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Nuclear Thermal Rocket Propulsion for Future Human Exploration Missions presented by

Nuclear Thermal Rocket Propulsion for Future Human Exploration Missions presented by Dr. Stanley K. Borowski Chief, Propulsion and Controls Systems Analysis Branch at the Future In-Space Operations (FISO) Colloquium. Wednesday, June 27, 2012. 1.

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Nuclear Thermal Rocket Propulsion for Future Human Exploration Missions presented by

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  1. Nuclear Thermal Rocket Propulsion for Future Human Exploration Missions presented by Dr. Stanley K. BorowskiChief, Propulsion and Controls Systems Analysis Branch at the Future In-Space Operations (FISO)Colloquium Wednesday, June 27, 2012 1

  2. Nuclear Thermal Rocket (NTR) Concept Illustration(Expander Cycle, Dual LH2 Turbopumps) NTR: High thrust / high specific impulse (2 x LOX/LH2 chemical) engine uses high power density fission reactor with enriched uranium fuel as thermal power source. Reactor heat is removed using H2 propellant which is then exhausted to produce thrust. Conventional chemical engine LH2 tanks, turbopumps, regenerative nozzles and radiation-cooledshirt extensions used -- “NTR is next evolutionary step in high performance liquid rocket engines” During his famous Moon-landing speech in May 1961, President John F. Kennedy also called for accelerated development of the NTR saying this technology “gives promise of some day providing a means of even more exciting and ambitious exploration of space, perhaps beyond the Moon, perhaps to the very end of the solar system itself.” Ceramic Metal(Cermet) Fuel NERVA-derivedCarbide Fuel NTP uses high temperature fuel, produces ~525 MWt (for ~25 klbf engine) but operates for < 85 minutes on a round trip mission to Mars (DRA 5.0) 2

  3. Larger Coresfor Higher Thrust System Baseline for NERVA Program Higher PowerFuel Elements Tech Demo • 20 NTR / reactors designed, built and tested at the Nevada Test Site – “All the requirements for a human mission to Mars were demonstrated” • Engine sizes tested • 25, 50, 75 and 250 klbf • H2 exit temperatures achieved • 2,350-2,550 K (in 25 klbf Pewee) • Isp capability • 825-850 sec (“hot bleed cycle”tested on NERVA-XE) • 850-875 sec (“expander cycle”chosen for NERVA flight engine) • Burn duration • ~62 min(50 klbf NRX-A6 - single burn) • ~ 2 hrs(50 klbf NRX-XE: 27 restarts / accumulated burn time) • -----------------------------* NERVA: Nuclear Engine for Rocket Vehicle Applications Rover / NERVA* Program Summary (1959-1972) The smallest engine tested, the 25 klbf “Pewee” engine, is sufficient for human Mars missions when used in a clustered engine arrangement The NERVA Experimental Engine (XE) demonstrated 28 start-up / shut-down cycles during tests in 1969. 3

  4. “Heritage” Rover / NERVA Homogeneous Thermal ReactorFuel Element and Tie Tube Bundle Arrangement Hexagonal FE: 0.75 in across the flats; 35 – 52 in length with 19 coolant channels CVD-coated UC2 Particles 4

  5. Key Elements of the NERVA NTR Engine 0.75’’ ~35-52’’ 5

  6. Performance Characteristics for Small & Full Size NERVA-Derived Engine Designs – Composite Fuel State-of-the-Art “Pewee”Engine Parameters Ref: B. Schnitzler, et al., “Lower Thrust Engine Options Basedon the Small Nuclear Rocket Engine Design”, AIAA-2011-5846 6

  7. NOTE: Figure depicts performance regions typically shown for the various fuel options. Fuels can be operated at lower temperature levels to extend fuel life & / or increase engine operational margins. Also, by reducing the fuel loading, higher operating temperatures & specific impulse values are achievable to improve performance 7

  8. g SUN g Trajectory Options for Human Mars Missions • Opposition-Class Mission Characteristics(Used in “90-Day” / SEI Mars Studies) • Short Mars stay times (typically 30 - 60 days) • Relatively short round-trip times (400 - 650 days) • Missions always have one short transit leg (eitheroutbound or inbound) and one long transit leg • Long transit legs typically include a Venus swing-by and a closer approach to the Sun (~0.7 AU or less) • This class trajectory has higher DV requirements • NOTE: Short orbital stay missions will most likely be chosen for initial human missions to Mars & its moons, Phobos and Deimos Outbound Surface Stay Inbound • Fast-Conjunction Class Mission Characteristics (Used in DRM 4.0 & DRA 5.0) • Long Mars stay times (500 days or more) • Long round trip times (~900 days) • Short “in-space” transit times (~150 to 210 days each way) • Closest approach to the Sun is 1 AU • This class trajectory has more modestDV requirements than opposition missions 8

