1. Solid State Aircraft
Phase I Project
NIAC Fellows Conference
October 23-24, 2002
NIAC Headquarters
Atlanta, Georgia
Anthony Colozza
Northland Scientific / Ohio Aerospace Institute
Cleveland, Ohio
2. Solid State Aircraft
Team Members
• Mr. Anthony Colozza (PI): NSI/OAI
• Mr. Phillip Jenkins: OAI
• Dr. Mohsen Shahinpoor University of
New Mexico
• Mr. Teryn Dalbello: University of
Toledo/ICOMP
• Mr. Curtis Smith: OAI
3. Solid State Aircraft Artist Concept Drawing
The aircraft concept is to
integrate three unique types of
materials (thin film solar arrays, thin film lithium
batteries and an ionic polymer metal composite) to produce an
aircraft that has no moving parts, can fly at high altitudes, is easily
deployable and has applications on Earth, Venus and Mars
4. Aircraft Operation
•The aircraft operates by collecting and
Thin Film Array converting sun light to electricity
through a thin film photovoltaic array.
•This electricity is then stored in a
battery.
•At specified intervals the energy is
discharge to the anode and cathode
grids to set up an electric field about
the IPMC (synthetic muscles )material
•This electric field causes the IPMC to
Thin Film move thereby causing a flapping
Battery
motion of the wing.
Cathode •This flapping motion produces lift and
Grid thrust for the aircraft.
•The electric field generated by the
IPMC grids is controllable, therefore the
Material
Anode Grid shape and motion of the wing is
controllable on each flap.
5. Aircraft Construction & Control
•The unique structure combines airfoil, propulsion, energy production
and storage and control.
•To control the motion of the wing a control grid will be used. This
grid will enable various voltages to be sent to different sections of the
wing, thereby causing varying degrees of motion along the wing
surface. The amount of control on the wing will depend on the fineness
of this control grid. A central processor will be used to control the
potential of each of the sections.
PV Array
•This control enables the wing to flap, Battery
Anode
provide differential lift (which is used IPMC
for steering), and alter the camber Cathode
of the wing to maximize lift
under a given operational condition. Each Grid Location is Individually Controllable
6. Thin Film
Photovoltaic
Array
Light Weight: Active
material is on the order
of 1 to 2 microns thick
Highly Flexible: Ideal for the
flexing and motion of a flapping wing
Substrate: Can be made of most materials,
presently the best candidate is Kapton (or other
polymers). Potentially the Battery or IPMC can be
utilized as the substrate
Specific Power: 1 kW/kg near term, 2 kW/kg projected
8. Thin Film Battery/Capacitor
Characteristics
•Rechargeable, Lightweight and Flexible
•Configurable in any series / parallel combination
•Rapid charging / discharging capability
•Can be charged / discharged 1000s of times with
little loss in capacity
–Enables long duration flight times
ITNES sample battery
• Long shelf life with little self discharge
– Ideal for stowage during interplanetary transit
• Operate over a wide temperature range
– Enables the batteries to operate under various environmental
conditions
• The batteries have the capability to provide high pulse currents
– Ideal for short duration power loading such as flapping the wings
9. Battery Construction & Operation
Types of Lithium ion thin-film batteries differ in the cathode
material they use.
–Ex. Magnesium Oxides, Cobalt Oxides, Yttrium Oxide
The battery is produced by depositing (through sputtering or
evaporation techniques) the various material layers that make
up the components (cathode, electrolyte, anode and current
collector) onto a substrate.
Oak Ridge Battery Design Cross Section (15 micron thick)
10. Ionic Polymer-Metal Composite (IPMC)
• This is the core material of the
aircraft. It provides the propulsion
and control for the vehicle.
• The IPMC material has the unique
capability to deform when an
electric field is present across it.
The amount and force of the
QuickTime™ and a deformation is directly related to
decompressor
are needed to see this picture. the strength of the electric field.
• The deformation is not permanent
and returns to its original shape
once the electric field is
eliminated.
• The material can be manufactured
in any size and initial or base
shape.
11. IPMC Material
Constructed of an Ion Exchange
Membrane that is surface coated with
a conductive medium such as
Platinum
Placement of the electrodes can be used to
tailor the bending of the material to any Metal Electrodes
shape
C o n ta c t E le c tro d e
+
P IE M
- P t P a r tic le s
The material will P t E le c tro d e
O FF
bend toward the anode P IE M
side of the electrodes
+
P IE M
-
ON
12. IPMC Motion
Under an electric field the ion exchange membrane
Enables the migration of ions which allows water molecules and
Hydrated cations to migrate toward the negative pole.
