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Ae105c Term Project Design Review MA’AM June 7, 2007

Ae105c Term Project Design Review MA’AM June 7, 2007. Mars Advanced Attack Mission. Teams. Trajectory Paul Hebert Ashley Moore Jack Ziegler Structures and Configuration Devvrath Khatri Francisco L ópez Jiménez Celia Reina Romo Thermal Annamarie Askren Philipp Boettcher Angie Capece

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Ae105c Term Project Design Review MA’AM June 7, 2007

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  1. Ae105c Term Project Design ReviewMA’AMJune 7, 2007 Mars Advanced Attack Mission 1

  2. Teams • Trajectory • Paul Hebert • Ashley Moore • Jack Ziegler • Structures and Configuration • Devvrath Khatri • Francisco López Jiménez • Celia Reina Romo • Thermal • Annamarie Askren • Philipp Boettcher • Angie Capece • Olive Stohlman 2

  3. Review Board Greg Davis Rob Manning Jay Polk Marco Quadrelli Mike Watkins Paul Dimotakis 3

  4. Mission Overview and Requirements If JPL could land 200 kg on MARS, can we land 20 kg?

  5. Outline • Introduction • Analytical Work • Trajectory • Thermal • Structures & Configuration • Conclusion

  6. Mission Overview and Requirements Mission Statement: Design a trajectory, thermal system, and structure system for a Mars Network Lander that safely delivers a 20 kg payload to Nili Fossae. Level 1 requirements 6

  7. Payload Overview and Requirements The payload allows visualization of the Martian environment, analysis of Mars’ atmosphere and weather patterns, as well as the on-going search for water on Mars. 7

  8. Design Assumptions • Used previous work (MER, Pathfinder) as design starting point • Scaled MER in mass and volume to match our design requirements • Verified our model results with data from these missions • Used first-order approximations to model the design and make design decisions 8

  9. System Evolution • Satisfy the requirements • Reliable and affordable Viking / MER heritage • Delta II launch vehicle: maximum aeroshell diameter ≈ 2.5 m General ideas for the design 9

  10. Final Configuration BACKSHELL PARACHUTE AND CANISTER PAYLOAD PAYLOAD PROTECTION HEATSHIELD 10

  11. Final Configuration 0.845 m 70 deg 1.10 m 11

  12. EDL Sequence 12

  13. EDL Sequence 13

  14. Final Configuration Payload Crushable material Payload Separation Bolts 14

  15. Notional Design of Instrument Setup • Partially Isotropic Payload • Telescoping Weather Boom • Visible Wavelength Camera with Mirror System • Surface Water Detection System Cut View of P/L 15

  16. Trajectory Paul Hebert, Ashley Moore, Jack Ziegler Presented by: Jack Ziegler 16

  17. Trajectory Overview • Tasks • Model entry vehicle using Matlab, separation to impact • Disperse nominal trajectory for landing ellipse • Supply to thermal team • As Functions of time: • Altitude • Angle of attack • Free stream velocity • Density Profile • Dynamic pressure • Supply to structures team • Terminal velocity and orientation at impact • Body accelerations in flight • Landing Ellipse Requirements • Along-track error +/- 50km • Cross-track error +/- 5km • Receivables • Vehicle shape • Center of mass, inertias • Aerodynamic models • Parachute size • Heat shield mass 17

  18. Dynamic Model Assumptions • Aerodynamic coefficients (Lift, Drag, Moments) • Hypersonic Model (M>5) • Cd: Supersonic and transonic models • Subsonic Mach # gaps, using MER data • Parachute • Deploy at Mach 1.8, • ->Cd 0.4, radius 7.5m • (linearly increases in 2 secs) • Mass loss • Drop 45 kg mass at • Mach 0.95 • Spherical planet model (neglect oblatness, rocks, mountains, etc.) • Rigid atmosphere/planet model • (atmosphere moves with planet) • No wind gusts • Two density models • (avg. Mars day) • Simple exponential: h=125-55km, Table: h<55km • Uniform gas properties (C02) 18

  19. Basic Code Structure 19

  20. Nominal Non-Dispersed Initial Conditions Altitude: H = 125 km Latitude:  = 13.67o Longitude:  = 73.8o Bank Angle = 0o Flight Path Angle = -10o Heading Angle = 0o X velocity: U = 4.5 km/s Y velocity: v = 0 km/s Z velocity: W = 0 km/s Roll rate: p = 0 deg/s Pitch rate: q = 0 deg/s Yaw rate: r = 0 deg/s • 12 Dynamic • Degrees of Freedom • Translation • Orientation • Linear Velocity • Angular Velocity Enters Atmosphere from S  N Polar Orbit Similar velocity &flight path angle of Viking Missions 20

  21. Dispersion Motivation and Assumptions Created Dispersion Analysis (Monte Carlo) - Models the uncertainties and errors - 500 random simulations Dispersed (perturbed) initial states and parameters - Values similar to MER dispersion Random Gaussian Dispersion - Gaussian distributions - Centered at nominal with 3 max Created landing ellipse - Obtained from all dispersed locations of impact. Notes: Did not calculate the 3 ellipse (enclosed all points) Approximate size of ellipse calculated from all impact locations projected onto a plane in the vicinity of Nili Fossae 21

