1 Titan EDD Fall Semester 12/02/2008 Andrew Welsh Jon Anderson Nick Delucca Steve Hu Travis Noffke Pawel Swica CDR Andrew Welsh

Team Lead Hours Worked: Team Member Andrew Welsh

Agenda 3 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 3Andrew Welsh

4 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

5 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

6 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

7 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

8 Requirements Major Tasks (Added): Detailed aerial vehicle selection EDD animation Bad weather simulation Andrew Welsh

9 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 Dep. Team: Jon Anderson, Steve Hu Andrew Welsh

10 Approach Research Aerial Vehicle Selection Descent Method Selection Entry Method Selection Previous Research Our Research and Calculations Final Product Andrew Welsh

11 Program Plan Gantt Chart Andrew Welsh

12 Program Plan Task list with responsible engineers, status, hours worked, total hours worked Andrew Welsh

13 Design Walkthrough

14 References htmhttp:// htm 2. tanExplorer.htmhttps:// tanExplorer.htm Andrew Welsh

Pawel Swica15 Pawel Swica Entry/Integration Hours Worked: Team Member

Pawel Swica16 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 Swica17 Relevant Equations

Pawel Swica18 Orbit Diagrams

Pawel Swica19 Orbit Diagrams

Pawel Swica20 Orbit Diagrams

Pawel Swica21 Resulting trade study

Pawel Swica22 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 Swica23 Relevant Equations

Pawel Swica24 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 4. Entry heating graph Figure 5. Aeroshell model

Pawel Swica25 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 º down at entry interface (1000 km) with 5º margin in either direction Modeling in simulink roughly agreed with these values

Pawel Swica26 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 Swica27 Relevant Calculations

Pawel Swica28 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 Swica29 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

Vpython Video Pawel Swica30

Team Member Nick De Lucca Titan Atmosphere Entry Simulation 113 Hours Nick De Lucca31

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 correspondence. Nick De Lucca32

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 Lucca33

Newtonian Aerodynamics Formulas: Nick De Lucca34

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 Lucca35

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 Lucca36

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 Lucca37

Results Total time taken: 3947 seconds Peak deceleration: 86.6 Maximum heating rate: 146 W Total heat Transferred: J Nick De Lucca38

Results Nick De Lucca39

Results Nick De Lucca40

Results Nick De Lucca41

Results Stephen Hu42

Results Nick De Lucca43

Methods for improvement Non-Newtonian aerodynamics CFD for the heat shield Nick De Lucca44

References Dynamics and Thermodynamics of Planetary Entry. W.H.T. Loh. Prentice-Hall Space Technology Series Kazeminejad et al. Temperature Variations in Titan's Upper Atmosphere: Impact on Cassini/Huygens. Annales Geophysicae 23. pp Stephen Hu45

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: /19/200846Travis Noffke

Primary Goals 1.Provide deceleration force to top of aeroshell 2.Clearance of unnecessary system components 3.Minimize payload descent stability disruption 11/19/200847Travis Noffke

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 11/19/200848Travis Noffke 10 meters Payload Containment

System Events 1.Aeroshell separation mechanism fires 2.Mortar fires deployment bag 3.Parachute inflates 4.Deceleration on top aeroshell 5.Complete separation 11/19/200849Travis Noffke Payload Containment V payload V top

Design Tasks Geometry Selection Parachute Characterization Sizing Opening Forces Loading Mass Ratio Ballistic Coefficient Material Selection Canopy Design 11/19/200850Travis Noffke Conical Ribbon 1.5 m Diameter nom Kevlar© 29

Drag Generation and Time 11/19/2008Travis Noffke51 v terminal = 12 m/s c D = 0.50 Altitude = 8 km Atmospheric Density (ρ)= kg/m^3

Opening Forces 11/19/200852Travis Noffke

Separation Mechanics 11/19/200853Travis Noffke Figure 5. Separation Spring Concept Example of separation mechanism Commonly used in spacecraft Performance must meet requirements One of several methods

Canopy Structure 11/19/200854Travis Noffke

Canopy Material Kevlar© 29 Highest strength-weight ratio Superior tensile strength Space tested Used in heritage systems 11/19/200855Travis Noffke

