Design and Wind Tunnel Testing of Cal Poly's AMELIA 10 ...
Transcript of Design and Wind Tunnel Testing of Cal Poly's AMELIA 10 ...
27TH
INTERNATIONAL CONGRESS OF THE AERONAUTICAL SCIENCES
DESIGN AND WIND TUNNEL TESTING OF CAL POLY’S AMELIA 10 FOOT SPAN HYBRID WINGBODY
LOW NOISE CESTOL AIRCRAFT
Kristina K. Jameson*, David D. Marshall*, Robert Ehrmann*, Eric Paciano*, Rory Golden*, and Dave Mason**
*California Polytechnic State University, San Luis Obispo, **PatersonLabs
Keywords: Poweredlift, Circulation Control, CFD database, NFAC
Abstract
California Polytechnic Corporation, Georgia Tech Research Institute (GTRI), and DHC Engineering collaborated on a NASA NRA to develop and validate predictive capabilities for the design and performance of Cruise Efficient, Short TakeOff and Landing (CESTOL) subsonic aircraft. In addition, a large scale wind tunnel effort to validate predictive capabilities for aerodynamic performance and noise during takeoff and landing has been undertaken.
The model, Advanced Model for Extreme Lift and Improved Aeroacoustics (AMELIA), was designed as a 100 passenger, N+2 generation, regional, CESTOL airliner with hybrid blended wingbody with circulation control. The model design was focused on fuelsavings and noise goals set out by the NASA N+2 definition. The AMELIA is 1/13 scale with a 10 ft wing span. PatersonLabs was chosen to build AMELIA and The National FullScale Aerodynamic Complex (NFAC) 40 ft by 80 ft wind tunnel was chosen to perform the nine week long large scale wind tunnel test in the summer of 2011.
1 Introduction
With the very recent advent of NASA’s
Environmentally Responsible Aviation Project
(ERA)[1], dedicated to designing aircraft that
will reduce the impact of aviation on the
environment, there is a need for research and
development of methodologies to minimize fuel
burn, emission, and a reduction in community
noise produced by regional airlines. ERA is
specifically concentrating in the areas of
airframe technology, propulsion technology,
and vehicle systems integration all in the time
frame for the aircraft to be at a Technology
Readiness Level (TRL) of 46 by the year of
2020 (deemed N+2). The proceeding project
looking into similar issues was led by NASA’s
Subsonic Fixed Wing Project and focused on
conducting research to improve prediction
methods and technologies that will produce
lower noise, lower emissions, and higher
performing subsonic aircraft for the Next
Generation Air Transportation System.
The work provided in this investigation was
an NRA funded by Subsonic Fixed Wing
Project starting in 2007 with a specific goal of
conducting a large scale wind tunnel test along
with the development of new and improved
predictive codes for the advanced poweredlift
concepts incorporated into the wind tunnel
model in conjunction with the verification of
these codes with the experimental data obtained
during the wind tunnel test. Poweredlift
concepts investigated are Circulation Control
(CC) wing in conjunction with over the wing
mounted engines to entrain the exhaust and
further increase the lift generated by CC
technologies alone.
There are a number of papers in the past
few years presenting computational studies of
CC technologies. Most of them have focus on
2D studies[211] While there are a number of
excellent 2D experimental datasets available for
such CFD validation[1215], the same is not
1
JAMESON, MARSHALL, EHRMANN, PACIANO, GOLDEN, MASON
true for 3D experimental data[16]. This effort
aims to address this short fall by creating a
comprehensive and relevant 3D database for
current and future 3D simulations. Experimental
measurements included in the database will
include forces and moments, surface pressure
distributions, local skin friction, boundary and
shear layer velocity profiles, farfield acoustic
data and noise signatures from turbofan
propulsion simulators. This paper focuses on
designing and developing a model with NASA’s
N+2 goals for less environmental impact as well
as the fabrication of a 10 foot span wind tunnel
model to be used to create the 3D validation
database for numerical simulations. The
resulting design is the Advanced Model for
Extreme Lift and Improved Aeroacoustics
(AMELIA) and is the subject of the rest of this
paper.
2 Advanced Model for Extreme Lift and Improved Aeroacoustics (AMELIA)
2.1 Summary of NASA’s N+2 Design Goals
NASA is committed to identify solutions that
meet improvements for noise, emissions, and
energy usage (fuel burn). They have classified
the N+2 design metrics as a 40% reduction in
fuel consumption, progress towards 42 dB
lower noise levels, a 70% decrease in emissions,
and a 50% reduction in field length performance
over current generation aircrafts. Theoretically
the aircraft should reach a Technology
Readiness Level (TRL) of 46 by the year 2020.
Dave Hall at DHC Engineering did the
conceptual design of four separate
configurations to address the N+2 goals with a
down selection to one favorable configuration.
This yielded the Advanced Model for Extreme
Lift and Improved Aeroacoustics (AMELIA).
2.2 Conceptual Designs Considered
Four CESTOL configurations were
developed for consideration for the large scale
wind tunnel test. The first concept,
Configuration 1, has a more conventional
appearance with a high aspect ratio wing, a tee
tail, and twin engines configured over the wing,
as can be seen in Fig. 1. The trailing edge wing
surface aft of the engine is designed for turning
the flow to create poweredlift.
Fig. 1. Conceputal design of Configuration 1 consisting of
a high aspect ratio wing, a teetail, and twin engines
configured over the wing.
The second concept, Configuration 2, is a
hybrid blendedwingbody. The design utilizes
circulation control at the leading and trailing
edges of the wing. This configuration utilizes
over the wing engine locations for noise
shielding and engine exhaust entrainment
purposes. A structural rudder and Vtail were
also employed in this design. Configuration 2 is
shown in Fig. 2.
Fig. 2. Conceputal design of Configuration 2 consisting of
circulation control wings and over the wing engine
mounting for enhanced powered lift.
