1. Development of a Twisting Wing Powered By a Shape
Memory Alloy Actuator
Joshua S. Herrington∗
, Logan H. Hodge∗
, Christian A. Stein∗
, Yogesh Babbar†
,
Robert Saunders†
and Darren J. Hartl ‡
Texas A&M University, College Station, Texas 77843.
James Mabe§
The Boeing Company, Seattle, Washington 98124.
A variably twisting wing has many beneficial qualities that can improve aircraft per-
formance for a variety of flight conditions. Beneficial variable wing twist features include
a means to reduce induced drag in cruise conditions, to increase lift performance, and to
increase roll control. However, the actuation hardware usually required to twist large scale
wings, especially given their structural stiffness, presents a substantial limitation. Through
the use of shape memory alloy (SMA) torque tubes and other associated active spars, such
structural deformations can be enabled. In this work, we consider the development of a
benchtop and wind tunnel testing platform for potentially assessing the response of a vari-
ety of SMA-based wing twisting concepts. Using additive manufacturing (3D printing), we
have designed, built, and tested a small scale prototype. The SMA actuated twisting wing
developed herein consists of a rapid prototype shell that was specifically designed using a
finite element approach to maintain stiffness in bending while reducing torsional rigidity.
Using LabView and a simple PID controller, we have shown that an SMA torque tube
was able to drive and steadily maintain the spanwise linear twist in the wing under both
benchtop and wind tunnel conditions. This prototype will allow future assessment of new
control schemes, new SMA actuator materials, and new structural configurations toward
the development of flight-capable self-twisting wings.
Nomenclature
α Root angle of attack, degrees
θ Wing twist angle with respect to root chord, degrees
CD Coefficient of drag
CL Coefficient of lift
Ktotal Total stiffness constant, in-lbs/degree
I. Introduction
Fixed wing aircraft have a set span-wise twist schedule that is optimized to make the wing as efficient as
possible given the expected mission profile. This usually implies compromises between each of the different
flight maneuvers of ascending, descending, and cruise flight for the aircraft, since a fixed twist schedule
can only be optimized for a single coefficient of lift design point.1
Having a wing that is reconfigurable
in a torsional (twist-wise) manner will allow the wing to achieve the optimized twist distribution for any
coefficient of lift corresponding to a given point in the flight envelope.1
∗Undergraduate Student, Department of Aerospace Engineering, AIAA Student Member
†Graduate Research Assistant, Department of Aerospace Engineering
‡TEES Research Assistant Professor, Department of Aerospace Engineering, AIAA Member
§Associate Tech Fellow, The Boeing Company
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2. Induced drag is a major contributor to an aircraft’s overall drag at subsonic speeds, which leads to a
reduction in fuel efficiency. Induced drag, as its name implies, is an adverse force created by the presence of
lift generated by a body. An elliptical wing can be shown to be the most efficient wing planform (i.e. the
planform with the least induced drag); however, fabrication of such a wing is an expensive endeavor. For
this reason, most airplane wings have a rectangular or tapered planform.2
Generally, these two planform
configurations have a higher induced drag that can be reduced by introducing a span-wise twist. This
variation of local effective angle of attack can be optimized for any given planform to achieve a desired
coefficient of lift while minimizing the induced drag on the wing.1,3
Modern transport-class aircraft wings,
for example, are often designed for optimum performance during cruise, where most of the flight occurs and
most of the fuel is consumed. However, an optimized twist schedule for cruise conditions is generally not
optimized for takeoff and landing. In order to best reduce the induced drag across the entire flight envelope,
a wing that can reshape itself throughout the flight is desired.4
This reshaped wing could also increase lift
performance5
and increase roll control.6
This goal of structural adaptivity through morphing has led Boeing
and others to the development and demonstration of a range of lightweight, novel, and robust actuation
systems that enable controllable shape change of aerostructures.3,7
The active material component implemented in the active wing twist system is a shape memory alloy
(SMA). SMA actuators are compact, robust, and lightweight with a high energy density, and they can
be sized over a wide range of force and displacement applications. These are phase transforming alloys
capable of generating and fully recovering substantial strains (∼ 5%) under large loads (300MPa, 40ksi).8
Common structural implementations of this strain generation/recovery include tube twisting, beam bending,
or tensile contraction. Conventional aircraft reconfiguration methods such as flaps and slats require heavy
and complex actuation systems and their associated structures. Given their very small installation volumes,
SMA components represent an ideal enabling technology for the continuous morphing of aero-surfaces,3,9
including the variable twist of an adaptive wing. One such development was the DARPA “SMART wing”,
which successfully demonstrated continuous wing twist, but did not progress beyond wind tunnel testing at
16% scale.10–14
The use of a highly energy-dense embedded actuator can both enable a new generation of
compact morphing test articles and demonstrate the capability of controlled SMA actuators in this role.
