio
Mean
Tip
chord,
deg
Incidence
at
table
in ...................................................
.... .....,.,
deg
at
chord,
See
6
23
, ....................................
in ................................................................
Dihedral,
Sweep
0.52
......................................................
in ...................................
chord,
Twist,
11.48
................................................................
section
aerodynamic
Root
270
...............................................................
Taper-ratio
Airfoil
44.1
29
16
......,..,.....,,,..,...,.,,,..,,,,,,....,.,,.,...,,..,
0
.,..,..........,,...,......,,..........o..,,,..,,,.,,,,o.,..,o,.
root,
leading
deg
....................
edge,
deq
4
. ..................................
...................................................
0
3.2
Powerplant:
Manufacturer
Model
..................................................
Takeoff
and
Revolutions
Propeller
Number
Diameter
66
Avco
.....................................................................
maximum
per
(fixed
of
,
maximum
power,
hp
....................................
.............................................
Corp.
I0-320
160
2800
pitch):
blades
in
continuous
minute,
Lycoming
•
io,oQoQIooQeleOOOlileQl•oleollooloolieoooloeQtlleeelo,eol,ool
•
°oele0+eeoOlOleeeollloooeoloooooloooleeloeelOllllesloleeOlllleo
2
60
ii
TABLE
8.-
AIRFOIL
DESIGN
WINGS
COORDINATES
OF
(a)
(x/c)u
irl
0.00000
.00194
.00388
.00775
.03000
.OGO00
.09000
12000
15000
18000
21000
24000
27000
.30000
.33000
.36000
.39000
.42000
.45000
.48000
.51000
.54000
.57000
.60000
.63000
.6600O
.69000
.72000
.75000
.78000
.81000
.84000
.87000
,90600
,93000
.96000
.99000
1.00000
;,Li
(z/c)
BIPLANE
Forward
U
FOR
L
o.ooooo
00977
01353
01857
03318
04349
05054
05609
06070
06465
06810
.07109
07357
07562
07717
078_2
07880
07884
07833
O7733
07574
.07364
.07101
.06787
.OG42G
.0G023
.05578
.05097
04585
04054
03504
02942
02384
01837
01310
.00814
.00360
.00221
AND
AFT
winq
(x/c)
o ooooo
FORWARD
RACER
.00194
.00388
.00775
.03000
.OGO00
.09000
.12000
.15000
.18060
.21000
.24000
.27000
.30000
.33000
.36060
.39000
.42000
,45000
.48000
.51000
,54000
.57000
.GO000
.o3000
.66000
.69000
.720O0
.75000
.78000
.81000
.84000
.87000
,90000
.93000
.96000
.99000
1.00000
(z/c)
L
0 00000
00934
01267
01690
02698
- 0316}
- 03364
-.03465
-.03523
-.C3566
-.03605
-;03640
-.03674
-.03705
-.03729
-.03740
-.03744
-.03725
-.03690
-.03632
-.03550
-.63446
-.03314
-.03159
-.02977
-,02775
-.02550
-.02316
-.02054
-.01791
-.01523
-.01260
-.01004
-.00767
-,00558
-.00380
-.00252
-.00221
67
t
...............................
._ ........ , _,, ;_...u. ............
L. "Z,',
TABLE
Z
8.-
(b)
:c
Concluded
Aft- winq
(x/c)i_
( z/c)u
(x/c)L
(z/c)L
000000
00288
00575
01151
02998
05996
08994
11992
14990
17988
20986
23984
,26982
,29980
,32978
,35977
.38975
,41973
,44971
,47969
,50967
,_3965
,56963
,59961
,62959,65957
,68955
.71953
.74951
.77949
,80947
,83945
,86943
,89941
,92939
,95937"
,98935
1,00000
0 00000
01168
01600
02158
03159
04022
04581
,05001
.05346
,05645
,05910
,06146
,06347
,06514
,06646
,06744
.06802
,068t9
,06790
,06721
.06600
.064._3
,06226
,05967
.05668
,05329
,04949
.04534
.04097
,03631
.03148
.02659
.02164
.01675
.01203
.00760
,00357
.00224
0,00000
,00288
.00575
,01151
,02998
,05996
,08994
,11992
,14990
,17988
.20986
0,00000
-,01145
-,01554
-,02072
-,02929
-,03585
-,03942
-.04183
-.04362
-.04506
-,04627
-,04730
-,04816
-,04885
-,04932
-,04960
-,04960
-,04932
-,04874
-.04782
-.04681
-,04511
-,04322
-,04109
-.03861
-,03585
-,03286
-,02969
-,0264!
