Solid state aircrafts

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

+

+

+

+

+

+

+-

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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0.0 50.0 100.0 150.0 200.0

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)

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Vernal Equinox (March 21st)Summer Solstice (June 22)Autumnal Equinox (September 23rd)Winter Solstice (December 22nd)

Latitude 80° N

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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


angle 23.5°

Power Production: MarsAvailable Power Throughout the Day (Hours)

0° Latitude

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Time (hours)

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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

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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


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

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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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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)

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∑

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





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

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urat

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(s)

1 km Earth, 53 km Venus, O 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

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equi

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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

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32 km Earth, 79 km Venus, 1.4 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

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(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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Flap

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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

Solid state aircrafts

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