1. Solid State Aircraft
Phase II Project
NIAC 5thAnnual Meeting
November 5-6, 2003
Georgia Centers for Advanced
Telecommunications Technology
Atlanta, Georgia
Anthony Colozza
Northland Scientific / Ohio Aerospace Institute
Cleveland, Ohio

2. Solid State Aircraft Project
Team Members
•Mr. Anthony Colozza (PI): NSI/OAI
•Mr. Phillip Jenkins:OAI
•Dr. Mohsen ShahinpoorUniversity of New Mexico
•Mr. Teryn Dalbello:University of Toledo/ICOMP
•Mr. Curtis Smith:OAI
•Dr. Carol KoryConsultant
•Dr. Iakkattukuzhy IsaacUniversity of Missouri-Rolla

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, iseasily deployable and has applications on Earth, Venus and Mars

4. 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 thin film lithium 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 ArrayThin Film BatteryCathode GridAnode GridIPMC Material

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 is 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 is 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 ControllablePV ArrayBatteryCathodeAnodeIPMC

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 projected024681012141619761978198019821984198619881990199219941996CdS/Cu2SCuInSe2CuGaSe2CuInS2CdTea-SiCASCADES

7. 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

8. Battery Construction & Operation
Types of Lithium ion thin-film batteries differ in the cathode material they use. –Ex. Magnesium Oxides, Cobalt Oxides, Yttrium OxideThe 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)

9. 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 decompressorare needed to see this picture.

10. 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 shape
Metal ElectrodesPIEM- + ONOFF- + Pt ElectrodePIEMContact ElectrodePt ParticlesPIEM
The material will
bend toward the anode
side of the electrodes

11. IPMC Motion + + + + + + ++ side chainhydrated cation-Na(H2O)4+waterfixed anionmobile 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 moveFor the IPMC material to operate it must be sufficiently HydratedLeakage and operation in dryenvironments may require sealing or redesign of the material for efficient long term use

12. IPMC Material CharacteristicsYoung's Modulus, EUp to 2 GPaShear Modulus, GUp to 1 GPaPoisson's ratio, νTypical: 0.3-0.4Power density (W/mass)Up to 100 J/kgMax force density (Cantilever Mode)Up to 40 Kgf/KgMax displacement/strainUp to 4% linear strainBandwidth (speed)Up to 1 kHz in cantilever vibratory modefor actuationsUp to 1 MHz for sensingResolution (force and displacementcontrol) Displacement accuracy down to 1 micronForce resolution down to 1 mgEfficiency (electromechanical)Up to 6 % (frequency dependent) foractuationUp to 90% for sensingDensityDown to 1.8 g/cm3Deflection (cm) vs VoltageRoom Temperature100°CRoom Temperature100°CPowerConsumption (W) vs Voltage

13. Material IntegrationGoal: To produce a method for combining the solar array, batteryand IPMC materials into a single flexible structure. •One of the key aspects of the SSA’s feasibility is the integration of the three main component materials•To continue to establish the concept feasibility, material integration, which was not specifically address under Phase I, will be a major part of the Phase II program.

14. Near Term Integration Plan
Goal: To provide a wing section for testing under laboratory conditions
•Off-the-Shelf components will be used as much as possible.
•The array & IPMC will be assembled using adhesives, lamination or other readily available means.
•The control grid will be constructed integral with the IPMC material.
•Potential adhesion methods will be tested on coupon samples of the thin film array, IPMC, wiring and electrodes.
•Due to the development stage of thin film batteries, these initial integration tests will only include the array and IPMC materials.
•External batteries or capacitors will be used during this initial testing.
•The various integration methods will be tested for adhesion and bending capability.
•The near term integration method that provides the greatest flexibility and adhesion will then be used in the construction of a large wing section.