  22. Description of the Monte-Carlo code Values taken mostly from “MER EDL Trajectory Analysis”, N. Desai 22

  23. Dispersion Results Impact Velocity Distribution Number of Cases Impact Velocity 23

  24. Dispersion Results 24

  25. Altitude - Quick descent in low density region - Increasing deceleration with exponential increase in density 25

  26. Velocity of Free Stream Nominal Impact Velocity = 24 m/s Parachute deploys 26

  27. Orientation Note the switch from fast to slow oscillations at parachute deployment 27

  28. States from Numerical Integration Output from Matlab ode15s (“stiff” adaptive time step integrator) 28

  29. Aerodynamic Drag and Lift Cd fairly const. for M > 5 Increases near Transonic 0.4 for parachute note spike when parachute deploys CL function of Angle of attack 29

  30. Aerodynamic Side Force and Moments 30

  31. Latitude and Longitude Approaching the Landing Site: Nili Fossae (latitude=20.93N, longitude=74.35E) 31

  32. Trajectory Relative to Inertial Frame Time = 0 to 50 sec Non-Rotating Inertial Frame at center of planet Entry Vehicle displayed at constant time intervals 32

  33. Trajectory Relative to Inertial Frame Time = 50 to 100 sec 33

  34. Trajectory Relative to Inertial Frame Time = 100 to 150 sec Note the rapidly decreasing flight path angle 34

  35. Landing Ellipse • 500 Randomly Dispersed Trajectories • Nili Fossae 47.7 km 20.6 km 35

  36. Summary of Nominal Trajectory Results Impact Velocity V = 24 m/s t=~150 secs Max in flight acc. ax = 20g, ay = 10g, az = 60g (Off b/c of hypersonic aero. moments invalid) Landing site Altitude: h = 1 km Latitude:  = 74.35 Longitude:  = 20.93 Landing Ellipse Major axis: 47.7 km Minor axis: 20.6 km • Dispersion References • Desai, P. N. and Knocke, P. C. “Mars Exploration Rovers Entry,Descent, and Landing Trajectory Analysis.” AIAA. • Desai, P. N., Schoenenberger, M. and Cheatwood, F. M. “Mars Exploration Rover Six-Degree-Of-Freedom Entry Trajectory Analysis,” Proceedings of the AAS/AIAA Astrodynamics Specialists Conference, August 3-7, 2003, Big Sky Resort, Big Sky, MT, AAS 03-642. 36

  37. Thermal Annamarie Askren, Philipp Boettcher, Angie Capece, Olive Stohlman Presented by: Olive Stohlman 37

  38. Thermal Overview Requirement: Maintain the P/L between 5˚C and 35˚C Approach: Determine the dynamic and integrated heat loads Design the ablator Volume Shape Materials Determine the heat load to the structure and P/L Prevent inner shell materials from melting Sustain attachment mechanism (bondline temperature) 38

  39. Assumptions • Reliance on previous work (MER, Pathfinder, …) for cross-checking our results • 1-D numerical analysis to determine conductive heating of payload 39

  40. Model Comparisons • Regan • (1984) • Corning • (1964) • Hankey • (1988) • Sutton • (1971) • Tauber (radiative)

  41. Convective and Radiative Heat Loads Pathfinder 4.5 km/s 5.6 km/s Our Design Convective Heat Rate: 260,000 W/m2 Radiative Heat Rate: 20,000 W/m2 Used chart above to verify analysis by cross-checking with MPF 41

  42. Mars Re-entry Mission Data Wright, M.J., Edquist, K.T., Hollis, B.R., Brown, J.L., Olejniczak, J., “A Review of Aerothermal Modeling for Current and Future Mars Entry Missions,” Draft paper for submission to AIAA Journal of Thermophysics and Heat Transfer – February 2007

  43. Ablation Design Ablation Material: SLA561V* Effective Heat of Ablation: 5.41 x 107 J/kg Material Density: 264.3 kg/m3 Specific Heat Cp: 1.16 x 103 J/(kg-K) Nominal Ablation Thickness: 1.6 cm (using Sutton) Constant Ablation Profile *TPSX Materials Properties Database. NASA. http://tpsx.arc.nasa.gov/ 43

  44. 1-D Heat Conduction Analysis Assumes: Simplified payload-heatshield interface Upper bound on heat load Sutton + 30% Entire structure begins at payload minimum temperature (5˚C) Lower than true value of heat of ablation (80% of theoretical values) for SLA 561 Unknown: • Actual temperature of ablation of SLA 561 • Assumed 4000 K (Quartz: 2700 K) • Char layer behavior • Assumed no char layer left 44

  45. 1-D Heat Conduction Analysis Finite Element Model Confirms glue line temperature under 250˚C (220˚C), payload temperature under 30˚C 1-D Position vs. Temperature 1-D Position vs. Temperature (Zoom) Temperature [K] Temperature [K] Distance from Payload [m] Distance from Payload [m] 45

  46. Thermal Final Design Total Heat Load 1.1 x 107 J Ablator Material: SLA561 Thickness: 1.6 cm Safety Factor: 2 Mass: 4.3 kg Bondline Temperature 230 ºC Payload Temperature 6 ºC 46

  47. Structures & Configuration Devvrath Khatri, Francisco López Jiménez, Celia Reina Romo Presented by: Francisco López Jiménez 47

  48. Impact calculation Honeycomb as crushable material: uniform, predictable and efficient *Hexcel energy absorption systems Assumptions: Stroke 80 % of honeycomb thickness Type Cross-Core: multi-directional energy absorption Precrushing to eliminate peak load Constant force: 700 g 48

  49. Impact calculation Calculation kinetic energy energy absorbed before impact by crushable Nominal 3-sigma Note: solution for elastic (harmonic response) = 49

  50. Impact calculation Final design crushable material 1/4 -5052-5.2 CRIII Aluminum HexWeb (1300 kPa) 10 cm thickness 200 % margin for nominal case in stroke 25 % margin for 3-sigma case in stroke maximum load 700g 40 % margin 50

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