Solidstateaircraft

Solid State Aircraft Phase II Project

NIAC 5th Annual Meeting November 5-6, 2003 Georgia Centers for Advanced Telecommunications Technology Atlanta, Georgia Anthony Colozza Northland Scientific / Ohio Aerospace Institute Cleveland, Ohio

Solid State Aircraft Project Team Members • Mr. Anthony Colozza (PI): • Mr. Phillip Jenkins: • Dr. Mohsen Shahinpoor • Mr. Teryn Dalbello: • Mr. Curtis Smith: • Dr. Carol Kory • Dr. Iakkattukuzhy Isaac

NSI/OAI OAI University of New Mexico University of Toledo/ICOMP OAI Consultant University of Missouri-Rolla

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 Thin Film Array

Thin Film Battery Cathode Grid IPMC Material Anode Grid

•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.

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.

PV Array

•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.

Battery Anode IPMC Cathode

Each Grid Location is Individually Controllable

Thin Film Photovoltaic Array Light Weight: Active material is on the order of 1 to 2 microns thick 16

Highly Flexible: Ideal for the flexing and motion of a flapping wing

CdS/Cu2S

14

CuInSe2 CuGaSe2 CuInS2

12

CdTe a-Si CASCADES

10

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

8

6

4

2

0

Specific Power: 1 kW/kg near term, 2 kW/kg projected

1976

1978

1980

1982

1984

1986

1988

1990

1992

1994

1996

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

The battery is produced by depositing (through sputtering or evaporation techniques) the various material layers that make up the components (cathode, electrolyte, anode and current collector) onto a substrate.

Oak Ridge Battery Design Cross Section (15 micron thick)

Ionic Polymer-Metal Composite (IPMC) •

•

QuickTime™ and a decompressor are needed to see this picture.

•

•

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.

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 Electrodes

C o n ta c t E le c tro d e

+ P IE M

The material will bend toward the anode side of the electrodes

P t E le c tro d e

-

P t P a r tic le s

OFF

P IE M

+ P IE M

ON

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

+ +

+ +

+ s ide cha in

hydra te d ca tion-Na (H2O)4

+ +

wate r

fixe d a nion

mobile ca tion

IPMC Material Characteristics Young' s Modulu s, E Shear Modulu s, G Poisson's ratio, ν Power density (W/mass) Max force density (Cantilever Mod e) Max displacement/strain Bandwidth (speed) Resolution (force and di splacement control ) Efficiency (electrom echanical) Density

Up to 2 GPa Up to 1 GPa Typ ical: 0.3-0.4 Up to 100 J/kg Up to 40 Kgf/Kg Up to 4% linear strain Up to 1 kHz in cantil ever vib ratory mode for actuations Up to 1 MHz for sensing Displacement accuracy down to 1 micron Force resolut ion down to 1 mg Up to 6 % (frequency depend ent) for actuation Up to 90% for sensing Down to 1.8 g/cm 3

Room Temperature

100°C

Deflection (cm) vs Voltage

100°C

Room Temperature

Power Consumption (W) vs Voltage

Material Integration Goal: To produce a method for combining the solar array, battery and 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.

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.

Long Term Integration Strategy Goal: To provide a scheme to producing a fully integrated composite structure from the three main materials and subsequent components.

•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 electrodes and 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 material compatibilities, adhesion characteristics, material thickness and deposition and lithography techniques, their capabilities and requirements.

Wing Control •The placement of electrodes and the shape of the electric field they generate dictates the wing motion resistor

Electrode IPMC sample

a voltage across a small resistor

Signal generator (LabView) with an amplifier

a voltage applied to the sample

Scope

Schematic Design of the I-V Measurement Set-Up

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

Analysis of the electric field generation for various electrode placements

Devise an electrode layout & associated voltage strength / polarity to achieve a given wing shape

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

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.