15. •The potential of construction the SSA by depositing each component material layer onto the subsequent layer is being investigated. •This method would use the IPMC material as the main substrate. •The battery layer and solar cell layer as well as the electrodesand wiring would be build directly onto the IPMC through a combination of deposition and lithography techniques. •Some of the main issues that are to be examined are the materialcompatibilities, adhesion characteristics, material thickness and deposition and lithography techniques, their capabilities and requirements. Goal: To provide a scheme to producing a fully integrated composite structure from the three main materials and subsequent components. Long Term Integration Strategy

16. Wing Control
•The placement of electrodes and the shape of the electric field they generate dictates the wing motion Schematic Design of the I-V MeasurementSet-UpSignal generator (LabView) with an amplifierresistorScopea voltage applied to the samplea voltage across a small resistorIPMC sampleElectrode

17. Electrode Placement & Operation
This is one of the more critical & challenging tasks to be undertaken
•For efficient flight the wing must be capable of flapping, twisting & changing the camber of the airfoil along the wing length
•This is accomplished by the placement and operation of the electrodes that control the IPMC material motion
Devise an electrode layout &
associated voltage strength /
polarity to achieve a given
wing shape
Analysis of the electric
field generation for various
electrode placements
Devise a means of controlling
electrode voltage from a single
processor & power source
Produce a wing test section
to validate the electric field
analysis & control scheme

18. 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.

19. Venus Environment
•Rotation Period (Day) of Venus is Longer the
Revolution Period (Year) Potentially Enabling
Continuous FlightEarth Surface Density•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

20. 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

21. 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

22. 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

23. Wing Aerodynamics
Like all flapping wing flyers in nature the solid state aircraftwill 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= ρVc μ 0.00E+001.00E+052.00E+053.00E+054.00E+055.00E+056.00E+051.2251.0070.660.4140.0890.0190.004Atmospheric Density (kg/m^3) 2.12 kg/m^2 Wing Loading2.62 kg/m^2 Wing Loading3.12 kg/m^2 Wing Loading04.54102040Earth Altitude (km) 5256617085Venus Altitude (km) 014Mars Altitude (km) Wing Design: Initial assumptions for the Phase I analysis were for a curved flat plate airfoil rectangular wing planformUnder Phase II a specialized wing design will done which should significantly increase performance.

24. Effect of Re# on Aircraft Performance00.20.40.60.811.21.4024681012Angle of Attack (α, degrees) Re# 420,000Re# 128,000Re# 84,000Re# 42,00000.010.020.030.040.050.060.070.080.090.100.20.40.60.811.21.4Lift Coefficient (Cl) Re# 105,000Re# 84,000Re# 42,000
Flight Reynolds number based on chord length for a flat plate airfoil

25. SD8000-PT -0.200.20.40.60.811.21.4-6-4-20246810121416Angle of Attack (alpha) SD8000-PT -0.200.20.40.60.811.21.400.010.020.030.040.050.06CdAirfoil Selection: Initial airfoil selected is the SD8000-PTAttributes:Good low Reynolds number characteristicsThin cross section minimizes massPerformance data is available at low Reynolds numbersOngoing Aerodynamic Analysis & DesignTo Scale1: 5 Ratio (X to Y) Additional airfoils will be examined as alternatives2-D performance estimates will be generated under specific flight conditions using airfoil analysis tools (MSES, XFOIL) and CFD

26. CFD Airfoil Analysis
CFD analysis is ongoing to provide performance data
on the wing at various angles of attack and Reynolds
numbers representative of the estimated SSA flight
regime and wing motion. This data will be used in an
iterative process to aid in the wing design and establishing
the desired wing motion

27. Nature Inspired Configuration
Nature Has A Way Of Finding The Optimum
•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 shape will be used as the model.
•Initial Chord distribution is based on the Pteranodon wing shape.
•Chord distribution will be optimized through CFD analysis for the design flight conditions of the SSA.

28. Initial Wing Layout Model
Cross Section at Wing RootLeading Edge ViewUpper Surface View

29. Determination of Optimum Wing Motion
•There are two main wing motions that must be defined.
Flap Motion (bending)
Twist Motion
•CFD analysis is being used to examine the wing motion and determine flap motion that provides optimal performanceTwist AngleTwist
BendingBending Angle

30. Aerodynamic & CFD Analysis Goals
•The aerodynamic analysis is to provide a performance estimate (lift & drag) of the SSA under various operational conditions.
•Through an iterative process this analysis will be used to determine the optimum wing geometry and motion.
•Ultimately the CFD analysis is to provide a full 3D flow field over the SSA using the optimized wing geometry and motion for each potential operating location.
•This will validate the estimated SSA performance & provide a tool for further detailed evaluation of the concept.

