Active Aeroelastic Wing

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Nov 16, 2013 (3 years and 6 months ago)

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Active Aeroelastic Wing


Sarah Chandrasekar

Department of Engineering, Calvin College

Engineering 315
Final Paper

Professor Ribeiro



Abstract
:
Aero elastic Wing will soon take the place
of today's wing control. These changes ought to
generate savings in

weight and increases in
reliability, employing technologies that have been in
development and testing for years. The research and
development required for developing MAVs and
related systems is technically challenging and
requires a number of technologica
l advances that
may benefit a broad range of aerospace
applications. The development of a vehicle with
aero elastic wing could also promote development
of component technologies and help to support an
emerging growth market for micro aerial vehicles.



1.
0 Intr
oduction

The Air Force calls it “back to the future”

taking an
element of the Wright brothers’ original aircraft, a
feature deliberately engineered out of modern planes,
and putting it back, literally. In this case, it means
restoring preproduction w
ings on a Navy F/A
-
18

wings that had been replaced because they twisted in
flight. The

active aeroelastic wing (AAW) aircraft is a
unique joint research effort involving the Navy, the Air
Force Research Lab at Wright
-
Patterson AFB, NASA
-
Dryden, and the Boe
ing Phantom Works. Funding for
the Navy plane comes from AFRL’s Air Vehicles
Directorate and NASA’s Office of Aerospace
Transportation Technology, with Boeing performing
the F/A
-
18 modifications under contract to the Air
Vehicles Directorate.

For his hist
oric first flight on
December 17, 1903, Orville Wright used the
movements of his hips in the airplane’s “saddle”

in
which he lay prone

to twist or warp either the left or
right wingtip. This provided flight control without the
use of ailerons or flaps. Whi
le such aeroelastic warping
is inherent in the wings of modern high
-
speed aircraft,
engineers have done everything possible to

counteract
it, from physically stiffening the wings to
incorporating other control surfaces.

The degree of twisting involved is
actually rather small

less than 4°. AAW technology you can provide a weight
-
competitive wing, reduce drag, improve range, and reduce
fuel consumption, because you have a more
aerodynamically efficient wing with an increased aspect
ratio
1
.


Identify 3 axes
of rotation


Aircraft fly in three dimensions, and they move in
directions other than straight and level.

The axis that
extends lengthwise (nose through tail) is called the
longitudinal axis
, and rotation about this axis is called
roll.
The axis that exte
nds crosswise (wingtip through
wingtip) is called the lateral axis
, and rotation about this
axis is called pitch.
The axis that passes vertically
through the center of gravity (when the aircraft is in
level night) is called the vertical axis
, and rotation
about
this axis is called yaw.



Figure 1

Role of Aircraft wings in different Maneuvers

Basic flight maneuvers include climbs, descents, turns,
and combinations of these.

Generally, the basic flight
maneuvers are started from what is called straight and
l
evel flight.

Straight and level flight (also called
controlled flight
) is a flight condition where the wings
are kept level and the altitude and heading constant.
Power setting is maintained at 55 percent to 75 percent
of available power. If speed is desir
ed a higher setting is
required, however if fuel needs to be saved then a lower
setting is required. Straight and level flight is a series of
slight adjustments or corrections in pitch, yaw, and roll
to keep the wings level and

heading and altitude
constan
t.

Climbs are a combination of power and "up
elevator."

The amount of power used determines
whether the climb is steep or shallow. In order to use all
available power, the climb angle must be as steep as
possible. This is called the best angle of climb, b
ut it is a
short
-
term climb. A sustained climb at this angle can
overheat the engine.

The third basic maneuver is the
turn
.
Turns are gentle, medium, or steep; and they may
be made when climbing, descending, or while not
gaining or losing altitude
. Causing

the airplane to turn
requires smooth coordination of aileron, rudder, and
elevator controls; in other words, pressure on the control
wheel and the rudder pedal should be applied
simultaneously. The moment a wing begins to rise in a
banked turn, it experie
nces more drag because of the
lowered aileron and its higher angle of attack. A
simultaneous application of rudder compensates for this
additional drag by making the airplane also rotate about
its vertical axis.


Figure 2
: Elements of a turn


2.
0

History

When Orville Wright first took to the air on Dec. 17,
1903, he didn't have ailerons or flaps to control his
airplane. Instead, the Wright brothers had chosen to twist
or "warp" the wingtips of their craft in order to control its
rolling or banking motion.
Rather than using one of the
craft's two control sticks to make the wingtips twist, they
had devised a "saddle" in which the pilot lay. Cables
connected the saddle to the tips of both wings. By moving
his hips from side
-
to
-
side, the pilot warped the wingti
ps
either up or down, providing the necessary control for the
Wright Flyer to make turns.

