CDR
Titan EDD
Fall Semester
12/02/2008
Andrew Welsh
Jon Anderson
Nick Delucca
Steve Hu
Travis Noffke
Pawel Swica
Andrew Welsh
1
Team Member
Andrew Welsh
Team Lead
Hours Worked: 108
Andrew Welsh
2
Agenda
Titan EDD CDR Agenda:
•Introduction and Overview – Andrew Welsh
• Entry Position and Entry Capsule – Pawel Swica
• Entry Simulation and Atmospheric Profile – Nick Delucca
• Parachute and Parachute Deployment – Travis Noffke
• Airship and Airship Deployment – Jon Anderson
• Helicopter and Helicopter Deployment – Steve Hu
• FMEA – Andrew Welsh
Andrew Welsh
3
Overview
Purpose:
• Design a system to insert an aerial vehicle into
Titan’s atmosphere capable of exploring the ethane
lake on Titan’s south pole.
Ontario Lacus (Credit: Right image - NASA/JPL/University
of Arizona Left image - NASA/JPL/Space Science Institute)
Andrew Welsh
4
Overview
Mission Description:
• 2018 tentative launch date
• Aerial vehicle four year operational lifetime
• Aerial vehicle: Helicopter, Airship, Fixed Wing
• Capable of exploring Ontario Lacus
Andrew Welsh
5
Requirements
1. Design an Entry, Descent, and Deployment (EDD) process
and select an aerial vehicle based of work done elsewhere.
2. EDD system capable of delivering the aerial vehicle on or
near the surface of Titan.
• Heating Constraints
• Deceleration Constraints
3. Aerial vehicle must at a minimum be able to explore
Ontario Lacus.
4. Successful aerial vehicle deployment in a configuration
capable of beginning its exploration mission.
Andrew Welsh
6
Requirements
Major Tasks (Given):
• Review previous work done in this area
• Top level trade study for EDD configuration
• Develop or modify an entry and descent phase simulation
• Design a baseline entry and descent phase
• Identify the aerial exploration vehicle
• Design aerial vehicle separation method
• Integrate all systems into a package that will take aerial
vehicle from insertion to employment
• Formula derivations
Andrew Welsh
7
Requirements
Major Tasks (Added):
• Detailed aerial vehicle selection
• EDD animation
• Bad weather simulation
Andrew Welsh
8
Approach
Team Lead:
Andrew Welsh
Integration Team:
Travis Noffke, Pawel Swica, Steve Hu
Entry Team:
Nick Delucca,
Pawel Swica
Descent Team:
Travis Noffke,
Andrew Welsh
Andrew Welsh
Dep. Team:
Jon Anderson,
Steve Hu
9
Approach
Research
Aerial Vehicle
Selection
Descent Method
Selection
Previous
Research
Entry Method Selection
Our Research and
Calculations
Final Product
Andrew Welsh
10
Program Plan
Gantt Chart
Andrew Welsh
11
Program Plan
Task list with responsible
engineers, status, hours
worked, total hours
worked
Andrew Welsh
12
Design Walkthrough
Andrew Welsh
13
References
1.
2.
