Mars Sample Return Implementation Discussions

Transcription

1 Mars Sample Return Implementation Discussions November 4,

2 Agenda Topics Context of Mars Sample Return architecture* Composition of multi-element Mars Sample Return campaign* Key technical/technological challenges Cost assessments Summary *Mars Sample Return is conceptual in nature and is subject to NASA approval. This approval would not be granted until NASA completes the National Environmental Policy Act (NEPA) process 2

3 Why Mars Sample Return Now? Scientific impetus Significant improvements of understanding of Mars as a system from our past, current, and planned exploration missions Sample return is necessary to achieve the next major step in our understanding of Mars and Solar System exploration Compelling high-priority sites for sample return have been identified Engineering readiness MEP investments have developed many capabilities critical for return of samples from Mars (e.g., rover mobility, skycrane EDL system, aerobraking) Key remaining technical challenges are identified/understood, with technology plans defined Multi-flight-element campaign Multi-flight-element campaign involves flight element concepts that are similar to our existing implementation experience Allows analogies for scope and cost estimation Reduces the number of technical challenges per element Provides a resilient program and science robustness with ability to exercise contingency Based on the past decade of Mars exploration, NASA is poised to embark on a campaign of missions leading to the return of Martian samples in the 2020ʼs 3

4 MEPAG Next Decade-SAG/iMARS MSR Science Requirements* Sample diversity Similarities and differences between samples is as important as the absolute characterization of a single sample. Identifying and selecting samples that are different from each other is crucial. Minimum necessary sample size/mass The optimal sample size for rock samples is about 8-10 grams For regolith samples, a larger sample size is better. Minimum necessary number of samples Rock samples: 20 Regolith samples: 4 Dust sample (if collectable): 1 Gas sample: 1 Sample preservation needs Samples must be labelled (to link to field context) Retain pristine nature of samples prior to arrival on Earth (avoid excess heating, organic and inorganic contamination.) Samples would require packaging to ensure that they do not become mixed *SOURCE: ND-SAG (2008), imars (2008) 4

5 Functional Steps Required to Return a Scientifically Selected Sample to Earth Sample Caching Rover (MAX-C) Mars Sample Return Lander * Mars Sample Return Orbiter * * Launch from Earth/Land on Mars Retrieve/Packag e Samples on Mars Capture and Isolate Sample Container Select Samples Return to Earth Launch Samples to Mars Orbit Acquire/Cach e Samples ** Land on Earth ** Sample Canister On Mars Surface Orbiting Sample (OS)* in Mars Orbit Orbiting Sample (OS) On Earth * * Retrieve/Quaranti ne and Preserve Samples on Earth Assess Hazards Sample Science Mars Returned Sample Handling (MRSH) Facility *Artist s Rendering ** Launching orders of MSR orbiter and lander can be reversed Sample Science 5

6 Architectural Trade Space L a r g e M e n u o f O p t i o n s Launch Mode Earth to Mars Approach Nav. MOI Mode Entry Mode Descent Mode Landing Mode Telecom Network Surface Ops Lander Rover Planetary Protection Outbound Inbound MAV Details Rendezvous Mode Mars to Earth Single LV Two LV Three LV Ballistic Radio Chemical Direct Guided Low L/D Propulsive Base to Orbiter Solar Solar Chemical Load Reduction Encapsulation Unguided MOR Ballistic SEP Optical SEP From Orbit Guided Mid L/D Hazard Avoidance Base to Rover RPS RPS Biological Load Reduction Thermal Guided LPR SEP Hybrid Aerocapture Viking Chute Technology Pallet Base to Earth Subsurface to 2m Subsurface to 2m Sense to Abort Active Last Stage DSR Hybrid Bioshield Aerobraking Beyond Viking Chute Technology Air Bag Rover to Orbiter Subsurface to 10m Mobility to 1 km Passive Last Stage Historical studies Joint NASA/CNES study US industry team studies Workshops on required technology investments International Mars Architecture for Return of Samples (imars) working group Landing to 1 km accuracy Rover to Earth Sample ID and Assay Mobility to 10 km Landing to 10 km accuracy Orbiter to Earth Sample Xfer to MAV/OS Sample ID and Assay Passive Only Post-MAV mission Sample Xfer to MAV/OS (Control bioburden and prevent recontamination) (Break the chain and sense leak to abort) Gel/liquid/solid OS Orbiter to MAV Orbiter to OS (assume MAV on lander) Post-MAV Mission Both Active and Passive Total System Entry Mode Direct STS Tug (STS likely implies tug also) 6

