45.1

VERTICAL LIFT PLANETARY AERIAL VEHICLES: THREEPLANETARY BODIES AND FOUR CONCEPTUAL DESIGN CASES

Larry A. YoungEdwin W. Aiken

Army/NASA Rotorcraft DivisionAmes Research Center

Moffett Field, CA

Abstract

NASA Ames Research Center has been studying the feasibility of vertical lift aerial vehicles to support planetaryscience and exploration missions. Besides Earth, it appears that there are three planetary bodies within our solarsystem where vertical flight might not only be theoretically feasible, but would also have unique missioncapabilities that no other platform (ground-based, aerial, or orbital) could provide. Several vertical lift vehicleconfigurations might be applicable for planetary science missions. This paper presents a few representativeconceptual design cases and the design challenges inherent in their development. Finally, more detailedcomments are directed to the issues inherent in developing a NASA ‘Mars Scout’ mission employing the use of aMartian autonomous rotorcraft.

Introduction

Humankind’s understanding of the universe hasundergone tremendous advances over the lastfew decades. Robotic missions to planetarybodies within the solar system have beenparticularly instrumental in achieving thisunderstanding. But, planetary science is at acrossroads. A new generation of roboticexplorers – with substantial improvements inautonomy, mobility, power/energy availability,and instrumentation sophistication - is requiredto make further advances. Successfuldevelopment of a new generation of roboticexplorers, including all of the attendanttechnologies for their operation, will also aid inthe ultimate transition from robotic to humanexploration of the solar system.

Recent research has focused on the feasibility ofdeveloping vertical lift aerial vehicles that couldaid in the exploration of various planetarybodies in our solar system. Specifically, theutility of vertical lift aerial vehicles to supportmissions to Mars, Titan (a moon of Saturn), andVenus is being studied. Recent advances inautonomous system technology,microelectronics, ultra-lightweight structuralmaterials, innovative power systems, and low-Reynolds number, compressible flowaerodynamics have been instrumental in

Presented at 27th European Rotorcraft Forum,Moscow, Russia, September 11-14, 2001.Copyright 2001 by KAMOV Company.

establishing the conceptual viability of verticallift planetary aerial vehicles.

Table 1 summarizes a few of the importantgeophysical and atmospheric properties ofMars, Titan, and Venus. The correspondingproperties for Earth are also provided as areference.

Table 1–Planetary Description (Ref. 1)

MeanRadius(km)

Gravity(m/s2)

MeanSurfaceAtmos.Temp.(o K)

MeanSurfaceAtmos.Pressure

(Pa)

MeanSurfaceAtmos.Density(kg/m3)

Atmos.Gases

Earth 6371 9.82 288.2 101,300 1.23 N2 78%O2 21%

Mars 3390 3.71 214 636 1.55x10-2CO2

95%N2 2.7%Ar 1.6%O2 0.1%

Titan 2575 1.354 94 149,526 5.55N2 65-

98%Ar<25%CH4 2-

10%

Venus 6052 8.87 735.3 9.21x106 64.79CO2

96%N2 3.5%

Achieving vertical flight for Mars, Titan, andVenus will not be easy to accomplish.Nonetheless, preliminary work to date has been

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4. TITLE AND SUBTITLE Vertical Lift Planetary Aerial Vehicles: Three Planetary Bodies and FourConceptual Design Cases

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7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) Army/NASA Rotorcraft Division,Army Aviation and MissileCommand,Aeroflightdynamics Directorate (AMRDEC), Ames ResearchCenter,Moffett Field,CA,94035

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14. ABSTRACT NASA Ames Research Center has been studying the feasibility of vertical lift aerial vehicles to supportplanetary science and exploration missions. Besides Earth, it appears that there are three planetary bodieswithin our solar system where vertical flight might not only be theoretically feasible, but would also haveunique mission capabilities that no other platform (ground-based, aerial, or orbital) could provide. Severalvertical lift vehicle configurations might be applicable for planetary science missions. This paper presents afew representative conceptual design cases and the design challenges inherent in their development.Finally, more detailed comments are directed to the issues inherent in developing a NASA ‘Mars Scout’mission employing the use of a Martian autonomous rotorcraft.

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45. 2

promising (Refs. 2-12). Development of suchvehicles will be a tremendous engineeringundertaking – both in terms of technical risk andscientific payoff.

Yet, despite the technical challenges for verticallift planetary aerial vehicles, the most difficulthurdle for the development of these vehicles islikely to be both perceptual and programmaticin nature. Applying rotary-wing technologies toplanetary science applications will require thedevelopment of cross-cutting technologies thatmust bridge the interests of a diverse group ofresearch and technical communities --“Strategic Enterprises” -- within NASA. It willbe difficult, but not impossible, to bridge thesedisparate interests/requirements enabling thesuccessful launch of vertical lift planetary aerialvehicles. Initiating such an engineering programwould enable a wholly new technical discipline,leading to a truly “revolutionary” new approachto planetary science data gathering.

Four conceptual design cases will now bediscussed that will illustrate the technicalopportunities and design challenges of verticallift planetary aerial vehicles. Finally, a notionalNASA “Mars Scout” mission will be discussedin the context of a Martian autonomousrotorcraft performing aerial survey andsampling flights in conjunction with in-situanalysis science investigations back at alander/primary-base.

Mars Coaxial Helicopter

Mars has been described as the most terrestrialof all the other planetary bodies in the solarsystem. And yet it is clear, from anaeromechanics perspective, that Martianrotorcraft will be very different from theirterrestrial counterparts. Martian autonomousrotorcraft will have very large lifting-surfacesand will be required to have ultra-lightweightconstruction (Fig. 1). Further, Mars rotorcraftwill have a unique combination of low Reynoldsnumber and compressible flow aerodynamics,require new types of propulsion systems, andrequire high levels of vehicle autonomy. Someearly work and discussion on Martian verticallift vehicles can be found in Refs. 2-3 and 8-9.

(Disk Loading = 4 N/m^2)

0

500

1000

1500

2000

2500

3000

3500

5 10 15 20 25 30

Vehicle Mass (kg)

Rot

or S

haft

Pow

er

(Wat

ts)

0

0.5

1

1.5

2

2.5

3

3.5

Rot

or R

adii

(m)

Power (Hover Figure of Merit = 0.4)

Rotor Radius

Fig. 1 – General Sizing Trend for an IsolatedMars Rotor

Recent work at NASA Ames has focused on acoaxial helicopter configuration for early Marsexploration missions (Fig. 2).

Fig. 2 -- A Coaxial Helicopter Configuration forMars Exploration

Such a vertical lift aerial vehicle could aid inNASA’s “search for water” and “hunt for life”astrobiology objectives for Mars. A Marscoaxial helicopter would be more compact –and therefore more easily transportable fromEarth – than other alternate rotorcraftconfigurations.

