Using LEO Depots to Enable Dedicated Interplanetary Smallsat Missions FISO Working Group Telecon Presentation – 28 November 2018 Jonathan Goff – Altius Space Machines, Inc. Mike Loucks and John Carrico – Space Exploration Engineering, Inc. FISO Telecon 11-28-18 Using LEO Depots to Enable Dedicated Interplanetary Smallsat Missions 1 Company Overviews Altius Space Machines (www.altius-space.com) • Satellite Servicing/Space Tug Development and (future) Operations • Satellite Servicing and In-Space Refueling Interfaces • Contract R&D and Prototype Development and Testing • Space Robotics and Spacecraft/Launch Vehicle Mechanisms • Electropermanent Magnetic Systems • Cryogenic Fluid Systems • Advanced Space Techonologies • Space Environmental Testing (TVAC, Vibe, HALT/HASS) Space Exploration Engineering (www.see.com) • Mission Concepts & Planning • Satellite Analysis & Operations • Technical Due Diligence • Astrogation and Training FISO Telecon 11-28-18 Using LEO Depots to Enable Dedicated Interplanetary Smallsat Missions 2 Entering an Age of Interplanetary SmallSat Missions (credit: NASA JPL) • Supporting traditional missions • More frequent missions to less-popular destinations • Relay/Observation Constellations • Ubiquitous solar system exploration missions FISO Telecon 11-28-18 Using LEO Depots to Enable Dedicated Interplanetary Smallsat Missions 3 Interplanetary Smallsat Transportation Challenges/Opportunities • Current Options for Interplanetary Smallsat Delivery • Interplanetary Hitchhiking • Integrated SEP or Chemical Propulsion • SEP or Chemical Kick Stages • Interplanetary hitchhiking missions are more affordable but constrain destinations options, integrated propulsion or kick stages give more flexibility, but at significantly higher $/kg of net delivered payload • LEO smallsat launcher depots may provide the best of both worlds • Fast smallsat missions with modest $/kg of net delivered payload • Much larger dedicated smallsat missions for a given launcher size • Dedicated missions to where you want to go, when you want to go • LEO-based smallsat launcher depots can also be a great pathfinder for larger or deeper-space depots • Get significant experience before trying to make cryogenic depots work in lunar or Martian orbit FISO Telecon 11-28-18 Using LEO Depots to Enable Dedicated Interplanetary Smallsat Missions 4 What is a Smallsat Launcher Depot? • LEO-based refueling station that refuels a smallsat launcher mission stack (smallsat launcher upper stage, kick stage, and possibly spacecraft) prior to a beyond-LEO departure FISO Telecon 11-28-18 Using LEO Depots to Enable Dedicated Interplanetary Smallsat Missions 5 Oasis Smallsat Launcher Depot Key Elements Smallsat Launcher Upper Stage w/ Kick Stage and Payload Repurposed Upper Stage Depot Kit • Primarily a LOX/Kerosene depot • Launch as a secondary payload (e.g. on a Cygnus delivery • Short-term LH2 storage, w/ LH2 used as a coolant for other fluids mission) plus repurposing tanks gives the maximum depot • Plus Storable propellants and GHe storage capacity for the minimum launch costs • “Depot Kit” launched as a secondary payload + repurposed upper stage tankage • >40mT LOX/Kero capacity depot, w/ ~10-12mT initial propellant load, launched for <<$50M • LOX and LH2 storage based on repurposed upper stage • Kero and storable/pressurant storage, fluid controls, depot avionics, • ISS co-orbiting depot location can enable very low-cost power/propulsion, and capture/refueling robotics all in depot kit • Depot kit takes over operations from upper stage on-orbit depot resupply options FISO Telecon 11-28-18 Using LEO Depots to Enable Dedicated Interplanetary Smallsat Missions 6 Oasis Depot Launch/Resupply CONOPS • For the initial deployment: • For resupply: • Oasis depot is launched as a secondary payload on an ISS • After an ISS cargo vehicle is dropped off on orbit, its upper resupply mission (Cygnus on Vulcan/Centaur V?) stage performs a burn to rendezvous with the depot • After ISS cargo vehicle departs, upper stage moves depot to • The stage either performs a direct rendezvous with the depot, desired operating orbit or a small prox-ops tug is used to bring it to the depot • Once in the desired orbit, the upper stage hands off fluid • The stage then transfers most of its excess propellant to the controls to the depot kit depot • The depot kit uses LH2 to subcool LOX and GHe and then to • The stage then departs the depot and performs a disposal burn maintain temperatures FISO Telecon 11-28-18 Using LEO Depots to Enable Dedicated Interplanetary Smallsat Missions 7 Oasis Interplanetary Smallsat Launch CONOPS • Smallsat Launcher Refueling Operations: • Interplanetary Departure Operations: • The launch vehicle + empty kick stage + payload are • The upper stage