7 Current Multi-Element Architecture Approach 7

8 Composition of Multi-element Mars Sample Return Campaign 8

9 Multi-Element Campaign For Returning Samples from Mars Is A Resilient Approach * * * * MAX-C (Caching Rover) *Artist s Rendering Mars Sample Return Orbiter Science Robustness Mars Sample Return Lander Allows adequate surface operation duration for scientific sample selection and acquisition Technical robustness Spreads technical challenges across multiple elements: reduces number of engineering challenges per element Keeps landed mass requirements within MSL EDL capability: use MSL EDL as work horse Sample caching rover will allow us to fly a less-complicated fetch rover on sample return lander: reduced landed mass with improved margin Programmatic Robustness Involves concepts similar to our implementation experience Spreads budget needs and reduces peak year program budget demand Provides strategic resilience to individual mission/engineering failure: only need to redo failed step Mars Returned Sample Handling 9

10 MAX-C (Caching Rover) Overview Science Capability Mass Allocation (Launch/Entry/Landed) Major Mission/Spacecraft Attributes Remote and Contact Science (Color stereo imaging, macro/micro-scale mineralogy, micro-scale organic detection/characterization, micro-scale imaging) Coring and Caching Rock Samples for Future Return 4025 / 3300 / 975* kg Launch Vehicle (Baseline) Atlas V 531 Power/Energy per Sol Cruise: 1250 W Solar Surface: ~1600 WHrs/sol Solar Cruise ACS Entry Vehicle Diam. / Parachute Diam. Landing System Rover Mast Height / Wheelbase Ground Clearance/Wheel Diam. Data Return per Sol (2-week average) Data Storage Science Payload Mass Motor Architecture Traverse Capability (Design Distance) Flight Software Surface WEB Thermal Range/ Design Surface Design Lifetime Stable Spinner (MSL Design) 4.5 m / 21.5 m Skycrane throttled monoprop with landing pallet ~1.6 m / ~1.5 m ~0.35 m / ~0.32 m ~250 Mbits UHF (w/tgm); MER/MSL-class Xband DTE 32 Gbits ~15 kg instruments ~68 kg including coring/caching/mast/arm Brushless hybrid distributed electronics 20 km MSL-based -40C to +50C / CO2 gap insulation, RHUs, supplemental htrs 500 Sols MAX-C delivered by a derivative of the MSL Cruise/EDL system (cruise stage not shown) Skycrane lands MAX-C rover on landing pallet * MAX-C Rover ~300 kg (with component maturity-based uncertainty) * * Artist s Rendering * * * Landed mass includes an allocation of 675 kg for MAX-C Rover, Landing Pallet/Structure, and system-level mass margin plus an allocation of ~300 kg for possible additional payload. Sample Canister On Mars Surface 10

11 MSR Orbiter Functions Insert into Mars orbit/aerobrake Rendezvous with OS in 500km orbit. Capture, transfer and package OS into EEV Return to Earth, including Earth swingby Release EEV for entry Divert into a non-return trajectory If Before MSR Lander Monitor critical events of EDL and MAV launch Provide telecomm relay for lander and rover Team-X Design No staging required Alternate Design Separate prop stage that separates after Trans Earth Injection Features Over twice the propellant needed by typical Mars orbiters. Uses bi-prop systems flown on previous Mars missions UHF Electra relay system used for surface relay and OS beacon reception. Orbiter mass quite dependent on mission parametrics of specific launch and return years. Design to envelop several EEV opportunities. Orbiter Rendezvous systems 20 Systems Capture/Sample Transfer 40 Propellant Avionics kg Power 130 Structures/Mechanisms 330 Cabling 40 Telecom 30 Propulsion 170 Thermal 40 Misc. Contingency 100 TOTAL (40% margins) TOTAL Orbiter Systems Mass 940 kg 2280 kg 3260 kg 11

12 Sample Capture/Earth Entry Vehicle (EEV) Capture Basket concept testing on NASA C-9 zero-g aircraft Orbiting Sample (OS) container with battery operated UHF low-duty cycle beacon (as backup) Detection and rendezvous systems OS is released in to a 500km circular orbit by the MAV Optical detection from as far as 10,000km. Autonomous operation for last 10s of meters OS has a battery operated, very low duty cycle UHF beacon for coarse location as backup. Capture System Capture basket concept designed Prototype demonstrated on a NASA C-9 zero-g aircraft flight campaign. Technical progress on Orbital Express improves confidence for Strawman EEV design 0.9m diameter, 60 sphere-cone blunt body Design responsive to stringent Earth planetary protection requirements Self-righting configuration No parachute required Hard landing on heatshield structure, with crushable material surrounding OS EEV design capitalizes on design heritage Extensive aero-thermal testing and analysis Wind tunnel tests verified self-righting UTTR drop test reached terminal velocity Strawman landing site based in US 12