The third aircraft configuration was inspired
by the extensive research now underway to
develop a true BlendedWingBody aircraft,
BWB, by the Boeing Company[17]. This
aircraft concept is a significant departure from
the first two aircraft designs. The turbofan
engines are embedded within the very thick
wing root as the wing blends into the fuselage,
as shown in Fig. 3. The exhaust discharges
2
SIGN AND WIND TUNN
OT SPAN HYBRID WIN
zle at the
tent is not
is nozzle
increased
generating
consisting of
termed the
may be
a vertical
aft wings
to improve
rmation at
se vertical
sails than
tr
aft wing
planform
ave a high
ystem is a
e mounted
consisting of
hannel wing
D
F STO
D no
trailing edge of the vehicle. The i
only to produce thrust through t
throughout the flight but to creat
flow circulation around the aircraft
3 Conceputal design of Configuration 3
The final design, seen in Fig. 4, is
. It
Wing with
member joining the fore an
at the outer span points. The intent is
local air flow and mitigate shock f
high subsonic Mach numbers. Th
members are more like wingtip
winglets and act structurally as
forward wing sweeps aft and th
sweeps forward forming a diamon
shape in the top view. Both wings
propulsion
sized geared turbofan engi
Conceputal design of Configuration 4
wing design with
AMELIA
ch design, it w
n 1
N+2 design.
r hand was too
thin an N+2 time
unnel test, being
A utilized a
y with circulation
uration
riate
tigations into this
span model based
st locations.
s project, it was
st suited for the
s a rendering of
ter many design
tion 2
s
Configuration 2
wind tunnel test
difications were
alteration to the
ng system of the
as chosen as the
odynamic forces
s ability to take
Direct mounting
ough the aft end
the flow around
ody mount was
ent location with
EL TESTING OF CAL P
BODY LOW NOISE C
2.3 Configuration 2 Become
After close consideration of e
apparent that Configurati
ntional to be an
Configuration 4 on the oth
advanced to be considered w
frame. A large scale wind
conducted by a competing N
model that was similar to
it was a blended wing bo
Confi
appro
inve
showed that a 10’
would n
our perspective t
consulting all involved in th
onfiguration
was b
5 sho
a
Configur
AMELIA Design Featur
In order to utilize the
geometry in a large scale
design m
needed. The most significan
geometry came in the mount
wind tunnel model. A sting
ideal method to measure ae
for i
intrusively
of the model to the sting th
raised concerns with disturbin
the beaver tail. An under
designed to provide an attach
DEESIGN AND WIND TUNNEL TESTING OF CAL PO OLY’S AMELIA 10 FO OOT SPAN HYBRID WINGBODY LOW NOISE CE ESTOL AIRCRAFT
through a high aspect ratio 2D nozzzle at the
trailing edge of the vehicle. The in ntent is not
only to produce thrust through thhis nozzle
throughout the flight but to create e increased
flow circulation around the aircraft generating
additional lift.
Fig.3 Conceputal design of Configuration 3 consisting of
a blended wing body and embeded engines.
The final design, seen in Fig. 4, is termed the
DiamondWingBody (DWB). It may be
thought of as a JoinedWing with a vertical
structural member joining the fore andd aft wings
at the outer span points. The intent is to improve
local air flow and mitigate shock fo ormation at
high subsonic Mach numbers. The ese vertical
members are more like wingtip sails than
winglets and act structurally as sstruts. The
forward wing sweeps aft and the e aft wing
sweeps forward forming a diamond d planform
shape in the top view. Both wings h have a high
aspect ratio. The chosen propulsion s system is a
mediumsized geared turbofan engin ne mounted
in a channel wing.
Fig. 4. Conceputal design of Configuration 4 consisting of
high aspect ratio diamondwing design with cchannel wing
mounted engines.
2.3 Configuration 2 Becomes s AMELIA
After close consideration of ea ach design, it was
apparent that Configuratio on 1 was too
conventional to be an N+2 design.
Configuration 4 on the othe er hand was too
advanced to be considered wi ithin an N+2 time
frame. A large scale wind ttunnel test, being
conducted by a competing NR RA utilized a test
model that was similar to CConfiguration 3 in
that it was a blended wing bod dy with circulation
control wings[18]. Config guration 2 was
considered to be at the approp priate level for the
N+2 time frame. Further inves stigations into this
design also showed that a 10’ span model based
on Configuration 2 would no ot exceed the load
limits of all our perspective te est locations. After
consulting all involved in thiis project, it was
decided that Configuration 2—the hybrid
blended wing body—was be est suited for the
AMELIA test. Figure 5 show ws a rendering of
Configuration 2 in flight af fter many design
iterations.
Fig. 5. Rendered image of Configura ation 2 the Hybrid
Blended Wing Body in flight.
2.4 AMELIA Design Feature es
In order to utilize the Configuration 2
geometry in a large scale wind tunnel test
setting, many design mo odifications were
needed. The most significant t alteration to the
geometry came in the mountiing system of the
wind tunnel model. A sting w was chosen as the
ideal method to measure aerrodynamic forces
and moments, mainly for it ts ability to take
measurements nonintrusively. . Direct mounting
of the model to the sting thr rough the aft end
raised concerns with disturbingg the flow around
the beaver tail. An underb body mount was
designed to provide an attachm ment location with
3
t
vertically
scussed in
e inverted
e negative
eded to be
s to extend
three view
the blade
oved and
mpennage
use
e majority
main focus
ustic and
strakes,
e previous
order to
d testing.
el via off
h
tilizes an
e see Ref.
radius flap
cost and
º, 60º and
to be
anical flap
e 90º flap
e to issues
he flaps of
modified
GOLDEN, MASO
minimal flow disturbance. The mou
with a clamshell blade that extend
from the sting tip. For purposes d
subsequent sections the model will
for a portion of the test, as a result t
angle of attack limit of the sting n
ded. The blade mount also serv
shows a
drawing of the model mounted to
attachment with empennage re
The tail
view bec
t
testing, due to the fact that the
of the testing is on aeroac
aerodynamic measurements. Th
in t
tured in
supplement subsequent research a
These surfaces attach to the mo
wi
The selected configuration
[plea
with a dual
2
order to minimize
complexity dual radius flaps of 0º, 3
propose
manufactured, as opposed to a mec
where deflection can be varied. T
deflection was later changed to 80º d
of the flap.
design were als
for each wing, in
flow disturbance
p surface.
the 80° flap,
showing the inter
attachment and tail
del.
low pressure air
pipe supplies the
ower the
units, while the
e low pressure air
ontrol slots at the
ed
ng, however two
in order to study
n entrainment of
trol flow. Height
ted using faired
rent heights [see
l engine height
act as pressure
pressure air is fed
n
experiment it’s
el be highly
re the maximum
instrumentation
t wing and was
ry CFD res
atic of the model
cement of the
ACIANO,
be a single continuous flap
order to reduce the amount o
from discontinuities of the fl
, wit
Cut away of the AMELIA
blad
empennage on the manufacturing m
supplies the necessary high an
to the model. The larger blue
ir required to
Propulsion Simulator (TPS)
pipe provides t
necessary for the circulation
piping systems will be discus
onfiguratio
will be mounted over the w
mounting heights will be teste
the effects of engine height
exhaust in the circulation co
adjustments will be compl
two diff
for the optim
These pylons als
vessels within which the high
A Instrumentati
As a CFD validation
imperative that the mo
instrumented in order to capt
amount of data. Almost al
placement occurs on the le
chosen based on prelimin
span sche
illustrating the relative pl
JAMESON, MARSHALL, EHRMANN, PACIANO, GOLDEN, MASON
minimal flow disturbance. The moun nt is faired
with a clamshell blade that extends s vertically
from the sting tip. For purposes di iscussed in
subsequent sections the model will b be inverted
for a portion of the test, as a result th he negative
angle of attack limit of the sting ne eeded to be
extended. The blade mount also serve es to extend
this negative limit. Figure 6 shows a three view
drawing of the model mounted to the blade
attachment with empennage rem moved and
relevant dimensions shown. The tail e empennage
is not shown in the threeview beca ause it will
not be attached to the model during th he majority
of the testing, due to the fact that the main focus
of the testing is on aeroaco oustic and
aerodynamic measurements. The e strakes,
structural rudder and Vtail seen in th he previous
figure were manufactured in order to
supplement subsequent research an nd testing.