then performs the burn to enter the launched into LEO in a phasing orbit close to the depot phasing orbit of the 3-burn departure maneuver • The launch vehicle upper stage then performs the • The kick stage and payload then separate from the direct rendezvous maneuvers (or a tug can perform upper stage, and the kick stage performs the final two them) maneuvers to place the payload into its final trajectory • The depot captures and refuels the upper stage, and • At that point, if desired the payload can separate from fuels the kick stage to the proper levels of propellant the kick stage for the planned mission FISO Telecon 11-28-18 Using LEO Depots to Enable Dedicated Interplanetary Smallsat Missions 8 3-Burn Departures Methodology Rationale • In order to hit a desired hyperbolic departure vector, you must perform your final departure burn centered on a point at the “locus of periapses”, with your plane lined up so it intersects with the departure vector • In LEO orbital planes slowly precess due to the oblateness of the earth, causing the depot RAAN to drift by 5deg/day westward (for an ISS-like orbit) • The odds of the depot orbital plane lining up optimally for a traditional single-burn departure during an interplanetary departure window is not very good • The authors invented a 3-burn departure method to deal with this problem • Inspired by a departure methodology proposed by Kirk Sorensen to enable an equatorial MXER tether to send payloads to interplanetary destinations • This approach enables hitting departure declinations slightly higher than the inclination of the depot (depending on the departure C3). FISO Telecon 11-28-18 Using LEO Depots to Enable Dedicated Interplanetary Smallsat Missions 9 3-Burn Departure Methodology Overview 1. The spacecraft starts in the depot orbit (Red) 3. At apogee on the last phasing loop, the spacecraft performs a maneuver to cancel out an solar/lunar 2. Prior to the departure window, during an orbit that crosses the locus of periapses, the spacecraft performs a perturbations and to align its orbital trajectory (Pink) burn centered on the point where it intersects the locus, with the departure asymptote. entering a highly-elliptical phasing orbit (Green). The phasing orbit period is selected so that after an integer 4. When the spacecraft returns to perigee on the last loop, number of loops, the spacecraft returns to perigee at the it is aligned properly for its departure burn, entering the departure date. desired hyperbolic departure trajectory (White). FISO Telecon 11-28-18 Using LEO Depots to Enable Dedicated Interplanetary Smallsat Missions 10 3-Burn Departure Trajectory Design Methodology • Step One: Identify Ideal Departure Geometry • Departure time, Asymptote Ra and Dec, and Departure C3 (using pork-chop plots) • Angular extent of the locus of periapses – derived from C3 • Note: a 3-burn departure is always possible if absolute value of the declination is less than or equal to the depot inclination plus the angular extent of the locus of periapses • Step Two: Calculate Time/Orbit of Apogee Raise Maneuver • For departure declinations less than depot orbit inclination, this will occur on the last orbit prior to the departure window when the depot plane intersects the departure asymptote • For departure declinations within limits but greater than depot orbit inclination, this will occur at the last orbit prior to the departure date where the dot product of the orbit’s angular velocity vector and the departure asymptote are at a minimum • Step Three: Calculate Duration of Phasing Orbits and Number of Loops • The phasing loop duration x number of loops should equal the time between entering the phasing orbits and the desired departure date. Typically 1-6 loops is optimal. • More smaller loops tend to be less susceptible to lunar/solar perturbations, but more susceptible to nodal precession and tend to require more DV for plane changes. • Step Four: Calculate Plane Change/Perturbation Correction Maneuvers • This will typically require a targeting program to get right, and may require multiple small burns to avoid having the perigee dip too low. • Step Five: Calculate Final Departure Burn FISO Telecon 11-28-18 Using LEO Depots to Enable Dedicated Interplanetary Smallsat Missions 11 3-Burn Departure Penalties Quantification • In order to quantify the penalties associated with using the 3-burn departure methodology, a RAAN-sweep analysis was performed using a Mars departure • The DV for a single-burn departure was calculated (3679 m/s), and then the DV and phasing orbit duration for the 3-burn maneuver was calculated at different initial depot RAANs • The worst case DV penalty was 103 m/s and the worst case phasing duration was 52 days. • For human spaceflight missions the crew