AIAA 2001-4005
Stability and Control Properties
of an Aeroelastic Fixed Wing
Micro Aerial Vehicle
Martin R. Waszak and Luther N. Jenkins
NASA Langley Research Center
Hampton, VA 23681-2199
Dr. Peter Ifju
University of Florida
Gainesville, Florida
32611-6250
AIAA Atmospheric Flight Mechanics Conference
6-9 August 2001
Montreal, Canada
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STABILITY AND CONTROL PROPERTIES OF AN AEROELASTIC FIXED WING
MICRO AERIAL VEHICLE
Martin R. Waszak* and Luther N. Jenkins_
NASA Langley Research Center, Hampton, Virginia
Dr. Peter Ifju _
University of Florida, Gainesville, Florida
Micro aerial vehicles have been the subject of considerable interest and
development over the last several years. The majority of current vehicle
concepts rely on rigid fixed wings or rotors. An alternate design based on an
aeroelastic membrane wing concept has also been developed that has exhibited
desired characteristics in flight test demonstrations and competition. This paper
presents results from a wind tunnel investigation that sought to quantify
stability and control properties for a family of vehicles using the aeroelastic
design. The results indicate that the membrane wing does exhibit potential
benefits that could be exploited to enhance the design of future flight vehicles.
Introduction
Micro aerial vehicles, or "MAVs", are designated by
the Defense Advanced Research Projects Agency
(DARPA) as a class of aircraft with a maximum
dimension of 6 inches and are capable of operating at
speeds of 25 mph or less. m Developments in
miniaturized digital electronics, communications, and
computer technologies and strong support by DARPA
have moved the prospect of very small autonomous
flight vehicles from the realm of science fiction to
science fact. The goal is for these vehicles to provide
inexpensive and expendable platforms for surveillance
and data collection in situations where larger vehicles
are not practical. For example, they can be used for
battlefield surveillance or mapping the extent of
chemical/radiation spills or viral outbreaks. Equally
useful civil applications include use in search and
rescue operations, traffic/news coverage, and crop or
wildlife monitoring. Many potential uses would require
cooperative and collaborative control capabilities so
that large numbers of MAVs could be used to cover a
large operational area. In these types of applications
Senior Research Engineer, Dynamics and Control
Branch. Senior Member AIAA.
_ Research Engineer, Flow Physics and Control Branch.
$ Associate Professor, Department of Aerospace
Engineering, Mechanics, and Engineering Science.
Copyright © 2001 by tile American Institute of
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the copyright owner.
Figure 1 - photograph of Univ. of Florida MAV.
MAVs could be coordinated from a central base
station or used in collaborative swarms to collect and
transmit data.
The research and development required for
developing MAVs and related systems is quite
challenging and requires a number of technical advances
that may benefit a broad range of aerospace
applications. The development of a vehicle could also
foster development of component technologies and
may help to support an emerging growth market for
micro aerial vehicles.
An aeroelastic fixed wing micro aerial vehicle
concept has been developed by a team at the University
of Florida with a goal to design a vehicle that could
win the ISSMO (International Society of Structural
and Multidisciplinary Optimization) Micro Aerial
Vehicle Competition; a goal that was accomplished
[2 3]
each of the last three years. '
The vehicle exploits an innovative aeroelastic wing
with the ability to adapt to the atmospheric
disturbances and provide smoother flight thus
1
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providing a better surveillance platform and making the
vehicle easier to fly. This is accomplished via the
passive mechanism of adaptive washout. This
technique has been adapted from sailing vessels in
which adaptive washout is produced through twist of
the sail. This greatly extends the wind range of the sail
and produces more constant thrust (lift), even in gusty
wind conditions. Adaptive washout is produced in the
MAV through extension of the membrane and twisting
of the structure in response to changes in speed and
vehicle attitude causing changes in angle of attack
along the span. The effect is to reduce the response of
the vehicle to disturbances.
The benefits of the flexible membrane wing appear
substantial but have not yet been studied in detail. In
addition, the nature of mirco aerial vehicles in general
and the flexible wing concept in particular make
analysis and design of the vehicle quite challenging.
Despite this fact, the vehicle provides an excellent
basis upon which to develop and apply ongoing
research in dynamics and control, aeroservoelasticity,
multi-functional structures, mircoelectronics, measure-
ment and actuation systems, and many others.
NASA is collaborating with the University of
Florida to develop an understanding of the underlying
physical phenomena associated with the vehicle
concept with a goal of enhancing the vehicle design
and developing a capability for investigating
autonomous and collaborative control technologies.
A wind tunnel test was performed to provide data
with which to investigate the benefits of the aeroelastic
wing concept and to support related research. This
paper presents some of the key results of the wind
tunnel test and analysis of these results in the context
of stability and control. The data described here will
also be the basis for a dynamic simulation model
currently under development.