In this work, we consider the design, fabrication, and testing of a prototype twisting wing driven by
an internal SMA torsional actuation component previously developed by the researchers at The Boeing
Company. The wing was thus designed to match the capabilities of the tube (instead of applying an SMA
actuator to a pre-designed wing), thus creating a more system-optimized configuration. The purpose of
this effort was to demonstrate the use of an SMA actuator in a closed-loop control system to command
wing twist and systematically alter the aerodynamic qualities of the wing; using this same system, we will
continue to test other SMA component, sensor, heater, and control software options. The simple prototype
wing consisted of five major structural components: ribs, a passive torque tube, an SMA torque tube, a
mounting structure for use in a wind tunnel, and a 3D printed rapid prototype skin. In addition to these
pieces, seven key electrical components where attached to the prototype: a cartridge heater, a thermocouple
attached to the SMA torque tube, a torque sensor in between the SMA torque tube and the passive torque
tube, two inclinometers placed at 1/3rd
and 2/3rd
of the 36” span, a third inclinometer placed at the free
end (full span), and a fourth inclinometer attached to the end of the SMA tube to measure its rotation.
(a) Covered wing (wind tunnel configuration) (b) Uncovered wing, showing sensor locations (mounted
internally)
Figure 1: Prototype wing model fully configured.
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3. A second configuration that has only been tested in a preliminary manner employed an induction heating
system in lieu of the cartridge heater. The final configuration of the prototype wing is shown in Fig. 1a,
while the location of the inclinometers (tilt sensors) are shown in Fig. 1b.
II. Design
Initial sizing of the prototype wing was based off the Prototype-Technology Evaluation Research Aircraft,
or PTERA, manufactured by Area-I, Inc. The PTERA is a UAV that is a modular test bed for experimentally
demonstrating new aircraft technology at large scales and high speeds.15,16
It represents a future flight test
option for a similarly configured twisting wing section. The prototype had a 36 inch (91.4 cm) span, 12.8
inch (31.5 cm) chord, and incorporated a NACA 0017 airfoil section. A rectangular planform was chosen
because it offered simplicity in design and fabrication and matched the baseline PTERA configuration. It was
determined that a removable clamshell-style lid would be necessary to allow for easy access to the internal
components of the wing during both benchtop and wind tunnel testing.
For the purpose of design, the aerodynamic loading of the wing was estimated using the direct analysis
of a NACA 0017 airfoil in XFLR.17
This 2D analysis was compiled at a Reynold’s number of 4.02x105
, and
Mach number of 0.059. This analysis produced the coefficients of pressure at 199 points along the airfoil chord
that had a local angle of attack ranging from 0 degrees (at the root) to 10 degrees (at the tip). Using this
data, the coefficient of pressure was found for the entire wing by dividing the wing into 100 equally spaced
sections. Each section was assigned an angle of attack corresponding to a linear twist distribution with the
wing root at zero angle of attack and the tip at an angle of attack of 10 degrees. Such a 2-D analysis of this
airfoil yielded an overestimate of the highest aerodynamic loads applied to the wing by ignoring downwash
effects at the wing tip. This conservative estimate ensured that the designed wing structure would withstand
maximum loads.
With these basic parameters, a 3-D solid model was constructed and analyzed using Abaqus FEM soft-
ware18
to determine the minimum thickness of the wing shell. Due to the limited strength of ABS-M3019
(3-D
printing material), it was determined that the thickness of the shell should not be less than 0.1”. FEM anal-
ysis ensured that the shell could twist up to 10 degrees under the maximum load of the SMA tube without
buckling or permanently deforming. Because of the rigid nature of closed section structures, the original
iteration of the wing design required more torque to twist 10 degrees than the SMA tube could provide.
To address this problem, several spanwise slits were introduced to the shell model (See Fig. 1b). These
slits allowed the wing to be more compliant in torsion while remaining stiff in bending. The final design
was 3-D printed using ABS-M30 material and the slots were covered with fiber-reinforced pressure sensitive
polyethylene tape (See Fig. 1a).
The mounting hardware, ribs and most of the internal structure were custom-designed to meet design
requirements and goals. As previously mentioned, the design was based around the known performance of the
SMA tube; the torque sensor and a swivel joint were commercial off-the-shelf items. The ribs were machined
from aluminum and were designed to fit inside the 3-D printed shell with several attachment points to ensure
Figure 2: Exploded view of prototype wing assembly.
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4. a rigid structure was formed. The wind tunnel balance mounting structure was also machined from aluminum
due to weight constraints of the wind tunnel load cell and included an access hole to insert and remove a
cartridge heater, which was nested inside the SMA tube. The passive torque tube and the interface pieces
between the ribs and the torquing apparatus were machined from steel due to their high stress requirements
and the desire for small strains. Figure 2 shows an exploded view of the final assembly. Note the relative
simplicity of this fully functional twisting wing.