This internal movement of water molecules is responsible for creating
Internal strains within the material which enable it to move
- +
+
+
+ For the IPMC material to
+
operate it must be sufficiently
+ +
+
Hydrated
+
Leakage and operation in dry
+
+
+
environments may require
sealing or redesign of the
+
material for efficient long term
use
+ +
side chain + water fixed anion mobile cation
hydrated cation-Na(H2O)4
13. IPMC Material Characteristics
Young' s Modulu s, E Up to 2 GPa
Shear Modulu s, G Up to 1 GPa
Poisson's ratio, ν Typ ical: 0.3-0.4
Power density (W/mass) Up to 100 J/kg
Max force density (Cantilever Mod e) Up to 40 Kgf/Kg
Max displacement/strain Up to 4% linear strain
Bandwidth (speed) Up to 1 kHz in cantil ever vib ratory mode
for actuations
Up to 1 MHz for sensing
Resolution (force and di splacement Displacement accuracy down to 1 micron
control ) Force resolut ion down to 1 mg
Efficiency (electrom echanical) Up to 6 % (frequency depend ent) for
actuation
Up to 90% for sensing
Density Down to 1.8 g/cm 3
14. Solid State Aircraft Applications
There is sufficient solar intensity for this aircraft to operate
on Earth, Venus or Mars
Because of its projected relatively small mass and flexibility, the aircraft
is ideal for planetary exploration. These characteristics allow the aircraft
to be easily stowed and launched at a minimal cost. Potentially, a fleet of
these aircraft could be deployed within a planet’s atmosphere and used
for comprehensive scientific data gathering, as an quickly deployable
quiet observation platform or as a communications platforms.
15. Venus Environment
• Rotation Period (Day) of Venus is Longer the
Revolution Period (Year) Potentially Enabling
Continuous Flight
• Atmosphere is mainly Carbon Dioxide (96.5%)
Also contains trace amounts of corrosive
Compounds (Hydrochloric, hydrofluoric &
Sulfuric Acids)
Earth Surface • Atmospheric Density Equals Earth Surface
Density at ~50 km
• Incident Solar Intensity is ~2600 W/m2
• Very high wind speeds above the cloud tops
~ 100 m/s
• Clouds On Venus Extend Upwards
to ~64 km
16. Mars Environment
• The atmosphere on Mars is very thin. At the Surface the density is similar to
30 km on Earth
• The atmosphere is composed mostly of Carbon Dioxide
• The temperature on Mars is on average much colder then on Earth
Although at certain times of the year and locations the temperature
will rise above freezing, most of the time temperatures are well
below the freezing point of water.
• The gravitational force on Mars
(3.57 m/s2)is about 1/3 what it is on Earth.
• Solar intensity at Mars is ~590 W/m2
• There are few clouds but dust storms are fairly common
17. Earth Environment
• Gravitational Force 9.81 m/s2
• Solar Intensity 1352 W/m2
• Atmospheric composition is approximately 80%
Nitrogen, 20% Oxygen
Wind speeds generally increase from
the surface up to a maximum around
the top of the Troposphere (Jet Stream)
The majority of Earth’s weather occurs within
the Troposphere which extends to approximately 12 km
18. Lift and Thrust Generation
The propulsion force and lift generation of the aircraft
are accomplished by the flapping of the wings.
By altering the shape and angle of
attack of the wing the amount of lift
and the direction of this lift force
can be controlled
This is the same method birds use to
generate lift and thrust
The lift and lift vector generated can
Vary between each wing as well as
along the wing span itself.
This provides a significant amount of
control and provides a means for
maneuvering
19. Wing Aerodynamics
Like all flapping wing flyers in nature the solid state aircraft will operate
within a low Reynolds number flight regime. This is due mainly to its
required low wing loading and the potential for high altitude operation,
where the air density is low.
6.00E+05
ρVc
2.12 kg/m^2 Wing Loading
Re =
µ
2.62 kg/m^2 Wing Loading
5.00E+05 3.12 kg/m^2 Wing Loading
4.00E+05
Wing Assumptions:
3.00E+05
2.00E+05
Curved flat plate airfoil
Rectangular wing planform
1.00E+05
0.00E+00
These are conservative
1.225 1.007 0.66 0.414 0.089
Atmospheric Density (kg/m^3)
0.019 0.004
estimates. Wing and
52 56 Venus Altitude (km) 61 70 85 aircraft performance can be
0 4 Earth Altitude (km) 10 20 40
increased by optimizing
Mars Altitude (km) 0 4.5 14
the wing design for a
specific flight regime.