31. Structural Analysis & Design•The geometry and motion of the wing (established by the aerodynamic analysis) is used as the basis for the structural analysis. •The structural analysis will determine the forces and stresses on the wing. •From these the necessary IPMC material thickness will be determined. •The thickness will vary from root to tip. •The structural analysis is an iterative process with the aerodynamic analysis Stresses Due to Motion
Wing
Motion &
Geometry
Structural
Sizing &
Loading
Lift
Generation

32. Environmental Considerations
on the Structure
•The SSA must be capable of withstanding the environmental conditions at the various proposed operational locations.
•Resistance to corrosive atmospheric compounds (ex, sulfuric acidon Venus).
•Resistance to water evaporation from IPOC within arid environments.
•Low temperature operation.
•Erosive effects of dust (especially for Mars operation).
The selection and evaluation of coatings to resist these environmental effects is being examined.

33. Internal Structure•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

34. Sizing Analysis
During Phase I an analysis was performed to determine the feasibility
of the SSA concept and establish the range of operation on the planets
of interest. This analysis will be refined and expanded in PhaseIIPower ProductionPower ConsumptionIncoming solar fluxTime of year LatitudeMotion of the wing DragLift Generation Flapping frequencyMaximum flap angle traversed Length of the wingInduced drag Profile dragAltitude Flight SpeedWing Geometry

35. Power Production
The amount of power available to the aircraft is based on the
Environmental conditions it is flying withinOutput power will vary based on the Latitude of flight (φ) Time of year (δ) Time of day (θ) Available power also dependson theAtmosphere attenuation (τ) Solar cell efficiency (η)

36. 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) 0204060801001201401601800.050.0100.0150.0200.0Time (Earth Days) Available Power (W/m2) 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)

37. 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° N01020304050607080901000.005.0010.0015.0020.00Time of Day (Hours) Power Available (W/m2) Vernal Equinox (March 21st) Summer Solstice (June 22) Autumnal Equinox (September 23rd) Winter Solstice (December 22nd) Latitude 80° N01020304050607080901000.005.0010.0015.0020.00Time of Day (Hours) Power Available (W/m2) 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°

38. Power Production: Mars
Available Power Throughout the Day (Hours) 0° Latitude010203040506005101520Time (hours) Available Power (W/m2) Vernal Equinox (Day 1) Summer Solsitce (Day 167) Autimal Equinox (Day 333) Winter Solstice (Day 500) 80° Latitude051015202505101520Time (hours) Available Power (W/m2) 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°

39. 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.

40. Wing Force Due to Motion
Force versus Distance Traveled for Various Wing Lengths, Maximum Flap Angles and Flap Frequencies
The power requiredto move the wing is the area under the force vs distance 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. 0.000010.00010.0010.010.111000.511.522.533.5Distance Travled (m) Force (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

41. Drag Due to Lift and Velocity11010010001.2251.0070.660.4140.0890.0190.004Atmospheric Density (kg/m^3) 2.12 kg/m^2 Wing Loading2.62 kg/m^2 Wing Loading3.12 kg/m^2 Wing Loading04.54102040Earth Altitude (km) 5256617085Venus Altitude (km) 014Mars Altitude (km) 11.522.533.540.030.330.630.931.231.531.832.132.432.73Length Along Wing (m) Lift Force (N) Cycle 1s" Cycle 4s" Cycle 2sCycle 3sDf= 12 ρVi2(cf2S+cdS) 0iΣ Required Speed for FlightLift Generated
The drag is dependent on the flight velocity which in turn sets the lifting capacity and the lift to drag characteristics of the airfoil.
The aerodynamic drag is due to the
generation of lift and the movement of the aircraft through the atmosphere
Lift generation depends on flight speed, flapping frequency & flap angle

42. Analysis Method
The energy consumed during the flap has to equal the energy collected
during the total flap and glide cycle. Power Available Lift GeneratedPower Consumed(Motion & Drag) Inputs Flapping Frequency Flap AngleAircraft SizeAltitudeFlight 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.