III. Results: Benchtop Testing
Prior to the wind tunnel testing, benchtop tests were performed to assess the controlled twist capabilities
of the prototype wing. A simple PID controller was implemented in LabView to control the voltage supplied
to the cartridge heater, which in turn supplied heat to the SMA tube. Four inclinometers, a torque sensor, and
a thermocouple were used to collect twist, SMA tube torque, and SMA tube temperature data, respectively,
throughout the runs. One inclinometer was used to measure the amount of twist achieved by the SMA
tube, and the other three inclinometers measured the amount of twist at 1/3rd
, 2/3rd
and full span.a
The
inclinometer at full span was used as the feedback mechanism in the PID controller. This controller in
LabView allowed for effective control of the angle of attack at the free end, within 0.5 degrees of the user
input, on a bench top test stand.
Preliminary benchtop testing allowed the system to be fine-tuned while simultaneously determining the
limits of wing twist and SMA output torque. Figure 3a shows the frame of reference used throughout this
work; the same reference frame was used in the wind tunnel with the lift and drag acting in the directions
shown. Figure 3b shows the stiffness results of the full wing.
(a) Relationship of lift, drag, α, and θ. (b) Torque versus twist angle, θ.
Figure 3: Wing reference frame and torsional stiffness.
The actual results from benchtop testing supported expected theoretical results arising from FEA. Fig-
ure 3b shows the linear relationship between torque and wing twist angle, as expected in an elastic structure.
From this plot, the stiffness of the entire model was estimated at Ktotal = 2.9 in-lbs/degree (0.33 Nm/degree).
The result of twisting the prototype wing to an angle of attack at the free end (θ) of 10 degrees in
a benchtop scenario is shown in Fig. 4a. The PID controller allowed the user to input the desired angle
of attack of the free end at any given time, potentially allowing the wing to adjust to changing ambient
conditions if necessary. A controlled incremental step of 2 degrees up in free end angle of attack (6 to 10
degrees) is shown in Fig. 4b. During this process, the PID controller was allowed to settle the free end angle
of attack before the next step was commanded.b
aNote that, due to the compliance of the passive torque tube and the use of a circumferential slotted tube connector, the
SMA tube twist and full span twist were not identical.
bNote that PID parameter tuning was not rigorously addressed, and the observed overshoots are considered acceptable for
the tests at hand.
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5. (a) PID-controlled free end twist angle of 10 deg with
additional inclinometers (1/3rd
and 2/3rd
spanwise loca-
tions and SMA tube mounted)
(b) PID-controlled step up of free end twist angle with
additional inclinometers (1/3rd
and 2/3rd
spanwise loca-
tions and SMA tube mounted)
Figure 4: Benchtop test results.
IV. Results: Wind Tunnel Testing
After benchtop testing, the wing was installed in a 3’ X 4’ (0.91 m X 1.22 m) wind tunnel to obtain
experimental measurements of lift and drag changes in response to controlled changes in induced linear twist.
The wing was first successfully tested at 32.8 ft/s (10.0 m/s). Its performance was then assessed at
various angles of attack, α, ranging from -7.2 to +5.0 ± 0.1 degrees at 64.3 ft/s (19.6 m/s) with the addition
of controlled, variable, span-wise linear twist. The angle of attack was varied while the wing tip angle was
held constant at a 4, 6 and 8 degrees positive twist relative to the root chord line, for a total of three sweeps.
Throughout the sweeps, the wing tip angle (θ) was actively controlled using the LabView-implemented
closed-loop PID controller. Sensor input from an inclinometer measured the free end angle of attack and the
root angle of attack was taken from a rotary encoder attached to the wind tunnel mounting structure. The
wing lift and drag versus angle of attack were collected using the wind tunnel load cell. Figure 5 shows data
from the three sweeps.
(a) Wing coefficient of lift vs. angle of attack with
varying wing twist (θ = 4, 6, 8)
(b) Wing coefficient of drag vs. angle of attack with
varying wing twist (θ = 4, 6, 8)
Figure 5: Wind Tunnel Test Results
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6. Preliminary data (toward future developments) was also collected using an alternate heating method
employing the effects of induction. The induction heater system consisted of low-gauge copper wire wrapped
tightly around the exterior of the SMA tube; this coil extended most of the length of the tube. To further
reduce cycle time, the SMA torque tube was cooled by introducing compressed air into its central channel,
which was hollow once the unused cartridge heater was removed. Figure 6a shows the drastic reduction
in cycle time due to the induction heating/compressed air cooling system. The wing was set at zero angle
of attack and the tip twist was set to 6 degrees positive twist from the root chord line using the alternate
heating and cooling systems. Because of the high rate of heat transfer of the induction heating process, no
attempt to control the system was made at this time, and joint Boeing/TAMU efforts are ongoing. This set
of data is included herein primarily to demonstrate the benefit of induction heating with respect to reduced
cycle time, and thus motivated future studies. The solid line in Fig. 6a represents a typical cycle using the
cartridge heater; Fig. 6b shows a corresponding hysteresis loop with points 1 and 2 corresponding to the
maximums of Fig. 6a.