20. Effect of Re# on Aircraft Performance
1.4
1.2
Flight Reynolds number
1
based on chord length
0.8 Re# 420,000
Re# 128,000
for a flat plate airfoil
Re# 84,000
Re# 42,000
0.6
0.4
0.2
0.1
0
0 2 4 6 8 10 12
0.09
Angle of Attack ( α, degrees)
Re# 105,000
0.08 Re# 84,000
Re# 42,000
0.07
0.06
0.05
0.04
0.03
0.02
0.01
0
0 0.2 0.4 0.6 0.8 1 1.2 1.4
Lift Coefficient (Cl)
21. CFD Airfoil Analysis
CFD analysis is ongoing to provide lift coefficient
and drag coefficient data for a thin curved airfoil at
Angles of attack and Reynolds numbers representative
Of the estimated SSA flight regime and wing motion
22. Sizing Analysis
An analysis was performed to determine the feasibility of the
SSA concept and establish the range of operation on the planets
of interest.
Power Production Power Consumption
Incoming solar flux Motion of the wing
Time of year Drag
Latitude Lift Generation
Flapping frequency Induced drag Altitude
Maximum flap angle traversed Profile drag Flight Speed
Length of the wing
Wing Geometry
23. Power Production
The amount of power available to the aircraft is based on the
Environmental conditions it is flying within
Output power will vary
based on the
Latitude of flight (φ)
Time of year (δ)
Time of day (θ)
Available power also depends
on the
Atmosphere attenuation (τ)
Solar cell efficiency (η)
24. Power Production: Venus
There is no variation between daily and yearly power profiles
because of the very long day length (equal to 263 Earth days) which
is longer then the Venus year (equal to 244 Earth days)
180
160 0° Latitude
10° Latitude • Solar cell efficiency 10%
• Solar cell fill factor 80%
20° Latitude
140
30° Latitude
40° Latitude
120 50° Latitude • Horizontal solar array
60° Latitude
Available Power (W/m2)
100
70° Latitude
80° Latitude
• Atmospheric
80
90° Latitude Attenuation 25%
• Mean Solar Intensity above
60
atmosphere 2620 W/m2
40
• Longitude 0°
20 • Maximum declination
0
angle 3°
0.0 50.0 100.0 150.0 200.0
Time (Earth Days)
Available Power Throughout A Day (Earth Days)
25. Power Production: Earth
Available Power Throughout the Day (Hours)
100
Latitude 0° N
90
Vernal Equinox (March 21st)
•Solar cell efficiency 10%
80
Summer Solstice (June 22)
Autumnal Equinox (September 23rd) • Solar cell fill factor 80%
Winter Solstice (December 22nd)
70
• Horizontal solar array
Power Available (W/m2)
60
• Atmospheric
50
Attenuation 15%
40
30
20 100
Latitude 80° N
10 90
Vernal Equinox (March 21st)
Summer Solstice (June 22)
0 80
Autumnal Equinox (September 23rd)
0.00 5.00 10.00 15.00 20.00
Winter Solstice (December 22nd)
Time of Day (Hours) 70
Power Available (W/m2)
60
• Mean Solar Intensity above 50
atmosphere 1353 W/m2 40
• Longitude 0° 30
• Maximum declination 20
angle 23.5° 10
0
0.00 5.00 10.00 15.00 20.00
Time of Day (Hours)
26. Power Production: Mars
Available Power Throughout the Day (Hours)
60
0° Latitude
50 Vernal Equinox (Day 1)
Summer Solsitce (Day 167)
•Solar cell efficiency 10%
Autimal Equinox (Day 333)
Winter Solstice (Day 500)
• Solar cell fill factor 80%
• Horizontal solar array
40
Available Power (W/m2)
30
• Atmospheric
Attenuation 15%
20
25
10
80° Latitude
Vernal Equinox (Day 1)
0 20 Summer Solsitce (Day 167)
0 5 10 15 20 Autimal Equinox (Day 333)
Time (hours)
Available Power (W/m2) Winter Solstice (Day 500)
15
•Mean Solar Intensity above 10
atmosphere 590 W/m2
• Longitude 0° 5
• Maximum declination
angle 24° 0
0 5 10 15 20
Time (hours)
27. Power Consumption due to Motion
Motion of the wing consists of a The forces generated by the wing
flapping rate and maximum motion are due to the acceleration and
angle traversed during the flap deceleration of the wing mass.