These surfaces attach to the mod del via off
blocks.
Fig. 6 Threeview drawing of AMELIA wit th stingblade
attachment with the tail surfaces removed.
The selected configuration uutilizes an
optimized supercritical airfoil [pleas se see Ref.
1920 for further details] with a dual radius flap
at the trailing edge[Please see Ref. 21 1 for further
details]. In order to minimize cost and
complexity dual radius flaps of 0º, 300º, 60º and
90º deflections were proposed d to be
manufactured, as opposed to a mech hanical flap
where deflection can be varied. Th he 90º flap
deflection was later changed to 80º du ue to issues
with the manufacturing of the flap. T The flaps of
the Configuration 2 design were alsoo modified
to be a single continuous flap for each wing, in
order to reduce the amount of f flow disturbance
from discontinuities of the flaap surface. A cut
away view of the model, withh the 80° flap, is
shown in Fig. 7.
Fig 7. Cut away of the AMELIA showing the internal
structure along with the stingblade e attachment and tail
empennage on the manufacturing mo odel.
The piping seen in Fig. 7 exiting the blade
supplies the necessary high andd low pressure air
to the model. The larger blue pipe supplies the
high pressure air required to p power the Turbine
Propulsion Simulator (TPS) units, while the
smaller amber pipe provides th he low pressure air
necessary for the circulation c control slots at the
leading and trailing edges. TThe details of the
piping systems will be discusssed in subsequent
sections.
As shown in Configuration n 2 the TPS units
will be mounted over the wi ing, however two
mounting heights will be tested d in order to study
the effects of engine height oon entrainment of
exhaust in the circulation con ntrol flow. Height
adjustments will be comple eted using faired
structural pylons of two diffe erent heights [see
Ref. 2223 for the optima al engine height
location]. These pylons also o act as pressure
vessels within which the high pressure air is fed
to the TPS units.
2.5 AMELIA Instrumentatio on
As a CFD validation experiment it’s
imperative that the moddel be highly
instrumented in order to captu ure the maximum
amount of data. Almost alll instrumentation
placement occurs on the lef ft wing and was
chosen based on prelimina ary CFD results.
Figure 8 is a halfspan schem matic of the model
illustrating the relative plaacement of the
4
DESIGN AND WIND TUNNEL TESTING OF CAL POLY’S AMELIA 10 FOOT SPAN HYBRID WINGBODY LOW NOISE CESTOL AIRCRAFT
pressure ports and unsteady pressure
transducers. The model is instrumented with
280 static pressure ports in five chordwise
groups and one spanwise group (highlighted in
red in Fig. 8). Five static pressure ports are
located on the right wing in order to verify
symmetry in the pressure distribution.
Eight unsteady pressure Kulites exist on
the model (shown as blue circles in Fig. 8). As
aeroacoustics is one of the main focuses of the
large scale test, five Kulites are located in line
with the engine assembly. The remaining three
Kulites are located at the estimated passengers
head level, on the wing blend to gain a better
understanding of the implications of over the
wing mounted engines and passenger comfort
level. Six stationary microphones will be placed
underneath model approximately 5 ft above the
tunnel floor (in order to be sufficiently out of
the boundary layer) to measure farfield noise
generated by the poweredlift design. A 70
element stationary microphone array will also
be placed directly underneath the wing below
one of the propulsion simulators to determine
the noise signature associated with TPS unit.
Fig. 8. Halfspan schematic of AMELIA illustrating the
relative location of the pressure ports and unsteady
pressure transducers on the left half of the model.
In order to monitor the health of the TPS
units, each unit is instrumented with seven
rakes, each consisting of 4 total pressure probes
and a thermocouple. The design and amount of
these rakes were based on discussions with
engineers from industry who have previous
experience with TPS units.
Each of the circulation control plenums is
instrumented with 3 total pressure probes (the
plenums can be seen in Fig. 7), while a cross
correlation rake measures boundary layer
interactions at the plenum exit or slot. Further
discussion of the design and theory behind the
cross correlation rake can be found in Section 5.
The majority of the right wing is free of
instrumentation in order for the local shear
stress measurements to be obtained. In order to
take high fidelity measurements, the surface
must be of number 2 finish (mirror like) or
better. For this reason the model will be painted
with a black Imron coating.[24] Further
discussion of the requirements of the technique
will be discussed in Section 4.
2.6 Internal Design of AMELIA
The model is supplied using two separate air
systems for the propulsion simulators and the
blown slots. Figure 9 shows a complete cut
away schematic of the model showing the
internal piping for both the high and low
pressure systems.
Fig. 9. A complete cutaway schematic of AMELIA
showing the 8 in flow thru balance and the internal piping
for both the high and low pressure systems.
The high pressure air system (600 psi
maximum determined by the limits of the flow
thru balance) is supplied thru the NASA
provided sting into the fabricated strut. The air
passes up the strut and through the 8in flow thru
balance. On the front of the balance is a flow
control system that controls the air flow to the
left and right hand TPS units. The airflow is
controlled using conical plugs that can be
remotely controlled while the tunnel is running.
The conical plugs are driven using MMP 24vdc
gear motors, and use linear potentiometers for
position feedback. The plugs can be positioned
5
. The TPS
steel pipes
at feed the
lete piping
the high
sting
air system
roximately
air to the
ading and
trol wing.
lled using
pots for
m can
trol room.
cts on the
d by a pipe
e pipe
for the low
for system
ment.
GOLDEN, MASO
100% mass flo
stainless
that attach to wing mounted pylons t
e 10 shows the com
and mass flow control plenum fo
pressure air system along with the
A schematic of the high pressur
(ap
supply th
at the l
trialing edges for the circulation co
The low pressure system is contr
, with rotary
feedback. The flow to each plen
controlled individually from the co
adverse eff
balance the low pressure plenum is f
that has bellows at each end of t
pressure system along with the bellows
blade attac
est
y
e Aerodynamic
0 ft win
e week long large
e NFAC offered
rge wind tunnels
with the most
model could be
llows for cleaner
ed aerodynamic
l could supply the
ow rate necessary
nd the turbofan
enough such that
he CCW wings
or of the tunnel
r
was acoustically
ic and acoustic
e performed
AC’s cost and
time frame and
cale schematic of
g in the 40 ft by
th the associated
mounted on the
ind tunnel model
nnel.