Fig. 3 shows first-order estimates of theforward-flight performance of a range of Marscoaxial helicopters sized from 10 to 50 kg. Theperformance estimates for these small coaxialhelicopters assumes that the rotor tip Machnumber is held constant at 0.65 and the diskloading is 4 N/m2. A very conservative inducedpower constant and mean blade profile dragcoefficient was used for the rotor performanceestimates in Fig. 3. Performance estimates forthe coaxial helicopter configuration conform tothe methodologies noted in Refs. 13-14. Similaraerodynamics and rotor performancecharacteristics are noted for rotary-wing microair vehicles (Ref. 15). As can be readily seen,the rotor profile drag is a major contributor tothe overall rotor power. There is almost a

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negligible ‘power bucket’ for these vehicles.This performance estimate conservatismaccounts for: 1. the high profile drag typical ofvery low Reynolds number airfoils; 2. the use ofa bluff cross sectional shape for the inboardblade spar (out to 40% blade span); 3. the effectof large blade-root cutouts to allow for rotorblade fold and telescoping (for vehicletransport/deployment); 4. the high inducedlosses for low aspect ratio rotor blades withlarge blade root cutout. To achieve significantimprovements in rotor profile power it will benecessary to use an improved low-drag inboardspar design (using a streamlined spar cross-section and reducing the blade root cutout) andimproved low Reynolds number airfoils. Theabove rotor changes, though, will likely affectthe volume of the stowed vehicle during transitto Mars. It is expected that substantialimprovements in the rotor profile drag andvehicle parasite drag – as compared to thevalues used in these initial performanceestimates - will be achieved as planned detailedcomputational and experimental investigationsare made into Mars rotorcraft configurations.

0

1 0 0 0

2 0 0 0

3 0 0 0

4 0 0 0

5 0 0 0

6 0 0 0

7 0 0 0

8 0 0 0

9 0 0 0

0 1 0 2 0 3 0 4 0F o r w a r d -F lig h t S p e e d ( m /s e c )F o r w a r d -F lig h t S p e e d ( m /s e c )F o r w a r d -F lig h t S p e e d ( m /s e c )F o r w a r d -F lig h t S p e e d ( m /s e c )

Tota

l Sha

ft P

ower

(Wat

ts)

Tota

l Sha

ft P

ower

(Wat

ts)

Tota

l Sha

ft P

ower

(Wat

ts)

Tota

l Sha

ft P

ower

(Wat

ts)

V e h ic le M a s s = 5 kg1 0 kg2 5 kg5 0 kg

Fig. 3 – Mars Coaxial Helicopter PerformanceEstimates

Fig. 4 is an illustrative weight trend plot for thesmall, high-power density, electric motorspotentially suitable for Mars rotorcraftconfigurations in the under 100kg weight class(Refs. 16-17). Given the singular nature ofvertical lift planetary aerial vehicles, derivingweight estimates for key vehicle subsystems is adifficult but crucial design challenge. Weighttrend methodologies (for example, Refs. 18-21)used in conventional rotorcraft preliminarydesign can only provide general insight at best.

y = 0.0004xR2 = 0.8792

0

0.5

1

1.5

2

2.5

3

3.5

0 2000 4000 6000 8000

Motor Power (Watts)

Elec

tric

Mot

or M

ass

(kg)

Fig. 4 – Illustrative (Current Technology) SmallElectric Motor Weight Trend

Fig. 5 shows first-order estimates for vehiclerange as a function of fuel/energy-source weightfraction for a 10 kg vehicle, at a forward flightspeed of 40 m/s. Three families of curves areshown in the figure: range estimates usingbattery technology, estimates for fuel cells, andpropulsion from a hydrazine-based Akkermanengine (Ref. 22). An Akkerman engine is amonopropellent-based propulsion system and,therefore, should operate satisfactorily in thecarbon-dioxide-dominated atmosphere of Mars.It has been successfully used on high-altitude,long endurance terrestrial experimental aircraft.

Many factors must be accounted for in thepropulsion system used for a Mars rotorcraft.Though hydrazine-based Akerrman enginetechnology promises the greatest range benefitsfor such vehicles, having a “clean” non-volatile(and non-toxic) energy source for such vehicleshas much merit. Even among the various fuel-cell technology choices (non-regenerativeversus regenerative systems and different typesof reactants) each will have their relativeadvantages and disadvantages. Environmentalcontamination from fuel-cell by-products (fromnon-regenerative systems which expel/exhaustthe fuel-cell products) can not be allowed tobias the science mission measurements beingmade. (For example, water vapor ‘exhaust’from a hydrogen and oxygen non-regenerativefuel-cell could clearly contaminate the ‘searchfor water’ measurements/results.) Finally,though solar power may seem to promise avirtually inexhaustible energy source for a Marsrotorcraft (where its batteries or fuel-cells arerecharged by a lander’s solar array panels), thisis an overly optimistic viewpoint. In reality, theduration of a mission is just as likely to bedetermined by the amount of operationalresources available on Earth as any other factor.Additionally, there are practical limits as to howlong a solar array can deliver power efficientlyon Mars due to dust adhesion/accumulation on

45. 4

the solar array panels. Trade studies betweenhydrazine propulsion and fuel-cell systems for aMars vertical lift aerial vehicle merit continuedinvestigation.

0

50

100

150

200

250

0 0.1 0.2 0.3 0.4Fuel/Energy-Source Weight Fraction

Forw

ard-

Flig

ht R

ange

, (km

) Hydrazine Akkerman Engine(SFC=1kg/MJ), Breguet Eq.

Lithium Ion Battery (150Wh/kg)

Fuel Cell (300 Wh/kg)

Fig. 5 – Mars Coaxial Helicopter RangeEstimates

The flight dynamics of a Martian rotorcraft willbe quite different compared to its terrestrialcounterparts. The rotors for a Martian rotorcraftwill have very low Lock numbers and will havevery low aerodynamic damping. The rotorblades will also likely have relatively low valuesof torsion and bending stiffness because of theirlarge blade planform area and ultra-lightweightstructure. Yaw control for a Mars coaxialhelicopter configuration will be maintained bydifferential rotor collective (resulting indifferential torque) instead of relying on fixedtail surfaces as is done with most conventionalterrestrial coaxial helicopters.

Mars Tiltrotor

A tiltrotor is a particularly attractiveconfiguration (Fig. 6) for Mars exploration. Atiltrotor represents a good compromise betweenhover performance and cruise range/endurance.

(a)

(b)

Fig. 6 -- A Mars Tiltrotor: (a) helicopter-modein vertical climb over Valles Marineris; (b)

airplane-mode

Fig. 7 presents some initial sizing estimates fora small (10 kg) autonomous Mars tiltrotorconfiguration. Fig. 7 shows the trend of rotorsize as a function of rotor mean lift coefficientand tip Mach number. A notional rotor designpoint of Mtip =0.7 and CL =0.4 is noted on thefigure. As expected the resulting proprotors arequite large. One of the biggest issues for theMars tiltrotor configuration is that thedeployment of a tiltrotor on the surface of Marswill be fairly complicated, and will requireastronaut-assisted assembly or some type ofautonomous assembly process on the landerplatform.

