MARVEL – Mars Aerial Research Vehicle and Environmental LaboratoryMAE 4351/ Aerospace Vehicle Design II, Summer 2016Aurora Concepts

Conceptual design of an Aerial Platform for Mars: the Mars Bush Plane

Demonstrate that flying in Mars is possible Take-off and landing capability (re-usable) Technology demonstrator for future manned missions to Mars

Account for current/future space industry capability Technology Readiness Level (TRL) First generation Mars ground infrastructure Autonomous deployment from Aeroshell / Space Capsule Incorporate into promising commercial/scientific plans

Investigate: Past and present Mars aircraft concepts Mars atmosphere and topography Launch vehicle capabilities and constraints

Top: Luis Salas, Juan Lopez, Nic Dwyer, Ismael SanabriaBottom: Ryan Manns, Yasir Rauf, Justin Kenna, Ian Maynard

4) TEAM STRUCTURE & ROLES3) AURORA CONCEPTS TEAM1) STUDY OBJECTIVES

5) Multidisciplinary Analysis 7) PARAMETRIC SIZING PROCESS

8) GEOMETRY AND WEIGHTS

6) PREVIOUS AIRCRAFT CONCEPTS 9) PERFORMANCE

10) AERODYNAMICS

11) PROPULSION AND ENERGY

12) STRUCTURES 13) STABILITY & CONTROLS

2) BUSINESS CASESpaceX

Simplicity, reliability, and cost effectiveness

April 8 2016: Falcon 9 successfully lands on drone ship

NASA contracted them for 12 robotic supply flights for $1.6 billion

Falcon Heavy is expected to have its first demo flight in December

First crewed mission to Mars in 9 years

NASA Currently working on the Space

Launch System (SLS) SLS will provide access to deep

space SLS Block 1B will be ready by

2017

Taken from: B. Hill & S. Creech, “NASA’s Space Launch System: A Revolutionary Capability for Science,” NASA, July 2014. [Online]. Available: https://www.nasa.gov/sites/default/files/files/NAC-July2014-Hill-

Creech-Final.pdf. [Accessed 05 July 2016]

Astroplane (1978)Taken from: Developmental Sciences Inc, “A Concept Study of a Remotely Piloted Vehicle for

Mars Exploration,” NASA CR-157942, City of Industry, CA, 1978

MAGE (1999)Taken from: J. McGowan, “A Mars Airplane,” JMcGowan.com, 03 December

1999. [Online]. Available: http://www.jmcgowan.com/marsplane.html. [Accessed 12 June 2016]

ARES (2003)Taken from: M. Guynn, M. Croom, S. Smith, R. Parks and P. Gelhausen, "Evolution of a Mars Airplane Concept for the ARES Mars Scout Mission," AIAA Unmanned Unlimited

Systems, Technologies, and Operations, San Diego, CA, 2003.

Loftin-based Sizing Methodology, adapted for low wing loading aircraft

Determination of aircraft weight (Wg), wing area (S), and power (P) based on mission requirements and historical trends

Mission requirements: Payload Mass: 250 kg

(Pilot+Space Suit+Life Support Systems+Miscellaneous)

Stall Speed: 70 m/s (155 mph) Max. Speed: 150 m/s (335 mph) Short Takeoff-Landing or VTOL Range: 400 km

Feasible region

Twin Vertical Tail

Two 4-bladed Propellers

Wing Tip Dihedral

Tadpole Fuselage for 1 Person

Collapsible Tail Boom

High AR Wing

Challenges: Atmospheric composition is 95% CO2 leading to air

breathing propulsion to be unusable Rocket Propulsions takes a large weight fraction to

go long distances Propeller propulsion has issues producing thrust in

a low density atmosphereVariable Description Value UnitsSw Wing planform area 110 m^2ARw Wing aspect ratio 17.6 -bw Wing span 44 mTRw Wing taper ratio 0.8 -CRw Wing root chord 2.5 mCTw Wing tip chord 2 mCMACw Wing mean aerodynamic chord 2.5 mALLEw Leading edge sweep - degAL25w Quarter chord sweep - degTCw Average chord sweep - -TWISTw Wing twist angle 8 deg

WING DESCRIPTIONVariable Description Value Units

Sh Horizontal tail area 15.21 m^2ARh Horizontal tail aspect ratio 3 -Bh Horizontal tail span 6.755 mCRh Horizontal tail root chord 2.25 mTCh Horizontal tail thickness ratio 0.08 -ALCh Distance from wing to HT MAC 10 m

Variable Description # Value UnitsSv Vertical tail area 2 5.88 m^2ARv Vertical tail aspect ratio 1.5 -Bv Vertical tail span 2 2.613 mCRv Vertical tail root chord 2 2.25 mTCv Vertical tail thickness ratio 0.101 -ALCv Distance from wing toVHT MAC 10 m

HORIZONTAL TAIL DESCRIPTION

VERTICAL TAIL DESCRIPTION

Variable Description Value UnitsALFUS Fuselage length 13.27 mHFUS Fuselage max height 1.92 mWFUS Fuselage max width 1.2 mDMAX Fuselage max equivalent diameter 1.92 m