AMELIA in the
ounted and inverted
with the far
nary acoustic array.
ACIANO,
Large Scale Wind Tunnel
Wind Tunnel Test Facili
Sca
Complex (NFAC) 40 ft by
was chosen to perform the ni
scale wind tunnel test. T
several benefits over other l
,
significant being: the 10 foo
measurements of the produ
forces and moments, the tunn
high pressure air at the mass f
to operate the CCW slots
simulators, the tunnel is large
nwash created by
would not impinge on the fl
thus creating cleaner f
measurements, the tunnel
treated such that aerodyna
measurements could
e N
schedule fit within Cal Poly’
2 shows a
the model inverted on the sti
80 ft NFAC wind tunnel w
acoustic measurement device
r relative to the
placed in the center of the t
shows the relative scale of
12. A schematic of AMELIA
on the sting in the NFAC alon
es and the stati
JAMESON, MARSHALL, EHRMANN, PACIANO, GOLDEN, MASON
to provide from 0100% mass flow w. The TPS
units are supplied through stainless steel pipes
that attach to wing mounted pylons th hat feed the
TPS units. Figure 10 shows the comp plete piping
and mass flow control plenum forr the high
pressure air system along with the stingblade
attachment.
Fig. 10. A schematic of the high pressuree air system
internal to AMELIA.
The low pressure system (app proximately
100 psi) will be used to supply the e air to the
plenums that feed the slots at the leeading and
trialing edges for the circulation con ntrol wing.
The low pressure system is contro olled using
24vdc gear motors, with rotary pots for
feedback. The flow to each plenu um can be
controlled individually from the con ntrol room.
To isolate and minimize adverse effeects on the
balance the low pressure plenum is fe ed by a pipe
that has bellows at each end of th he pipe, as
shown in Fig. 11.
Fig. 11. An internal schematic of the pipingg for the low
pressure system along with the bellows for system
isolation downstream of the stingblade attachhment.
3 Large Scale Wind Tunnel T Test NFAC
3.1 Wind Tunnel Test Facilit ty
The National FullScal le Aerodynamic
Complex (NFAC) 40 ft by 8 80 ft wind tunnel
was chosen to perform the nin ne week long large
scale wind tunnel test. Th he NFAC offered
several benefits over other la arge wind tunnels
across the United States, with the most
significant being: the 10 foott model could be
mounted on a sting which aallows for cleaner
measurements of the producced aerodynamic
forces and moments, the tunne el could supply the
high pressure air at the mass fl low rate necessary
to operate the CCW slots aand the turbofan
simulators, the tunnel is large enough such that
the downwash created by tthe CCW wings
would not impinge on the flooor of the tunnel
thus creating cleaner fa arfield acoustic
measurements, the tunnel was acoustically
treated such that aerodynam mic and acoustic
measurements could bbe performed
simultaneously, and the NFFAC’s cost and
schedule fit within Cal Poly’ss time frame and
budget. Figure 12 shows a s scale schematic of
the model inverted on the stin ng in the 40 ft by
80 ft NFAC wind tunnel wi ith the associated
acoustic measurement devices s mounted on the
tunnel floor relative to the w wind tunnel model
placed in the center of the tu unnel. Figure 13
shows the relative scale of AMELIA in the
NFAC.
Fig. 12. A schematic of AMELIA m mounted and inverted
on the sting in the NFAC along g with the farfield
stationary microphones and the statioonary acoustic array.
6
DESIGN AND WIND TUNNEL TESTING OF CAL POLY’S AMELIA 10 FOOT SPAN HYBRID WINGBODY LOW NOISE CESTOL AIRCRAFT
Fig. 13. Isometric view of AMELIA mounted on the sting
in the NFAC.
3.2 AMELIA Test Matrix
The proposed text matrix for the model includes
calibrations while mounted on the sting balance
of both model and acoustic instrumentation,
static tests of all blown features on the model, a
Reynolds number sweep, a dynamic pressure
sweep, a turbofan simulator sweep, and a CCW
sweep.
After the completion of the preliminary
sweeps, eight to ten critical tests points will be
identified from the experimental test data
obtained. During the critical tests points all
experimental instrumentation will be utilized.
Acoustically, the farfield measurements along
with a traversing 30 degree sideline
measurements will be made. Also the 70
element stationary array placed under one of the
turbofan simulators will be utilized to
characterize the noise signature beneath the
wing. The aerodynamic forces, moments, static
pressure measurements, and unsteady pressure
measurements will simultaneously be
conducted. Then the model will be inverted and
the same critical test points will be repeated.
Inverting the model offers several advantages in
obtaining experimental measurements. The oil
interferometry requires less lighting in the large
wind tunnel and high power lights can be shown
directly on the reflective surface. The 70
element stationary array can be utilized to
identify any hot spots created by the TPS units.
The cross correlation rake will also be utilized
while the model is inverted.
Once the critical test points have been
completed, the model will be returned to right
side up and alpha (up to +20 to 5 degrees) and
beta (+20 to 20 degrees) sweeps will be
conducted at three different tunnels speeds. The
majority of the tests will be conducted at lower
tunnel speeds in order to correctly match the
necessary design thrust coefficients (limited by
the TPS thrust output). Only farfield acoustic
measurements, aerodynamic forces and
moments, and static pressure measurements will
be conducted for this portion of the test
allowing a significant number of test points to
be investigated to create the database for current
and future CFD validation efforts.
4 Oil Interferometry
The FringeImage Skin Friction (FISF)
technique, also known as oil interferometry, was
chosen for the large scale wind tunnel test to
measure local skin friction because both
magnitude and direction can be determined
from a single image. The FISF technique has the
advantage of maturity and reliability which
becomes significant due to the difficulties of
obtaining measurements in the NFAC due to its
sheer size and the amount of time between
tunnel shut down and the point where
photographs of the model can be obtained.
4.1 FISF Technique
The FISF technique was developed by Monson
et al.[25] The theory behind the technique is that
a single relationship can relate the thickness of
the oil drop at a single location to the skin
friction magnitude and direction. The oil
thickness is measured via photographic
interferometry. Data reduction is completed
with CXWIN4GG, a PC application developed
by Zilliac.[26]
The FISF technique has a few key steps in its
process to obtain the crucial photographs
necessary to determine surface skin friction. The
process is as follows: A drop of silicone oil of
known viscosity is placed on the model surface.
Once the oil is applied to the surface, the air
flow begins, causing the oil to spread and thin.