13 MSR Lander System Artist s concept MSL Cruise and EDL system in Hi-bay Strawman approach uses Atlas 551 MSL system with MSR Lander in-place of Curiosity Rover Cruise Stage Backshell w/parachute Descent Stage (Skycrane) Lander w/mav & Rover Heatshield MSR lander Capitalizes on reuse of MSLʼs cruise & entry, descent and landing (C&EDL) systems (also planned for MAX-C). MSR lander pallet delivered to the surface via the MSL skycrane system (soft-touchdown). Updated Team-X study (Oct 2009) verifies lander platform with MAV and rover fit within the volume constraints of the MSL system. 13

14 Mars Ascent Vehicle* Payload Fairing Orbiting Sample (OS) Avionics Compartment Star 13A SRM TVC Controllers Stretched Star 17A SRM SRM Igniters All figures are artist s concepts TVC Actuators 2.5 m Launches 5kg Orbiting Sample (OS) into 500+/- 100 km orbit, +/-0.2deg Kept thermally stable in an RHU augmented thermal igloo. Continuous telemetry for critical event coverage during ascent. Strawman approach: Solid two-stage launch vehicle. Uses standard solid rocket motors (SRMs) 3-axis monoprop system for control. 300kg (40% margin; including OS) Incremental testing and Earth-based high-altitude testing part of technology investments NASA investigating technology investment options Will explore various solid and liquid propellant system approaches. * 01-02: 3 MAV industry studies; Team X/MSFC red-team reviewed the strawman approach; continuing workshops in confirmed assessments 14

15 MSR Lander Platform Utraflex (stowed) HGA LGA UHF MAV Bio-Thermal Barrier Sample Handling Mechanism Lander Arm Bio-Barrier Extended deck surface not shown * Fetch Rover (old design) Ultraflex Solar Array (2x) Supports and protects MAV in thermal igloo with 10s of RHUs. Sample transfer of rover-cache, local regolith and atmosphere. Cache loaded with Phoenix-like arm/scoop/camera; also can collect contingency sample. Seals canister, packages in OS and transfers to MAV. 1 Earth year life Lander WEB * MAV Erector Egress Ramp Lander arm *Artist s concept 15

16 Fetch Rover Requirements Retrieves sample cache within ~3-months of surface mission operation One Earth-year design lifetime 2009 design study concept features Similar to MER design, but with repackaged MSL avionics Enhanced autonomous driving (increased image processing) adopted from MAX-C 1-DOF arm for pickup, demonstrated on rover in Mars-yard UHF link relies on orbital relay (or lander while in proximity) 155 kg (40% margin) Fetch Rover MER 16

17 MSL EDL Feed Forward To MSR MSL EDL system will deliver ~1000 kg rover to Mars surface for ʻ11 launch Can be used for MAX-C in 2018 * and the MSR-L in MSR-Lander 3 major sub-elements at 40% mass margin adds to 995kg Mars Ascent Vehicle: 300 kg Platform that supports MAV: 540 kg Fetch Rover: 155 kg MSL EDL performance is relatively constant between 2011 MSL and ʻ22-ʼ26 Arrival seasons result in similar atmospheric conditions: can accommodate the MSR-Lander If needed, MSL EDL system capability can be increased to accommodate minor addition to landed mass Options include: increased entry body L/D, increase parachute deployment Mach number, mission design modifications The 3-flight-element approach improves resilience to the EDL-landed mass picture A key reason for evolution into 3-flight-element approach MSLʼs 10km radius landing precision can be improved to ~6-7km to reduce rover traverse V α L g D * 2018 is a more favorable opportunity than ʼ11 for EDL 17

18 Surface Rendezvous Examples Go-to site example MAWRTH??? Sample-locally site example Traverse distance for MAX-C depends on landing site: Go-to site requires up to 20km traverse: will bring sample back to near center of landing ellipse Sample-locally site requires less than 20km traverse Fetch rover is required to traverse up to ~12km Use of MSL design capability supports above traverse distances 18

19 Mars Returned Sample Handling Facility Artist s concept Artist s concept of an SRF Functions Contain samples as if potentially hazardous, equivalent to bio-safety level-4 (BSL-4). Keep samples isolated from Earthsourced contaminants Provide capability to conduct biohazard test protocol as a prerequisite to release of samples from containment. Industry studies performed to scope facility and processes (2003) NASA Draft Test Protocol used as basis 3 architectural firms, with experience in bio-safety, semi-conductor and food industries. Sample handling approaches ranged from glove-box to all-robotic. Current costs estimates and scope are based on industry studies and comparison to existing BSL-4 facilities. Could function as sample curation facility after hazard assessment. 19