FUSELAGE DESCRIPTION

Wing 259.46Empennage 36.26Power Plant 150.07

Batteries 210Rocket Propellantr 300

Fuselage 105.32Landing Gear 59.6973

Fixed Equipment 96.91Payload 250

Total 1467.717

Main Components Mass (kg)

Altitude (m)Stall Velocity

(m/s)

Velocity min Drag/ max Range

(m/s)

Velocity min Pwr/ max Endurance

(Hr)0 62.666 81.949 62.268

250 63.342 82.833 62.940500 64.026 83.727 63.619750 64.717 84.631 64.306

1000 65.415 85.544 65.0001250 66.121 86.467 65.7011500 66.835 87.400 66.4101750 67.556 88.343 67.1262000 68.285 89.296 67.8512250 69.021 90.260 68.5832500 69.766 91.234 69.3232750 70.518 92.218 70.0703000 71.279 93.212 70.826

Challenges: Low atmospheric pressure and density High Mach number Airfoil must have high L/D characteristics Airfoil must perform in low Re environments

Airfoil CLmax Cm,c/4 ReE214 1.25 -0.11 2x10 5̂E423 2 -0.25 2x10 5̂

FX 63-137 1.75 -0.17 2x10^5LA2573A 1.86 0.02 2x10 5̂

S1223 2.23 -0.29 2x10 5̂

0 0.5 1 1.5 2 2.5 3Reynolds Number 105

0

0.5

1

1.5

2

2.5

CL m

ax

CLMAX

vs Re for FX 63-137 and S1223 Airfoils

FX 63-137S1223

0 0.5 1 1.5 2 2.5 3Reynolds Number 105

0

10

20

30

40

50

60

70

80

90

100

L/D

max

L/D max vs Re for FX 63-137 and S1223 Airfoils

FX 63-137S1223

-0.53

0

0.5

20

C L

2

1

Variation of CL with Re and for FX 63-137 Aerofoil

Reynolds Number

105

1.5

10

, deg

2

1 00 -10

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1x/c

-0.1

0

0.1

0.2

y/c

S1223 High Lift Low Re Airfoil

0 0.2 0.4 0.6 0.8 1x/c

-0.1

0

0.1

0.2

y/c

FX 63-137 Airfoil Low Re Airfoil

Airfoil Selection

0 0.5 1 1.5 2 2.5 3Reynolds Number 105

0

0.02

0.04

0.06

0.08

0.1

0.12

C DM

IN

CDMIN

vs Re for FX 63-137 and S1223 Airfoils

FX 63-137S1223

Lift distribution at V=100 m/s at α=5°

Cp over wing at V=100 m/s at α=5°

Wing Planform Shape

MAC = 2.5 m Ensures Re does not drop

below critical along the wing Large wing area = 110 m2

Low wing loading & stall speed High aspect ratio = 17.6

High CLmax

Good climb performance Wing dihedral

Recovery from gust loading Sweeped winglet

Less induced drag and higher efficiency

Complete Aircraft Aerodynamics

-6 -4 -2 2 4 6, deg

-0.5

0.5

1

1.5

C L

CL vs. for MARVEL at 100 m/s

no flapsflapup 10°flapdown 10°

-6 -4 -2 2 4 6, deg

10

20

30

40

CL/C

D

CL/CD vs. for MARVEL at 100 m/s

-6 -4 -2 2 4 6, deg

-0.25

-0.2

-0.15

-0.1

-0.05

0.05

0.1

CM

Pitching Moment vs. for MARVEL at 100 m/s

System Description: Rocket-assisted takeoff and landing

Hydrazine or CO2 as propellant Propeller-driven cruise flight

Two four-bladed propellers, 4.25 m in diameter Airfoil: NACA 65 Siemens electric engine (2400 rpm with 95% efficiency)

Ultra Light Lithium Sulfur batteries provide energy (450 Wh/kg) Recharged either by solar cells or RTG

Propeller Design + Drivetrain

Ultra Light Batteries

Gear Box Weight 18.10 kgWeight of Prop 9.90 kgWeight of Controller 9.52 kgWeight of Motor 45.96 kgTotal weight 157.44 kg

System Mass

Gear Box Effi ciency 0.95Prop Effi ciency 0.75Controller Effi ciency 0.99Motor Effi ciency 0.97Total Efficiency 0.68

System Efficiency

Diameter, m 4.25Number of Blades 4

RPM 800.00

Thrust, N 511.25Power, kW 59.34Effi ciency 0.75Overall Effi ciency 0.68

Location of EnginesIn line with Vertical Tail

Propeller Design And Outputs

Rocket-Assisted Takeoff/Landing

Challenges: Low atmospheric density and low Re

results in large control surfaces Constrained by volume limitations

imposed by aeroshell Optionally-piloted aircraft

Fly-by-wire/Three-axis autopilot Ensure static stability during regular and

critical flight conditions Thrusters and control surfaces

locations

Static Margin estimated to be ~ 9%

Low thrust + Low atmospheric density = excessively large ground roll. Since no infrastructure is available on Mars, it is necessary to have short (or vertical) takeoff/landing TO roll: 17.25 km Landing roll: 10.4 km

16% of the aircraft mass is rocket fuel Hydrazine Isp: 230 s CO2 Isp : ~130 (theoretical)

Shear, Moment, and Deflection analysis of wing during flight