43. 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

44. Aircraft Sizing
The initial feasibility study was perfo
rmed 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/m2
Payload Specific Mass
2.00 kg/m2
IPMC Specific Mass
0.75 kg/m2
Battery Specific Mass
0.12 kg/m2
Solar Cell Specific Mass
10%
Solar Cell Efficiency
45°
Maximum Flap Angle
0.008
Wing Friction Coefficient
8
Aspect Ratio
3.12 kg/m2
Wing Loading
1 km to 7 km
Mars
1 km to 35 km
Earth
53 km to 82 km
Venus

45. Sizing Results0204060801001201401600123456Flap Duration (s) 1 km Earth, 53 km Venus, O km Mars 3 km Earth, 55 km Venus, 0 km Mars5 km Earth, 57 km Venus, 0 km Mars10 km Earth, 61 km Venus, 0 km Mars15 km Earth, 65 km Venus, 0 km Mars20 km Earth, 70 km Venus, 0 km Mars3 m Wingspan SSARequired power in W/m2of wing area were generated for a range of flap durations over a number of altitude levels3 m Wingspan00.511.522.533.544.550123456Flap Duration (s) Glide Duration (s) 1 km Earth, 53 km Venus, O km Mars 3 km Earth, 55 km Venus, 0 km Mars5 km Earth, 57 km Venus, 0 km Mars10 km Earth, 61 km Venus, 0 km Mars15 km Earth, 65 km Venus, 0 km Mars20 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.

46. 100 m Wingspan0204060801001201401600510152025303540Flap Duration (s) Specific Power Required (W/m^2) 25 km Earth, 74 km Venus, 0 km Mars30 km Earth, 78 km Venus, 1 km Mars32 km Earth, 79 km Venus, 1.4 km Mars35 km Earth, 82 km Venus, 5 km Mars100 m Wingspan SSASizing Results100 m Wingspan051015202530350510152025303540Flap Duration (s) Glide Duration (s) 25 km Earth, 74 km Venus, 0 km Mars30 km Earth, 78 km Venus, 1 km Mars32 km Earth, 79 km Venus, 1.4 km Mars35 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

47. 0204060801001201400510152025303540Earth Altitude (km) Glide Duration (s) 3 m Wingspan6 m Wingspan12 m Wingspan20 m Wingspan50 m Wingspan100 m Wingspan250 m WingspanGlide duration goes to zero with increasing altitude
Sizing Results Optimal Operation
Flap 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 altitude0204060801001201401601802000510152025303540Earth Altitude (km) Required Power (W/m^2) 3 m Wingspan6 m Wingspan12 m Wingspan20 m Wingspan50 m Wingspan100 m Wingspan250 m Wingspan0510152025303540450510152025303540Earth Altitude (km) Flap Duration (s) 3 m Wingspan6 m Wingspan12 m Wingspan20 m Wingspan50 m Wingspan100 m Wingspan250 m WingspanFlap duration remains constant until the glide Duration reaches zero then begins to decrease
Required power increases exponentially with
increasing altitude

48. Sizing Analysis Summary0204060801001201401601802000510152025303540Earth Altitude (km) 3 m Wingspan6 m Wingspan12 m Wingspan20 m Wingspan50 m Wingspan100 m Wingspan250 m Wingspan525761667074788285012Mars Altitude (km) Venus Altitude (km) Venus Maximum Available Power Earth Maximum Avaailable PowerMars Maximum Available Power

49. Continuation of Performance Analysis
•The performance analysis, based on a basic energy balance between energy availability from the solar array & energy consumed by the vehicle due to wing motion & systems operation, conducted under Phase Iis being expanded.
•Changes in the wing design produced through the aerodynamic analysis will be incorporated into the analysis.
•Variation in array output due to wing motion & temperature effects are being considered.
•The power required to fly under various wind conditions is beingincorporated into the analysis. This ability to overcome the wind speed will enable the SSA to station keep over a specific location.
•Continuous day / night operation is being evaluated. The abilityto store energy in altitude and glide throughout the nighttime period is being examined.