20 Technical/Technological Challenges for Multi-element Mars Sample Return Campaign 20

21 Technology Development Difficulty Qualitative Assessments High Medium Low Sample Acquisition Round Trip PP Precision Landing Hazard Avoidance (if needed) Mobility Technology MAX-C Mission Difficulty High Medium Low MAV Back PP MSR-Lander High Medium Back PP Rendezvous and Sample Capture EEV Low Technology MSR-Orbiter 21

22 MAX-C Mission Technologies Round- Trip PP Avoid false positive life detection Reduce number of viable organisms Sterilize components that come in contact with samples Use clean-assembly process for assembling sterilized components Develop bio-barriers to avoid recontamination Develop analytical tools to estimate contamination risks during and post landing Rover Technology Increase average drive speed to ~200 m/sol MER record for one sol is 224m Sensor processing and algorithms for autonav/visual odometry reduce average drive speed to ~30m/sol MAX-C drive speed will be ~65m/sol when using MSL avionics Using co-processors and faster algorithms 200m/sol is achievable for MAX-C (and fetch rover) Precision Landing Decrease MSL landing ellipse radius to 6-7 km Use MSL EDL architecture, including guided entry Reduce entry attitude initialization error prior to entry Use range trigger for deployment of the parachute Sampling and Caching Obtain core samples Industry and JPL have developed coring technologies in the past, but not yet with a system level demonstration Characteristics needed include: low-mass/lowpower rock coring; core break-off, retention, ejection; and bit change-out In 2009, ND-SAG sample requirements were used to develop four independent concepts for sample acquisition and encapsulation by ATK, ASI, JPL, and Honeybee Robotics These studies indicate feasible solutions exist that can be developed and matured by MAX-C PDR 22

23 MSR-L Mission Technologies Back PP Assure containment of all returned Martian samples and flight hardware exposed to Martian material, until they could be tested for possible biohazards * Breaking-the-chain of contact including dust mitigation Sealing & leak detection Containment vessel that withstands impact Earth return targeting Meteoroid protection & breach detection for sample container and EEV heat shield Mars Ascent Vehicle (MAV) Launch 5kg OS to a 500 +/- 100 km Mars Orbit Critical technologies for solid MAV are: Thrust vector control (TVC) suitable for MSR environment Propellant compatibility with MSR/Mars environment A flight-like version to be designed, developed, and tested in flight-like conditions prior to PDR for Realistic landing tests for g loads High altitude tests for Mars conditions Alternative MAV systems to be considered such as liquid MAV system MAV technology and system development must start 7 years prior to MSR lander PDR Three years for trades and component technology development to TRL6 Four years for a flight-like MAV system development and a flight test * Back PP could be implemented on the lander and orbiter or just the orbiter 23

24 MSR-O Mission And MRSH Technologies Rendezvous & Sample Capture Track from 10s of thousands of km, rendezvous, and capture OS Sensor technology examples: Optical cameras to detect OS from 10,000 km (ONC already flown on MRO) Beacon as a possible backup to cameras Mechanisms: Capture mechanism (a prototype developed and tested in 0-g) Integrated optical sensors and gimbals MSR Rendezvous and Capture GN&C System Techniques demonstrated by Orbital Express Earth Entry Vehicle Safely deliver OS to Earth surface In the 2000 timeframe, NASA developed a detailed conceptual design of the MSR Earth Entry Vehicle This design was supported by wind tunnel and impact testing Key development includes heat shield material development/selection l Mars Returned Sample Handling Assure Containment and Prevent Contamination During Ground Processing Assure Containment and Prevent Contamination During Ground Processing Sample transfer from landing site to Sample Receiving Facility Biological safety combined with sample protection Ultra-clean sample manipulation Sample sterilization techniques, if required, prior to release of samples to research institutions l 24