http://www.sciencedaily.com/releases/2008/07/0807301
40726.htm
https://www.aem.umn.edu/courses/aem4331/fall2008/Ti
tanExplorer.htm
Andrew Welsh
14
Team Member
Pawel Swica
Entry/Integration
Hours Worked: 101
Pawel Swica
15
1
Initial Orbit Calculation
•First thing done was orbit calculation to reach Titan
•The final entry speed was 3.6 km/s
•However, the orbity relied on a large change in velocity
at earth orbit
•Though entry speed was better, published materials
detailed how an ion engine powered by solar cells would
make launch less expensive
Pawel Swica
16
Relevant Equations
Pawel Swica
17
Orbit Diagrams
Pawel Swica
18
Orbit Diagrams
Pawel Swica
19
Orbit Diagrams
Pawel Swica
20
Resulting trade study
Pawel Swica
21
Heat Shield
• To get an idea of heating, an attempt was made to get an
equation that we could put into simulink to get heating during
entry
• While mostly successful the results were off by some fudge
factor
• We went to Professor Candler for assistance
•
10/14 notes from meeting with Candler
– Detailed modeling of entry heating unrealistic given our level of
experience
– Best approach would be to tweak results of previous publications to
fit our conditions (emphasis on Laub paper)
Pawel Swica
22
Relevant Equations
Pawel Swica
23
Heat Shield Results
• Results were taken from Laub paper for mass and material
• SRAM material and less radiative heating gave up to 100 kg
mass savings from previous study
Figure 5. Aeroshell model
Figure 4. Entry heating graph
Pawel Swica
24
Entry Corridor
•
•
Our gathered materials gave no indication of what the entry corridor
was or how it could be determined
10/24 notes from meeting with Candler
– No easy way to determine entry corridor
– Hunt through references to find entry angle used
– Also can plug angles into simulink to see corridor
• Found entry corridor information in AAS 06-077
• 50º down at entry interface (1000 km) with 5º margin in either
direction
• Modeling in simulink roughly agreed with these values
Pawel Swica
25
Calculation of Landing Point
• Lastly the point where the probe is expected to land needed to
be calculated
• Outside searches proved fruitless, however using the given
initial conditions calculation was possible and successful
Pawel Swica
26
Relevant Calculations
Pawel Swica
27
Conclusive Results
• Ontario Lacus lies at latitude 72º and the downstream distance
given by simulink is 1200 km
• Based on the final results of the calculations, the probe can land
as close as 181 km from Ontario Lacus. This is a very
manageable distance for the probe to traverse
• However, if the planet is facing the wrong way, this distance
could become almost 1800 km, in which case reaching the lake
would depend heavily on the durability of the probe
• This depends on timing the approach just right, which is beyond
the scope of our analysis
Pawel Swica
28
Vpython Model
• To verify results a Vpython orbit script was
modified to match the precise conditions
given by the calculations
• To ensure accuracy, starting point is about
180 Titan radii out
Pawel Swica
29
Vpython Video
Pawel Swica
30
Team Member
Nick De Lucca
Titan Atmosphere
Entry Simulation
113 Hours
Nick De Lucca
31
Atmospheric Profile
• Needed Information
– Density
– Temperature
• Sources
– From 50 km to 1000
km data pulled from
plots generated by
others
– Sea level to 50 km
data from email
correspondence.
Nick De Lucca
32
Titan Entry Simulation
• Goals:
– Determine flight characteristics as a function of
time
– Analyze heating
• Simulation Method:
– Newtonian aerodynamics
– Attempted heating calculation
– Given original version of simulation by Professor
Garrard
Nick De Lucca
33
Newtonian Aerodynamics
• Formulas:
GMm
Fg 
R2
1
2 m
FD    v
2
B
FB  Vg
B
m
CD A
c  20.05 T
Nick De Lucca
34
Simulation Methods and
Parameters
• Polar Coordinate System
– Using Velocity and Flight Path angle as reference