50. Communications System
•Getting data & information to and from the SSA is a critical function.
•The unique configuration of the SSA & its thin light flexible structure pose a unique challenge for the communications system.
Power
Environment
Payload & Mission
UWB
Fixed Band (ka, radio etc.)
Data Requirement
SSA Layout
Mass
Trailing
Integrated
Coating
Antenna Placement
SSA Layout
Flexibility Requirement Data Transfer Capability
Wire
Flexible
Thin Film
Planar
Antenna Type
Design & Selection Factors
Options
Issue

51. Communications to satellite with transparent metal antennaCommunications or data collection through sounding to surface with thin film antenna

52. Thin Film Antenna
•The thin film antenna holds the most promise for use on the SSA
•It is thin lightweight and flexible which fits well with the SSArequirements
•The antenna could be mounted over the complete bottom side of the SSA. This would provide a large receiving and transmitting area.
•Bottom side mounting would enable transmission to & from the surface for both communications & science data collection.
•There is the potential for making this type of antenna transparent which would all it to be mounted on top of the solar array for data transfer to an orbiting spacecraft.
•Ultimately, if the thin film antenna can meet the SSA mission needs, there is the potential of having the antenna constructed as another material layer on the SSA which would provide a truly integrated composite vehicle.

53. Ongoing Communications System Design & Analysis
A design and analysis of the communications system is underway. The goal is to have the communications system fit seamlessly into the SSA design and potentially double as a science data collection tool. This effort includes:
•Performing a system (transmitter, receiver & antenna) design forvarious types of potential communications / data collection systems.
•Determining the mass, range, power requirements and reliability of each identified system.
•Evaluate operation under the three potential operational environments (Venus, Earth & Mars)
•Evaluate the potential of using the system for science data generation.
•This includes determining what frequencies are required to collect atmospheric and surface composition and structure information.
•The ability to generate these with the communications system andor what modifications would be necessary for this to occur.
•The determination of mass, power requirements and data resolution capabilities for science data return.

54. Mission Analysis
•Utilizing the SSA capabilities established from the performance & communications analysis & vehicle design work, an evaluation of the potential missions the SSA may be capable of carrying out will be made.
•Critical factors will be flight duration, flight altitude, payload capacity, payload power availability & data transfer capability.
•From these various types of science data gathering & observing equipment will be assessed.
•The mission assessments will be performed for all three potential planets of operation.
•Both individual & multiple aircraft operations will be considered.
•For multiple aircraft missions, a fleet of SSAs can be considered which could either perform a unique task or provide an infrastructure such as a communications network.
•Specific aspects of the missions are also being considered theseinclude:
•Stowage while in space transit (for Venus and Mars missions)
•Deployment
•Navigation
•Altitude Control
•Landing

55. Science & Payload
Science & other equipment will be evaluated for use on the SSA. This evaluation will take into consideration the instruments mass, volume, powerrequirements and operational concerns such as vibration as well as the potential for miniaturization.
Potential science & data gathering includes:
•Camera, high resolution and context
•Atmospheric measurements: temperature, pressure etc.
•Magnetic field measurements.
•Communications relay transmitter / receiver
•Atmospheric sounding with various frequencies (dependent on the capabilities of the communications system)
•Beacon
In addition to specific payloads, the systems on board the SSA will be evaluated to determine if any dual use potential exists. For example the communications antenna may be used to send receive signals at various frequencies in order to collect data in a sounding fashion.

56. Science: High Resolution Imagery
•Detailed images can be taken on a regional scale at high resolution
•Vertical structures (canyon, mountain) can be imaged at various angles
•Imagery can be used for surveillance, mapping or geological characterization of the planet.

57. Science: Atmospheric Sampling & Analysis
•Examine the Atmosphere Both Vertically & Horizontally (Temperature & Pressure)
•Sample Atmospheric Trace Gases
–Determine Concentrations of Trace Gases and Reactive Oxidizing Species
–Examine the Correlation with the Presence of Active Oxidizing Agents and Absence of Organics in Martina Soil
•Investigation of Dust within the Atmosphere Sample Long Lived Airborne Dust in the Atmosphere (Size, Distribution, Electrostatic Charging etc.)

58. Science: Magnetic Field Mapping
•Magnetic field mapping over a region at high resolution.
–Resolution from orbit is too low and coverage from a rover or lander is too limited
•Mars has a very unique magnetic field distribution characterized by regions of very strong magnetic fields and regions of no magnetic fields
•Mapping of the magnetic fields can give insightinto the tectonic history of the planet & investigate its geology and geophysics

59. Deployable Observation / Communications NetworkCommunications Between AircraftCommunications / Data Transfer to SatelliteCommunications / Observation with Surface