-,02296
-,01951
-,01611
-,01283
0097"
-.00G96
-.00460
-.00276
-.00224
t%
"lCl_Oh
26982
29980
32978
,35977
38975
41973
.44971
,47969
.50967
.53965
.56963
,5996I
62959
65957
68955
71953.
74951
77949.......................
80947
,83945
,86943
,89941
,92939
,95937
,98935
1.00000
-- .
(L
68
®
i
'I
TABLE
Gross
weight,
9.-
ib
PHYSICAL
CHARACTERISTICS
OF
GATES
LEARJET
MODEL
28/29
..............................................................
15
000
Wing:
Area,
ft 2
Span,
in ....................................................................
Aspect
ratio
Taper
(main
chord
Mean
(main
sweep
at
Winglet
wing),
in ..................................................
Root
in ..................................................
83
........................................................
0
.......
leading
edge,
deg
..o ...............................................
17
in ..................................................................
chord,
chord,
Area
(projected
ratio
ratio
Sweep
at
Twist
(leading
on
ft 2
20.76
............................................
vertically
projected
geometry)
6
.......................
2.33
.................................................................
leading
at
angle
Airfoil
9.99
in ..................................................
vertically),
(based
Taper
Incidence
28.53
in ...............................................................
aerodynamic-chord,
Aspect
44.9
in ..............................................................
Mean
thrust_
edge,
edge
root
deg
outward
(leading
(wingle_
section
Powerplant:
Manufacturer
Model
Rated
112
43.8
(upper):
Length,
Cant
0.39
in ...................................................
deg
deg
6.48
.....................................................
chord,
(washout),
Dihedral,
506.4
wing)
wing),
aerodynamic
Twist
Tip
(main
chord
274.3
.................................................................
ratio
Root
Tip
...................................................................
tip
0.35
...................................................
within
edge
canted
toed
out)
lower
out),
40-percent
deg
span),
deg
40
..............
I
..............................
-2
.........................................
.................................
LS(I)-0413
.....................................................
{b'_[[_[i_[_[_[[_i_[_i_[[_[[[i[i[_[_[[_[[_[_[[_[_[[_.C.
thinned
15
to
General
t/c
=
0.08
Electric
J610-8A2950
69
*
_mlllmmrw---
• ,°
TABLE
Gross
ight
we
Wing:Area,
Ib
,
ft 2
Span,
Aspect
10.-
PHYSICAL
CHARACTERISTICS
OF
CESSNA
4066
eleoleeeeee&eelellel_eeeleeleeeeeeeeeleloeloleeeeeeeleeee-oeeee.
....................................................................
175
ft
........
ratio
[[[[[[[[.[[i[[[[[[[[[[[[[i[[.[[[[[[[[[[[[[[.[[[[[_[[[.[_[[[[[[[[[
Taper
ratio
Airfoilsection:
Root
(main
wing)
.....................................................
chord,
Mean
441
7.72
......................................................
chord,
(washout),
Dihedral,
deg
Incidence
at
Sweep
642A215
NACA
641A412
. ......................
(a
=
0.5)
(a
=
0.5)
70.8
in ................................................................
aerodynamic
Twist
0.70
NACA
Tip
......................................................
Root
chord,
in ........................................
Tip
P-210•CENTURION
at
deg
61
. ........................................................
3.0
...............
root,
deg
leading
50
in ...................................................
+ .................................................
2.6
.....................................................
edge,
deg
.....................................
..
1.5
, .............
0
Powerplant:
Manufacturer
Model
....................................
Takeoff
and
maximum
Revolutions
Propeller
per
Number
c6ntinuous
minute,
(constant
Manufacturer
of
blades
McCauley
in ...............................
310
, .........
Accessories
Div.,
Cessna
Aircraft
2700
Co.
3
80
....................................................................
ratio
Root
.....
. ..................................
wing)
chord,
deg
Incidence
at
. ............
.................................
in .........................
......
in ................
chord,
root,
leading
deg
edge,
0005
56
33
. ....................
45.5
, ........................
......
deg
0009
NACA
................................
in ..............................
0.58
NACA
+ ...............................................
..........................................
at
. .......
. ...........................
aerodynamic
Twist,
156
3.5
..............................................
.....................................................
......
chord,
48
.................................................................
Taper
ratio
(main
Airfoil
section:
7O
...................
tail:
Aspect
Sweep
....
.................................................
Span,
Mean
..............
Corp.
TSIO-540-P
.............................................................
ft 2
Tip
hp
Motors
.............
....................................
Area,
Tip
Continental
speed):
in ................
Horizontal
power,
maximum
.......................
Diameter,
Root
Teledyne
.... .................................................
. .........................
...............
, .... + ......
, ............................
