Overview of the System Accommodation and Mission Design for a Fission Powered Titan Saturn System Mission (TSSM) Robert L. Cataldo1, Steve R. Oleson1, Young H. Lee2, 1NASA Glenn Research Center, 21000 Brookpark Rd, Cleveland, OH 44135, 2Jet Propulsion Laboratory, 4800 Oak Grove Dr, Pasadena, CA 91009

In 2008, NASA’s Jet Propulsion Laboratory (JPL) Team X developed a conceptual design for a mission that would explore the Saturn’s moon Titan using an orbiter, a Montgolfier balloon, and a lander. Powered by five (4 plus one spare) Advanced Stirling Radioisotope Generator (ASRG) systems, the orbiter would carry over 100 kg of science instruments and 800 kg of Montgolfier/lander payloads to the Saturn system in seven years. Maneuvering the Titan Saturn System Mission (TSSM) [1] vehicle to get to Titan’s orbit utilized three different propulsive techniques: a solar electric propulsion stage (SEP) with Earth and Venus flybys to ‘pump-up’ the orbit to reach Saturn (which is jettisoned after the last Earth flyby), a bipropellant chemical system used to capture and maneuver around the Saturn gravity well, and an aerocapture campaign which would greatly reduce the chemical propellant mass needed for Titan capture, and would provide additional atmospheric science. The NASA Glenn Research Center’s COMPASS Team participated in the 2014 NASA Nuclear Power Assessment Study (NPAS) and was tasked to determine if a mission solution existed if the 2008 TSSM proposed ~500 W ASRG systems were replaced with a higher power 1 kW fission power system (FPS), using either Stirling or thermoelectric power conversion systems[2]. The ground rules stated that the reactor could not be used to power electric propulsion (in order to show how reactor power without electric propulsion could be implemented), a fixed boom should be used (to avoid the risk of deployable booms), and the baseline launch vehicle should be the Atlas 551 to be consistent with the baseline TSSM concept and in addition, to minimize nuclear launch costs since the Atlas LV has nuclear launch legacy having been utilized for the most recent radioisotope powered missions.

Mission Accommodations With adding the reactor power system, two major configuration changes were made to the TSSM 2008 concept: a deployable 2.25 meter antenna and a 4.5 meter diameter drag flap for Titan aerobraking. Increasing the power level from ~500 We to 1.0 kWe allowed for the use of a smaller antenna and a two-fold increase in data rate(TWTA capability from 35 We RF to 140 We RF). The reduced antenna size provided the option of relocating the antenna to accommodate the desired FPS system location on top of the spacecraft where the HGA was previously located and still fit within the Atlas V 5 meter fairing. The same instruments and two ESA supplied probes, Titan Montgolfier and lander, were maintained. Elimination of the original 4 m HGA allowed for the placement of the fission power system attached to the ~ 8.5 meter boom. Higher power also allowed simplification of the thermal system by adding electric heaters for the propellant tanks that originally used waste heat from the ASRG’s axillary cooling loop feature. Due to the increase in mass of the FPS compared to the ASRGs, additional Xenon propellant was required: ~1000 kg and 1500 kg for the Stirling and TE options, respectively. The SEP stage tanks were also sized accordingly. An additional ion engine was also included. The additional TSSM concept used the 4 meter HGA as the drag surface for Titan orbit aerobraking. Since this was no longer an option, a 4.5 meter drag plate was utilized for both FPS options to maintain the original Titan capture timeline and heating loads. Due to the more aft center of gravity configuration, the orbiter flight path during aerobraking is reactor first. The heat load on the radiator panels is similar to the thermal environment posed during the Venus flyby, ~ 2500 W/m2. The ESA probes are separated prior to any aerobraking maneuvers allowing the drag plate implementation. In order to remain close to the original transit time of 13 years and remaining on the Atlas, the increase in delta-v is solved by employing a Earth spiral out regimen with a duration of 2 years for the Stirling and 3 years for the TE option requiring an increase in solar array power from 15 kWe to 19 kWe. The higher power array also supplies sufficient power during the EVEE part of the trajectory for both SEP and orbiter operations. This provides the option of reactor start up after the last Earth flyby, about 5 years after launch. Two benefits are realized in that, 1) reactor operating time is reduced resulting in a 50% reduction in radiation fluence and subsequent reduction in shied mass, and 2) eliminates potential safety concerns with an Earth flyby of an operated reactor.

Stirling FPS Shield Cone Radiation protection for spacecraft components and instruments are accomplished by a mass-optimization of separation distance and shielding thickness. At a 10 m separation distance the 11 degree shield provides a 4.5 m diameter 25 kRad, 15 year total ionizing dose. An additional zone was added to the shield to increase the cone angle from ~11º to ~15º thus reducing potential backscatter off the antenna protruding beyond the shadow cone. Spot shadow shielding would be used for components requiring additional protection. 80 deg

Comparison of 2008 TSSM Study and 2014 FPS Concepts Red: Increased value; Green: Decreased value Comparison of 2008 TSSM Study and 2014 FPS Concepts

Power Conversion Technology Stirling Engine TE Module

1 kWe Stirling 1 kWe TE Reactor Core Reactor Shield 42 47 Primary (Na) HP Reactor Shield 40 56 30 Stirling Convertors 302 50 47 43 TE Convertors 394 Secondary (H2O) HP Primary (Na) HP 99 Radiator 240 3.2 m2 5.2 m2 105 Dimensions in cm. 187

1 kWe TE 1 kWe Stirling 13 kWt Reactor Thermal Power Eight 1.6 m long Na Primary Heat Pipes Eight 150W ASRG-derived Stirling convertors 1.1 m OD 1-Sided Cylindrical Radiator 3.2 m2, 400º K radiator 24% BOM efficiency 3.0 m Total System Length 406 kg 1 kWe TE 13 kWt Reactor Thermal Power Eighteen 4 m long Na Primary Heat Pipes 18x21 SKD/LaTe/Zintl TE Modules 1.9 m OD 1-Sided Conical Radiator 5.2 m2, 475º K radiator 8.0% BOM Efficiency 3.9 m Total System Length 604 kg

Stirling FPS TSSM Trajectory Departure C3 0 km2/s2 Earth Departure Mass 7,500 kg SEP ΔV 3,391 m/s Xe Prop 598.8 kg Total Trip Time 3,763 d Thrusting Time 891.6 d Saturn Orbit Insertion (SOI) ΔV 592.3 m/s

Orbiter SEP

Shield Reactor Fixed Boom SEP Stage 2.5 m Antenna 4.5 m dia. Drag Plate

Case 3 TE – heavier reactor/shield more radiators Approach: Use Case 2 MEL and estimate changes with heavier reactor on top Mission: need to start lower C3 (-22 km^2/s^2, apogee 23,000 km) than case 2, longer trip time (spiral out up to 3.2 yrs) Layout: Same size boom, thicker shield, boom area sufficient additional radiators Need to watch cg at launch- checking Boom needs to carry heavier mass at launch (~20% heavier) Cg further forward for aerobraking (good) Will height violate large shroud - yes Propulsion: More xenon, more chemical propellant Xenon: ~1800 kg, add another NEXT engine string Chemical: ~3000 kg C&DH: Same AD&CS: Bigger wheels?, bigger drag flap Communications: Same Thermal for orbiter/SEP Stage: Same? More spiral time / van allen radiation Mechanical: Heavier boom Reactor/shield grows from 4.3 kWth to 13 kWth, Shield size grows in diameter from 40cm to 56 cm Radiator area grows from 3.2m^2 to 5.2 m^2 Xenon load increased to ~1800 kg, added 1 NEXT Engine string