directions
• Atmospheric Modeling
– Two exponential profiles for density
– Three linear temperature profiles
Nick De Lucca
35
Thermodynamic Analysis
• Original Plan: Model using Simulink
– Too complicated to manage within time allowance
and with our current
• Current Method
– Adapt the results of others
– Scale for our ballistic coefficient
– Determine location of peak heating by normalizing
Nick De Lucca
36
Simulation Scope
• Three Total Simulations:
– 1000 km to 8 km: entry capsule
– 8 km to 5.6 km heat shield with inflating airship
• Time variant ballistic coefficient
• Buoyancy
– 5.6k km and below heat shield with helicopter
Nick De Lucca
37
Results
•
•
•
•
Total time taken: 3947 seconds
m
Peak deceleration: 86.6 s 2
Maximum heating rate: 146 W
Total heat Transferred: 22552 J
Nick De Lucca
38
Results
Position
1000
900
800
700
Altitude
600
500
400
300
200
100
0
0
200
400
600
Down Range (km)
Nick De Lucca
800
1000
1200
39
Results
Position
8.5
8
Altitude (km)
7.5
7
6.5
6
5.5
5
1174
1174.05
1174.1
1174.15
1174.2
1174.25
Down Range (km)
Nick De Lucca
1174.3
1174.35
1174.4
40
Results
1000
900
800
700
Altitude
600
Position
500
Normalized Heating
400
Normalized Deceleration
300
200
100
0
0
200
400
600
Down Range (km)
800
Nick De Lucca
1000
1200
41
Results
Velocity
7
6
Velocity (km/s)
5
4
3
2
1
0
0
200
400
600
800
1000
Time (s)
Stephen Hu
1200
1400
1600
1800
2000
42
Results
Time Dependencies
1200
1000
Distance (km)
800
Down Range
600
Altitude
Normalized Heating
400
Normalized Deceleration
200
0
0
200
400
600
800
1000
Time (s)
Nick De Lucca
1200
1400
1600
1800
2000
43
Methods for improvement
• Non-Newtonian aerodynamics
• CFD for the heat shield
Nick De Lucca
44
References
• Dynamics and Thermodynamics of Planetary
Entry. W.H.T. Loh. Prentice-Hall Space
Technology Series. 1963.
• Kazeminejad et al. Temperature Variations in
Titan's Upper Atmosphere: Impact on
Cassini/Huygens. Annales Geophysicae 23.
pp1183-1189. 2005
Stephen Hu
45
Titan Entry Descent and Deployment:
Descent Phase
Parachute Decelerator System
for Aeroshell Separation
Travis Noffke
•Decelerator System Definition
•Parachute Characterization and Mechanics
•Aeroshell Separation Analysis
•System Integration
•Hours: 104
11/19/2008
Travis Noffke
46
Primary Goals
1. Provide deceleration force to top of aeroshell
2. Clearance of unnecessary system components
3. Minimize payload descent stability disruption
11/19/2008
Travis Noffke
47
Requirements
1. Deploy decelerator at altitude
which allows for airship
inflation prior to final
separation and deployment
2. 10 meter clearance between
top aeroshell and leading
payload within 5 seconds
3. Maintain stable descent of
each payload component to
respective deployment phase
11/19/2008
Travis Noffke
10 meters
Payload Containment
48
System Events
1.
2.
3.
4.
5.
Aeroshell separation mechanism fires
Mortar fires deployment bag
Parachute inflates
Deceleration on top aeroshell
Complete separation
V top
Payload Containment
V payload
11/19/2008
Travis Noffke
49
Design Tasks
Geometry Selection
Parachute Characterization
• Sizing
• Opening Forces
Conical Ribbon
1.5 m Diameter nom
• Loading
• Mass Ratio
• Ballistic Coefficient
• Material Selection
• Canopy Design
11/19/2008
Kevlar© 29
Travis Noffke
50
Drag Generation and Time
v terminal = 12 m/s
c D = 0.50
Altitude = 8 km
Atmospheric
Density (ρ)= 3.910 kg/m^3
11/19/2008
Travis Noffke
51
Opening Forces
11/19/2008
Travis Noffke
52
Separation Mechanics
Example of separation mechanism
Commonly used in
spacecraft
Performance must
meet requirements
One of several methods
Figure 5. Separation Spring Concept
11/19/2008
Travis Noffke
53
Canopy Structure
11/19/2008
Travis Noffke
54
Canopy Material
Kevlar© 29
•
•
•
•
Highest strength-weight ratio
Superior tensile strength
Space tested
Used in heritage systems
11/19/2008
Travis Noffke
55
Further Studies
• Separation System