0
, ......
,.,.
. ......
-3.6
8
-
TABLE11.- PHYSICAL
CHARACTERISTICS
OF BEECH24RSIERRA
Gross weight, ib ...............................................................
Wing:
Area, ft 2 ........................
Span,
ratio
146
Twist
7.34
..................................................................
deq
Incidence
at
1.0
......................................................
chord,
in .....................................
(washout),
Dihedral,
at
393
.................................................................
ratio
Airfoil
section
Mean
aerodynamic
Sweep
, ...........................................
in .....................................................................
Aspect
Taper
2750
deg
....................................
.......................
root,
leadinq
deg
edge,
NACA
............
632A415
_. 52.8
. ....................
-2.0
•.........................................
6.5
.......................................................
deg
3.0
...................................................
0
Powerplant:
Manufacturer
..................................................
Model
.................................................................
Takeoff
and
Revolutions
Pr6peller
maximum
per
(constant
continuous
minute,
maximum
power,
hp
Avco
.....................................
200
..............................................
2700
speed):
Manufacturer
Number
of
...............................................
Hartzell
blades
.............................................................
Diameter,
in .................................................................
Airfoil
section
Chord
&t 0.25d,
Lycoming
Corp.
IO-360-AIB6
Co.
-2
76
I' .........................................................
in ...........................................................
r
Propeller
Clark
.
Y
6.5
71
TABLE12.- PHYSICAL
CHARACTERISTICS
OF BELLANCA
SKYROCKET
II
Gross weight, ib ..............................................................
4100
wing:
Airfoil
..............................................................
Area, ft 2 .......................
. ............
. ..............................
Span, ft ....................................................................
Aspect ratio ................................................................
Taper ratio .................................................................
Root chord, in, .............................................................
Tip
chord,
Mean
in ...................................
aerodynamic
Inciden&e,
chord,
deg
NACA632-215
182.6
35.0
6.7
0.57
80.2
, ...........................
45.9
in ..................................................
64.6
..............................,...........................o...
2
Oihedral
deg,
Twist,
Sweep
_eg
at
2
................................................................
leading
edge,
deg
...
..................................................
3
2.8
Powerplant:
Manufacturer
....................................
Teledyne
Model
..................................................................
Maximum
continuous
Revolutions
Propeller
per
(constant
Manufacturer
Model
Number
power,
minute,
hp
maximum
Continental
.................................................
435
.............................................
3460
speed):
...............................................
Hartzell
........
. ................................................
of blades
.......
% ....................................................
Diameter,
in.
Motors
Corp.
GTS10-520F
....................
Propeller
Co.
HC-H3YN-1RF/F8475-4
...........................................
3
82
;_ i_
Revolutions
"_.
72
per
minute,
maximum
............
, ................................
2270
ORIGINAL PAG_ |g
OF POOR QUALITY
Fs 118.6)
1
I
I
FS(O)
FS (99)
BL (56.6)
Elevator
'
FS.(121}
BL(99,5)
90_
BL{
==::==_
I
f_f
FS(O)
Figure
I.- Geometric
characteristics
Of VariEze
in inches.
__.
airplane.
_
Dimensions
are
?3
®
i:
OfllOINAL PA_£ I_
OP' POOR
OLIALITY
I
_r
i
+
.................................
1+,.I ,-[r.
i.+, t+l, i p, I r + _+
........
ORIRINALPAGE19
OP'POORQUALITY
!i
75
,----7"I __ -_---qnlmwmnm_-_
>-
I
76
OIIlalNAL PAQ| IR
OF t_00_ QUALITY
oJ
_J
°_
_e
_J
0
0
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L-81-8426
(a) Top
Figure
19.-
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boundary-layer
R = o,62s × 1o6 ft -_;
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26,-
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103
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Inboard
wake
36.-
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probe
station;
of
predicted
Bellanca
130
Rc
=
9.04
pressure
Skyrocket
II
x
10 6 ;
C_
distribution
(right
wing).
= 0.288;
M
=
0.31.
with
transition
for
ORIGINAL PAGE IS
OF POOR QUALITY
-.8
".4
Cp 0
.4 -
FLIGHTMEASURED TRANSITION
PREDICTED TRANS ITION (Granville
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1.2
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Outboard
wake
probe
station;
Figure
Rc
36.-
=
8.39
x
106;
C£
=
0.254;
M
=
0.31.
Concluded.
l_?_,
.
131
ORIGINAL PAGE IS
OF POOR QUALITY
CHORDREYNOLDSNUMBER, R
C
10 9 8
.0200 -
I I
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7
6
.5
4 X 106
I
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0
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F1 FIXED TRANSITION, (X/C)t =5_._
36)
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[_[_'-1"1"__/_
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132
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NATURALTRANSIT ION
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Figure
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..