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 11/19/2008Travis Noffke56

Additional Information (Backup Slides) RocketsParachutes Aero-control Surfaces Complexity:HighLowMed Cost:HighLowHigh Risk:HighLowMed Efficiency:LowHigh AvailabilityMedHighMed EffectivenessHigh Med 11/19/2008Travis Noffke57 Deceleration Method Trade Study

Parachute Geometry Study 11/19/2008Travis Noffke58 Additional Information (Backup Slides) Characterization Conical RibbonDisc-Gap-Band Reliability:High Mass:LessMore Average Oscillation Angle±3°±10 to 15° Drag Coefficient Range.5 to.6.52 to.58 Opening Force Coefficient1.05 to Performance Comparison Conical RibbonDisc-Gap-Band Reliability: + + Mass: + - Average Oscillation Angle: + - Drag Coefficient Range: + + Opening Force Coefficient: + -

Cluster Single Conical Ribbon Parachute Reliability:High Difficulty:MedLow Redundancy:HighLow StabilityHigh+High- Cost:HighLow Mass:HighMed 11/19/2008Travis Noffke59 Additional Information (Backup Slides) Parachute Cluster Configuration Trade Study Andy Welsh

Rocket Assisted Separation Parachute Separation Complexity:HighMed Cost:HighLow Risk:MedLow Effectiveness:HighMed- 11/19/2008Travis Noffke60 Additional Information (Backup Slides) Parachute Avoidance Trade Study Andy Welsh

11/19/2008Travis Noffke61 Additional Information (Backup Slides) Used Derivations:

11/19/2008Travis Noffke62 Additional Information (Backup Slides)

Jon Anderson Hours Worked: Team Member Jon Anderson

Outline 64 Outline: Objective Vehicle selection Airship Design Design constraints Assumptions General design Performance Deployment Enabling technologies Recommendation and conclusion 64Jon Anderson

65 Objective Jon Anderson 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.

66 Performance Jon Anderson Float mass195 Kg Operational Cruse Velocity2.5 m/s Max Velocity2.98 m/s Min Climb/Descent Rate *50 m/min Range36200 km Service Ceiling5 km Absolute Ceiling40 km Estimated Lifetime *150 days Length13.83 m Width3.45 m Volume34.47 m Mass He Needed (10% reserve)27.4 kg

67 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

68 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

69 Vehicle Selection Jon Anderson CategoryHelicopterAirship Helicopter & Airship Combination Mass320 kg490 kgUNK kg Technology Development Needs MediumLowerHigh Operational Risk MediumLowMedium Environmental Tolerance MediumHighMedium Surface Capability HighMediumHigh Mission Completion Probability Medium High Deployment Ability MediumHigh

70 Design Constraints Jon Anderson Communication payload Extra redundancy – orbiter and helicopter Science payload Propulsion subsystem Mass assumptions – initial starting value Power subsystem MMRGT

71 Assumptions Jon Anderson 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

72 Reynolds # and Drag vs Velocity Jon Anderson

73 Power Required/Available vs Velocity Jon Anderson

74 Inflation time/percent vs Lift Jon Anderson

75 Deployment Jon Anderson Airship inflation immediate Both bayonets and main envelope Changing ballistic coefficient Separate via explosive shearing bolts Immediately max velocity

76 Enabling Technologies Jon Anderson 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

77 Recommendation and Conclusion Jon Anderson High Altitude Design Detailed data bandwidth analysis Hull/system optimization Experments

78 Questions? Jon Anderson

79 Backup slides - Mass Jon Anderson ComponentMass (kg)Mass after 20% Margin (kg) Subsystem Power2nd Generation MMRTG Battery - 12 A h lithium Turbomachinery Turbine Compressor Piping Electric Motor Alternator Total PropulsionPropeller, axel, case* Total Science InstrumentsHaze and Cloud Partical Detector33.6 Mass Spectrometer1012 Panchromatic Visible Light Imager Total CommunicationX-Band Omni - LGA SDST X-up/X-down X-Band TWTA UHF Transceiver (2) UHF Omni UHF Diplexer (2)11.2 Additional Hardware (switches, cables, etc.)67.2 Total ACDSSun Sensors IMU (2)910.8 Radar Altimeter Antennas for Radar Altimeter Absorber for Radar Altimeter Air Data System with pressure and temperature56 Total2024