0

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1.5

2

2.5

3

3.5

4

4.5

5

0 0.1 0.2 0.3 0.4 0.5 0.6

Rotor Mean Lift Coefficient

Mtip = 0.50.60.7Design Point

Fig. 7 – Mars Tiltrotor Rotor Size Estimates(Total Vehicle Mass = 10 kg)

Fig. 8 shows wing planform area as a functionof maximum wing lift coefficient and the end-of-conversion Mach number (airspeed at whichthe wing, versus the rotors, carries all thevehicle lift). Three considerations constrain thewing sizing effort: first, there is a maximumadvance ratio to which the rotors can flyedgewise in helicopter-mode (because of highvibratory loads); second, maximum wing liftcoefficient is significantly lower for the low

45. 5

Reynolds number regime typical of flight in theMartian atmosphere; third, there is a minimumwing stiffness required for aeroelastic stability(particularly for ultra-lightweight structures). Itis beyond the scope of this paper to addressthese design considerations in other than aqualitative sense.

0

5

10

15

20

25

0.4 0.6 0.8 1 1.2

Maximum Wing Lift Coefficient

Win

g Pl

anfo

rm A

rea

(m^2

)

End of ConversionMach # = 0.1Mach # = 0.15

Mach # = 0.2

Design Point

Fig. 8 – Mars Tiltrotor Wing Area SizeEstimates (Vehicle Mass = 10 kg)

Fig. 9 shows preliminary range estimates (usingthe Breguet range equation) of the 10 kg Marstiltrotor configuration, assuming propulsion isprovided by an Akkerman hydrazine pistonengine, for various vehicle L/Ds and fuelfractions. The specific fuel consumption (SFC)constant used for the Akkerman hydrazinepiston engine is 1.0 kg/MJ (Ref. 22). A typicalvalue for conventional terrestrial tiltrotoraircraft is L/D ~ 7.

As shown in Fig. 9, a Mars tiltrotor usinghydrazine piston engine propulsion will be amedium-range planetary aerial vehicle. In orderto improve vehicle range, in addition toimproving L/D efficiency of the aircraft, thepropulsion system SFC must be improved.This will necessitate developing alternatepropulsion systems having improved SFC –perhaps those involving propellants generatedby in-situ production techniques.

0

100

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300

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500

600

0.15 0.2 0.25 0.3 0.35 0.4 0.45

Fuel Fraction

Ran

ge (K

m)

L/D = 6 L/D = 8

L/D = 10 L/D = 12

Fig. 9 – Breguet Range Estimates for a MarsTiltrotor

Besides potentially having range and enduranceadvantages over other Mars rotorcraftconfigurations, Mars tiltrotors will also likely beable to operate at higher altitudes (~1km aboveground level). Many geologically interestingsites on Mars may only be accessible with aMars tiltrotor versus a helicopter configuration.

Deserving considerable follow-on analysis isaeroelastic/whirl-flutter stability for tiltrotorvehicles constructed of ultra-lightweight andlow stiffness structures. A considerable amountof creativity may well be required to insuresatisfactory cruise speeds with acceptablestability margins for such radically differentvehicles and structures compared to theirterrestrial counterparts.

More discussion related to Mars rotorcraft, withrespect to the NASA Mars Scout program, canbe found in the Appendix.

Titan Tilt-Nacelle VTOL

Several types of rotorcraft, or alternativelypowered lift vehicles, could be developed foraerial exploration of Titan (Refs. 4-6, 7, 12).Such vehicles will likely have electricpropulsion driving their rotors or fans. Inparticular, ducted fan configurations such as tilt-nacelle aircraft are perhaps well suited for Titan(Fig. 10). Use of electric propulsion inconjunction with a lander-based power sourcewill maximize the number of flights (and,therefore, remote sites that can be visited andsamples and measurements made). Ducted fanaerial vehicles would inherently be more robustduring take-off or landing in an unknown,

45. 6

potentially hazardous, environment as comparedto conventional rotors.

(a)

(b)

Fig. 10 -- A Titan Tilt-Nacelle VTOL: (a) take-off; (b) cruise.

Fig. 11 shows a first-order estimate of hovertotal shaft power for a notional Titan tilt-nacelleVTOL vehicle having two ducted fans that canpivot at the wing tips (similar in configuration tothe Doak VZ-4). A conservative shroud thrustfraction of 0.3 (i.e., 30% of the total thrust isprovided by the duct/nacelle aerodynamics inhover) is used in the hover performanceestimate. The hover performance and fan sizingestimates are for a disk loading of 600 N/m2, afan blade tip Mach number of 0.7, and a fanblade solidity of 0.25. A Titan VTOL’s ductedfans will be very small and consume very littlepower as a result of the high atmosphericdensity and low gravity field for Titan.

Initial mission concepts being studied at NASAAmes would employ a lander-based architecturewhere small ducted fan tilt-nacelle vertical take-off and landing (VTOL) aircraft could use thelander as a primary base site. The lander wouldservice and support (including battery/fuel-cellrecharging) the vertical lift aerial vehicles.This lander/aerial vehicle power source willinevitably be nuclear in nature (RadioisotopeThermoelectric Generators (RTGs) or Stirlingcycle reactors); because of the great distance ofTitan from the sun, and its atmospheric haze,solar power is not an acceptable alternate powersource. RTG units -- in the under 250 Wattclass -- are a proven technology. This size ofRTG units have been demonstrated in previousouter planetary missions.

0

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Vehicle Mass, kg

Tota

l Sha

ft Po

wer

(Bot

h D

ucte

d Fa

ns),

Wat

ts

0

0.01

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0.05

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0.1

Fan

Bla

de R

adiu

s, m

Hover PowerFan Blade Radius

Fig. 11 – Ducted Fan Hover Performance forTitan Vehicle

Fig. 12 shows range estimates for a 50 kg Titantwin tilt-nacelle/ducted-fan VTOL vehicle,assuming power matching between the hoverand cruise design points. The range estimatesare based on the estimated power from Fig. 11,with reasonable drive train and electric motorefficiencies applied. The cruise speed isassumed to be 50 m/s. These range estimatesassume minimum hover/loiter time. The TitanVTOL cruise speed is relatively low to reflectthe higher atmospheric density of Titan.

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0 0.1 0.2 0.3 0.4

Energy Source Weight Fraction

Tota

l Ran

ge, k

m

Battery (150 Wh/kg)Fuel Cell (300 Wh/kg)

Fig. 12 – Range Estimates for a Twin DuctedFan Titan VTOL

By 2004, the joint NASA, ESA, and Italianspace agency Cassini space mission will reachSaturn’s orbit and release the Huygens probe(descending via parachute) into Titan’satmosphere. The Huygens atmospheric probeand the complementary Cassini observationswill provide invaluable insights into theatmospheric chemistry/properties of Titan.

