Solid State Aircraft
Anthony ColozzaNorthland Scientific / Ohio Aerospace Institute
Cleveland, Ohio
NIAC Fellows Conference October 23-24, 2002
NIAC HeadquartersAtlanta, Georgia
Phase I Project
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
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
Aircraft Operation•The aircraft operates by collecting and 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 move thereby causing a flapping motion of the wing.
•This flapping motion produces lift and thrust for the aircraft.
•The electric field generated by the grids is controllable, therefore the shape and motion of the wing is controllable on each flap.
Thin Film Array
Thin Film Battery
Cathode Grid
Anode Grid
IPMC Material
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.
•This control enables the wing to flap, provide differential lift (which is used for steering), and alter the camber of the wing to maximize lift under a given operational condition.
Each Grid Location is Individually Controllable
PV Array
Battery
Cathode
Anode
IPMC
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
Thin Film Solar Array Historical Performance Trends
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1976 1978 1980 1982 1984 1986 1988 1990 1992 1994 1996
CdS/Cu2SCuInSe2CuGaSe2CuInS2CdTea-SiCASCADES
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
Battery Construction & OperationTypes 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)
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 deformation is directly related to 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.
QuickTime™ and a decompressor
are needed to see this picture.
IPMC MaterialConstructed 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 shape
Metal Electrodes
P IE M-
+
O N
O F F
-
+
P t E le c tro d e
P IE M
C o n ta c t E le c tro d e
P t Pa r tic le s
P IE M
The material will bend toward the anode side of the electrodes
IPMC Motion
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side chain hydrated cation-Na(H2O)4+ water fixed anion mobile cation
+
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+-
Under an electric field the ion exchange membraneEnables 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 dryenvironments may require sealing or redesign of the material for efficient long term use
IPMC Material CharacteristicsYoung' s Modulu s, E Up to 2 GPaShear Modulu s, G Up to 1 GPaPoisson's ratio, ν Typical: 0.3-0.4Power density (W/mass) Up to 100 J/kgMax force density (Cantilever Mode) Up to 40 Kgf/KgMax displacement/strain Up to 4% linear strainBandwidth (speed) Up to 1 kHz in cantilever vib ratory mode
for actuationsUp to 1 MHz for sensing
Resolution (force and di splacementcontrol )
Displacement accuracy down to 1 micronForce resolut ion down to 1 mg
Efficiency (electrom echanical) Up to 6 % (frequency depend ent) foractuationUp to 90% for sensing
Density Down to 1.8 g/cm3
Solid State Aircraft Applications
There is sufficient solar intensity for this aircraft to operateon 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.
Venus Environment
• Rotation Period (Day) of Venus is Longer the Revolution Period (Year) Potentially Enabling Continuous Flight
Earth Surface
• Atmosphere is mainly Carbon Dioxide (96.5%)Also contains trace amounts of corrosive Compounds (Hydrochloric, hydrofluoric & Sulfuric Acids)
• 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
Mars Environment
• The atmosphere on Mars is very thin. At the Surface the density is similar to30 km on Earth
• The atmosphere is composed mostly of Carbon Dioxide
• The temperature on Mars is on average much colder then on EarthAlthough 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
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 withinthe Troposphere which extends to approximately 12 km
Lift and Thrust GenerationThe 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 canVary between each wing as well as along the wing span itself.
This provides a significant amount of control and provides a means for maneuvering
Wing AerodynamicsLike 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.
Re =ρVcµ
0.00E+00
1.00E+05
2.00E+05
3.00E+05
4.00E+05
5.00E+05
6.00E+05
1.225 1.007 0.66 0.414 0.089 0.019 0.004
Atmospheric Density (kg/m^3)
2.12 kg/m^2 Wing Loading2.62 kg/m^2 Wing Loading3.12 kg/m^2 Wing Loading
0
4.5
4 10 20 40Earth Altitude (km)
52 56 61 70 85Venus Altitude (km)
0 14Mars Altitude (km)
Wing Assumptions:
Curved flat plate airfoilRectangular wing planform
These are conservative estimates. Wing and aircraft performance can be increased by optimizing the wing design for a specific flight regime.
Effect of Re# on Aircraft Performance
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Angle of Attack ( α, degrees)
Re# 420,000Re# 128,000Re# 84,000Re# 42,000
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Lift Coefficient (Cl)
Re# 105,000Re# 84,000Re# 42,000
Flight Reynolds number based on chord length for a flat plate airfoil
CFD Airfoil AnalysisCFD analysis is ongoing to provide lift coefficientand 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
Sizing AnalysisAn analysis was performed to determine the feasibility of theSSA concept and establish the range of operation on the planetsof interest.
