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Strategies for mars network missions via an alternative entry, descent, and landing architecture

10 th International planetary probe workshop. Strategies for mars network missions via an alternative entry, descent, and landing architecture. 17-21 June, 2013; San Jose State University, CA, United States. Sarag J. Saikia, Blake Rogers, James M. Longuski

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Strategies for mars network missions via an alternative entry, descent, and landing architecture

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  1. 10th International planetary probe workshop Strategies for mars network missions via an alternative entry, descent, and landing architecture • 17-21 June, 2013; San Jose State University, CA, United States Sarag J. Saikia, Blake Rogers, James M. Longuski School of Aeronautics and Astronautics, Purdue University

  2. Mission concept and architecture • GOAL: Deliver four Mars Phoenix-class landers with a minimum separation of 3,000 km via a single launch from Earth

  3. Launch and cruise configuration Payload mass of 60 kg (x4) Flight system mass of 1380 kg + 2% reserve (x2)

  4. Single-event drag modulation

  5. Ballistic Coefficient Analysis

  6. Interplanetary Trajectory Trajectory Constraints • Atlas V 541 Launch Vehicle • Maximum launch V∞ for a mass of 1382 kg ≈ 7 km/s • Maximum entry speed of 6 km/s • Reduces the heating rates and heat loads of EDL • Corresponds to a maximum arrival V∞ of ≈ 3.5 km/s

  7. Interplanetary Trajectory 7 Day Separation: Low Thrust

  8. Range separation and divert capability Global Reach Capability

  9. Monte carlo results: Range Separation Landing Error Flight System 1 Flight System 2

  10. Stagnation-point heating rate

  11. Deceleration, release

  12. Monte carlo results Spacecraft Range and range separation Distribution

  13. Monte carlo results Primary Spacecraft Range distribution

  14. Monte carlo results Integrated Heat load: Primary Spacecraft

  15. Monte carlo results Integrated Heat load: Secondary Spacecraft

  16. Other Potential applications • Primary heat load for all the cases is < 2200 J/cm2 • Primary landing 3-σerror for all the cases is < ±10 km

  17. conclusions • Low-Thrust Propulsion represents an attractive ‘augmentation’ for any future mission to Mars • Benign aerothermal environments, reduced heat rates and loads • Very low ballistic coefficient achievable: no supersonic decelerator (parachute) required • Increased risks of separation: flight systems, spacecraft from a flight system • Mass increase: due to extra spacecraft adapter; Decrease due to reduction in cruise stages and supersonic parachutes • Other potential applications of multiple spacecraft lander/orbiter missions • Single Atlas V 541 launch required • Incorporation of the guidance on the second will reduce the landing error

  18. acknowledgment • Thanks to the IPPW10 student organizing committee for providing the ‘generous’ scholarship to attend the workshop

  19. Questions?

  20. backup

  21. Back up Instruments Mission Mass Breakdown • Mass Breakdown • Payload mass of 60 kg (same as Phoenix mission) • Mass of flight system is 1380 kg + 2% reserve

  22. Monte carlo results Integrated Heat load % TPS mass is estimated using an empirical formula based on previous probe missions Slightly high TPA mass for primary, and lower for secondary Total range separation requirement is the determinant of % TPS mass of secondary For low range requirements (<500km) secondary needs no TPS mass at all!

  23. Analysis of Drag Modulation Combine with the previous slide #4

  24. Drag skirt options Heat Shield Extension Rigid Deployable Decelerator (ADEPT) Hypersonic Inflatable Aerodynamic Decelerator (HIAD)

  25. Uncertainty analysis: Monte carlo Simulation uncertainly model parameters and input

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