– Nominal functionality test
– Drop testing
• Parachute Decelerator System
– Deployment Test
– Drop Test
• Stability Analysis
– Test body and flow measurements for appropriate Re
11/19/2008
Travis Noffke
56
Additional Information (Backup Slides)
Deceleration Method Trade Study
Complexity:
Cost:
Risk:
Efficiency:
Availability
Effectiveness
11/19/2008
Aero-control
Rockets Parachutes
Surfaces
High
Low
Med
High
Low
High
High
Low
Med
Low
High
High
Med
High
Med
High
High
Med
Travis Noffke
57
Additional Information (Backup Slides)
Parachute Geometry Study
Characterization
Reliability:
Mass:
Average Oscillation Angle
Drag Coefficient Range
Opening Force Coefficient
Performance Comparison
Reliability:
Mass:
Average Oscillation Angle:
Drag Coefficient Range:
Opening Force Coefficient:
11/19/2008
Conical Ribbon
High
Less
±3°
.5 to .6
1.05 to 1.3
Conical Ribbon
+
+
+
+
+
Travis Noffke
Disc-Gap-Band
High
More
±10 to 15°
.52 to .58
1.3
Disc-Gap-Band
+
+
58
Additional Information (Backup Slides)
Parachute Cluster Configuration Trade Study
Single Conical
Cluster Ribbon Parachute
Reliability:
High
High
Difficulty:
Med
Low
Redundancy: High
Low
Stability
High+
HighCost:
High
Low
Mass:
High
Med
Andy Welsh
11/19/2008
Travis Noffke
59
Additional Information (Backup Slides)
Parachute Avoidance Trade Study
Rocket Assisted Parachute
Separation
Separation
Complexity:
High
Med
Cost:
High
Low
Risk:
Med
Low
Effectiveness:
High
MedAndy Welsh
11/19/2008
Travis Noffke
60
Additional Information (Backup Slides)
Used
Derivations:
11/19/2008
Travis Noffke
61
Additional Information (Backup Slides)
11/19/2008
Travis Noffke
62
Team Member
Jon Anderson
Hours Worked: 118
Jon Anderson
63
Outline
Outline:
•
•
•
•
•
•
•
•
Objective
• Vehicle selection
• Airship Design
Design constraints
Assumptions
General design
Performance
Deployment
Enabling technologies
Recommendation and conclusion
Jon Anderson
64
Objective
Goal – Vehicle Selection: Conducted trade studies and
vehicle selection process to determine the best possible
vehicle to complete the science mission.
Goal - Airship: The primary mission of the airship is to
function as a relay between the orbiter and the helicopter.
The secondary mission of the airship is to function as a
reserve platform capable of carrying out the science
mission should the helicopter become inoperable.
Jon Anderson
65
Performance
Float mass
Operational Cruse Velocity
Max Velocity
Min Climb/Descent Rate *
Range
Service Ceiling
Absolute Ceiling
Estimated Lifetime *
Length
Width
Volume
Mass He Needed (10% reserve)
Jon Anderson
195 Kg
2.5 m/s
2.98 m/s
50 m/min
36200 km
5 km
40 km
150 days
13.83 m
3.45 m
34.47 m
27.4 kg
3
66
Vehicle Selection
Mass (lower is better): This category physically rates the aerial
vehicles on their expected mass.
Technology Development Needs (Lower is better): This
category qualitatively rates the aerial vehicles on the amount of
research and development needed to make the design option
feasible.
Operational Risk (Lower is better): This category qualitatively
rates the aerial vehicles on the “risk” associated with deploying
and operating on Titan.
Environmental Tolerance (Higher is better): This category
qualitatively rates the aerial vehicles on their ability to withstand
the environment and correct faults.
Jon Anderson
67
Vehicle Selection
Surface Capability (Higher is better): This category
qualitatively rates the aerial vehicles on their ability to
interact with the Titan surface. While all aerial vehicle
options can move close to the surface, only the helicopter
and combination can physically land on the surface.
Mission Completion Probability (Higher is better): This
category qualitatively rates the aerial vehicles on their
ability to complete the mission
Deployment Ability (Higher is better): This category
qualitatively rates the aerial vehicles on their deployment
methods.