Power Production Power Consumption
Incoming solar fluxTime of year Latitude
Motion of the wing DragLift Generation
Flapping frequencyMaximum flap angle traversed Length of the wing
Induced drag Profile drag
Altitude Flight Speed
Wing Geometry
Power ProductionThe 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 dependson the
Atmosphere attenuation (τ)Solar cell efficiency (η)
Power Production: VenusThere is no variation between daily and yearly power profilesbecause of the very long day length (equal to 263 Earth days) which is longer then the Venus year (equal to 244 Earth days)
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Time (Earth Days)
Ava
ilabl
e Po
wer
(W/m
2)
0° Latitude10° Latitude20° Latitude30° Latitude40° Latitude50° Latitude60° Latitude70° Latitude80° Latitude90° Latitude
• Solar cell efficiency 10%• Solar cell fill factor 80%• Horizontal solar array• Atmospheric
Attenuation 25%• Mean Solar Intensity above
atmosphere 2620 W/m2
• Longitude 0°• Maximum declination
angle 3°
Available Power Throughout A Day (Earth Days)
Power Production: Earth
•Solar cell efficiency 10%• Solar cell fill factor 80%• Horizontal solar array• Atmospheric
Attenuation 15%
Available Power Throughout the Day (Hours)Latitude 0° N
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Time of Day (Hours)
Pow
er A
vaila
ble
(W/m
2)
Vernal Equinox (March 21st)Summer Solstice (June 22)Autumnal Equinox (September 23rd)Winter Solstice (December 22nd)
Latitude 80° N
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Time of Day (Hours)
Pow
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vaila
ble
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Vernal Equinox (March 21st)Summer Solstice (June 22)Autumnal Equinox (September 23rd)Winter Solstice (December 22nd)
• Mean Solar Intensity above atmosphere 1353 W/m2
• Longitude 0°• Maximum declination
angle 23.5°
Power Production: MarsAvailable Power Throughout the Day (Hours)
0° Latitude
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Time (hours)
Ava
ilabl
e Po
wer
(W/m
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Vernal Equinox (Day 1)Summer Solsitce (Day 167)Autimal Equinox (Day 333)Winter Solstice (Day 500)
80° Latitude
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Time (hours)
Ava
ilabl
e Po
wer
(W/m
2)
Vernal Equinox (Day 1)Summer Solsitce (Day 167)Autimal Equinox (Day 333)Winter Solstice (Day 500)
•Solar cell efficiency 10%• Solar cell fill factor 80%• Horizontal solar array• Atmospheric
Attenuation 15%
•Mean Solar Intensity above atmosphere 590 W/m2
• Longitude 0°• Maximum declination
angle 24°
Power Consumption due to Motion
The forces generated by the wing motion are due to the acceleration and deceleration of the wing mass. These forces vary along the wing length
Motion of the wing consists of a flapping rate and maximum angle traversed during the flap
Wing Force Due to Motion
Force versus Distance Traveled for Various Wing Lengths, Maximum Flap Angles and Flap Frequencies
The power required to move the wing is the area under the force vsdistance traveled curve.
The distance traveled varies along the wing length to a maximum at the tip.
Power consumption can be reduced by tapering the wing so there is less mass at the tip. Thereby reducing the force needed for motion. 0.00001
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Distance Travled (m)
Forc
e (N
)
Length 1m, Angle 40°, Cycle 1sLength 1m, Angle 40°, Cycle 4sLength 1m, Angle 60°, Cycle 1sLength 1m, Angle 60°, Cycle 4sLength 3m, Angle 40°, Cycle 1sLength 3m, Angle 40°, Cycle 4sLength 3m, Angle 60°, Cycle 1sLength 3m, Angle 60°, Cycle 4s
Drag Due to Lift and Velocity
1
10
100
1000
1.225 1.007 0.66 0.414 0.089 0.019 0.004
Atmospheric Density (kg/m^3)
2.12 kg/m^2 Wing Loading2.62 kg/m^2 Wing Loading3.12 kg/m^2 Wing Loading
0
4.5
4 10 20 40Earth Altitude (km)
52 56 61 70 85Venus Altitude (km)
0 14Mars Altitude (km)
1
1.5
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2.5
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3.5
4
0.03 0.33 0.63 0.93 1.23 1.53 1.83 2.13 2.43 2.73
Length Along Wing (m)
Lift
Forc
e (N
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Cycle 1s"Cycle 4s"Cycle 2sCycle 3s
D f =12
ρVi2 (c f 2S+ cdS)
0
i
∑
The aerodynamic drag is due to the generation of lift and the movement of the aircraft through the atmosphere
The drag is dependent on the flight velocity which in turn sets the lifting capacity and the lift to drag characteristics of the airfoil.