Vehicle Description
The University of Florida MAV (UFMAV)
incorporates a high mounted wing and low mounted
cruciform tail attached to a tapered fuselage with
rectangular cross section (see figure 1). The fuselage is
a truss-like design constructed of a graphite/epoxy
material covered with a thin transparent monofihn
membrane. A more detailed description of the vehicle
and its construction can be found in reference 3.
Table 1 summarizes the pertinent geometric and mass
properties of the vehicle.
A unique aspect of the vehicle is its flexible
membrane wing. The cambered wing structure is
constructed of unidirectional carbon fiber prepreg
laminate forming a leading edge spar and chordwise
ribs or battens. A membrane material is bonded to the
Table 1 - UFMAV geometric and mass properties.
Empty Weight
Wing Area
0.12 lbs
2
19.8 in
Span 6 in
Mean Chord 3.3 in
Moments of Inertia:
Ixx
Iyy
Izz
Ixz
2
0.086 lb in
2
0.23 lb in
2
0.21 lb in
2
0.037 lb in
spar and batten. Several membrane materials with
varying stiffness properties have been investigated
and three of these will be described in this paper: a 4
mil thick flexible latex membrane, an inextensible
monofilm membrane (the same material used in the
fuselage construction), and a stiff graphite sheet.
The maximum dimension (including length and
wing span) of the vehicle is six inches. The wing area
is approximately 19.8 square inches. The root chord is
4.25 inches and the mean chord is 3.3 inches. The
wing camber of the unloaded wing is approximately
6.5 percent of the root chord with the maximum
camber occurring at approximately 30 percent chord
and is uniform across the span. The wing is mounted
at an incidence of approximately nine degrees where the
wing incidence is defined as the angle between the root
chord line and the longitudinal axis of the fuselage.
Control is accomplished using two independently
controlled elevons that are actuated using small rotary
servos. A small gas engine normally provides
propulsion, but an electric motor was substituted
during the wind tunnel test to better control propeller
speed and reduce operational complexity. The propeller
was three inches in diameter with a pitch of 1.25.
Wind Tunnel Test
The wind tunnel test was conducted in the Basic
Aerodynamics Research Tunnel (BART) at NASA
Langley Research Center. [4'51The purpose of the test
was to collect a variety of data to aid in the study of
the dynamics and control properties of the UFMAV
concept. The data consist of aerodynamic force and
moment data measured with an external 6-component
strain gauge balance, static wing deformation data from
a projection moir6 interferometry (PMI) system, [61
dynamic wing deformation data from a high speed
videogrammetry system, [71 and digital video of flow
visualization using smoke and helium bubbles.
Figure 2 depicts the UFMAV mounted in the BART.
These data were collected for a rigid wing and three
different batten/membrane arrangements over a range of
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Figure2- UFMAVmountedinBART.
(a) (b)
T
(c)
Figure 3 - batten arrangement for UFMAV wings:
(a) one-batten, (b) two-batten, (c) six-batten.
operating conditions determined by dynamic pressure,
power setting, vehicle attitude, and control surface
deflection. The different batten arrangements are
depicted in figure 3. More flexibility and larger
membrane stretch characterize the one-batten design.
The two-batten design is, by comparison, stiffer and
exhibits less membrane stretch under aerodynamic
load. Both wings were tested using a 4 rail latex
membrane. The six-batten wing was covered with an
inextensible monofilm membrane that further
increased the stiffness of the wing and exhibited less
membrane deformation and vibration. The rigid wing
was constructed of a two-batten frame covered with a
graphite sheet.
Three dynamic pressure values were considered,
1.0 psf, 1.6 psf, and 2.0 psf, and represent a range of
speeds over which the actual vehicle operates. Two
power settings were considered: power off (characterized
by pinning the propeller) and the power associated with
longitudinal trim (i.e. thrust offsets axial aerodynamic
force, lift offsets weight, and zero pitching moment).
Vehicle attitude is represented by fuselage incidence
angle (i.e., angle of attack) and sideslip angle. The
angle of attack was varied between -5 and +42.5
degrees in 2.5 degree increments (though PMI data
were only collected up to 30 degrees). The sideslip
angle was varied between-5 and +5 degrees in 2.5
degree increments with the vehicle angle of attack fixed
at zero degrees. Control inputs are characterized by
symmetric and antisymmetric elevon deflections.
Symmetric inputs were varied between -25 and +25
degrees in 5 degree increments. Antisymmetric inputs
were varied between-20 and 20 degrees in 5 degree
increments.
The static aerodynamic data were collected using a 6-
component strain gauge balance and resolved into lift,
drag, side force, pitching moment, rolling moment,
and yawing moment. Examples of the force and
moment data collected during the test are presented in
the Appendix (Figures A1 - A11).
Structural deformation data were obtained using two
techniques: projection moir6 interferometry (PMI) [6]
and high-speed videogrammetry. E71PMI was used to
collect mean static deformation over a large fraction of
the wing surface and the variance of the motions about
the mean shape. Videogrammetry was used to collect
dynamic deformation of selected points on the surface
of the wing but only at a few selected operating
conditions. The wing deformation data can be used to
determine the manner and degree to which the wing
(spar, battens, and membrane) deforms under
aerodynamic loading.
