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Development of an Earth Smallsat Flight Test to Demonstrate Viability of Mars Aerocapture

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Development of an Earth Smallsat Flight Test to Demonstrate Viability of Mars Aerocapture SmallSat Aerocapture Demonstration Development of an Earth SmallSat Flight Test to Demonstrate Viability of Mars Aerocapture Michael Werner / Bryce Woollard / Anirudh Tadanki / Swapnil Pujari / Dr. Robert Braun Space Systems Design Laboratory, Georgia Institute of Technology Rob Lock / Adam Nelessen / Ryan Woolley Jet Propulsion Laboratory, California Institute of Technology Aerocapture Benefits & Control Flight System Overview TrajectoryTrajectory Analysis Modeling Benefits of Aerocapture Aeroshell Design L Trajectory Modeling • Minimal propellant requirements • Aeroshell and drag skirt made from machined R • End-to-end simulation models • Potential large savings in mass and cost Example Aerocapture (“A/C”) Mass Savings aluminum vehicle trajectory and Skirt Jettison • Studies have shown conceptual & • PICA selected as TPS material uncertainties Entry Exit Mission A/C Mass (kg) Non-A/C Mass (kg) % Increase r technical viability ψ • 3-DOF integration of Venus Low-Circular Orbit 5078 2834 79 β β2 atmospheric equations of 1 Lack of flight demonstration inhibits Venus Elliptical Orbit 5078 3542 43 Example Trajectory Outputs Aeroshell & Drag Skirt Properties motion Nominal Atmospheric Trajectory Plots use of aerocapture on actual missions Mars Low Circular Orbit 5232 4556 15 Location Dimension Value • Key outputs: Altitude Jupiter Low Circular Orbit 2262 N/A Enabling Drag Skirt Outer Radius (R) 25 cm Aerocapture Control: Drag Modulation • Deceleration (km) • Ballistic coefficient changes used to alter Saturn Circular Orbit 494 N/A Enabling Blunted Cone Angle 60° • Heat Rates Inner Radius (r) 10 cm vehicle’s trajectory Titan Circular Orbit 2630 691 280 Foreshell • Orbit uncertainties Nose Radius 5.625 cm • Simple to implement compared to Deceleration Uranus Elliptical Orbit 1966 618 218 Shoulder Radius 1 cm traditional lifting methods: (g’s) Neptune Elliptical Orbit 1680 180 832 Length (L) 22.7 cm • No CG offset required Backshell Backshell Taper (ψ) 5.3° • Simple avionics algorithms Adapted from: Hall et al, “Cost-Benefit Analysis of the Aerocapture Mission Set”, Journal of Spacecraft & Rockets, 2005 Heat Rate So easy, a SmallSatcan do it (W/cm2) Discrete Single-Event Drag Modulation Drag Skirt & Jettison Mechanism Planet-Relative Velocity (km/s) • Drag skirt broken into 4 quadrants • Jettison accomplished via 3 staggered M3 Control Scheme pyrotechnic bolts per quadrant • Jettison timing controls amount of End-to-End Monte Carlo Results: 10000 Samples, Target = 1760 km • Spring loaded at interface to ensure safe aerocapture ΔV detachment • Guidance algorithm: numerical predictor-corrector • Only atmospheric accelerometer measurements required • Monte Carlo results show robustness to uncertainties in entry FPA, atmosphere, and burns Number of Samples Mission Concept Static Stability Static Margin Timeline Pre-Jettison Configuration Time (s) Static Margin (cm) • CG Location: 7.65 cm aft of nose vertex Monte Carlo Summary Purpose: Develop SmallSat mission concept for flight test demonstrating aerocapture at Earth 0 N/A • Backshell is shadowed from hypersonic flow by drag Parameter Value Philosophy: Prioritize simplicity & affordability Final Apogee Altitude (km) skirt 50 63.42 Mean 1746.0 km Requirements: Pre-Jettison • Ample static margin and stability 1. Aerocapture ΔV: ~2 km/s 100 58.99 3σ Error 169.4 km Jettison guidance algorithm is able to 2. Successful raise maneuver 150 58.64 Minimum 1596.1 km effectively target desired orbit Post-Jettison 3. Telemetry data return Maximum 1813.7 km • 150 7.98 Deliverable: Fully-documented mission concept in anticipation of future SmallSat proposal opportunities CG Location: 10.6 cm aft of nose vertex • Exposure of backshell to flow leads to reduced static 200 8.52 Mission Timeline stability Post-Jettison 250 11.34 Conclusions # Event Details 5. Atmospheric Interface 300 -3.31 1 GTO Period: 10 hrs 4. Descent Orbit Takeaways 350 -243.88 • A SmallSat mission featuring single-event jettison drag modulation is a promising architecture to Separation from 2 1 hr before apogee demonstrate a ~2 km/s aerocapture maneuver at Earth Host 9. Final Orbit • Structural and mission design analyses have led to a <25 kg flight system featuring COTS hardware Perigee Lowering Nominal Flight System 3 ΔV: ~10 m/s • Trajectory analyses and Monte Carlo results show robustness of system design to uncertainties Maneuver (PLM) 8. Perigee Raise • Aerojet MPS-120XL Hydrazine Propulsion 3. Perigee Maneuver • Sensonor IMU 4 Descent Orbit Duration: 5 hrs Lowering • ISIS On-Board Computer Atmospheric Altitude: 125 km Maneuver 5 • GomSpace EPS & Batteries Interface Velocity: 10.3 km/s 6. Aerocapture 7. Post Aerocapture Orbit • Astrodev UHF Radio For more information: 6 Aerocapture ΔV: ~2000 m/s • GomSpace UHF Dipole Antenna Post-Aerocapture Michael Werner 7 60 km x 1750 km 2. Separation Orbit 1. Geosynchronous [email protected] Transfer Orbit (GTO) from Host Perigee Raise 8 ΔV: ~40 m/s Maneuver (PRM) GTO to LEO A <25 kg flight system based on COTS CubeSat 9 Final Orbit 200km x 1750 km Aerocapture ΔV: 2 km/s hardware meets all mission requirements.
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