Venus Environment •Rotation Period (Day) of Venus is Longer the Revolution Period (Year) Potentially Enabling Continuous Flight

Earth 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

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

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

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

Wing Aerodynamics Like all flapping wing flyers in nature the solid state aircraft will operate within a low Reynolds number flight regime. This is due mainly to its required low wing loading and the potential for high altitude operation, where the air density is low. Ď Vc Re =

Âľ

6.00E+05

2.12 kg/m^2 Wing Loading 2.62 kg/m^2 Wing Loading

Wing Design:

3.12 kg/m^2 Wing Loading

5.00E+05

4.00E+05

Initial assumptions for the Phase I analysis were for a curved flat plate airfoil rectangular wing planform

3.00E+05

2.00E+05

1.00E+05

0.00E+00 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

Mars Altitude (km)

0

4.5

14

Under Phase II a specialized wing design will done which should significantly increase performance.

Effect of Re# on Aircraft Performance 1.4

Flight Reynolds number based on chord length for a flat plate airfoil

1.2

1

Re# 420,000

0.8

Re# 128,000 Re# 84,000 Re# 42,000

0.6

0.4

0.2

0.1 0 0

2

4

6

Angle of Attack (

8

10 0.09

12

Îą, 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

Lift Coefficient (Cl)

1

1.2

1.4

Ongoing Aerodynamic Analysis & Design Airfoil Selection: Attributes:

Initial airfoil selected is the SD8000-PT Good low Reynolds number characteristics Thin cross section minimizes mass Performance data is available at low Reynolds numbers

Additional airfoils will be examined as alternatives 2-D performance estimates will be generated under specific flight conditions using airfoil analysis tools (MSES, XFOIL) and CFD

To Scale

1: 5 Ratio (X to Y)

1.4

1.4

SD8000-PT

SD8000-PT 1.2

1.2

1

1

0.8

0.8

0.6

0.6

0.4

0.4

0.2

0.2

0

0 -6

-4

-2

0

2

4

6

8

10

12

14

0

16

0.01

0.02

0.03

-0.2

-0.2 Angle of Attack (alpha)

Cd

0.04

0.05

0.06

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

Nature Inspired Configuration

•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.

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 Wing Layout Model Cross Section at Wing Root

Leading Edge View

Upper Surface View

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 performance Bending

Twist

Bending Angle Twist Angle

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.

Structural Analysis & Design Stresses Due to Motion

Lift Generation

•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

Wing Motion & Geometry

Structural Sizing & Loading

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 acid on 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.

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

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 Phase II Power Production

Power Consumption

Incoming solar flux Time of year Latitude

Motion of the wing Drag Lift Generation

Flapping frequency Maximum flap angle traversed Length of the wing

Induced drag Profile drag

Altitude Flight Speed Wing Geometry

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

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 20° Latitude

140

30° Latitude 40° Latitude

Available Power (W/m2)

120

50° Latitude 60° Latitude 70° Latitude

100

80° Latitude 90° Latitude

80

60

40

20

0 0.0

50.0

100.0

150.0

• 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°

200.0

Time (Earth Days)

Available Power Throughout A Day (Earth Days)

Power Production: Earth Available Power Throughout the Day (Hours) 100

Latitude 0° N

•Solar cell efficiency 10% • Solar cell fill factor 80% • Horizontal solar array • Atmospheric Attenuation 15%

90 Vernal Equinox (March 21st) Summer Solstice (June 22)

80

Autumnal Equinox (September 23rd) Winter Solstice (December 22nd)

60 50 40 30 20

100

10

90

Latitude 80° N Vernal Equinox (March 21st)

0

Summer Solstice (June 22)

80

0.00

5.00

10.00

15.00

Time of Day (Hours)

• Mean Solar Intensity above atmosphere 1353 W/m2 • Longitude 0° • Maximum declination angle 23.5°

Autumnal Equinox (September 23rd)

20.00

Winter Solstice (December 22nd) 70

Power Available (W/m2)

Power Available (W/m2)

70

60

50

40

30

20

10

0 0.00

5.00

10.00 Time of Day (Hours)

15.00

20.00

Power Production: Mars Available Power Throughout the Day (Hours) 60

0° Latitude

•Solar cell efficiency 10% • Solar cell fill factor 80% • Horizontal solar array • Atmospheric Attenuation 15%

Vernal Equinox (Day 1)

50

Summer Solsitce (Day 167) Autimal Equinox (Day 333) Winter Solstice (Day 500)

30

20

25

10

80° Latitude Vernal Equinox (Day 1)

0

20

0

5

10

15

Summer Solsitce (Day 167)

20

Autimal Equinox (Day 333) Winter Solstice (Day 500)

Time (hours)

•Mean Solar Intensity above atmosphere 590 W/m2 • Longitude 0° • Maximum declination angle 24°

Available Power (W/m2)

Available Power (W/m2)

40

15

10

5

0 0

5

10

15

Time (hours)

20

Power Consumption due to Motion Motion of the wing consists of a flapping rate and maximum angle traversed during the flap.