Comparisons
characteristics
of
I
632-215 (REF. 37)
I
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1
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.6
.8
1.0
SECTION LIFT COEFFICIENT,Ce
Inboard
wake
flight-measured,
for Bellanca
probe
I
I
1.2
I
station.
wind-tunnel
Skyrocket
II
measured,
and
predicted
air_oil
(riqht
wing).
]
1.4
ORIGINAL PAGE I_
OF POOR QUALITY
L
CHORD REYNOLDSNUMBER, Rc
10 98
7
6
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DRAG
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wake
37.-
probe
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I
1,2
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1.4
, iC
ORI_NAL PAC_E19
OF POOR QUALn'v
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.05
.04
.03
SUPERCRITICAL,
R
INSECTS
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"1
2.2-hrFLIGHT
Vc = 178 knots
.02
SEA LEVEL
SUBCRITICAL INSECTS
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0
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l
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.12
X/C
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INSECTHEIGHT SCALE
0 .02.04.06
(in.)
PREDICTED CRITICAL
EXCRESCENCE HEIGHT
h = 25000ft;
- ..__'-_----_
//
Vc =258 knots_//
h = SEA LEVEL; Vc = 178 knots---"
Figure
38.-
Insect
contamination
pattern
accumulated
I
on
in
Bellanca
Skyrocket
II
NLF
wing,
fl.ight.
,,,
134
®
ORIGINAL pAGE IS
OF pOOR QUALrrV
.7-
(
O INSIDE SLIPSTREAM
[] OUTSIDESLIPSTREAM
SOLID SYMBOLS: FIXEDTRANSITION,
.6-
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ABOVE
SURFACE,
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-
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.9
1.0
BOUNDARY-LAYER-PROFI
LE, u/u e
Fiqure
39.-
Effect
of
profiles,
propeller
slipstream
s/c = 28.7 percent;
on Bellanca
SkMrocket
II boundary-layer
n = 1806 rpm;V
= 178 knots.
4
'li _,
135
®
ORIGINAl,.P_';k_ _
OF POOR Ou_LnY
HOT-FILM SIGNALS
BLADE-[..........
PASSING
]
J
FREQUENCY
i.-
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NLF GLOVE
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TRANSITION
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0 .1 .2 .3 .4 .5 .6 .7
TIME, sec
Fiqure
40.-
In-flight,
slipstream
n
=
150
on
laminar
hot-film
boundary
measured,
layer
time-dependent
(T-34C
airplane).
effeots
R
=
of
1.5
.I
.8
propeller
X
10 6
ft-1;
rpm.
I'
h .
136
®
ORIGI/NAL
PA_'%_ 1_1
OF POOR
QUALITY
× 10-3
_:_: i 4. :_:_::-::::
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:-:.
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:::--:I
RELATIVE
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F...-_
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(..:
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0
40
_
4
8 12 16 20 24 28 32 36 40 44 48 52
DISTANCEALONG SURFACEFROMLEADINGEDGE, in.
(a)
5O
_--f-fi-: --P_--L::_:___::_-L--Z__.::.L-::Z.L:L__-_
King
Cobra
(filled
and
sanded
wing,
1950).
circa
-3
x 10
,I
i
'
i_
RELATIVE
GAUGE 30
READING,
20
in.
,
I
I
I
. UPPER SURFACE :__
i
I
I
',
I
-- MEASURED
----CALCULATEDMEAN
I_
i-
,%
_
_
_
"'
'1
"
1O
]
0
(b)
Figure
41.-
-,
2 4 6 8 10 12 14 16 18 20 22 24 26 28 30 32 34 36 38
DISTANCE ALONG SURFACE FROM LEADINGEDGE, in.
Shyrocket
Indicated
II
surface
(as-pr0duced
waviness
composite
for
Bellanca
wing,
circa
Skyrocket
1970).
II
and
King
Cobra.
137
®
ORI_NAL
PAGE
OF
QUALITY
POOR
IS
Winglet
experimental
transition
I 29 °.
b/
Winglet root, RE) = 51
Experimental
transition
I
I
I.
I
I
I
I
I
66
,|
__
F_
(a) VariEze.
Figure
42, _ Comparison
of experimental
transition
criterion,
data
with
spanwise
contamination
138
®
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................
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................
ORIGINAL
PA_,_ Ig
OF POOR
QUALi_
Winglet
experimental
transition
Winglet tip, R8 = 33
Winglet root, R0 = 36
(b)
Figure
Long-EZ
42. &
..........
Concluded.
139
®