Flow visualization was collected using digital video.
Two methods were used: smoke flow and helium
bubbles. The flow visualization data provide insight
into the nndedying flow phenomena and can be
correlated with the aerodynamic and structural data.
Analysis
Additional wind tunnel data is needed before a
complete quasi-static aerodynamic database will be
available. The current data are sufficient to identify
various aspects of the UFMAV's aerodynamic
behavior. Some analysis results based on the available
data are described below. The preliminary results
characterize various aspects of aerodynamic
performance, stability and control, and static aeroelastic
behavior.
Aerodynamic Performance
Aerodynamic performance characterized by L/D is
summarized in figure 4. These results represent L/D of
the UFMAV with the propeller restrained from rotation
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Rigid (Graphite) + Streamlined Fuselage
3.04"0I-- fI ,=.._1_11llltll __ .,,"• r_. • 6-Batten (Monofilm)-X"0". " ,"2-Batten1-Batten (Latex)(Latex) 3.04.0 + Baseline Fuselage
L/D 2.0 _l_.mL._ IJD 2.0
1.0 1.0
0.0 • . i . i . i . i . i
-10 0 10 20 30 40 50 0.0 -10 0 10 20 30 40 50
Angle of Attack (deg)
Figure 4 - L/D for various wing configurations,
(q = 1.6 psf, prop pinned).
(i.e., pinned) for several wings with varying levels of
stiffness. The maximum L/D of approximately 3.0 is
relatively independent of wing configuration.
However, maximum L/D occurs at incidences at
approximately 7.5 degrees for the rigid wing and
roughly 10 degrees for the other wing configurations.
It is interesting to note that a comparable rigid fixed
wing micro aerial vehicle, Aerovironmenrs Black
Widow, has twice the maximum L/D of the vehicle
described herein. [8]
An alternate fuselage configuration was assessed to
determine the effect of streamlining on vehicle L/D.
Figure 5 depicts the variation in L/D over a range of
angles of attack for the baseline fuselage and a
streamlined fuselage. Both fuselages had the same wing
installed at the same incidence angle. The streamlined
configuration had a 20 percent greater maximum L/D
than the baseline configuration. There is a clear benefit
to streamlining the fuselage that can be exploited in
future designs.
Figure 6 depicts the lift curves for the various wing
configurations. For small angles of attack all the
wings demonstrate similar lift characteristics with the
stiffer wings having slightly higher lift coefficient.
However, it is clear that the membrane wings stall at
much higher angles of attack than the rigid wing. In
fact, the most flexible wing configuration has double
the stall angle of the rigid wing configuration (35
degrees and 15 degrees, respectively, for the prop
pinned case). This could be a key factor in enhancing
the range of operation and agility of micro aerial
vehicles.
While these results are similar to the results for
other low aspect ratio, low Reynolds number wings
presented in reference 9 there are important differences.
At low angles of attack the aeroelastic wings behave
like rigid wings with similar aspect ratio. The lift
curve slope for the UFMAV is approximately 2.9 with
the prop pinned. The lift curve slopes of similar rigid
wings studied in Reference 9 at comparable Reynolds
Angle of Attack (deg)
Figure 5 - L/D versus angle of attack for streamlined
and baseline fuselage configurations,
(q = 1.6 psf, prop pinned, 6-batten wing).
Rigid (Graphite)
2.0 • •13. • 6-Batten (Monofilm)
[" -. -_.-. 2-Batten (Latex) X X_
[ --X-.I-Batten(Latex) x,;(X' i.
1.5 /" _ W1.11,0. ,
C L 1.0
0.50
0.0
-10 0 10 20 30 40 50
Angle of Attack (deg)
(a) q=l.6 psf, prop pinned
Rigid (Graphite)
3.0
I . .171.. 6-Batten (Monofilm)
• .._). • • 2-Batten (Latex) E _ O"_41
2.2 --_t ill: _.
C L 1.5
0.75
0.0 _, i . i . i . i . i . i
-10 0 10 20 30 40 50
Angle of Attack (deg)
(b) q=l.6 psf, trim power
Figure 6 - lift coefficient vs. angle of attack for
configurations with varying wing stiffness.
number and aspect ratio (Re=70,000, AR=2) are
approximately 2.9 as well. However, these wings
have stall angles between 12 and 15 degrees. The stall
angles of the aeroelastic wings are between 30 and 45
degrees (i.e., stall angle of the vehicle plus the wing
incidence angle) and are similar to that of much lower
aspect ratio rigid wings (AR=0.5 to 1.0). However,
the very low aspect wings exhibit lower lift curve
slopes of 1.3 to 1.7. The aeroelastic wings appear to
exhibit the stall behavior similar to rigid aspect ratio
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