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.

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 vs distance traveled curve.

10

1 0

0.5

1

1.5

2

2.5

0.1 Length 1m, Angle 40°, Cycle 1s Force (N)

Length 1m, Angle 40°, Cycle 4s

3

3.5

The distance traveled varies along the wing length to a maximum at the tip.

Length 1m, Angle 60°, Cycle 1s 0.01

Length 1m, Angle 60°, Cycle 4s Length 3m, Angle 40°, Cycle 1s Length 3m, Angle 40°, Cycle 4s

0.001

Length 3m, Angle 60°, Cycle 1s Length 3m, Angle 60°, Cycle 4s

0.0001

0.00001 Distance Travled (m)

Power consumption can be reduced by tapering the wing so there is less mass at the tip. Thereby reducing the force needed for motion.

Drag Due to Lift and Velocity 1000

Required Speed for Flight

2.12 kg/m^2 Wing Loading 2.62 kg/m^2 Wing Loading 3.12 kg/m^2 Wing Loading

The drag is dependent on the flight velocity which in turn sets the lifting capacity and the lift to drag characteristics of the airfoil.

100

10

i

1 2 D f = ∑ Ď Vi (c f 2S + cd S) 2 0

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

Lift Generated 3.5

Cycle 1s" Cycle 4s" Cycle 2s

Lift Force (N)

The aerodynamic drag is due to the generation of lift and the movement of the aircraft through the atmosphere

Cycle 3s

3

2.5

2

Lift generation depends on flight speed, flapping frequency & flap angle

1.5

1 0.03

0.33

0.63

0.93

1.23

1.53

Length Along Wing (m)

1.83

2.13

2.43

2.73

Analysis Method The energy consumed during the flap has to equal the energy collected during the total flap and glide cycle.

Inputs Flapping Frequency Flap Angle Aircraft Size Altitude

Power Available

Lift Generated

Power Consumed (Motion & Drag)

Flight Speed Flap 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 determine the optimal glide to flap ratio for a give aircraft configuration under a specific flight condition

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 following assumptions Wing Loading

3.12 kg/m2

Aspect Ratio

8

Wing Friction Coefficient

0.008

Maximum Flap Angle

45째

Solar Cell Efficiency

10%

Solar Cell Specific Mass

0.12 kg/m2

Battery Specific Mass

0.75 kg/m2

IPMC Specific Mass

2.00 kg/m2

Payload Specific Mass

0.25 kg/m2

Potential flight altitude ranges investigated in the analysis for each of the planets of interest Venus

53 km to 82 km

Earth

1 km to 35 km

Mars

1 km to 7 km

Sizing Results

160 1 km Earth, 53 km Venus, O km Mars 140

3 m Wingspan SSA

3 km Earth, 55 km Venus, 0 km Mars 5 km Earth, 57 km Venus, 0 km Mars

120

10 km Earth, 61 km Venus, 0 km Mars

Required power in W/m2 of wing area were generated for a range of flap durations over a number of altitude levels

15 km Earth, 65 km Venus, 0 km Mars

100

20 km Earth, 70 km Venus, 0 km Mars 80 60 40 20 0 0

1

2

3

4

5

Flap Duration (s)

5

6

1 km Earth, 53 km Venus, O km Mars

4.5

3 km Earth, 55 km Venus, 0 km Mars 4

10 km Earth, 61 km Venus, 0 km Mars

3.5

15 km Earth, 65 km Venus, 0 km Mars Glide Duration (s)

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.

5 km Earth, 57 km Venus, 0 km Mars

3

20 km Earth, 70 km Venus, 0 km Mars

2.5

3 m Wingspan

2 1.5

The minimum specific power occurs at smaller flap durations (higher flapping frequencies) as the flight altitude is increased.