The air continues to flow for a given time,
continually thinning the oil. When sufficient
time has elapsed (220 minutes, depending on
oil viscosity), the air flow is turned off. A quasi
monochromatic light source is then indirectly
applied to the surface by use of a large diffuse
7
JAMESON, MARSHALL, EHRMANN, PACIANO, GOLDEN, MASON
reflector. Light is reflected from the surface of
the model and oil. Two specific light rays are
reflected and separated by the thickness of the
oil, shown in Fig. 14. Once the oil has thinned,
the oil height linearly varies where constructive
and destructive interference occurs, causing
light and dark fringes on the oil surface. Skin
friction is proportional to the spacing distance,
Δs, on the fringes which is directly related to the
thickness of the oil. The relationship for skin
friction is as follows:
C f = w
q
τ
∞ =
2 o o n s
q t
µ
λ
Δ
∞ ( )cos rθ
(1)
where Cf is the skin friction coefficient, τw is the
wall shear stress, q∞ is the freestream dynamic
pressure, no is the oil index of refraction, µo is
the oil viscosity, λ is the wavelength of the light
source, t is the duration of time the oil flow was
exposed to air flow, and θr is the light refraction
angle through the airoil interface. Eq. 1 holds
for zero pressure gradient and shear stress
gradients. Further details on the oil flow
technique and the theory behind it is covered by
Naughton and Sheplak.[27]
Fig. 14. A schematic of the basic FISF setup highlighting
the oil flow and fringe pattern on a droplet of oil[27].
4.2 Application of FISF to AMELIA
In order to successfully apply the FISF
technique, the fringes on a model need to be
clearly visible. Fringe visibility is based upon
the surface finish of the model. An ideal surface
is extremely smooth with consistent and durable
optical properties. Based on a study by
Zilliac[28] the best fringes appeared on high
flint content SF11 glass manufactured by Schott
Glass of Germany, which is an impractical
material for a wind tunnel model. A practical
surface finish for a model would be mirror like,
which can be achieved by nickel plating the
model surface. Acceptable substitutes have been
made utilizing polished stainless steel or black
Mylar sheets applied to the model surface.
Black Mylar sheets offer the most cost effective
solution for oil flow testing. However, at higher
speeds and long run times Mylar would begin to
peal along the edges, ruining any data
downstream. The continual Mylar reapplication
to the model would prove time consuming and
impractical due the model height in the NFAC
for the AMELIA test. Mylar is also difficult to
apply to a 3D surface, usually resulting in small
wrinkles in the Mylar distorting the skin friction
measurements. Polished stainless steel and
nickel plating have the durability that Mylar
lacks. Another alternative to nickel plated
aluminum is aluminum painted with black
Imron. Difficulties arose in the capability of
plating the fuselage of AMELA due to its sheer
size. For these reasons, the entire model will be
painted with Imron, also allowing future
experiments with AMELIA to investigate other
regions besides the wing.
In order to properly view the fringes, a
monochromatic light source must be reflected
off a diffuse reflector. In large wind tunnel
applications, the tunnel walls have been used as
the reflector.[29] Unfortunately, the NFAC
tunnels walls are composed of a matte metal
mesh covering a deep, perforated acoustic liner,
rendering the walls unable to sufficiently light
the tunnel. The next option is to build a reflector
which encompasses the portion of interest on
the model. A small hole would be cut in the
reflector allowing a camera to capture the fringe
spacings; a setup such as this is shown in Fig. 4.
AMELIA has a highly curved blended wing
which causes additional difficulty in the image
processing. Due to the models highly polished
and curved surfaces, the camera will see
reflections from a large area of the wind tunnel.
Therefore, if the wind tunnel is being used as
the diffuse reflector, large areas of the tunnel
need to be white. Since this would be costly in
the NFAC, it is necessary to use a curved
diffuse reflector. This will ensure the model is
uniformly lit allowing for accurate fringe
8
DESIGN AND WIND TUNNEL TESTING OF CAL POLY’S AMELIA 10 FOOT SPAN HYBRID WINGBODY LOW NOISE CESTOL AIRCRAFT
spacing identification and does so with fewer
lights. The type of diffuse reflector used on the
wing blend is also shown in Fig. 15.
The angle at which the light enters the
camera can greatly affect the skin friction
measurement, especially at large angles such as
leading edges or the blended wing portion of the
model. Zilliacs CXWIN4GG software utilizes
single camera photogrammetry to determine the
angle of the light reflecting off the oil on the
surface of the wind tunnel model. This is made
possible by using fiducial marks over the model
surface. The camera captures an image
encompassing the entire wing with several
fringes over the wing surface. Within that image
are multiple fiducial marks with known
locations on the model coordinate system.
Zilliacs software completes the photogrammetry
by matching the known fiducial marks locations
(both a pixel xy coordinate system as well as
the model coordinate) to a given set of model
points. This allows the software to calculate the
light incident angle at any visible point on the
model.
Due of the size of the NFAC, a special
procedure has been devised to ensure accurate
fringe production. Normally the tunnel
transients are short, resulting in little error from
the startup and shutdowns. However, the NFAC
requires a minimum of 5 minutes to startup and
shutdown which can introduce unacceptable
error into the skin friction measurements if the
incorrect viscosity of oil is chosen for the test.
In order to ensure recording accurate fringes, the
model will be at a high angle of attack during
the tunnel startup allowing separated flow (low
to no shear) over the wing. Once the tunnel
freestream has been reached, the model will be
positioned to the angle of attack of interest. At
this point, slot blowing and the turbine
simulators will be started as well. The model
will remain at this test condition for a minimum
of 20 minutes. Once the oil is sufficiently
spread, the slot blowing, turbine simulator, and
tunnel freestream will be turned off, while the
model is once again pitched upward to cause
separation over the wing allowing the fringes to
be unaffected by the shutdown procedure. Once
the flow has stopped, the diffuse reflector and
camera will be brought into the tunnel. The
model will be inverted for the skin friction
measurements, allowing for optical access to the
suction side of the airfoil without having to be
physically placed above the model. A diffuse
reflector will be held up to the model, lighting
it, while a second person will capture the fringe
spacings in two images. This process is time
consuming, but worthwhile to ensure quality
data for CFD validations to be made against.
For this reason, only eight to ten key test
conditions will be investigated with oil
interferometry for CFD validation. Please refer
to Ref. 30 for further investigation of the FISF
technique for the preparation of the AMELIA
test.
Fig. 15. The FISF solution used for the wing blend. The
dark spot is due to no light being reflected by the camera
lens.