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Fundamental insights into pre-biotic organicchemistry may result from the exploration ofTitan. With the anticipated success of theCassini/Huygens mission there may be anopportunity to take advantage of the excitementunderlying this adventure to advocate possiblefollow-on missions – including those employingTitan VTOL vehicles.

Venus Hybrid Airship

Of the three planetary bodies besides Earthwhere it theoretically is feasible to design andfly vertical lift aerial vehicles, Venus will likelypose the greatest challenge. The extremely highatmospheric densities near that planet’s surface(plus the near-Earth-magnitude of itsgravitational field) would suggest that abuoyant, or semi-buoyant, vehicle mightrepresent the most practical design forexploration of Venus (Fig. 13). The airframe ofa Venusian hybrid-airship would be a rigid hull,capable of sustaining substantial pressuredifferentials across (interior/exterior) the hullsurface.

Venus’ high surface temperatures also posetremendous challenges for aerial vehicle design.Though active and passive technologies existfor thermal management of planetary sciencehardware, extended operation of such hardwarenear Venus’ surface is currently problematicwith today’s technology. This will mean, forexample, that ‘waste heat’ will have to beminimized by keeping the power required forflight to an absolute minimum (thusnecessitating buoyancy fractions greater than75%).

Fig. 13 -- A Notional Venusian Hybrid Airshipwith Twin Hulls and Tandem Tilting Propellers

Fig. 14 shows first-order estimates of a notionalVenus hybrid-airship’s hull size. The resultsshown in this figure assumes a hybrid-airshipbuoyancy fraction of 0.9 and a propulsionenergy-source (batteries, fuel cells, etc.) weight

fraction of 25%. Helium is assumed as thehybrid-airship lifting gas. A thin skin oftitanium alloy is assumed for the hull. Hullskin thickness using titanium alloys ranges from0.5 to 1 mm thick for vehicle mass from 10 to50 kg. A similar analysis for low-altitudeballoons for exploration of Venus’ atmospherehas been previously proposed (Ref. 23).

0

1

2

3

4

5

6

7

8

0 20 40 60

Vehicle Mass, kgVehicle Mass, kgVehicle Mass, kgVehicle Mass, kg

Fig. 14 – Hull(s) Size Estimate

Fig. 15 shows a first-order estimate of the hoverperformance and sizing of a tandem propellercombination (sandwiched between twin airshiphulls) that could be used to take-off and landfrom Venus’ surface. The performance andsizing estimates shown in the figure assume theairship buoyancy fraction of 0.9 (therefore, thetwo propellers have to lift only 10% of vehicleweight in hover), a tip Mach number of 0.1, a200 N/m2 disk loading, and a solidity of 0.4 forthe propellers.

0

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10 20 30 40 50

Vehicle Mass, kg

Prop

elle

rs),

Wat

ts

0

0.02

0.04

0.06

0.08

0.1

0.12

0.14

0.16Pr

opel

ler R

adiu

s, m

Shaft PowerRadius

Fig 15 – Hybrid-Airship Propeller HoverPerformance and Sizing Estimates

In the near-term, aeromechanics work on aerialvehicles for exploration of Venus might benefitfrom collaborative work with naval researchersinvestigating undersea submersible robots.

45. 8

Planetary science missions to Venus, thoughperhaps not as frequent or of as great publicinterest as Mars and outer planet missions,nonetheless will ultimately test the capabilitiesof vertical lift planetary aerial vehicles to gain atrue sense of Earth’s “sister” planet.

Concluding Remarks

Over five hundred years ago, Leonardo DaVinci envisioned flight by means of a verticallift aerial vehicle. This progenitor of thehelicopter inspired generations, including theearly pioneers of the rotorcraft industry andresearch community. Over the last six decadesthe helicopter has found, among otherapplications, great utility in aiding in terrestrialexploration. Preliminary design studies by bothNASA and U.S. universities have establishedthe theoretical feasibility of vertical flight inextraterrestrial atmospheres. Much workremains. Nonetheless, with the new twenty-firstcentury, there lies the opportunity to inspire thewhole of humankind with the full potential ofrotorcraft, by demonstrating vertical flight onother planetary bodies in our solar system.

Acknowledgments

The contributions and encouragement of Dr.Geoffrey Briggs, Director for the NASA AmesCenter for Mars Exploration (CMEX) isgratefully acknowledged as being critical to thepromotion of the vertical lift planetary aerialvehicle concept. Thanks are also extended toMr. George Price (formerly of) and Mr.Christopher Van Buiten of Sikorsky Aircraft,Mr. Rhett Flater and Ms. Kim Smith of the AHSInternational, Drs. Virginia Gulick and RoccoMancinelli of the SETI Institute, and Ms. KellySnook of the NASA Ames Space ProjectsDivision. Finally, the contributions of Mr.Michael Derby, Mr. Jose Navarrete, and Drs.Wayne Johnson and Roger Strawn of theArmy/NASA Rotorcraft Division at AmesResearch Center are also gratefullyacknowledged.

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European Rotorcraft Forum, Rome, Italy,September 14-16, 1999.

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17. Small Electric Motor Weight/Power Data:(http://www.astroflight.com)

18. Stepniewski, W.Z. and Shinn, R.A.,“Soviet Vs. U.S. Helicopter WeightPrediction Methods,” 39th Annual Forumof the American Helicopter Society, St.Louis, MO, May 9-11, 1983.

19. Stepniewski, W.Z., “Some Weight Aspectsof Soviet Helicopters,” 40th Annual Forumof the American Helicopter Society,Arlington, VA, May 16-18, 1984.

20. Vega, E., “Advanced Technology Impactson Rotorcraft Weight,” 40th Annual Forumof the American Helicopter Society,Arlington, VA, May 16-18, 1984.

21. Smith, H.G., “Helicopter Structural WeightPrediction and Evaluation – Theory VersusStatistics,” 26th Annual Forum of theAmerican Helicopter Society, Washington,DC, June 16-18, 1970.

22. Akkerman, J.W. “HydrazineMonopropellent Reciprocating EngineDevelopment” NASA ConferencePublication 2081, 13th AerospaceMechanisms Conference, Proceedings of aSymposium held at Johnson Space Center,Houston, TX, April 26-27, 1979.

23. Nishimura, J., et al, “Venus Balloons atLow Altitudes,” Advances in SpaceResearch, Vol. 14, No. 2, Great Britain,1994.

24. NASA Mars Scout Program and MarsExploration Program/Payload AnalysisGroup (MEPAG) white paper:

(http://spacescience.nasa.gov/an/marsscoutsworkshop/mepag.pdf)

25. Healey, A. "Mars Explorer," HelicopterWorld, Shephard Publishing Group,London, England, December 1999.

26. Thompson, B., “Full Throttle to Mars,”Rotor & Wing, Phillips BusinessInformation, LLC, Potomac, MD, March2001.

27. University of Maryland Design Proposal(http://www.enae.umd.edu/AGRC/Design00/MARV.html).

28. Georgia Institute of Technology DesignProposal(http://www.ae.gatech.edu/research/controls/projects/mars/reports/index.html).