1 0.5 0 0

1

2

3 Flap Duration (s)

4

5

6

Sizing Results

160

25 km Earth, 74 km Venus, 0 km Mars 30 km Earth, 78 km Venus, 1 km Mars

140

100 m Wingspan SSA

32 km Earth, 79 km Venus, 1.4 km Mars

Specific Power Required (W/m^2)

120

35 km Earth, 82 km Venus, 5 km Mars

100

As the altitude increases the vehicle must flap more frequently or continuously to maintain flight.

80

60

40

20

100 m Wingspan 0 0

5

10

15

20

25

30

35 35

40

25 km Earth, 74 km Venus, 0 km Mars

Flap Duration (s)

30 km Earth, 78 km Venus, 1 km Mars

30

As the flap duration increases (lower frequency) the glide duration approaches zero

35 km Earth, 82 km Venus, 5 km Mars

25

Glide Duration (s)

For all vehicle sizes the glide duration decreases with increasing altitude

32 km Earth, 79 km Venus, 1.4 km Mars

20

15

100 m Wingspan 10

5

0 0

5

10

15

20 Flap Duration (s)

25

30

35

40

Sizing Results Optimal Operation

140

3 m Wingspan

120

Flap duration, Glide Duration and Required Power as a Function of Altitude (Earth) for Various Size Aircraft

6 m Wingspan 12 m Wingspan 20 m Wingspan

Glide Duration (s)

100

50 m Wingspan 100 m Wingspan 250 m Wingspan

80

60

40

20

0 0

5

10

15

20

25

30

35

40

Earth Altitude (km)

Glide duration goes to zero with increasing altitude

Graphs shown are based on the flap duration that produced a minimum required power for a given size aircraft and flight altitude 200

45

180

40 3 m Wingspan 35

6 m Wingspan

12 m Wingspan

Required Power (W/m^2)

50 m Wingspan 100 m Wingspan 250 m Wingspan

25

12 m Wingspan

140

20 m Wingspan

30

Flap Duration (s)

3 m Wingspan

160

6 m Wingspan

20

15

20 m Wingspan 50 m Wingspan 100 m Wingspan

120

250 m Wingspan

100

80

60

10

40

5

20

0

0 0

5

10

15

20

25

30

35

Earth Altitude (km)

Flap duration remains constant until the glide Duration reaches zero then begins to decrease

40

0

5

10

15

20

25

30

35

40

Earth Altitude (km)

Required power increases exponentially with increasing altitude

Sizing Analysis Summary 200

3 m Wingspan

6 m Wingspan

12 m Wingspan

20 m Wingspan

50 m Wingspan

100 m Wingspan

250 m Wingspan

180

Venus Maximum Available Power

160 140 120

Earth Maximum Avaailable Power

100 80 60

Mars Maximum Available Power

40 20 0 0 52

5 57

10 61

15

20

25

Earth Altitude (km) 66 70 74 Venus Altitude (km) Mars Altitude (km)

30

35

40

78

82

85

0

1

2

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 I is 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 being incorporated 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 ability to store energy in altitude and glide throughout the nighttime period is being examined.

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.

Issue

Options

Design & Selection Factors

Antenna Type

Wire Flexible Thin Film Planar

SSA Layout Flexibility Requirement Data Transfer Capability

Antenna Placement

Trailing Integrated Coating

SSA Layout Mass

Data Requirement

UWB Fixed Band (ka, radio etc.)

Power Environment Payload & Mission

Communications to satellite with transparent metal antenna

Communications or data collection through sounding to surface with thin film antenna

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 SSA requirements •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.

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 for various 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 and or what modifications would be necessary for this to occur.

•The determination of mass, power requirements and data resolution capabilities for science data return.

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 these include: •Stowage while in space transit (for Venus and Mars missions) •Deployment •Navigation •Altitude Control •Landing

Science & Payload Science & other equipment will be evaluated for use on the SSA. This evaluation will take into consideration the instruments mass, volume, power requirements 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.

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.

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.)

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 insight into the tectonic history of the planet & investigate its geology and geophysics

Deployable Observation / Communications Network Communications / Data Transfer to Satellite

Communications Between Aircraft

Communications / Observation with Surface