The test aircraft chosen for
the AAW research is a modified F/A
-
18A obtained from
the U.S. Navy in 1999. Begun in 1996, the AAW flight
research program has completed
detailed design and the
wing modifications required for the program have been
completed. The test aircraft has been extensively
instrumented, and reassembly was completed by early
2001. Over the course of the year, the AAW test aircraft
was subjected to ex
tensive structural loads, wing stiffness
and vibration tests, installation of the initial control
software into the aircraft's research flight control
computer, systems checkout and flight simulation activity.
The first parameter identification flights in
the two
-
phase
flight test program are expected to begin in mid
-
2002 and
continue for about six months. These flights will be used
to measure the forces available from each surface to twist
the wing and control the aircraft. That will be followed by
a yearl
ong period of data analysis and control software
redesign to optimize the performance of the flexible wing.
The final phase of flight tests are expected to be flown in
2003, and will evaluate the handling and performance
qualities available from the flexib
le wing concept.
2

3
.0 Model for F14 Jet aircraft

The following diagram displays the top level of the
model of a flight controller for the longitudinal motion
of a Grumman Aerospace F
-
14 aircraft



Figure 3
: Model of F14 Jet aircraft

The model simulates th
e pilot's stick input with a
square wave having a frequency of 0.5 (radians per
second) and an amplitude of ± 1. The system outputs
are the aircraft angle of attack and the G forces
experienced by the pilot. The input and output signals
are visually monito
red by Scope blocks.


Figure

4:

Pilot G
-
Force Scope


The angle
-
of attack range extends from
-
10 to 90 deg,
with tunnel data used where it is available and
estimates where it is not. Linear interpolation of
tabulated data is used to model nonlinearities in

aerodynamic coefficients. Control surface effects are
linear in their respective deflections, and static lateral
-
directional effects are linear in sideslip angle. The
derivatives for both vary with angle of attack to the
limits of data presented in the re
ference and are
constant at higher angle. Angular rate derivatives are
constant throughout the angle of attack range.


Figure

5
: Angle of Attack



The different components of the model of the F14 are
described in the following sections, Controller,
Aircra
ft dynamics model and the Nz pilot calculation






3
.1 Controller


Figure 6
: Controller

Stick Input

The stick input
is the input from the pilot to control the
aircraft. Maneuvers are made possible through the
stick input
.


Figure 7
: Stick Input

Alpha (R
ad)

The Alpha is calculated in radians from the vertical
velocity w (ft/sec) once it is passed through a gain. The
vertical
velocity w (ft/sec) is explained in section
3
.2



Figure 8
: Alpha

q
(rad/sec)
(also called Pitch Rate)


Pitch rate is explained in de
tail in section
3
.2.




Figure 9
: Q (rad/sec)


Output

of the controller


The output of the controller is also called the Elevator
Command (deg) which one of the inputs to the Aircraft
Dynamics Model




Figure 10
: Output of the controller













3
.2
Aircraft Dynamics Model




Figure 11
: Aircraft Dynamics Model


The aircraft is subject to random gusts of wind. The
aircraft responds quite strongly to the gusts of wind,
thus making it difficult for the human pilot to maintain
control. The gusts of winds

can
differentiate

as
Vertical Gust

(ft/sec) and Rotary
Gust (
rad/sec). The
inputs of the Vertical Gust referred to as wGust is
given in Figure
:
12,

and the Rotary gust referred to as
qGust is given in
F
igure:

13
.


Vertical Gust (ft/sec)

Vertical

Gust is t
he component of gust winds that
exhibits vertical motion and is measured in ft/sec.



Figure 12
: Vertical Gust(ft/sec)




Ro
tary Gust (rad/sec)


Rotary Gust is the gust winds that exhibit rotary
motion and is measured in rad/sec




Figure 13
: Rotary Gust

(rad/sec)


Vertical Velocity, w (ft/sec)


The vertical velocity is the final output after the
Vertical Gust is passed through the transfer function

coupled with the pitch rate.




Figure 14
: Vertical Velocity (ft/sec)











Pitch Rate, q (rad/sec)


T
he pitch rate is the
output of the rotary gust passed
through a transfer function, gains

and coupled with the
value of the velocity rate.




Figure 15
: Pitch Rate (rad/sec)


2.3
Nz Pilot Calculation



Figure 16
: Control system for Nz Pilot Calculation


O
utput signal before gain


Figure 17

Pilot g Force (g)
(Output signal after gain)


This is one of the final outputs for the F 14 jet aircraft.
It is explained in detail in section 3.0



Figure 18
: Pilot Force

3.
0

Design requirements For the Pitch
Controlle
r

The next step is to set some design criteria

for the pitch
controller
. We want to design a feedback controller so
that the output has an overshoot of less than 10%, rise
time of less than 2 seconds, settling time of less than 10
seconds, and the steady
-
s
tate error of less than 2%. For
example, if the input is 0.2 rad (11 degre
e
s), then the
pitch angle will not exceed 0.22 rad, reaches 0.2 rad
within 2 seconds, settles 2% of the steady
-
state within
10 seconds, and stays within 0.196 to 0.204 rad at the
ste
ady
-
state.



Overshoot: Less than 10%



Rise time: Less than 2 seconds



Settling time: Less than 10 seconds



Steady
-
state error: Less than 2%

3.1 Transfer function


Figure 19
: Shows longitudinal, lateral and vertical axis

The longitudinal equations of moti
on of this aircraft
are the following
, assuming zero initial conditions.