29. Kroo, I. and Kunz, P., “Development of theMesicopter: A Miniature AutonomousRotorcraft,” American Helicopter Society(AHS) Vertical Lift Aircraft DesignConference, San Francisco, CA, January2000.

30. NASA Mars Pathfinder missioninformation:(http://www.jpl.nasa.gov/missions/past/marspathfinder.html)

31. NASA 2003 Mars Exploration Rovermission information:(http://www.jpl.nasa.gov/missions/future/marsexplorationrovers.html)

32. Association for Unmanned VehicleSystems, International (AUVSI) Web-Siteon the International Aerial RoboticsCompetition:http://avdil.gtri.gatech.edu/AUVS/IARCLaunchPoint.html

33. Carnegie Mellon University RoboticsInstitute web-site on the AutonomousHelicopter project:http://www.cs.cmu.edu/afs/cs/project/chopper/www/

Appendix -- Rotorcraft and the NASA ‘MarsScout’ Program

The most likely near-term candidate for avertical lift planetary aerial vehicle will be forthe exploration of Mars. The NationalAeronautics and Space Administration’s mostrecent revised Mars exploration plan includes

45. 10

the development of Mars Scout missions. MarsScout missions are intended to be competitivelyselected projects that complement the baselineMars program established within NASA. Aninitial solicitation has been circulated for MarsScout concept studies (Ref. 24). A formalAnnouncement of Opportunity is expected byfirst-quarter calendar year 2002. Both thebaseline Mars Exploration Program and theMars Scout missions are directed to meet thegoals and objectives detailed by the planetaryscience community’s Mars ExplorationProgram/Payload Analysis Group (MEPAG).A key feature of many of the MEPAGobjectives is the requirement for multiple anddiverse site investigations and samplingmissions. A Mars rotorcraft/scout wouldrepresent a satisfactory solution for thisrequirement (Fig. 16).

Fig. 16 – A Rotary-Wing Mars Scout

Many of the most interesting geological featureson Mars lie in terrains that are essentiallyunreachable by wheeled vehicles and currentlanding systems. Examples include theheadwaters of the newly discovered smallMartian gullies and the layered cliff faces alongthe walls of Valles Marineris. Yet in situexploration of these features is critical tounderstanding their formation and the role ofwater in Mars' present and past climate. Avertical lift planetary aerial vehicle (a Marsrotorcraft) would have the flexibility to takeoffnearby, transit to, then hover over and examinesuch high priority science targets. Unlike"single shot" fixed wing aircraft concepts, aMars rotorcraft scout offers the opportunity toperform multiple flights by recharging at thelander.

A notional Mars Scout mission would entaillanding on the Martian surface a suite of scienceinstruments to study the geology and organicchemistry of Martian stratigraphic outcrops,rock fragments, soil and dust and determine itspast water history and biological potential. Thelander would likely be a variant of the 2003Mars Exploration Rover (MER) lander andwould carry a rotorcraft to image and obtainspectral data for geological sites, and to acquire

samples from up to 10 km or so distance fromthe lander.

The feasibility of vertical flight in the Martianatmosphere has been established by designstudies by NASA Ames Research Center andindependent analyses performed by severaluniversity teams (Refs. 25-29). Work on theMars rotorcraft concept is transitioning frompreliminary system analysis to proof-of-concepttest article design, fabrication, and assessmentand fundamental experimental investigations ofthe unique aerodynamics of these vehicles. Inparticular, an isolated rotor configuration --designed to constraints compatible with flight inthe Martian atmosphere -- has been designedand fabricated and is currently undergoing pre-test preparation for hover testing in a NASAAmes environmental chamber. Complementarywork is also under way examining autonomoussystem technology and other critical enablingtechnologies for vertical lift planetary aerialvehicles.

The ultra-lightweight rotorcraft will operatelargely autonomously and will be targeted tosites of interest identified from available orbitalimaging and spectral data after the actuallanding site is accurately determined. Therotorcraft will acquire high-resolution imagingand spectral data and return small samples ofsoil and rock fragments from the designatedsites. The instrumentation carried by the landerwill include an optical microscope, an Infrared(IR) spectrometer and a Gas ChromatographMass Spectrometer (GCMS).

The notional Mars Scout mission wouldcapitalize on lander designs and scienceinstrumentation that have already beendeveloped and will, in addition, introduce newcapabilities in addressing NASA's "follow thewater” theme.

Science Goals and ObjectivesDetermining the mineralogy of the Martiansurface material is the first step in understandingMartian geochemistry. In situ analyses of theMartian surface material can provideinformation on the mineralogy and volatilecontent of Martian surface material needed tocharacterize their geochemical and petrologicnature. Knowing the mineralogy of a sample ofthe Martian surface material provides data onthe environment under which it was formed.This information can be used to better define theearly environment of Mars especially withrespect to the history of water. For example,clays and evaporitic salts require the presence ofwater for their formation; as a consequence, ifthey form part of the Martian surface material

45. 11

their presence would be evidence for water tohave been on the Martian surface for somelength of time. Acquisition of multiple samplesfrom a number of distributed sites is a keyelement of a Mars rotorcraft mission and willclearly enhance the understanding of thegeochemical evolution of Mars.

A Mars Scout rotorcraft could follow a specificflight plan over interesting terrain, for examplethe course of a small gully or along a specificcliff face selected from orbital images. Forwardand aft mounted cameras would provide targetspecific views unobtainable by fixed-wingaircraft or rovers. The rotorcraft will have thecapability to land at remote sites. Landing-legmounted instruments could include amicroscopic imager for measurement of graincharacteristics and sizes. If adequate payloadweight margins could be achieved for the Marsrotorcraft, a sample-collecting scoop could beintegrated into one landing leg to collect soilsamples at the remote site that can betransported back to the lander for furtheranalysis. Sites well suited to rotorcraftexploration include:• Valles Marineris• Young gullies• Headwaters of outflow channels and valley

networks• Basal scarp surrounding Apollinaris Patera

to search for hydrothermal spring depositsand explore sapping valleys.

Mission Description:•Prime Mission: 10-15 Sols (a Sol is one

Martian ‘day’) devoted to acquisition, and in-situ analysis, of soil and small rock samplesimmediately adjacent to the lander (using arobotic arm); 5-10 Sols for the set-up andcheckout of a vertical lift aerial vehicle (anultra-lightweight robotic Marsrotorcraft/helicopter) with the robotic arm; 1sol to execute a short flight/hop and return ofapproximately a hundred meters or so (withinline of sight of lander) to performdemonstration flight and initial sample returnrun; 20-30 Sols to perform a low altitudehigh-resolution aerial survey, of a radius ofseveral kilometers with respect to the landerusing the vertical lift aerial vehicle. Allpower to be provided by the lander solar arraypanels. Aerial vehicle to be rechargedbetween flights by the solar array panels (4-6Sols between aerial survey flights and 6-10Sols for time between sample return flights).