25 Cost Assessment for Multi-element Mars Sample Return Campaign 25

26 Current Cost Estimate Bases Three sets of costs are provide, all in constant $FY 15 and include Mission and flight system development and ATLO (including technology) Launch vehicles Mission and ground operations Science team costs for MAX-C (science costs for future elements to be determined with community input) First set of mission cost estimates are derived from JPL Institutional estimates (Team X) Team X cost estimating process employs representatives of the potential doing organizations The flight system WBS for each mission composed of many separate cost elements (e.g., propulsion, thermal, mechanical, etc.) Several additional WBS elements estimated by wrap algorithms (e.g. mission design) and look-up catalogues, including the NASA instrument cost model Recent MSL experience is taken into account Team-X results have been modified to include Development (phase A-D) reserves of 50%, phase E reserve of 25% Technology/advanced engineering development, also with reserves of 50% Second set of mission cost estimates are derived from analogy (enabled by multi-element approach) with recently implemented missions in the program (completed by program office) MSL and MRO are used for analogy Third set of mission cost estimates are taken from a 2008 independent cost estimate by Aerospace Corp., conducted for NASA Headquarters, recently updated Oct

27 MAX-C (Sample Caching Rover) Cost Estimate Basis 1. Team-X Session Oct Analog / Metrics Cost Estimate ($FYʼ15) ~ $2.1 B Comments Technology cost is ~ $120 M (included in $2.1 B) 2. Analogy with MSL (quantized to 0.5B) 3. Independent Aerospace Corp. Mid-size rover with ~65 kg instrument payload and caching system (MSL: 235 kg); utilize MSL EDL/Cruise Capability ~ $2.0 B $2.1 B Capitalize on MSL EDL/cruise systems; focus on technical capability of scientific sample selection/sampl e caching 27

28 Mars Sample Return Orbiter Cost Estimate Basis 1. Team-X Session Oct Analogy with MRO (quantized to 0.5 B) Analog / Metrics Orbiter will be MRO-class but needs to carry EEV, capture sample capsule in Mars orbit and propulsively inject into Earth return trajectory Cost Estimate ($FYʼ15) ~ $1.3 B ~ $1.5 B Comments Technology cost is ~ $160M (included in $1.3B) Major propulsion system for MOI and trans- Earth orbit injection 3. Independent Aerospace Corp. ~ $1.1 B 28

29 Cost Estimate Basis 1. Team-X Session Oct Analogy with MSL (quantized to 0.5B) Mars Sample Return Lander Analog / Metrics Fetch rover (MER-class) to retrieve cached sample; MAV and landing platform Cost Estimate ($FYʼ15) ~ $2.3 B ~ 3.0 B Comments Technology cost is ~$250M (included in $2.3B, also MAV cost estimate from previous industry input) Capitalize on MSL EDL/cruise systems; technology investments needs to start ~8 yrs prior to launch 3. Independent Aerospace Corp. ~$2.4 B 29

30 Cost Estimate Basis 1. MSRH cost derived from previous industry input Mars Returned Sample Handling (MRSH) Analog / Metrics Cost Estimate (FYʼ15) ~$0.5 B Comments Technology cost is ~ $35M (included in $0.5B) 2. Analogy with Bio Containment Labs (quantized to 0.5B) 3. Independent Aerospace Corp. Bio-safety labs/curation facilities ~ $0.5 B ~ $0.3 B 30

31 Multi-element Sample Return Cost Estimates (All In $FY 15) MAX-C MSR Elements Team-X w/ 50% A-D & 25% E Reserves Analogy with past missions (quantized to 0.5B) Independent (Aerospace Corporation) ~$2.1 B ~$2.0 B ~$2.1 B MSR-Orbiter ~$1.3 B ~ $1.5 B ~$1.1 B MSR-Lander ~$2.3 B ~ 3.0 B ~$2.4 B MRSH ~$0.5 B ~$0.5 B ~$0.3 B MULTI-ELEMENT MSR TOTALS ~$6.2 B ~$7.0 B ~$5.9 B Cost estimates have been rounded off to $0.1B 31

32 Notional Schedule For Multi-Element Campaign To Return Samples From Mars 32

33 Summary Strong scientific impetus for sample return Sample return is necessary to achieve the next major step in understanding Mars and the Solar System Compelling sites for sample return have been identified Engineering readiness for sample return MEP investments have developed many capabilities critical to sample return Key remaining technical challenges are identified/understood, with technology plans defined Resilient multi-flight-element approach based on multiple studies in past years and recent flight experience Science robustness Allows proper surface operation duration for scientific sample selection and acquisition Technical robustness Spreads technical challenges across multiple elements Keeps landed mass requirements within MSL EDL capability Programmatic robustness Involves concepts similar to our existing implementation experience Reduces the number of technical challenges per element Spreads budget needs and reduces peak year program budget demand Provides a resilient program and science robustness with ability to exercise contingency Multi-element MSR should not be viewed as an isolated (flagship) mission but as a cohesive campaign that builds on the past decade of Mars exploration Approach amenable to international partnership 33