  9. NTR Requires Less Propellant than Chemical Systems Mi = Mf exp (V / gE Isp) “Rocket Equation” where Mi is initial total mass of spacecraft in low Earth orbit (LEO), Mi = MSC (spacecraft) + MPL (payload) + Mprop (propellant) and Mf = spacecraft mass after a given amount of propellant has been expended in providing a given velocity increment (V) to the spacecraft, gE = Earth’s gravity = 9.80665 m/s2; and Isp = specific impulse (pounds of thrust generated per pound of propellant exhausted per second) NTR has 100% Higher Ispthan Chemical Propulsion Crew Return in MAV Mission Mass Ratio: (RM = Mi /Mf = exp (DV / gE Isp) NTR is only Propulsion Option besides Chemical to be Tested at Performance Levels needed for a Human Mission to Mars! 9

  10. NTR Crewed & Cargo Mars Transfer Vehicles (MTVs)for DRA 5.0: “7-Launch” Strategy 3 – 25 klbf NDR Engines(Isp ~906 s, T/Weng ~3.5) Saddle Truss / LH2 Drop Tank Assembly Payload Element ~65 t(6 crew mission) NOTE: Ares-V Core Stage LH2 Tank is 10 m D x ~44.5 m L; two LH2 tanks cut in ~half with 4 extra end domes provides tanks needed for crewed & 2 cargo MTVs “0-gE” Crewed MTV:  IMLEO ~336.5 t  3 Ares-V Launches Cargo Lander MTV:  IMLEO ~236.2 t  2 Ares-V Launches Common NTR “Core” Propulsion Stages AC/EDL Aeroshell, Surface PLand Lander Mass ~103 t  IMLEO = 808.9 t Habitat Lander MTV:  IMLEO ~236.2 t  2 Ares-V Launches NOTE: With Chemical IMLEO >1200 t (Ref: S.K. Borowski, et al., AIAA-2009-5308) 10

  11. NTR Crewed Mars Transfer Vehicle (MTV) Allows NEO Survey and Short Orbital Stay Mars / Phobos Missions 3 – 25 klbf NTRs DRA 5.0 Crewed MTV Options:• “4-Launch” in-line configuration • Ares-V: 110 t; 9.1 m OD x 26.6 m L • IMLEO: ~356.5 t (6 crew) • Total Mission Burn Time: ~84.5 min • Largest Single Burn: ~30.7 min • No. Restarts: 3 ------------------------------------------- • “3-Launch” in-line configuration • Ares-V: 140 t; 10 m OD x 30 m L • IMLEO: 336.5 t (6 crew) • Total Mission Burn Time: ~79.2 min • Largest Single Burn: ~44.6 min • No. Restarts: 3 Phase II Configuration NTP identified as the preferred propulsion option for DRA 5.0 Configuration Used in“7-Launch” Mars Mission Option (ESMD AA Cooke) United States’ National Space Policy (June 28, 2010, pg. 11) specifies that NASA shall: By 2025, begin crewed missions beyond the Moon, including sending humans to an asteroid. By the mid-2030s, send humans to orbit Mars & return them safely to Earth. (Ref: Mars DRA 5.0 Study, NASA-SP-2009-566, July 2009 ) 11

  12. Growth Paths for DRA 5.0 “Copernicus” NTR Crewed MTV using Modular Components Applications:• Fast Conjunction Mars Landing Missions – Expendable• “1-yr” Round Trip NEA Missions to 1991 JW (2027), 2000 SG344 (2028) and Apophis (2028) – Reusable • Propulsion Stage & Saddle Truss / Drop Tank Assembly can also be used as: • Earth Return Vehicle (ERV) / propellant tanker in “Split Mars Mission” Mode – Expendable • Cargo Transfer Vehicle supporting a Lunar Base – Reusable “Saddle Truss” / LH2Drop Tank Assembly Crewed Payload 3 – 25 klbfNTRs MMSEV replaces consumables container for NEO missions Common NTR “Core” Propulsion Stages Applications:• Fast Conjunction Mars Landing Missions – Reusable• 2033 Mars Orbital Mission 545 Day Round Trip Time with 60 Days at Mars – Expendable • Cargo & Crew Delivery to Lunar Base – Reusable “In-Line” LH2 Tank Options for Increasing Thrust: • Add 4th Engine, or• Transition to LANTR Engines – NTRs with O2 “Afterburners” Applications:• Faster Transit Conjunction Mars Landing Missions – Reusable• 2033 Mars Orbital Mission 545 Day Round Trip Time with 60 Days at Mars – Expendable • Some LEO Assembly Required – Attachment of Drop Tanks • Additional HLV Launches Transition to “Star Truss”with Drop Tanks to Increase Propellant Capacity 12