To
find the transfer function of the above system, we

also
need

to take the Laplace transform
.

W
hen finding a
transfer function, zero initial conditions must be
assumed



t
(
)
0.313


t
(
)
56.7

t
(
)

0.232

t
(
)





t
(
)
0.0139


t
(
)
0.426

t
(
)

0.0203

t
(
)





t
(
)
56.7

t
(
)



After doing some calculations we are able to calculate
the following transfer function


s
(
)

s
(
)
1.151
s
0.1774

s
3
0.739
s
2


0.921
s





These values are taken from the data from one of the
Boeing's commercial aircraft.

3.3

Matlab re
presentation and open
-
loop
response

Now, we are ready to observe the system characteristics
using Matlab. First, let's obtain an open
-
loop system to a
step input and determine which system characteristics
need improvement. Let the input (delta e) be 0.2 ra
d (11
degrees). Create an new
m
-
file

and enter the following
commands.

de=0.2;

num=[1.151 0.1774];

den=[1 0.739 0.921 0];

step

(de*num,den)


Running this m
-
file in the Matlab command window
should give you the plot

on the next page
.


From the plot, we see that the o
pen
-
loop response does
not satisfy the design criteria at all. In fact the open
-
loop
response is unstable.

3.2 PID Controller

In order to correct the problem with the open loop
response we can use a PID controller in a feedback
loop.

Using the following

equation we can determine the
value of the proportional (
K
p
)
, integral(
T
i
)

and
derivative(
T
d
)

K
p
T
i

T
d

K
p
s
2

T
i

T
d

s



After a lot of calculations we get the following
template with the transfer function



s
(
)

s
(
)
1.151
T
d

s
2





1.151
K
p
0.1774
T
d




s


0.1774
K
p


T
i

s
3
0.739
1.151
T
d



s
2


0.921
1.151
K
p


0.1774
T
d




s


0.1774
K
p


T
i



We use the following code in m
file in order to execute
the function in Matlab

de=0.2;

Kp=9;

Td=3
;


Ti=4;

numc=[1.151*Kd 1.151*Kp+0.1774*Kd
0.1774*Kp];

denc=[1 0.739+1.151*Kd
0.921+1.151*Kp+0.1774*Kd 0.1774*Kp];

t=0:0.01:10;

step

(de*numc,denc,t)

When we use only Kp = 9, w
e get the following graph



We see that both the overshoot and the settling time
need improvement. Thus, we use a derivative gain
Td=4. Then we get the following graph




We add the integral in order to get a smoother curve.
Ti=4. Then we get the followi
ng graph



3.

Fundament
al
s of Active

Aeroelastic

Wing
(AAW)

Figure 1 shows the functionality of the aeroelastic wing.


Figure
20
: Difference between conventional aileron and active
aeroelastic wing, while performing Right Roll command




Figure
21
: Str
uctural testing on AAW

The wings from Dryden's F
-
18 #840, formerly used in
the High
-
Alpha Research Vehicle (HARV) program,
have been modified for the AAW flight research
program and installed on the AAW test aircraft. Several
of the existing wing skin pane
ls along the rear section of
the wing just ahead of the trailing
-
edge flaps and
ailerons have been replaced with thinner, more flexible
skin panels and structure, similar to the prototype F
-
18
wings. Original F
-
18 wing panels were light and
flexible. Durin
g early F
-
18 flight tests, however, the
wings were observed to be too flexible at high speeds
for the ailerons to provide the specified roll rates. This
was because the high aerodynamic forces against a
deflected aileron would cause the wing to deflect in
the
opposite direction. In addition, the F/A
-
18's leading
-
edge flap has been divided into separate inboard and
outboard segments, and additional actuators have been
added to operate the outboard leading
-
edge flaps
separately from the inboard leading
-
edge s
urfaces. By
using the outboard leading
-
edge flap and the aileron to
twist the wing, the aerodynamic force on the twisted
wing will provide the roll forces desired. Now, a flexible
wing will have a positive control benefit rather than a
negative one. In add
ition to the wing modifications, a
new research flight control computer has been
developed for the AAW test aircraft, and extensive
research instrumentation, including more than 350 strain
gauges, have been installed on each wing.

References


[1]
ACTIVE AE
ROELASTIC WING: A NEW/OLD
TWIST ON FLIGHT

http://www.aiaa.org/aerospace/Article.cfm?issuetocid
=256&ArchiveIssueID=30

[2]
http://www.nasa.gov/centers/dryden/news/FactSheets
/FS
-
061
-
DFRC.html

[3
]
http://www.engin.umich.edu/group/ctm/examples/pit
ch/digPCSS.htm
l

[4
]
Modeling Aircraft Wing loads from Flight data using
Neural

Networks

www.dfrc.nasa.gov/DTRS/2003/PDF/H
-
2546.pdf

[6]
htt
p://www.simlabs.arc.nasa.gov/vms/controls.html

[7] MATLAB toolbox