5 Boundary and Shear Layer Velocity Profiles with a Micro Flow Measurement Device
5.1 Application of FISF to AMELIA
In 2001, NASA Glenn developed a
thermocouple boundary layer rake which has
the capability to measure 0.0025 inches from
the surface, four times closer than any state of
the art measurement before.[31] This was
achieved with the device shown in Fig. 16. This
device is constructed with a base made out of
aluminum, a constant thickness strut which
protrudes from the base, and the necessary
electrical wires corresponding to the number of
thermocouples present for a given design. The
strut is made of quartz, chosen for its insulative
9
. From the
re several
functions
en height
ual to the
e no strut
horseshoe
d at the
in NASA
ke.
undary layer
ctions and b)
a platinum
ture above
p, the air
difference
ownstream
n velocity
rake setup
xactly the
rve can be
calibration
ds number
Gustave
le rake’s
CCR) was
calibration
anticipated
conductive
er of the
mperature
tantaneous
ding edge
oltage ta
ons in the
the local
the loop
th velocity
correlation
e between
GOLDEN, MASO
and low thermal expansion propertie
re
pairs of thermocouples. The devic
based on the theory that for a gi
above the base, the velocity is e
velocity of the flow as if there we
present. It should be noted that no
found to be sh
intersection of the strut and base
r
Photographs of a thermocouple b
rake, a) shows the base and electrical conn
At the center of the strut, there is
loop which is heated to a temper
ambient. As air flows over the lo
downstream is heated. The voltage
between the upstream and
lated to kno
values obtained from a total pressur
next to the thermocouple rake at
same axial position. A calibration c
developed from this data. This
procedure is needed for every Reyno
Upon further discussions wit
Fralick, one of the thermocou
designers, a Cross Correlation Rake
suggested for use since an in situ
process would prove difficult in the
The CCR contains a
platinum loop located at the cen
substrate. This loop is heated to a t
above the ambient conditions. In
voltage differences on the loop’s le
and trailing edge are measured by
along the loop. The voltage fluctuat
loop are caused by fluctuations i
resistance. The resistance relates t
temperature, which relates directly w
Utilizing a cross
the ti
upstream and
n be calculated.
he upstream and
p is known, the
MELIA
both leading and
to vastly increase
effort will be to
ve the flap. The
shown in Fig.
his region is of
he jet’s boundary
er above the jet.
validation of the
ill be achieved
er detail
the CCR].
ay be tested to
ss on AMELIA.
at Cal Poly, with
bstrate at NASA
, then a trailing
ss Correlation Rake
the airfoil section is
ounting apparatus is
AMELIA test is
strate, platinum
voltage taps, as
ig. 18
IA flap and was
n
arge as possible
ACIANO,
voltage fluctuations on th
downstream voltage taps c
Since the distance between
downstream edges of the lo
flow velocity can be calculate
CCR to
The AMELIA model has
trailing edge blowing in order
the lift. Part of the validatio
acquire velocity profiles ab
CCR will be implemented as
chieve this.
particular interest because of
layer as well as the shear la
The design, construction, and
bsonic speeds
for furt
design and construction o
addition, transonic speeds
demonstrate the CCR readin
constructed
the platinum applied to the s
Glenn. The CCR will first b
plate for validation purpose
A rendering of the Cr
applied to the AMELIA flap wher
shown as in light blue and the
The CCR designed for th
composed of a quartz su
heating loop, and platinum
photograph in
was sized based on the AME
sized to 1.1 x 0.55 inches
thick. It needed to be as
JAMESON, MARSHALL, EHRMANN, PACIANO, GOLDEN, MASON
and low thermal expansion properties s. From the
base to the top of the strut, there aare several
pairs of thermocouples. The device e functions
based on the theory that for a givven height
above the base, the velocity is eq qual to the
velocity of the flow as if there werre no strut
present. It should be noted that no horseshoe
vortices were found to be she ed at the
intersection of the strut and base in NASA
Glenn’s testing of the thermocouple raake.[32]
Fig. 16. Photographs of a thermocouple bo oundary layer
rake, a) shows the base and electrical conne ections and b)
detail of platinum and gold films.[32]
At the center of the strut, there is a platinum
loop which is heated to a tempera ature above
ambient. As air flows over the loo op, the air
downstream is heated. The voltage difference
between the upstream and ddownstream
thermocouples is then related to know wn velocity
values obtained from a total pressure e rake setup
next to the thermocouple rake at eexactly the
same axial position. A calibration cu urve can be
developed from this data. This calibration
procedure is needed for every Reynollds number
intended for testing.
Upon further discussions with h Gustave
Fralick, one of the thermocoup ple rake’s
designers, a Cross Correlation Rake ((CCR) was
suggested for use since an in situ calibration
process would prove difficult in the anticipated
application. The CCR contains a conductive
platinum loop located at the centter of the
substrate. This loop is heated to a teemperature
above the ambient conditions. Ins stantaneous
voltage differences on the loop’s leaading edge
and trailing edge are measured by v voltage taps
along the loop. The voltage fluctuatiions in the
loop are caused by fluctuations in n the local
resistance. The resistance relates to o the loop
temperature, which relates directly wiith velocity
at that location.[33] Utilizing a cross correlation
script from MATLAB[34], the tim me between
voltage fluctuations on the e upstream and
downstream voltage taps ca an be calculated.
Since the distance between tthe upstream and
downstream edges of the loo op is known, the
flow velocity can be calculated d.
5.2 Application of CCR to A AMELIA
The AMELIA model has both leading and
trailing edge blowing in order to vastly increase
the lift. Part of the validation n effort will be to
acquire velocity profiles abo ove the flap. The
CCR will be implemented as shown in Fig. 17
in order to achieve this. TThis region is of
particular interest because of t the jet’s boundary
layer as well as the shear lay yer above the jet.
The design, construction, and validation of the
CCR at subsonic speeds wwill be achieved
[please see Ref. 35 for furthher detail on the
design and construction of f the CCR]. In
addition, transonic speeds mmay be tested to
demonstrate the CCR readine ess on AMELIA.
The substrate was constructed at Cal Poly, with
the platinum applied to the su ubstrate at NASA
Glenn. The CCR will first be e applied to a flat
plate for validation purposes s, then a trailing
edge of an airfoil.
Fig. 17. A rendering of the Cro oss Correlation Rake
applied to the AMELIA flap where e the airfoil section is
shown as in light blue and the mmounting apparatus is
shown in lime green.
The CCR designed for the e AMELIA test is
composed of a quartz subbstrate, platinum
heating loop, and platinum voltage taps, as
shown in the photograph in F Fig. 18. The CCR
was sized based on the AMELLIA flap and was
sized to 1.1 x 0.55 inches aand 0.043 inches
thick. It needed to be as llarge as possible
10
SIGN AND WIND TUNN
OT SPAN HYBRID WIN
or trailing
inch. The
ight above
his height
The total
ith the top
lower half
lides from
substrate.
thickness
leted CCR
h only the
ires for the
Rake of the
Correlation
l model
gn g
n of the
completed
ete
te
scale wind
the NFAC
ese results
D
F STO
without being too near to the slot
edge. The resulting length was 1.0
height was based upon the desired h
the flap to measure velocity profiles.
termined to be 0.12 inches
height of the rake is 0.040 inches,
half above the flap and the remainin
below the flap. Quartz microscope
Ted Pella, Inc. were utilized as th
y sold at
Figure 19 shows a com
in the flat plate mounting device wi
exposed surface and the associated
Photograph of the Cross Correlatio
Cros
researchers have
a design for a large scale wind tunn
compliance of NASAs N+2 des
fabricati
and at the time this paper was
nearly 90% comp
a projected delivery d
Finally, large
ned for
T
and acoustic
and
acknowledge the
ent awar
with technical
Clif Horne, who
ork. The authors
ch Center’s Fluid
, Gregory Zilliac
ta
IA. The authors
NASA Glenn,
or his time and
A's
)
erence, Atlanta GA,
edaisch J. Evaluation
ediction of
ntrol and
ciences Meeting and
m A. Computational
foil using FLUENT.
ontrol Technology
. Jones, Vol. 214 of
eronautics, chap. 21,
ics and Astronautics,
, Xiao X
ng in
ls.
y,
214 of Progress in
chap. 19, America
stronautics, Inc., pp.
