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.