Presentation Poster - Aurora Concepts
1
MARVEL – Mars Aerial Research Vehicle and Environmental Laboratory MAE 4351/ Aerospace Vehicle Design II, Summer 2016 Aurora 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 Sanabria Bottom: Ryan Manns, Yasir Rauf, Justin Kenna, Ian Maynard 4) TEAM STRUCTURE & ROLES 3) AURORA CONCEPTS TEAM 1) 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 CASE SpaceX 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 (W g ), 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% CO 2 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 atmosphere Variable Description Value Units Sw W ingplanform area 110 m ^2 ARw W ingaspectratio 17.6 - bw W ingspan 44 m TRw W ingtaperratio 0.8 - CRw Wingrootchord 2.5 m CTw W ingtip chord 2 m CM ACw W ingm ean aerodynam icchord 2.5 m ALLEw Leadingedge sw eep - deg AL25w Quarterchord sw eep - deg TCw Average chord sw eep - - TW ISTw W ingtwistangle 8 deg W ING DESCRIPTION Variable Description Value Units Sh Horizontal tail area 15.21 m^2 ARh Horizontal tail aspectratio 3 - Bh Horizontal tail span 6.755 m CRh Horizontal tail rootchord 2.25 m TCh Horizontal tail thicknessratio 0.08 - ALCh Distance from w ingto HTM AC 10 m Variable Description # Value Units Sv Vertical tail area 2 5.88 m^2 ARv Vertical tail aspectratio 1.5 - Bv Vertical tail span 2 2.613 m CRv Vertical tail rootchord 2 2.25 m TCv Vertical tail thicknessratio 0.101 - ALCv Distance from w ingtoVHTM AC 10 m HORIZONTALTAILDESCRIPTION VERTICALTAILDESCRIPTION Variable Description Value Units ALFUS Fuselage length 13.27 m HFUS Fuselage m axheight 1.92 m W FUS Fuselage m axw idth 1.2 m DM AX Fuselage m axequivalentdiam eter 1.92 m FUSELAGEDESCRIPTION Wing 259.46 Em pennage 36.26 Pow erPlant 150.07 Batteries 210 RocketPropellantr 300 Fuselage 105.32 Landing Gear 59.6973 Fixed Equipm ent 96.91 Payload 250 Total 1467.717 M ain C om ponents M ass (kg) Altitude (m ) Stall Velocity (m /s) Velocitym in Drag/m axRange (m /s) Velocitym in Pw r/ m axEndurance (Hr) 0 62.666 81.949 62.268 250 63.342 82.833 62.940 500 64.026 83.727 63.619 750 64.717 84.631 64.306 1000 65.415 85.544 65.000 1250 66.121 86.467 65.701 1500 66.835 87.400 66.410 1750 67.556 88.343 67.126 2000 68.285 89.296 67.851 2250 69.021 90.260 68.583 2500 69.766 91.234 69.323 2750 70.518 92.218 70.070 3000 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 A irfoil C Lm ax C m ,c/4 Re E 214 1.25 -0.11 2x10^5 E 423 2 -0.25 2x10^5 FX 63-137 1.75 -0.17 2x10^5 LA 2573A 1.86 0.02 2x10^5 S 1223 2.23 -0.29 2x10^5 0 0.5 1 1.5 2 2.5 3 Reynolds Number 10 5 0 0.5 1 1.5 2 2.5 C L max C L MAX vs Re for FX 63-137 and S1223 Airfoils FX 63-137 S1223 0 0.5 1 1.5 2 2.5 3 Reynolds Number 10 5 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 63-137 S1223 -0.5 3 0 0.5 20 C L 2 1 Variation of C L with Re and for FX 63-137 Aerofoil Reynolds Number 10 5 1.5 10 , deg 2 1 0 0 -10 0 0.2 0.4 0.6 0.8 1 x/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 3 Reynolds Number 10 5 0 0.02 0.04 0.06 0.08 0.1 0.12 C D MIN C D MIN vs Re for FX 63-137 and S1223 Airfoils FX 63-137 S1223 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 m 2 Low wing loading & stall speed High aspect ratio = 17.6 High C Lmax 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 C L vs. for MARVEL at 100 m/s no flaps flapup 10° flapdown 10° -6 -4 -2 2 4 6 , deg 10 20 30 40 C L /C D C L /C D 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 C M Pitching Moment vs. for MARVEL at 100 m/s System Description: Rocket-assisted takeoff and landing Hydrazine or CO 2 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 GearBoxW eight 18.10 kg WeightofProp 9.90 kg WeightofController 9.52 kg W eightofMotor 45.96 kg Total weight 157.44 kg System M ass GearBoxEfficiency 0.95 Prop Efficiency 0.75 ControllerEfficiency 0.99 M otorEfficiency 0.97 Total Efficiency 0.68 System Efficiency Diam eter, m 4.25 Num berofBlades 4 RPM 800.00 Thrust, N 511.25 Pow er, kW 59.34 Efficiency 0.75 Overall Efficiency 0.68 Location ofEngines In line w ith Vertical Tail PropellerDesign 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 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 I sp : 230 s CO 2 I sp : ~130 (theoretical) Shear, Moment, and Deflection analysis of wing during flight
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Transcript of Presentation Poster - Aurora Concepts
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