(a) Actuation cycle time implementing cartridge heater
(solid line) and induction heating (dashed line).
(b) Sample hysteresis loop of SMA torque tube. (Car-
tridge heater option; control excursion shown in grey)
Figure 6: Actuation cycle time and hysteresis loop (wind tunnel testing).
V. Discussion
The benchtop test results in Fig. 4 were used to determine the limits of controllability of the wing twist
and other properties of the system, such as cycle time, overshoot, and settling time. This data was important
to identifying the range of controllability and to determining a wind tunnel test matrix. It was found that
the wing could be efficiently twisted up to 10 degrees; beyond that point excessive heat and time were
necessary. The precision of the controllable wing tip twist angle θ was determined at ± 0.5 degrees. The
overshoot and settling time were not major driving factors in the PID controller design, but reasonable values
of approximately 25% and 225 seconds were chosen and deemed acceptable.
During wind tunnel testing, twist capability was not hindered. The wing twist remained controllable up to
8 degrees (the wing was not twisted to 10 degrees in the wind tunnel due only to time constraints). The twist
precision remained at ± 0.5 degrees, and overshoot, settling time, and cycle time were virtually unchanged
from the benchtop testing. Furthermore, the effectiveness of the simple control system was demonstrated in
a relevant environment. As the sweeps were executed and the root angle changed (thus changing the applied
aerodynamic loads), the control system ensured that the tip angle was constantly held at the specified twist
angle relative to the root chord.
Figure 5 shows that wing twist does in fact have an effect on the wing lift and drag. In Figure 5a, it
is seen that as wing twist increases, the CL vs. α curve shifts upward. Additionally, in Figure 5b, it is
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7. seen that CD tends to converge to a single value at lower angles of attack. This effect combined with the
increased lift from wing twist creates a desirable outcome that will be assessed further in ongoing studies.
Finally, the induction heating and compressed air cooling systems show that cycle time can be greatly
reduced, given the proper experimental hardware. However, it was observed in this configuration that
controllability was not as easily attained. This relationship is largely due to the fact that induction heating
(especially on such a small scale) is quite rapid and not as effective at delivering minor heat adjustments.
The traditional cartridge heater, on the other hand, can increment temperature by fractions of a degree.
This engineering tradeoff will need to be studied and better understood before induction heating can be fully
implemented into similar systems.
VI. Conclusion
As stated above, the main goal of this effort was to create a testing platform for demonstrating controllable
wing twist using an SMA torque tube with closed-loop feedback. The data presented in this short paper
indicates that controllable twist of a specially designed wing is possible and even effective. While the current
results may not allow determination of optimum flight configurations (as the focus of this project was not
aerodynamics), they do show that a controllable SMA actuator can alter wing aerodynamic performance on
demand and allow for future studies. The basic design of the system using the cartridge heater proved to
be effective for this type of simple twist actuation. The design also allowed for implementation of a different
heating/cooling system (the induction heater). In the case of the cartridge heater design, the simple PID
control system was highly effective.
Future developments of twisting wing systems will now be based off of this proven system; however, many
improvements are still to be made. For instance, longer wings will require advanced structural and material
design to resist bending while remaining compliant in torsion, and overall strength-to-weight ratio will need
to be increased to enable a flight-capable system.20
To this end, a fully composite-based alternate structure
has been designed20
and fabrication is planned. Further development of a control system for the induction
heating system will be necessary to fully implement the induction heater as a replacement for the slower
cartridge heater.
Going forward, it is expected that this prototype will allow future assessment of new actuation and
morphing control schemes, new SMA actuator materials, and new structural configurations toward the de-
velopment of flight-capable self-twisting wings. It will also enable the experimental assessment of such active
aeroelastic effects as flutter mitigation. Finally, it provides yet another demonstration of the advantages of
SMA torsional components as embedded and compact actuators toward the future development of mission-
adaptive wings across scales and flight platforms.
Acknowledgments
Finite element analysis was performed using Abaqus through a research license granted by Simulia. The
authors would also like to acknowledge engineers from Boeing Research & Technology and Mr. Nicholas Alley
from Area-I for their contributions and feedback regarding this work. Financial support was provided by
Mr. Don Ruhmann of The Boeing Company and Mr. Dale Cope of the Texas A&M Engineering Experiment
Station (TEES).
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