Lunette: A Global Network of Small Lunar Landers Leon Alkalai and John O. Elliott Jet Propulsion Laboratory California Institute of Technology LEAG/ILEWG 2008 October 30, 2008
Baseline Mission Initial design concept focused on site survey mission for first lunar outpost on rim of Shackleton Multiple stationary landers could characterize outpost neighborhood without need for mobility Potential landing sites 2 INITIALS-2
Baseline Mission Overview 3. Carrier stage performs braking burn 2. Multiple landers with carrier stage cruise to moon 1. Launch as secondary payload on EELV 4. Landers separate at termination of burn (~4 km alt.), perform final descent and landing 3 INITIALS-3
Flight System Overview Star 48V Motor ESPA Ring Landers (6) Standard ESPA ring used for lander attachment Star 48V motor used for braking Thrust vector controlled version of Star 48 C&DH and comm during cruise provided by two landers, 1 prime, one backup Lander RCS motors used for cruise ACS and TCMs Star tracker is only active element on ESPA ring 4 INITIALS-4
Lander Overview Stowed ESPA Interface Deployed Radiator/ Instrument Deck WEB Propulsion Lander architecture represents straightforward design making use of largely off-the-shelf components Packaging envelope fits within ESPA accommodation requirements (61x71x96 cm) 180 kg wet mass includes ~12 kg of science payload Standard environmental suite on each lander ~3 kg additional available for specialized instruments Warm electronics box (WEB) design incorporates RHUs for overnight survival heating Solar arrays optimized for polar location provide ~100W 5 INITIALS-5
Measurement Objectives Demonstrate components of ALHAT Software, optical landing instruments Perform photographic reconnaissance of site during descent Perform detailed photographic survey of site from surface Document lighting conditions over mission life Environmental characterization over time Illumination Seismometry Magnetometry Temperature Radiation Dust Regolith characterization 6 INITIALS-6
Baseline Lander Instrument Suite Dust Sensor PanCam Radiation/Biology Experiment (C-CAM) Temp Sensor Micro Seismometer (4X) Magnetometer Thermal-electric Conductivity Probe 7 INITIALS-7
Example Launch Configuration (Atlas V option) Primary Payload Primary Payload Separation Plane Primary Payload Adapter Fitting Lunette Spacecraft Lunette Separation Plane Circular Cylinder Spacer Lunette Payload Adapter Fitting LV Upper Stage 8 INITIALS-8
Post Sep Trajectories (Low Energy) and Fuel Usage High sep point 23.2 Nominal sep point Low sep point 21.9 22.0 22.1 22.7 23.0 Fuel Usage (kg): 9 INITIALS-9
Fuel budget for each lander Each lander has 38 kg of fuel Cruise usage allocation: 13 kg 1 kg for ACS (rate damping, slews, deadbanding) 12 kg for TCMs (~44 m/s) Consistent with expected needs, not much dv99 margin Landing allocation: 25 kg Worst-case from SRM dispersions ~23.2 (Low energy) and ~ 22.6 kg (Direct) Can probably improve L-E worst-case usage by lowering target altitude and/or changing approach azimuth ACS usage is negligible Can probably define alternate landing targets if fuel insufficient to reach desired target 10 INITIALS-10
Single Lander Variant Single lander can be launched to any point on lunar surface Uses same lander design with slight modification to accommodate non-polar operation Solar Array and thermal design changes allow operation at any latitude Remove Electra subsystem (unless needed for far-side orbiter link) TVC version of Star 27 SRM provides braking Total stack launch mass ~600 kg Candidate launch vehicles: Delta II SpaceX Falcon IX Capability to launch multiple individual landers on single EELV Six landers could be launched on Atlas 511 11 INITIALS-11
Single Lander Configuration Flight Landed Solar arrays 12 INITIALS-12
Ongoing Work Team currently working to refine single lander design Capable of Global operation at any latitude Takes science data from selected instruments continuously, including over lunar night Currently designing to support continuous data from seismometer and magnetometer instruments Data taken and stored over night, downlinked during daylight operations Revised single lander design not constrained by ESPA envelope More volume for payload accommodation Larger solar array area Greater battery capacity 13 INITIALS-13
Optimized Single Lander Flight System Design retains all features of ESPA-based Lunette Removal of envelope constraint allows simplification Fixed landing legs Easier packaging of payloads Greater surface area for solar arrays Solar Arrays hinged at top, deployed after landing Optimum deployment angle could be set pre-flight, depending on target latitude Equatorial Horizontal Mid-latitudes ~ 45 degrees from horizontal Polar landing site vertical panels 14 INITIALS-14
Science Mission Concepts 15 INITIALS-15
Touch the Ice Variation on the baseline mission Six landers with ESPA carrier launch as secondary payload Up to four landers targeted for permanently shadowed regions of crater Short-lived (~hours) landers investigate area for presence of water ice Two or more landers targeted for crater rim to provide telecom relay Landers continue environmental monitoring mission after completion of ice sensing phase DTE Link Sunlit Crater Rim Lander relay Permanently Shadowed Crater Floor 16 INITIALS-16
Lunar Seismic Network Launch up to six individual lander/srm stacks on one launch vehicle (Atlas V 511) Three landers could be launched on Falcon 9 Up to eight landers could be launched on an Atlas V 531 Landers target landing sites distributed over lunar surface at separation distances of 1000-2000 km Latitudinally grouped in three sets spaced around planet Landers form seismic network for 2-5 year mission 17 INITIALS-17