•Secondary Mission: 20-40 Sols devoted toremote-site soil/rock sampling mission flightsat a distance of several kilometers from thelander (over potentially hazardous terrain) viathe vertical lift aerial vehicle (most of this

mission time will be dedicated to rechargingor refueling the vehicle and in-situ analysis ofthe samples and communication of results toEarth). Note overall mission time will beaffected by which of the two primarypropulsion systems options are chosen for thevertical lift aerial vehicle.

•Science PayloadLander Instruments:

Microscopic imagerIR Spectrometer or Raman SpectrometerGas Chromatograph Mass Spectrometer

(GCMS)Wide-field optical camera for

documenting/tracking Mars rotorcraft take-off and landing; used also to guide landerrobotic arm positioning for soil/rock sampletransfer from the rotorcraft to the landerand to aid in the aerial vehicle set-up andrecharging.

Vertical Lift Aerial Vehicle InstrumentsForward- and aft-mounted optical cameras

for Guidance/Navigation and aerialsurvey images

Sun trackerAtmospheric temperature and pressure

sensors for flight readiness anddocumenting remote-site climatology

Landing-leg-mounted camera for soil/rocksample identification and leg-integratedsample probe/scoop positioning

Several vehicle health and flight safety,navigation and control transducers

IMU and assorted accelerometers for flightcontrol.

General Lander and Associated EquipmentDescription (Fig. 17a-e)A lander carrier with solar array petals similar inconfiguration of the 2003 MER and MarsPathfinder landers (Refs. 30-31); an in-situinstrument science module for processing andanalyzing soil and small rock samples; a roboticarm for sampling/transferring rock samples andfurther, assisting set-up, handling, and usage ofthe Mars rotorcraft; the vertical lift aerialvehicle itself, with a transport frame andauxiliary support equipment; lander missioncomputer and communication package.

(a)

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(b)

(c)

(d)

(e)Fig. 17a-e – Mars Rotorcraft Deployment

Primary Objectives•Examine mineralogical and biochemicalcharacteristics of soil and small rock samplesin support of scientific investigations for ‘Huntfor Water,’ ‘Search for Life,’ and thegeological evolution of the Martian surface.

•Perform low-altitude, high-resolution aerialsurveys of geologically interesting Martiansurface features in hazardous or otherwise

inaccessible terrain for rovers and landers;identify remote-sites for follow-on samplingmission flights

•Perform a technology/flight demonstration ofan autonomous vertical lift planetary aerialvehicle to support infrastructure developmentof a class of ‘astronaut agents’ that couldenhance safety and mobility (and, thereby,mission science return) for human explorationof Mars.

Secondary Objectives• Robotic Mars rotorcraft, after initial flightdemonstration and aerial surveys would hoverand land at geologically-interesting remotesites and use a sampling probe – such as ascoop – and pick-up small soil and rocksamples; digital cameras and image processingsoftware on the flight/mission computer wouldautonomously the most interesting samples toacquire. Recorded images will define thecontext (in relation to the surfacecharacteristics in the vicinity of the sample andthe morphology of the surrounding area) of theacquired samples. Samples would be returnedto lander and placed in the in-situ sampleanalysis hopper; the Mars rotorcraft would behooked up (with the lander robotic arm) tolander auxiliary systems for recharging.

• Aerial vehicle ‘Final Flight’ would be a one-way mission to maximize flight range distancefrom the lander primary-site. The Marsrotorcraft would carry a small science payloadin place of the sampling probe to the maximumrange remote-site. The science payload wouldfocus on climatology experiments tocomplement primary-site measurements.

Science Implementation

Crucial to the success of any MarsScout/Rotorcraft mission will be the formationof a strong project team that provides the criticalmulti-disciplined expertise and technology.Research and technical communities thatheretofore have not interacted with each otherwill have to form close, efficient workingpartnerships. This process of openingcommunication and team building has begunbetween planetary scientists, spacecraftdesigners and mission developers, and therotorcraft research community. But themagnitude of this task should not beunderestimated; the cost of planetaryexploration, coupled with the negative impact ofmission failure, is such that a long process ofconfidence-building between these disparatecommunities will be required.

Mission and Flight System Architecture

45. 13

The development of any type of planetary aerialvehicle will be a technically challengingenterprise, a vertical lift vehicle perhaps evenmore so. Not only are there significanttechnical issues to be overcome, but there areperceptual issues as well. Even in space itseems that the friendly rivalry of the fixed-wingversus the rotary-wing aircraft communitiescontinues to thrive. But, even worse,compounding this competitive jostling forattention and potential adoption in the Marsexploration program are the rover andballoon/aerostat (and the ‘hopper’ and multiplesmall lander) proponents.

To minimize overall real and perceived risk, anyMars Scout rotorcraft mission will have toattempt to balance the risk of unproven aerialvehicle technology by maximizing the use of‘heritage’ technology previously demonstratedwith flight hardware. Therefore, a Mars Scoutrotorcraft mission will likely model itself inmany ways after the Mars Pathfinder and the2003 Mars Exploration Rover (MER) missions.

A baseline Mars rotorcraft mass target should beassumed to be approximately 20 kg. At leastone-half hour of flight should be sustained, withhover and take-off and landing from the landerand a remote site location. The ability torecharge/refuel back at the lander will be anessential mission feature. Two differentpropulsion strategies should be examined inparallel – for risk mitigation -- in the conceptualand preliminary design stages of a MarsScout/Rotorcraft effort (fuel-cell versusAkkerman hydrazine engine). Further, becausemass is always a critical issue in spacecraftdesign, tradeoff studies should be made for theaerial vehicle – varying the vehicle mass from10 to 20 kg -- to examine the impact on missionperformance versus risk. Finally, design studiesand experimental investigations should continuethroughout the early stages to benchmarkcoaxial helicopter configurations against quad-rotor vehicle designs. Both vehicleconfigurations have considerable merit/potentialfor early robotic missions to Mars (Refs. 26-28).By pursuing parallel investigation of both aerialvehicle types in the early stages of a Mars Scoutdevelopment effort, a strong final missioncandidate design will likely emerge.

Table 2 is a preliminary ‘Science to MissionTraceability Matrix’ for this notional MarsScout rotorcraft mission. Information containedin this table is used by science team peers andreviewers, and mission planners, to assesswhether or not a mission candidate concept canmeet its identified goals and objectives.