These forces vary along the wing
length
28. Wing Force Due to Motion
Force versus Distance Traveled for Various Wing The power required to
Lengths, Maximum Flap Angles and Flap Frequencies move the wing is the
area under the force vs
10
distance traveled curve.
1 The distance traveled
0 0.5 1 1.5 2 2.5 3 3.5
varies along the wing
0.1
length to a maximum at
Length 1m, Angle 40°, Cycle 1s
Length 1m, Angle 40°, Cycle 4s
the tip.
Force (N)
Length 1m, Angle 60°, Cycle 1s
0.01 Length 1m, Angle 60°, Cycle 4s
Length 3m, Angle 40°, Cycle 1s Power consumption
Length 3m, Angle 40°, Cycle 4s
Length 3m, Angle 60°, Cycle 1s
can be reduced by
0.001
Length 3m, Angle 60°, Cycle 4s
tapering the wing so
there is less mass at the
0.0001
tip. Thereby reducing
the force needed for
0.00001 motion.
Distance Travled (m)
29. Drag Due to Lift and Velocity
1000
The aerodynamic drag is due to the
2.12 kg/m^2 Wing Loading
2.62 kg/m^2 Wing Loading
3.12 kg/m^2 Wing Loading
generation of lift and the movement
100
of the aircraft through the atmosphere
i
1
10 D f = ∑ ρVi 2 (c f 2S + cd S)
0
2
1
1.225 1.007 0.66 0.414 0.089 0.019 0.004
Atmospheric Density (kg/m^3)
52 56 Venus Altitude (km) 61 70 85
0 4 Earth Altitude (km) 10 20 40
4
Mars Altitude (km) 0 4.5 14
3.5
Cycle 1s"
Cycle 4s"
Cycle 2s
3 Cycle 3s
The drag is dependent on the flight Lift Force (N)
velocity which in turn sets the lifting 2.5
capacity and the lift to drag 2
characteristics of the airfoil.
1.5
1
0.03 0.33 0.63 0.93 1.23 1.53 1.83 2.13 2.43 2.73
Length Along Wing (m)
30. Analysis Method
The energy consumed during the flap
has to equal the energy collected
during the total flap and glide cycle.
Inputs
Flapping Frequency
Power
Flap Angle
Available
Aircraft Size
Altitude
Lift Power Consumed
Generated (Motion & Drag) •The analysis was an iterative
process
•For a given set of inputs the flight
speed and flap to glide ratio would
be calculated.
•If no solution existed then certain
Flight Speed
inputs would be varied until a
Flap to Glide Ratio
solution was found.
31. Operational Scheme
•The aircraft will fly in a manner
similar to an Eagle or other large
bird.
•The aircraft will glide for
extended periods of time and flap
its wings periodically to regain
altitude and increase forward
speed.
The ratio of glide time to flap time will
depend on the available power, power
consumption and flight conditions.
The analysis was performed to determine
the optimal glide to flap ratio for a give
aircraft configuration under a specific
flight condition
32. Aircraft Sizing
The initial feasibility study was performed to determine the capabilities of
the aircraft under the environmental conditions of the planets of interest.