  13. Reusable Crewed Near Earth Asteroid (NEA) Survey Mission Using NTR NEA Exploration (B) MMSEV returns to the ASV NEA Rendezvous Trans-Earth Injection NEA MMSEV InboundTransit (C) Outbound Transit (A) MMSEV detaches from ASV for close-up inspection / sample gathering sorties Trans-NEA Injection (TNI) LH2 Drop Tank Jettisoned LEO: 407 km circular Crewed NTR Asteroid Survey Vehicle (ASV) Earth Entry Velocity <12.5 km/sec Crew recovery using CEV 3 HLV Launches Direct Entry Water Landing Earth Glenn Research Center Candidate NEAs (TNI): (A/B/C) days After HEEO LEO insertion, CEV/SM separates from ASV and re-enters Initial ASV capture into a • 1991 JW (5/18/27): (112/30/220)• 2000 SG344 (4/27/28): (104/ 7/216) • Apophis (5/8/28): (268/ 7/ 69) HEEO: 500 km x 71,136 km MMSEV attached to ASV’s transfer tunnel Pre-Decisional, For Discussion Purposes Only 13

  14. Asteroid 1991 JW: D ~490 m Total DV = 7.188 km/s Mission Times Outbound 112 days Stay 30 days Return 220 days Total Mission 362 days Reusable NTR NEO Survey Mission to 1991 JW IMLEO ~316.7 t 6 crew • Asteroid departure • 10/7/2027 • V = 0.612 km/s Near Earth Asteroid Orbit MMSEV Earth Orbit  • Asteroid arrival • 9/7/2027 • V = 0.851 km/s • HLV Lift: ~140 t • 10 m OD x 30 m L • Total Mission Burn Time: ~73.8 min • Largest Single Burn: ~37.3 min • No. Restarts: 4 • Earth return to 500 km x 71,136 km HEEO • 5/14/2028 • V = 1.711 km/s • Earth departure from 407 km circular orbit • 5/18/2027 • V = 4.014 km/s JSC performed “NEO Accessibility Study” and presented results to ESMD AA on April 7, 2011.Findings: NTR outperformed chemical, SEP/Chemical and all SEP systems, allowing access to more NEOs over larger range of sizes and round trip times for fewer HLV launches. 14

  15. 2033 Mars Orbital Mission Using “Split Mission”Option (RT Time: 545 days with 60 days at Mars) Earth Return Vehicle (ERV):• IMLEO: ~237.4 t • HLV Launches: 2 • Total Mission Burn Time: ~64.2 min • Largest Single Burn: ~25.4 min • No. Restarts: 3 Outbound Crewed MTV: (6 crew) • IMLEO: ~251.1 t • HLV Launches: 3 • Total Mission Burn Time: ~47.7 min • Largest Single Burn: ~25.2min • No. Restarts: 3 Total Mission IMLEO: 488.5 t --------------------------------------------------- “All-Up” Crewed MTV: (6 crew) • IMLEO: ~429.4 t • HLV Launches: 4 • Total Mission Burn Time: ~111 min • Largest Single Burn: ~41.3 min • No. Restarts: 3 LEO Configuration Earth Return Vehicle (ERV)/ tanker; uses “minimum energy” outbound trajectory Consumables canistertransfer tunnel & DM jettisoned before TEI “Switch-over” LH2 drop tankjettisoned after TMI ERV / crewed PL R&D in Mars Orbit LH2 for Earth return in“core” propulsion stage Crewed PL element transferred to ERVfor trip back to Earth Outbound crewed MTV; useshigher energy trajectorieson “1-way” transit to Mars LEO Configuration Crew deliveryOrion/SM 32.2 m 26.1 m 24.8 m 8.9 m (Ref: S.K. Borowski, et al., 2012 IEEE Aerospace Conference, March 3-10) 15