Requirements on Notional Mission:Orbiter Not required; will

utilize pre-existingcommunicationassets and/or lander-based directcommunication withEarth

Launch Vehicle Delta II 7925-9.5Launch Date ~ June 2007Mission duration 90 Sols (upon

landing)Flight System Elements Cruise stage; Entry,

Descent, Landingsystem (EDL):Pathfinder/MER-style tetrahedronwith inflatableairbags

Requirements on Spacecraft Flight System:

Control method Spin stabilized; 2rpm cruise stage.

Instrument Power Minimuminstrumentation (andpower requirements)for trajectorycorrections andspacecraft healthmonitoring; nospacecraft scienceinstrumentation.

Special protection: Mars rotorcraft willbe composed ofmaterials and sub-systems that need tobe assessed for theirenvironmentalcompatibility withspacecraft cruisestage.

Radiation environment No RTGs required;solar and batterypower only.

EDL Maneuvering: None requiredbeyond matchingMER or PathfinderError Ellipses.

Requirements on Communications & DataSystem:

Data Volume (Mbytes per day): ~100Megabytes(per flight)

Number of data downlinks per day: 1Real time requirements: None

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Table 2 - Science-To-Mission Traceability Matrix

Science Driver InstrumentRequirement

Mission Requirement Flight SystemRequirement

Comm. and Ground Data SystemRequirement

Mission OperationsRequirement

Technology Requirement

1. MEPAG Goal I,Objective A,“Determine if LifeExists Today,”Investigation 2,3.,5, 6

GCMS (GasChromato-graph MassSpectrometer)

MicroscopicImager

1. Perform low-altitude,low-speed aerial surveyand select remote-siteswhere geologic formationswould suggest water wasonce existent;2. Acquire at multiple sitessoil and small samples toassess existence of clays,hemotites, and/orsedimentary rocks throughspectrometry;3. Through use of GCMS,assess potential of soilsample for containingorganic compounds and/orlevels of oxidants

EDL must be capableof delivering to theMartian surface a 20kgaerial vehicle; 20 kg ofscience analysispackage/station; and atetrahedral solar array‘petals’ for power; arobotic arm andsupport frame for set-up and recharging

1. Aerial survey digitalimages will comprise the largestfraction (~75%) of data transmittalto Earth; aerial and remote-site(near- and far-field) images willneed to be transmitted throughoutmission duration in order to providethe scientific community thecontextual background toaccompany the soil and rock sampleanalyses;2. Sophisticated softwarefor science analysis, dataprioritization and communication,and mission planning will berequired for both the lander sciencestation and the aerial vehicle.

1. Singleoperations shiftrequired forEarth/Landercommunication;2. Two-three‘off-days’ betweencomplete data setdownlink andinitiation of next aerialvehicle flight requiredfor science teampreliminary analysisand planning;

A. Heritage Instrumentation

B. Development of a ‘Mars Rotorcraft’

C. Develop In-Situ Handling &Processing Tools for the Lander SciencePackage/Station.

D. From an overall Mars program riskmanagement perspective, it wouldprobably be best to couple a ‘low risk’and a ‘high risk’ (such as one employinga Mars rotorcraft) during the same Marstransit window opportunity.

2. Goal I, Obj. B,“Determine if LifeExisted in thePast,” Investig. 1& 2

IR (Infra-Red)Spectrometer

MicroscopicImager

Through use ofmicroscopic imager androckpreparation/processingtools (grinding/slicing)assess rock samples forpaleobiology potential.

Sample handling andprocessing techniquesneed to be developedto transfer samplesfrom rotorcraft tolander science module.

--- ---

A. Microscopic imager and IRSpectrometer will be heritage from2003 MER missions.

B. Robotic arm will have partialheritage from Mars Polar Landerhardware

3. Goal I, Obj. C,“Assess Pre-BioticOrganicChemistry,”Investig. 1

GCMS

---

Cross-contaminationbetween samples mustbe minimized. Propercataloging, archiving,and/or disposition ofsamples must beprovided for.

Sophisticated data management toolswill be required to optimize ‘datafusion’ between the in-situ analysisresults for soil and rock samples andthe sample ‘context’ informationderived from the aerial survey andremote-site imagery.

--GCMS will have heritage dating back tothe Viking lander missions.

4. Goal III, Obj. A,“DeterminePresent State,Distribution, andCycling of Water,”Investig. 2

GCMS

MicroscopicImager

APXS (AlphaProton X-RaySpectrometer)

Through use of the APXSassess the morphology ofsmall rock samples fororigin (volcanic versussedimentary)

--- -----

APXS will have heritage technologydating from the Mars Pathfinder mission.

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New Technology, Infrastructure and RiskAssessment

Heritage systems and technology will be used asmuch as possible in this notional Mars Scout mission,and will include as a minimum: all lander-basedscience instrumentation, the lander andaeroshell/entry vehicle configurations, and thespacecraft system. New technology for this notionalMars Scout mission will primarily be in the form ofthe Mars rotorcraft.

The current NASA Technology Readiness Level(TRL) for a Mars rotorcraft vehicle, as a whole, isTRL=2. Analytical assessments have been made ofthe Mars rotorcraft concept over the past two yearsboth within NASA and other institutions (Refs. 2, 7,11, 26-29).

(a)

(b)

(c)

Fig. 18 – University-Proposed Mars RotorcraftConcepts; (a) Stanford Mars Mesicopter; (b)

University of Maryland MARV; (c) Georgia Instituteof Technology GTMARS

Through the co-sponsorship of Sikorsky Aircraft andNASA Ames, the American Helicopter Society,International conducted its Year 2000 universitystudent design competition on Mars rotorcraft. Thesehighly detailed design studies of the Mars rotorcraftconcept – based on a common set of designrequirements very much consistent with the notionalMars Scout mission outlined in this paper –effectively constitutes a set of independentreviews/assessments of the feasibility of the conceptby academic institutions (Refs. 26-28). In all cases,these academic AHS design competition participantsanalytically verified the feasibility of the Marsrotorcraft concept. Further, funding from the NASAInstitute of Advanced Concepts (NIAC) has beenprovided to university researchers (Ref. 29) forcomplementary work on a very small rotary-wingplatform which has Mars exploration potential,among other applications. See Fig. 18a-c.

Fig. 19 – Mars Rotor Hover Test Stand

Table 3 – Proof-of-Concept Mars Rotor Description

45.17

Number of Blades 4Rotor Diameter 2.438mBlade Root Cut-Out(To simulate bladetelescoping requiredfor storage/transport)

40% blade span

Disk Loading(Nominal ‘1G’)

4 N/m2

Tip Mach Number 0.65Blade Tip ReynoldsNumber

54,855

Thrust Coefficient, CT(Nominal ‘1G’)

0.0108

Mean Blade LiftCoefficient

0.4

Blade Chord 0.3048m (constant) from 40% radialstation outward

Rotor Solidity 0.191Blade Linear TwistRate

0 deg. out to 40% span;+2.4 to –2.4 deg. from 40 to 100%span.