R. J., Ahuja, K. K.,
evaluation of steady
lation control airfoil.
ontrol Technology
. Jones, Vol. 214 of
eronautics, chap. 22,
ics and Astronautics,
EL TESTING OF CAL P
BODY LOW NOISE C
lift
current
The authors would like to
NASA Research Announce
#NNL07AA55C,
monitors Craig Hange and
provided funding for this
wish to thank the Ames Resea
Mechanics Lab, in particula
and David Driver for the assi
the FISF technique to AME
would also like to than
especially Gustave Fralick
Overview of NA
viation (ER
Fundamental Aeronautics Con
, Nagib H and K
46th AIAA Aerospace
, AIAA, Reno NV, AIA
McGowan G and Gopalarathn
study of circulation control ai
ions of Circulation
edited by R. D. Joslin and G.
Progress in Astronautics and
American Institute of Aeronau
McGowan G, Gopalarathnam
odel
control airfo
Circulation Control Technolo
Joslin and G. S. Jones, Vol.
Astronautics and Aeronautics,
Institute of Aeronautics and
Liu, Y., Sankar, L. N., Englar,
and Gaeta, R. J. Computationa
and pulsed set effects on a circ
tion
edited by R. D. Joslin and G.
Progress in Astronautics and
American Institute of Aeronau
DEESIGN AND WIND TUNNEL TESTING OF CAL PO OLY’S AMELIA 10 FO OOT SPAN HYBRID WINGBODY LOW NOISE CE ESTOL AIRCRAFT
without being too near to the slot or trailing
edge. The resulting length was 1.00 0 inch. The
height was based upon the desired he eight above
the flap to measure velocity profiles. TThis height
was determined to be 0.12 inches. . The total
height of the rake is 0.040 inches, w with the top
half above the flap and the remaining g lower half
below the flap. Quartz microscope s slides from
Ted Pella, Inc. were utilized as the e substrate.
Microscope slides were only sold at aa thickness
of 1.0 mm. Figure 19 shows a comp pleted CCR
in the flat plate mounting device witth only the
exposed surface and the associated w wires for the
rake.
Fig. 18. Photograph of the Cross Correlation n Rake of the
first batch of rakes delivered to Cal Poly.
Fig. 19. Photograph of the completed Crosss Correlation
Rake mounted in the flat plate apparatus.
6 Conclusions
Cal Poly and GTRI researchers have completed
a design for a large scale wind tunne el model in
compliance of NASAs N+2 i goals. design
PatersonLabs has started fabricatioon of the
model and at the time this paper was completed
AMELIA was nearly 90% compl lete. The
model has a projected delivery da ate of late
September of 2010. Finally, large scale wind
tunnel test activities are planned for the NFAC
starting in the summer of 2011. Th hese results
will provide poweredlift and acoustic
validation data for current and future 3D
modeling efforts.
6 Acknowledgements
The authors would like to acknowledge the
NASA Research Announcem ment award under
Contract i#NNL07AA55C, with technical
monitors Craig Hange and Clif Horne, who
provided funding for this wwork. The authors
wish to thank the Ames Researrch Center’s Fluid
Mechanics Lab, in particular r, Gregory Zilliac
and David Driver for the assis stance in applying
the FISF technique to AMEL LIA. The authors
would also like to thank k NASA Glenn,
especially Gustave Fralick ffor his time and
assistance.
References
[1]Collier F. Overview of NAS SA's environmentally
responsible aviation (ERA A) project. 2009
Fundamental Aeronautics Conf ference, Atlanta GA,
oral presentation, 2009.
[2]Madugundi D, Nagib H and Ki iedaisch J. Evaluation
of turbulence models through prrediction of separated
flows with and without flow coontrol and circulation
effects. 46th AIAA Aerospace S Sciences Meeting and Exhibit, AIAA, Reno NV, AIAA A20080567, 2008.
[3]McGowan G and Gopalarathna am A. Computational
study of circulation control air rfoil using FLUENT.
Applications of Circulation CControl Technology, edited by R. D. Joslin and G. SS. Jones, Vol. 214 of
Progress in Astronautics and A Aeronautics, chap. 21,
American Institute of Aeronaut tics and Astronautics,
Inc., pp. 539554, 2006.
[4]McGowan G, Gopalarathnam A A, Xiao X, and Hassan
H A. Role of turbulence modeli ing in flow prediction
of circulation control airfoi ils. Applications of Circulation Control Technolog gy, edited by R. D.
Joslin and G. S. Jones, Vol. 214 of Progress in
Astronautics and Aeronautics, chap. 19, American
Institute of Aeronautics and A Astronautics, Inc., pp.
499510, 2006.
[5]Liu, Y., Sankar, L. N., Englar, R. J., Ahuja, K. K.,
and Gaeta, R. J. Computational l evaluation of steady
and pulsed set effects on a circu ulation control airfoil.
Applications of Circulation CControl Technology, edited by R. D. Joslin and G. SS. Jones, Vol. 214 of
Progress in Astronautics and A Aeronautics, chap. 22,
American Institute of Aeronaut tics and Astronautics,
Inc., pp. 557577, 2006.
11
JAMESON, MARSHALL, EHRMANN, PACIANO, GOLDEN, MASON
[6]Chang III P A, Slomski J, Marino,T, Ebert M P, and
Abramson J. Full Reynoldsstress modeling of
circulation control airfoil. Applications of Circulation Control Technology, edited by R. D.
Joslin and G. S. Jones, Vol. 214 of Progress in
Astronautics and Aeronautics, chap. 17, American
Institute of Aeronautics and Astronautics, Inc., pp.
445466, 2006.
[7]Paterson E G and Baker W J. RANS and detached
eddy simulation of the NCCR airfoil. Applications of Circulation Control Technology, edited by R. D.
Joslin and G. S. Jones, Vol. 214 of Progress in
Astronautics and Aeronautics, chap. 16, American
Institute of Aeronautics and Astronautics, Inc., pp.
421444, 2006.
[8]Baker W J and Paterson E G. Simulation of steady
circulation control for the general aviation circulation
control (GACC) Wing. Applications of Circulation Control Technology, edited by R. D. Joslin and G. S.