Jon Anderson
68
Vehicle Selection
Category
Helicopter
Airship
Helicopter &
Airship
Combination
Mass
320 kg
490 kg
UNK kg
Technology
Development
Needs
Medium
Lower
High
Operational
Risk
Medium
Low
Medium
Environmental
Tolerance
Medium
High
Medium
Surface
Capability
High
Medium
High
Mission
Completion
Probability
Medium
Medium
High
Deployment
Ability
Medium
Jon Anderson High
High
69
Design Constraints
•
Communication payload
• Extra redundancy – orbiter and helicopter
•
Science payload
•
Propulsion subsystem
• Mass assumptions – initial starting value
•
Power subsystem
• MMRGT
Jon Anderson
70
Assumptions
Mass Assumption:
• Needed initial estimate for mass of hull and structural
components
• Found fraction of weight for non-hull components vs
NASA
• Estimated initial weight
•
Designed airship, calculated final mass
•
Reiterated process with calculated mass
Jon Anderson
71
350
12000000
300
10000000
250
8000000
200
6000000
150
4000000
100
2000000
50
Reynolds Number
14000000
Drag (N)
Reynolds # and Drag vs Velocity
Reynolds
number
Drag
0
0
0
1
2
3
4
5
6
7
Velocity (m/s)
Jon Anderson
72
Power Required/Available vs Velocity
500
450
400
Power (W)
350
300
Power Required - Drag
250
200
Full System Operation - Power Available with
20% margin
150
Full Proplusion Operation - Power Available
with 20% margin
100
50
0
0
0.5
1
1.5
2
2.5
3
3.5
4
Velocity (m/s)
Jon Anderson
73
Inflation time/percent vs Lift
1.2
350
300
1
0.8
200
0.6
150
0.4
100
0.2
Inflation Time (sec)
250
50
0
0
0
20
40
60
80
100
120
Buoyant Force/Lift (N)
Jon Anderson
140
160
180
Lift
3 min inflation time
5 min inflation time
74
Deployment
Airship inflation immediate
• Both bayonets and main envelope
• Changing ballistic coefficient
• Separate via explosive shearing bolts
• Immediately max velocity
Jon Anderson
75
Enabling Technologies
Multi Mission Radioisotope Thermal Generator
•
•
•
•
Heat exchanger – not fins
Centrifugal turbine – low power/mass flow levels
Alternator – bearing system, no gears
Centrifugal Compressor
•
•
5 fold increase in power
Lower mass
Jon Anderson
76
Recommendation and Conclusion
High Altitude Design
Detailed data bandwidth analysis
Hull/system optimization
Experments
Jon Anderson
77
Questions?
Jon Anderson
78
Backup slides - Mass
Component
Mass (kg)
Mass after 20% Margin (kg)
Subsystem
Power
2nd Generation MMRTG
Battery - 12 A h lithium
Turbomachinery
Turbine
Compressor
Piping
Electric Motor
Alternator
17
0.47
3.94
0.9
0.9
0.716
1.08
1.08
20.4
0.564
4.728
1.08
1.08
0.8592
1.296
1.296
26.086
31.3032
Propeller, axel, case*
5.25
6.3
Total
5.25
6.3
3
10
1.3
3.6
12
1.56
14.3
17.16
0.114
2.7
2.1
9.8
1.5
1
6
0.1368
3.24
2.52
11.76
1.8
1.2
7.2
23.214
27.8568
0.9
9
4.4
0.32
0.38
5
1.08
10.8
5.28
0.384
0.456
6
20
24
Total
Propulsion
Science Instruments
Haze and Cloud Partical Detector
Mass Spectrometer
Panchromatic Visible Light Imager
Total
Communication
X-Band Omni - LGA
SDST X-up/X-down
X-Band TWTA
UHF Transceiver (2)
UHF Omni
UHF Diplexer (2)
Additional Hardware (switches, cables, etc.)