This initial analysis was based on the Potential flight altitude ranges
following assumptions investigated in the analysis for each
of the planets of interest
Wing Loading 3.12 kg/m2 Venus 53 km to
Aspect Ratio 8 82 km
Wing Friction Coefficient 0.008 Earth 1 km to
Maximum Flap Angle 45° 35 km
Solar Cell Efficiency 10% Mars 1 km to
7 km
Solar Cell Specific Mass 0.12 kg/m2
Battery Specific Mass 0.75 kg/m2
IPMC Specific Mass 2.00 kg/m2
Payload Specific Mass 0.25 kg/m2
33. 160
1 km Earth, 53 km Venus, O km Mars
Sizing Results
140 3 km Earth, 55 km Venus, 0 km Mars 3 m Wingspan SSA
5 km Earth, 57 km Venus, 0 km Mars
120
10 km Earth, 61 km Venus, 0 km Mars
100
15 km Earth, 65 km Venus, 0 km Mars
Required power in W/m2 of wing
20 km Earth, 70 km Venus, 0 km Mars
80
area were generated for a range of
flap durations over a number of
60
altitude levels
40
20
0
5
0 1 2 3 4 5 6
1 km Earth, 53 km Venus, O km Mars
Flap Duration (s) 4.5
3 km Earth, 55 km Venus, 0 km Mars
4 5 km Earth, 57 km Venus, 0 km Mars
Form the curves it can be seen that 3.5
10 km Earth, 61 km Venus, 0 km Mars
15 km Earth, 65 km Venus, 0 km Mars
for a given size aircraft there is a
Glide Duration (s)
3
20 km Earth, 70 km Venus, 0 km Mars
flap duration that produces a 2.5
minimum required specific power. 2
3 m Wingspan
1.5
The minimum specific power 1
occurs at smaller flap durations 0.5
(higher flapping frequencies) as the 0
0 1 2 3 4 5 6
flight altitude is increased. Flap Duration (s)
34. Sizing Results
160
25 km Earth, 74 km Venus, 0 km Mars
140 30 km Earth, 78 km Venus, 1 km Mars
32 km Earth, 79 km Venus, 1.4 km Mars 100 m Wingspan SSA
120 35 km Earth, 82 km Venus, 5 km Mars
Specific Power Required (W/m^2)
100
80 As the altitude increases the
60
vehicle must flap more frequently
40
or continuously to maintain flight.
20
100 m Wingspan
0
0 5 10 15 20 25 30 3535 40
Flap Duration (s) 25 km Earth, 74 km Venus, 0 km Mars
30 30 km Earth, 78 km Venus, 1 km Mars
32 km Earth, 79 km Venus, 1.4 km Mars
For all vehicle sizes the glide 25
35 km Earth, 82 km Venus, 5 km Mars
duration decreases with
Glide Duration (s)
increasing altitude 20
15
As the flap duration increases 100 m Wingspan
(lower frequency) the glide 10
duration approaches zero 5
0
0 5 10 15 20 25 30 35 40
Flap Duration (s)
35. 140
Sizing Results Optimal Operation
3 m Wingspan
120
Flap duration, Glide Duration and
6 m Wingspan
12 m Wingspan
20 m Wingspan
100
Required Power as a Function of
50 m Wingspan
Glide Duration (s)
100 m Wingspan
250 m Wingspan
80
Altitude (Earth) for Various Size Aircraft
60
40
Graphs shown are based on the flap
20 duration that produced a minimum
0
0 5 10 15 20 25 30 35 40
required power for a given size aircraft
Earth Altitude (km)
and flight altitude
Glide duration goes to zero with increasing altitude
45 200
40
180
3 m Wingspan 160 3 m Wingspan
35 6 m Wingspan 6 m Wingspan
12 m Wingspan 12 m Wingspan
140
20 m Wingspan 20 m Wingspan
Required Power (W/m^2)
30
50 m Wingspan 50 m Wingspan
Flap Duration (s)
100 m Wingspan 120 100 m Wingspan
250 m Wingspan 250 m Wingspan
25
100
20
80
15
60
10 40
5 20
0
0
0 5 10 15 20 25 30 35 40
0 5 10 15 20 25 30 35 40
Earth Altitude (km)
Earth Altitude (km)
Flap duration remains constant until the glide Required power increases exponentially with
Duration reaches zero then begins to decrease increasing altitude
36. Internal Configuration Options
•Skeletal structure where strips of
IMPC Material are used to produce
motion by contracting
•Limits control but may be lighter and
stronger then the continuous sheet
option
•Continuous Sheet of IMPC with
electrode gird
•Provides very fine control of
motion
37. Nature Inspired Configuration
• The Pteranodon is the largest animal that ever flew.
• It is the closest in size, weight and wing span to the SSA
• As an initial starting point for a more detailed wing design the Pteranodon
wing will be used as the model (nature has a way of finding the optimum)
38. Future Plans
– Key Items to Further Develop the SSA Concept
• Construct a small (~.5 m) wing section and demonstrate
the operating principle of the vehicle
– Integration of Photovoltaic array and IMPC
– Establish a control scheme of wing motion
• Perform a detailed wing and airfoil design optimized for
flapping flight under the specified operational Reynolds
number
– CFD and wind tunnel validation
• Perform a more detailed system study
– Examine variations in wing geometry, operation conditions and
component masses
– Evaluate mission potential & payload