Jones, Vol. 214 of Progress in Astronautics and
Aeronautics, chap. 20, American Institute of
Aeronautics and Astronautics, Inc., pp. 513537,
2006.
[9]Sahu J. Timeaccurate simulations of synthetic jet
based flow control for a spinning projectile.
Applications of Circulation Control Technology, edited by R. D. Joslin and G. S. Jones, Vol. 214 of
Progress in Astronautics and Aeronautics, chap. 23,
American Institute of Aeronautics and Astronautics,
Inc., pp. 579596, 2006.
[10] Zha G C and Paxton C D. Novel flow control method
for airfoil performance enhancement using coflow
jet. Applications of Circulation Control Technology, edited by R. D. Joslin and G. S. Jones, Vol. 214 of
Progress in Astronautics and Aeronautics, chap. 10,
American Institute of Aeronautics and Astronautics,
Inc., pp. 293314, 2006.
[11] McGowanG, Rumsey C. L, Swanson R C, and
Hassan H A. A threedimensional computational
study of a circulation control wing. 3rd AIAA Flow Control Conference, AIAA, San Francisco CA,
AIAA20063677, 2006.
[12] Owen F K. Measurement and analysis of circulation
control airfoils. Applications of Circulation Control Technology, edited by R. D. Joslin and G. S. Jones,
Vol. 214 of Progress in Astronautics and
Aeronautics, chap. 4, American Institute of
Aeronautics and Astronautics, Inc., pp. 105112,
2006.
[13] Cerchie D, Halfon E, Hammerich A, Han G, Taubert
L, Trouve L, Varghese P, and Wygnanshi I. Some
circulation and separation control experiments.
Applications of Circulation Control Technology, edited by R. D. Joslin and G. S. Jones, Vol. 214 of
Progress in Astronautics and Aeronautics, chap. 5,
American Institute of Aeronautics and Astronautics,
Inc., pp. 113166, 2006.
[14] MunroS E, Ahuja K K, and Englar R J. Noise
reduction through circulation control. Applications of
Circulation Control Technology, edited by R. D.
Joslin and G. S. Jones, Vol. 214 of Progress in
Astronautics and Aeronautics, chap. 6, American
Institute of Aeronautics and Astronautics, Inc., pp.
167190, 2006.
[15] Jones G S.Pneumatic Flap Performance for a two
dimensional circulation control airfoil. Applications of Circulation Control Technology, edited by R. D.
Joslin and G. S. Jones, Vol. 214 of Progress in
Astronautics and Aeronautics, chap. 7, American
Institute of Aeronautics and Astronautics, Inc., pp.
191244, 2006.
[16] Englar R J. Experimental Development and
evaluation of pneumatic poweredlift superSTOL
aircraft. Applications of Circulation Control Technology, edited by R. D. Joslin and G. S. Jones,
Vol. 214 of Progress in Astronautics and
Aeronautics, chap. 7, American Institute of
Aeronautics and Astronautics, Inc., pp. 191244,
2006.
[17] Warwick G. Boeing works with airlines on
commercial blended wing body freighter. Flight
International, 2007.
[18] Collier F, Zavala E, and Huff D. Fundamental
aeronautics program, subsonic fixed wing reference
guide. NASA.
[19] Lane K A and Marshall D D. A surface
parameterization method for airfoil optimization and
high lift 2D geometries utilizing the CST
methodology. 47th AIAA Aerospace Sciences Meeting and Exhibit, AIAA, Orlando FL, AIAA
20091461, 2009.
[20] Lane K A and Marshall D D.Inverse airfoil design
utilizing CST parameterization. 48th AIAA Aerospace Sciences Meeting and Exhibit, AIAA, Orlando FL,
AIAA20101228, 2010.
[21] Golden R and Marshall D D. Design and
performance of circulation control flap systems,"
48th AIAA Aerospace Sciences Meeting and Exhibit, AIAA, Orlando FL, AIAA 20101053, 2010.
[22] EnglarR J, Gaeta R J, Lee W J and Leone V.
Development of pneumatic overthewing powered
lift technology; part I: aerodynamic propulsive. 27th AIAA Applied Aerodynamics Conference, AIAA, San
Antonio TX, AIAA20093942, 2009.
[23] Gaeta R J, Englar R J and Avera M. Development of
pneumatic overthewing powered lift technology art
II: aeroacoustics. 27th AIAA Applied Aerodynamics Conference, AIAA, San Antonio TX, AIAA2009
3941, 2009.
[24] MacphersonL. The wet look that lasts: Imronthe
choice of truck manufacturers and commercial
refinishers for 32 years. Painting and Coating Industry, 2005.
[25] Monson D J, Mateer G G and Menter F R. Boundary
layer transition and global skin friction
measurements. SAE 932550, Society of Automotive
Engineers, Costa Mesa CA, 1993.
12
DESIGN AND WIND TUNNEL TESTING OF CAL POLY’S AMELIA 10 FOOT SPAN HYBRID WINGBODY LOW NOISE CESTOL AIRCRAFT
[26] Zilliac G G. The fringeimaging skin friction
technique PC application user's manual. NASA Technical Memorandum 1999208794, NASA Ames
Research Center, Moffett Field CA, 1999.
[27] Naughton J W, and Sheplak M. Modern
developments in shearstress measurement. Progress in Aerospace Sciences, Vol. 38, No. 1, pp. 515570,
2002.
[28] ZilliacG G. Further Developments of the fringe
imaging skin friction technique. NASA Technical Memorandum 110425, NASA Ames Research
Center, Moffett Field CA, 1996.
[29] Driver D M, and Drake A. Skinfriction
measurements using oilfilm interferometry in
NASA's 11foot transonic wind tunnel. AIAA Journal, Vol. 46, No. 10, pp. 24012407, 2008.
[30] Ehrmann R, Paciano E, and Jameson K K.
Application of the FISF technique to a blended, 2
foot wing section. 28th AIAA Applied Aerodynamics Conference, Chicago Il, AIAA20104830, 2010.
[31] Hwang D, Fralick G C, Martin L C, Wrbanek J D,
and Blaha C A. An innovative flow measuring
devicethermocouple boundary layer rake.
NASA/TM—2001211161, 2001.
[32] Kendall J M.Boundary layer receptivity to weak
freestream turbulence. NASA/CP—2007214667,
March 2007.
[33] Fralick G,Wrbanek J D, Hwang D, and Turso J.
Sensors for using times of flight to measure flow
velocities. NASA Technical Briefs, LEW179441,
2006.
[34] TheMathworks Inc., Natick, MA, MATLAB, 7th
ed., 2008.
[35] Ehrmann R, and Jameson K K. Design and
fabrication of a micro flow measurement device. 40th Fluid Dynamics Conference and Exhibit, Chicago Il,
AIAA20104622, 2010.
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