Blade Weight 0.35 kg per bladeFirst FundamentalElastic Modes

1.264 per rev – first flap mode;1.118 per rev – first lag mode;2.310 per rev – first torsion

Outer Blade SpanAirfoil Section

Eppler 387

Spar Section Circular tube with chordwise flat platestiffener (30% chord)

Blade Construction Airfoil fairing is milled foam withinternal cavities;Circular graphite tube spar acrosscomplete span of blade;45 deg. graphite chordwise flat platestiffeners from 5% to 40% station;fiberglass leading edge cap on outerblade section airfoil fairing

Rotor HubConfiguration

Rigid/cantilevered hub, withtension/torsion straps, dry contact pitchbearings, and pitch arms at 5% radialstation

A hover test stand, and a baseline proof-of-conceptrotor (see Fig. 19 and Table 3), have been fabricatedand are nearly ready for testing in a largeenvironmental chamber – which can simulate Marssurface atmospheric conditions. This proof-of-concept rotor, though not as yet an optimized design,has been designed and fabricated to many of theexacting requirements dictated for a flight vehicle –including ultra-lightweight construction and bladedynamic tuning for low structural loads and vibration.The rotor airfoil used for this proof-of-concept rotoris the Eppler 387, a well-known low Reynolds airfoil.Recent unpublished two-dimensional airfoil test datain compressible, near transonic, test conditions atNASA Langley has been acquired for this airfoil,demonstrating moderately high lift coefficient values(R. Campbell - private communication). Anadvantage of rotorcraft, versus any other aerialvehicle proposed for Mars exploration, is the abilityto conduct hover testing in existing ground-testfacilities; additionally, it is also the unique advantageof the Mars rotorcraft concept that the most severeaerodynamic performance operating condition is inhover rather than forward-flight. Upon completion of

planned hover testing in a largeenvironmental/vacuum chamber at NASA Ames, theTechnology Readiness Level for the basic vehicleshould increase to a TRL of 3, wherein test articleshave been fabricated and performance assessed. Theanalytical tools used to date in assessing the aerialvehicle performance will be significantly upgraded inthe near future by applying very sophisticatedrotorcraft modeling tools to perform comprehensiveanalyses in forward-flight (Fig. 20) and Navier-Stokes CFD predictions of the Mars rotorcraft inhover. Confidence in these CFD predictions will begained through validation against the experimentaldata resulting from the proposed proof-of-concepthover testing. Subsequent to the initial isolated rotorhover testing and the CFD work, a tethered ‘flight’ ofa stripped down proof-of-concept vehicle in theAmes environment chamber will be pursued. Thisvehicle, by necessity because of Earth’s highergravity, will have to be powered by ground-basedpower sources and flight controllers (among otherthings) but will represent a major step ahead in thedevelopment of a Mars rotorcraft.

Fig.20 – Mars Rotorcraft: Putting AdvancedComputational Analyses to the Test

The TRL for the autonomous system technology andflight navigation and control should be consideredTRL=3, given past work performed within NASAand within various academic institutions (Refs. 32-33). A study, resulting from a university grant issuedby NASA Ames to Carnegie Mellon University, hasrecently been completed examining from aconceptual design perspective the challenges andpotential of using vision-based navigation systems fora Mars rotorcraft; these preliminary results were veryencouraging. A complementary research programwithin NASA Ames, funded by the AutomatedReasoning element of the NASA Intelligent Systemsprogram, is currently underway and is likely tosignificantly aid in the development of a flightcontroller/mission computer and software for a Marsrotorcraft -- as well as other, terrestrial applications.

The propulsion technology (electric motors and fuelcells (primary option) or hydrazine -- aka Akkermanreciprocating engines – (as secondary, back-upoption) should be considered to be TRL=3 for pastwork performed by NASA, Industry, and academicinstitutions. Some very exciting innovative

45.18

propulsion system concepts were developed as aresult of the AHS student design competition.

Proposing the use of a rotary-wing aerial platform fora Mars Scout mission is not as mature a technicalapproach as many other concepts likely to beadvocated for Mars Scout missions. And yet, theMars rotorcraft concept offers such a tremendouspotential increase in mobility for Mars exploration,with a corresponding near-order-of-magnitudeincrease in mission productivity, that a modestinvestment now, for the future, should be justifiable.

Martian aerial scouts offer the potential todramatically expand the surface area of Mars that canbe explored in future missions. By flying overdifficult topography, aerial vehicles are capable ofcovering much more area than a rover in significantlyless time. The 2003 mission Mars Exploration Roverswill cover approximately 100 meters per Sol; a Marsrotorcraft could cover over twenty times that distanceper flight (assuming a seven day between-flight cyclefor vehicle recharging and data analysis/transmittal toEarth). By operating above the ground surface, thepotential line of sight of sensor systems also greatlyexpands. A Martian aerial scout flying at 100m AGLwould have a line of sight in excess of 25 kmcompared to the 5 km line of sight of a ground basedvehicle assuming flat terrain.

Powered-flight aerial vehicles are superior toballoons/aerostats in all respects, except maybe,simplicity. However, even with respect to theirconceptual simplicity, one has to acknowledge thatballoons, as represented by their terrestrialcounterparts, are not without their own unique failuremechanisms (for example, the early attempts to flythe erstwhile Ultra Long Duration Balloonexperiments). The ability to select an area ofinterest on the Martian surface, direct a poweredaerial vehicle to that location, and to survey andconduct experiments as desired is essential forsuperior scientific investigations of Mars. Having aballoon passively, uncontrollably, skirt across theplanet will be of modest benefit at best.

Vertical lift aerial vehicles – including rotorcraft --combine the exploration area advantage describedabove with the ability to takeoff and land inunprepared sites of scientific interest. Unlike “singleshot” fixed wing aircraft concepts, a vertical liftaerial scout offers the opportunity to performmultiple mission sorties by recharging at the landersite. A vertical lift aerial vehicle solution enablessample return missions. Samples could be gatheredfrom a wide radius to a lander/primary-base. Asdemonstrated on Earth, rotorcraft uniquely havesuperior low-speed handling qualities. RotorcraftMars scouts would enable low-speed, precisemovement in three dimensions allowing the craft toclosely study cliff walls or capture a 360° surface

view of large objects. Highly sloped terrain, possiblyresultant from erosion, can be thoroughly studied.This terrain will remain unexplored by groundvehicles or fixed wing aircraft concepts while arotorcraft can fly low to the ground, allowing greatimage detail. Low speed handling qualities maketakeoff and landing operations possible in unpreparedterrain. Finally, fixed-wing aerial vehicles sufferfrom substantial technical challenges in their releasefrom entry vehicles in descent, or launch/catapultingfrom ground-based assets. Even hypersonic rocket-propelled ‘fixed-wing’ aerial vehicles -- that are bothentry vehicle as well as aerial scout -- pose significanttechnical challenges; such hypersonic aerial vehicleshave very limited developmental heritage forterrestrial applications, let alone their readiness forplanetary exploration missions.