Total
ACDS
Sun Sensors
IMU (2)
Radar Altimeter
Antennas for Radar Altimeter
Absorber for Radar Altimeter
Air Data System with pressure and temperature
Jon Anderson
Total
79
Backup slides - Mass
C&DH
Flight Processor
Digital I/O - CAPI Board
State of Health and Attitude Control
Power Distribution (2)
Power Control
Mother Board
0.6
0.6
0.6
1.2
0.6
0.8
0.72
0.72
0.72
1.44
0.72
0.96
Power Converters (For Integrated Avionics Unit)
Chassis
Solid State Data Recorder
0.8
3.4
1.6
0.96
4.08
1.92
10.2
12.24
Airship Hull
Gondola*
Tail Section: 4 Fins and attachments*
Attitude Control
Helium Mass (Float at 5 km)
Inflation tank for Helium*
Bayonet fans and eqipment
4.57
8.4
8.4
4
29.95
19.17
5.5
5.484
10.08
10.08
4.8
35.94
23.004
6.6
Total
79.99
95.988
Inflight and during operation
8.27
9.924
Total
8.27
9.924
Total Airship Dry Mass
187.31
224.772
Total Aiship Float Mass
217.26
260.712
Total
Structure
Thermal
Jon Anderson
80
Backup slides
Power
Power Required after
Required (W) 20% Margin (W)
Component
Subsystem
Power
580 W Generated
Proplusion
Propeller/Engine
See Figure 2
See Figure 2
Total
See Figure 2
See Figure 2
Fans (2)
90
108
Total
90
108
Haze and Cloud Partical Detector
Mass Spectrometer
Panchromatic Visible Light Imager
20
28
10
Total
58
Bayonets
Science
Instruments
Communication UHF Transceiver
69.6
74.88
Total
74.8
Jon Anderson
89.76
81
Backup slides - Power
ACDS*
C&DH*
Sun Sensors
IMU
Radar Altimeter
Air Data System with pressure and temperature
0.56
22.2
37.6
7.72
Total
68.08
Flight Processor; >200 MIPS, AD750, cPCI
Digital I/O - CAPI Board
State of Health and Attitude Control - SMACI
Power Distribution
Power Control
Power Converters (For Integrated Avionics Unit)
Solid State Data Recorder
11.6
3.44
3.44
6.88
3.44
13.84
0.64
Total
43.28
Total Power Required without proplusion with all systems operating Straight and level
244.16
Total Power Available for Propulsion - Straight and level
335.84
Jon Anderson
82
References
1.
Ravi Prakash. Design of a Long Endurance Titan VTOL Vehicle.
Guggenheim School of Aerospace Engineering. 2006
2.
Jeffery L Hall. Titan Airship Explorer. JPL. 2002.
3.
Dr. Joel S. Levine. Titan Explorer: The Next Step in the Exploration of a
Mysterious World. NASA Langley Research Center. 2005
4.
Wolfram: The Mathematica Book, Wolfram Media, Inc., Fourth Edition,
1999
5.
Gradshteyn/Ryzhik: Table of Integrals, Series and Products, Academic
Press, Second Printing, 1981
Jon Anderson
83
Equations
Buoyancy and Volume equations:
B  (  atm   He ) g tianVAirship
 He 
patmM
RTatm
m  V
Shape and Surface Area equations:
4
V  ab 2 ,4 : 1
3

 3b 2  10b 2 2 
S  2 b 2  abr 1 
r 
56b 2



b
cos 1  
a
r
b2
1 2
a
Sources:
Stephen Hu
84
Equations
Drag and Reynolds number equations:
1/ 3
C DV , Hull
D 
Re 
l
.172 
d 

1. 2
d 
d 
 .252   1.302 
l
l
1/ 6
Re
1
U 2V 2 / 3C DV , Hull
2
V (2b)

Thrust and power available equations:
Powerrequired  DU foward
Stephen Hu
85
Diagram of Airship
Stephen Hu
86
Team Member
Stephen Hu
Deployment
Vehicle Selection
Helicopter Design
Hours Worked: 98
Stephen Hu
87
Helicopter Design
Introduction
General Characteristics
Constraints
Deployment
Conclusions
Recommendations
Stephen Hu
88
Introduction
1. Investigation of the surface and lakes of
Titan
2. VTOL capability
3. Dependable performance in hostile
