Solid state aircrafts


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Solid state aircrafts

  1. 1. Solid State Aircraft Phase I Project NIAC Fellows Conference October 23-24, 2002 NIAC Headquarters Atlanta, Georgia Anthony ColozzaNorthland Scientific / Ohio Aerospace Institute Cleveland, Ohio
  2. 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. 3. Solid State Aircraft Artist Concept Drawing The aircraft concept is to integrate three unique types of materials (thin film solar arrays, thin film lithiumbatteries and an ionic polymer metal composite) to produce anaircraft that has no moving parts, can fly at high altitudes, is easilydeployable and has applications on Earth, Venus and Mars
  4. 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 toThin Film move thereby causing a flappingBattery 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. 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. Thisgrid will enable various voltages to be sent to different sections of thewing, thereby causing varying degrees of motion along the wingsurface. The amount of control on the wing will depend on the finenessof this control grid. A central processor will be used to control thepotential 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. 6. Thin Film Photovoltaic ArrayLight Weight: Activematerial is on the orderof 1 to 2 microns thickHighly Flexible: Ideal for theflexing and motion of a flapping wingSubstrate: Can be made of most materials,presently the best candidate is Kapton (or otherpolymers). Potentially the Battery or IPMC can beutilized as the substrateSpecific Power: 1 kW/kg near term, 2 kW/kg projected
  7. 7. Thin Film Solar Array Historical Performance Trends16 CdS/Cu2S14 CuInSe2 CuGaSe2 CuInS212 CdTe a-Si CASCADES10 8 6 4 2 0 1976 1978 1980 1982 1984 1986 1988 1990 1992 1994 1996
  8. 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 timesITNES 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. 9. Battery Construction & OperationTypes of Lithium ion thin-film batteries differ in the cathodematerial they use. –Ex. Magnesium Oxides, Cobalt Oxides, Yttrium OxideThe battery is produced by depositing (through sputtering orevaporation techniques) the various material layers that makeup the components (cathode, electrolyte, anode and currentcollector) onto a substrate. Oak Ridge Battery Design Cross Section (15 micron thick)
  10. 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. 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 sThe material will P t E le c tro d e O FFbend toward the anode P IE Mside of the electrodes + P IE M - ON
  12. 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. 13. IPMC Material CharacteristicsYoung s Modulu s, E Up to 2 GPaShear Modulu s, G Up to 1 GPaPoissons ratio, ν Typ ical: 0.3-0.4Power density (W/mass) Up to 100 J/kgMax force density (Cantilever Mod e) Up to 40 Kgf/KgMax displacement/strain Up to 4% linear strainBandwidth (speed) Up to 1 kHz in cantil ever vib ratory mode for actuations Up to 1 MHz for sensingResolution (force and di splacement Displacement accuracy down to 1 microncontrol ) Force resolut ion down to 1 mgEfficiency (electrom echanical) Up to 6 % (frequency depend ent) for actuation Up to 90% for sensingDensity Down to 1.8 g/cm 3
  14. 14. Solid State Aircraft Applications There is sufficient solar intensity for this aircraft to operate on Earth, Venus or MarsBecause of its projected relatively small mass and flexibility, the aircraftis ideal for planetary exploration. These characteristics allow the aircraftto be easily stowed and launched at a minimal cost. Potentially, a fleet ofthese aircraft could be deployed within a planet’s atmosphere and usedfor comprehensive scientific data gathering, as an quickly deployablequiet observation platform or as a communications platforms.
  15. 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. 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. 17. Earth Environment • Gravitational Force 9.81 m/s2 • Solar Intensity 1352 W/m2 • Atmospheric composition is approximately 80% Nitrogen, 20% OxygenWind speeds generally increase fromthe surface up to a maximum aroundthe top of the Troposphere (Jet Stream) The majority of Earth’s weather occurs within the Troposphere which extends to approximately 12 km
  18. 18. Lift and Thrust GenerationThe propulsion force and lift generation of the aircraftare 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. 19. Wing AerodynamicsLike all flapping wing flyers in nature the solid state aircraft will operatewithin a low Reynolds number flight regime. This is due mainly to itsrequired 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. 20. Effect of Re# on Aircraft Performance1.41.2 Flight Reynolds number 1 based on chord length0.8 Re# 420,000 Re# 128,000 for a flat plate airfoil Re# 84,000 Re# 42,0000.60.40.2 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. 21. CFD Airfoil AnalysisCFD analysis is ongoing to provide lift coefficientand drag coefficient data for a thin curved airfoil atAngles of attack and Reynolds numbers representativeOf the estimated SSA flight regime and wing motion
  22. 22. 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 flux Motion of the wing Time of year Drag Latitude Lift GenerationFlapping frequency Induced drag AltitudeMaximum flap angle traversed Profile drag Flight SpeedLength of the wing Wing Geometry
  23. 23. Power ProductionThe amount of power available to the aircraft is based on theEnvironmental 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. 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° LatitudeAvailable 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. 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 arrayPower 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. 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 40Available 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. 27. Power Consumption due to MotionMotion of the wing consists of a The forces generated by the wingflapping rate and maximum motion are due to the acceleration andangle traversed during the flap deceleration of the wing mass. These forces vary along the wing length
  28. 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. 29. Drag Due to Lift and Velocity1000 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. 30. Analysis MethodThe energy consumed during the flaphas to equal the energy collectedduring 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. 31. Operational Scheme•The aircraft will fly in a mannersimilar to an Eagle or other largebird.•The aircraft will glide forextended periods of time and flapits wings periodically to regainaltitude and increase forwardspeed. 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. 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 rangesfollowing assumptions investigated in the analysis for each of the planets of interestWing Loading 3.12 kg/m2 Venus 53 km toAspect Ratio 8 82 kmWing Friction Coefficient 0.008 Earth 1 km toMaximum Flap Angle 45° 35 kmSolar Cell Efficiency 10% Mars 1 km to 7 kmSolar Cell Specific Mass 0.12 kg/m2Battery Specific Mass 0.75 kg/m2IPMC Specific Mass 2.00 kg/m2Payload Specific Mass 0.25 kg/m2
  33. 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 MarsForm 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 Marsfor a given size aircraft there is a Glide Duration (s) 3 20 km Earth, 70 km Venus, 0 km Marsflap duration that produces a 2.5minimum required specific power. 2 3 m Wingspan 1.5The minimum specific power 1occurs at smaller flap durations 0.5(higher flapping frequencies) as the 0 0 1 2 3 4 5 6flight altitude is increased. Flap Duration (s)
  34. 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 MarsSpecific 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. 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 WingspanGlide 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 altitudeGlide 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. 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. 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. 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