Analysis MethodThe energy consumed during the flap has to equal the energy collected during the total flap and glide cycle.
Power Available
Lift Generated
Power Consumed(Motion & Drag)
Inputs Flapping Frequency Flap AngleAircraft SizeAltitude
Flight SpeedFlap to Glide Ratio
•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 inputs would be varied until a solution was found.
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 determinethe optimal glide to flap ratio for a give aircraft configuration under a specific flight condition
Aircraft SizingThe initial feasibility study was performed to determine the capabilities of the aircraft under the environmental conditions of the planets of interest.
Potential flight altitude ranges investigated in the analysis for each of the planets of interest
This initial analysis was based on the following assumptions
0.25 kg/m2Payload Specific Mass2.00 kg/m2IPMC Specific Mass0.75 kg/m2Battery Specific Mass0.12 kg/m2Solar Cell Specific Mass10%Solar Cell Efficiency45°Maximum Flap Angle0.008Wing Friction Coefficient8Aspect Ratio3.12 kg/m2Wing Loading
1 km to 7 km
Mars
1 km to 35 km
Earth
53 km to 82 km
Venus
Sizing Results
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Flap Duration (s)
1 km Earth, 53 km Venus, O km Mars
3 km Earth, 55 km Venus, 0 km Mars
5 km Earth, 57 km Venus, 0 km Mars
10 km Earth, 61 km Venus, 0 km Mars
15 km Earth, 65 km Venus, 0 km Mars
20 km Earth, 70 km Venus, 0 km Mars
3 m Wingspan SSA
Required power in W/m2 of wing area were generated for a range of flap durations over a number of altitude levels
3 m Wingspan
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Flap Duration (s)
Glid
e D
urat
ion
(s)
1 km Earth, 53 km Venus, O km Mars
3 km Earth, 55 km Venus, 0 km Mars
5 km Earth, 57 km Venus, 0 km Mars
10 km Earth, 61 km Venus, 0 km Mars
15 km Earth, 65 km Venus, 0 km Mars
20 km Earth, 70 km Venus, 0 km Mars
Form the curves it can be seen that for a given size aircraft there is a flap duration that produces a minimum required specific power.
The minimum specific power occurs at smaller flap durations (higher flapping frequencies) as the flight altitude is increased.
100 m Wingspan0
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Flap Duration (s)
Spec
ific
Pow
er R
equi
red
(W/m
^2)
25 km Earth, 74 km Venus, 0 km Mars
30 km Earth, 78 km Venus, 1 km Mars
32 km Earth, 79 km Venus, 1.4 km Mars
35 km Earth, 82 km Venus, 5 km Mars100 m Wingspan SSASizing Results
100 m Wingspan
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Flap Duration (s)
Glid
e D
urat
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25 km Earth, 74 km Venus, 0 km Mars
30 km Earth, 78 km Venus, 1 km Mars
32 km Earth, 79 km Venus, 1.4 km Mars
35 km Earth, 82 km Venus, 5 km Mars
As the altitude increases the vehicle must flap more frequently or continuously to maintain flight.
For all vehicle sizes the glide duration decreases with increasing altitude
As the flap duration increases (lower frequency) the glide duration approaches zero
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Earth Altitude (km)
Glid
e D
urat
ion
(s)
3 m Wingspan6 m Wingspan12 m Wingspan20 m Wingspan50 m Wingspan100 m Wingspan250 m Wingspan
Glide duration goes to zero with increasing altitude
Sizing Results Optimal OperationFlap duration, Glide Duration and Required Power as a Function of Altitude (Earth) for Various Size Aircraft
Graphs shown are based on the flap duration that produced a minimum required power for a given size aircraft and flight altitude
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0 5 10 15 20 25 30 35 40Earth Altitude (km)
Req
uire
d Po
wer
(W/m
^2)
3 m Wingspan6 m Wingspan12 m Wingspan20 m Wingspan50 m Wingspan100 m Wingspan250 m Wingspan
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0 5 10 15 20 25 30 35 40Earth Altitude (km)
Flap
Dur
atio
n (s
)
3 m Wingspan6 m Wingspan12 m Wingspan20 m Wingspan50 m Wingspan100 m Wingspan250 m Wingspan
Flap duration remains constant until the glide Duration reaches zero then begins to decrease
Required power increases exponentially with increasing altitude
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
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)
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