environments
4. Able to last four months under constant
operation
Stephen Hu
89
Constraints
• Environment
–
–
–
–
Temperature
Wind
Solar Energy
Atmospheric Density
• Volume/Storage
– Diameter of Heat Shield
– Airship Storage
Stephen Hu
90
General Characteristics
Vehicle: Helicopter
Type: Coaxial
Number of Blades per rotor: 2 blades
Airfoil: NACA 0012
Stephen Hu
91
Deployment
• Post-Airship
Separation
• Generator
startup
• Freefall rotor
startup
• Heat Shield
separation
Stephen Hu
92
Mass and Power Constraints
Power Required to Hover (W)
1400
1200
1000
800
600
400
Power Required to Hover
200
Power Available
0
0
50
100
150
200
250
300
350
Mass (kg)
Stephen Hu
93
Blade Radius vs. Power Required
Power Required to Hover (W)
3000
m=155.4
2500
2000
m=186.5
1500
1000
500
0
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
Blade Radius (m)
Stephen Hu
94
Forward Velocity vs. Power Required
1400
Power Required (W)
1200
1000
800
600
m=155.4
400
m=186.5
200
Power Available
0
0
1
2
3
4
5
6
7
8
9
10
Forward Velocity (m/s)
Stephen Hu
95
Conclusions
Characteristics/Performance
Mass (kg)
Payload Mass (kg)
Expected
155.4
15.00
Contingency (20%)
186.5
18.00
Rotor Diameter (m)
Main Blade Chord (m)
Fuselage Length (m)
Fuselage Height (m)
Total Height (m)
1.364
0.091
2.56
0.77
1
1.426
0.095
2.56
0.77
Total Width (m)
Max Climb Rate (m/s)
Spinup Time (s)
Max Cruise Velocity (m/s)
0.8
2.292
7.961
7.46
Optimal Cruise Velocity (m/s)
Range (km)*
3.1
188.0
Altitude (km)
15
Stephen Hu
1
0.8
1.349
9.096
6.95
3.6
175.1
10
96
Recommendations
• More in-depth aerodynamic design
• Materials
• Payload
• Deployment
Stephen Hu
97
References
1. "The Vertical Profile of Winds on Titan." www.nature.com. 8
Dec. 2005. Nature: International Weekly Journal of Science.
<http://www.nature.com/nature/journal/v438/n7069/full/nature0
4060.html>.
2. Wright, Henry S. Design of a Long Endurance Titan VTOL
Vehicle. Georgia Institute of Technology.
<https://wiki.umn.edu/pub/titanedd/deployment/vtol_design_pa
per.pdf>.
3. Leishman, Gordon. Principles of Helicopter Aerodynamics.
Cambridge UP, 2006.
Stephen Hu
98
FMEA
• Aeroshell separation failure, probability low,
mission failure
• Helicopter failure, probability medium, some loss of
data, airship has some redundancy
•Airship failure, probability low, shorter data transfer
window, some instrument loss
• Parachute failure, probability low, possible mission
failure
Andrew Welsh
99
FMEA
•Helicopter one engine failure, probability low,
extremely decreased functionality
• Helicopter two engine failure, probability low,
helicopter failure
• Heat shield failure, probability low, mission failure
• Incorrect entry position, probability medium,
possible mission failure or wasted time for airship and
helicopter to reach intended position
Andrew Welsh
100
FMEA
Environmental, Societal, and Global
Impacts:
• Launch failure, environmental radiation
contamination or death, find alternate power source
or plant a tree
• Noisy launch, plant a tree
• Excessive exhaust from launch, plant a tree
Andrew Welsh
101
FMEA
• Life found on Titan causes panic, hide the truth,
prepare public for truth, military law
• Contaminate Titan with Earth organisms,
sterilization
• Scientific breakthroughs eliminate human jobs,
socialism
Andrew Welsh
102