80 Backup slides - Mass Jon Anderson C&DHFlight Processor Digital I/O - CAPI Board State of Health and Attitude Control Power Distribution (2) Power Control Mother Board Power Converters (For Integrated Avionics Unit) Chassis Solid State Data Recorder Total StructureAirship Hull Gondola* Tail Section: 4 Fins and attachments* Attitude Control44.8 Helium Mass (Float at 5 km) Inflation tank for Helium* Bayonet fans and eqipment Total ThermalInflight and during operation Total Total Airship Dry Mass Total Aiship Float Mass

81 Backup slides Component Power Required (W) Power Required after 20% Margin (W) Subsystem Power580 W Generated ProplusionPropeller/EngineSee Figure 2 TotalSee Figure 2 BayonetsFans (2)90108 Total90108 Science InstrumentsHaze and Cloud Partical Detector20 Mass Spectrometer28 Panchromatic Visible Light Imager10 Total CommunicationUHF Transceiver74.88 Total Jon Anderson

82 Backup slides - Power Jon Anderson ACDS*Sun Sensors 0.56 IMU22.2 Radar Altimeter37.6 Air Data System with pressure and temperature7.72 Total68.08 C&DH*Flight Processor; >200 MIPS, AD750, cPCI11.6 Digital I/O - CAPI Board3.44 State of Health and Attitude Control - SMACI3.44 Power Distribution6.88 Power Control3.44 Power Converters (For Integrated Avionics Unit)13.84 Solid State Data Recorder0.64 Total43.28 Total Power Required without proplusion with all systems operating - Straight and level Total Power Available for Propulsion - Straight and level335.84

83 References 1.Ravi Prakash. Design of a Long Endurance Titan VTOL Vehicle. Guggenheim School of Aerospace Engineering Jeffery L Hall. Titan Airship Explorer. JPL Dr. Joel S. Levine. Titan Explorer: The Next Step in the Exploration of a Mysterious World. NASA Langley Research Center Wolfram: The Mathematica Book, Wolfram Media, Inc., Fourth Edition, Gradshteyn/Ryzhik: Table of Integrals, Series and Products, Academic Press, Second Printing, 1981 Jon Anderson

84 Equations Stephen Hu Buoyancy and Volume equations: Shape and Surface Area equations: Sources:

85 Equations Stephen Hu Drag and Reynolds number equations: Thrust and power available equations:

86 Diagram of Airship Stephen Hu

Deployment Vehicle Selection Helicopter Design Hours Worked: Team Member Stephen Hu

Helicopter Design 88 Introduction General Characteristics Constraints Deployment Conclusions Recommendations 88Stephen Hu

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 89Stephen Hu

Constraints Environment –Temperature –Wind –Solar Energy –Atmospheric Density Volume/Storage –Diameter of Heat Shield –Airship Storage Stephen Hu90

General Characteristics Vehicle: Helicopter Type: Coaxial Number of Blades per rotor: 2 blades Airfoil: NACA Stephen Hu

Deployment Post-Airship Separation Generator startup Freefall rotor startup Heat Shield separation 92Stephen Hu

Mass and Power Constraints 93Stephen Hu

Blade Radius vs. Power Required 94Stephen Hu

Forward Velocity vs. Power Required 95Stephen Hu

96 Characteristics/PerformanceExpectedContingency (20%) Mass (kg) Payload Mass (kg) Rotor Diameter (m) Main Blade Chord (m) Fuselage Length (m)2.56 Fuselage Height (m)0.77 Total Height (m)1 1 Total Width (m)0.8 Max Climb Rate (m/s) Spinup Time (s) Max Cruise Velocity (m/s) Optimal Cruise Velocity (m/s) Range (km)* Altitude (km)15 10 Conclusions Stephen Hu

Recommendations More in-depth aerodynamic design Materials Payload Deployment 97Stephen Hu

References 1."The Vertical Profile of Winds on Titan." 8 Dec Nature: International Weekly Journal of Science.. 2.Wright, Henry S. Design of a Long Endurance Titan VTOL Vehicle. Georgia Institute of Technology.. 3.Leishman, Gordon. Principles of Helicopter Aerodynamics. Cambridge UP, Stephen Hu

99 FMEA Andrew Welsh 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

100 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 FMEA Andrew Welsh

101Andrew Welsh 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

102 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 FMEA Andrew Welsh