  16. Nuclear Thermal Propulsion -- 2033 600 day Mars Transfer VehicleCore Stage, In-line Tank, & Star Truss w/ 2 LH2 Drop Tanks Payload: DSH, CEV, Food, Tunnel, etc. Star Truss with 2 LH2 Drop Tanks, Port & Starboard Vehicle Mass (mt) / Parameters: Core Propulsion Stage Three 25.1 klbfNTRs In-line Tank Comm. Ant. Design Constraints / Parameters: • # Engines / Type: 3 / NERVA-derived • Engine Thrust: 25.1 klbf (Pewee-class) • Propellant: LH2 • Specific Impulse, Isp: 900 sec • Cooldown LH2: 3% • Tank Material: Aluminum-Lithium • Tank Ullage: 3% • Tank Trap Residuals: 2% • Truss Material: Graphite Epoxy Composite • RCS Propellants: NTO / MMH • # RCS Thruster Isp: 335 sec (AMBR Isp) • Passive TPS: 1” SOFI + 60 layer MLI • Active CFM: ZBO Brayton Cryo-cooler • I/F Structure: Stage / Truss Docking Adaptor w/ Fluid Transfer Mission Constraints / Parameters: • 6 Crew • Outbound time: 183 days (nom.) • Stay time: 60 days (nom.) • Return time: 357 days (nom.) • 1% Performance Margin on all burns • TMI Gravity Losses: 310 m/s total, f(T/W0) • Pre-mission RCS Vs: 181 m/s (4 burns/stage) • RCS MidCrs. Cor. Vs: 65 m/s (in & outbnd) • Jettison Both Drop Tanks After TMI-1 • Jettison Tunnel, Can & Waste Prior to TEI NTP Transfer Vehicle Description: NTP system consists of 3 elements: 1) core propulsion stage, 2) in-line tank, and 3) integrated star truss and dual drop tank assembly that connects the propulsion stack to the crewed payload element for Mars 2033 mission. Each 100t element is delivered on an SLS LV (178.35.01, 10m O.D.x 25.2 m cyl. §) to LEO -50 x 220 nmi, then onboard RCS provides circ burn to 407 km orbit. The core stage uses three NERVA-derived 25.1 klbf engines. It also includes RCS, avionics, power, long-duration CFM hardware (e.g., COLDEST design, ZBO cryo-coolers) and AR&D capability. The star truss uses Gr/Ep composite material & the LH2 drop tanks use a passive TPS. Interface structure includes fluid transfer, electrical, and communications lines. 16

  17. Artificial Gravity Bimodal NTR MTV Option:“Copernicus –B” • 3 – 25 klbf Cermet-fuel BNTRs (ESCORT)• Tex ~2700 K, Isp ~911 s, T/Weng ~5.52 • IMLEO ~330 t (6 crew) • Total Mission Burn Time: ~77.4 min • Largest Single Burn: ~43.5 min • No. Restarts: 3 • Copernicus – B is an AG version of DRA 5.0 ”0-gE” NTR crewed MTV that uses its BNTR engines to generate both high thrust & electrical power. No large Sun-tracking PVAs (~3.5 t) are required • Copernicus – B uses 3 – 25 kWe Brayton Rotating Units (~2.63 t) each operating at 2/3rd of rated power (~17 kWe ) to produce the 50 kWe needed to operate the MTV • Brayton units are located within the propulsion stage thrust structure that also supports an ~71 m2 conical radiator mounted to its exterior • Vehicle rotation at 3.0 – 5.2 rpm provides a 0.38 – 1gE AG environment for the crew. A Mars gravity field is provided on the outbound mission leg. On the inbound leg, the rotation rate is gradually increased to help the crew readjust to Earth’s gravity level Vehicle rotation about its center-of-mass provides AG environment for the crew out to Mars and back 17

  18. NTP provides high thrust (10’s of klbf) with an ~100% increase in Isp over LOX/LH2 chemical propulsion (from 450 to 900 s) • NTP can transition to higher temperature binary, then ternary carbide fuels (Isp ~950 - 1050 s) • • “Bimodal” engines (BNTR) can produce modest electrical power (~15-25 kWe) to run the space-craft eliminating large Sun-tracking PVAs andallowing AG to improve crew health and fitness • The NTP engine can also be outfitted with an “LOX-Afterburner” nozzle and propellant feed system allowing supersonic combustion down-stream of the nozzle throat thereby enabling variable thrust and Isp operation depending on the O/H mixture used • The “LOX-Augmented” NTR (LANTR) can utilize extraterrestrial sources of H2O, ice to extend the range of human exploration throughout the Solar System without the need for very advanced, lower TRL systems • Coupling higher power BNTRs with EP in “hybrid” ~1.0 MWe BNTEP system offers performance comparable to low , 10 MWe “all NEP” system NTR Options Exist for Power Generation, Thrust Augmentation & Hybrid Propulsion Nuclear Thermal Propulsion: “The Next Evolutionary Step” in High Performance Liquid Rocket Propulsion Aerojet / GRC Non-Nuclear O2 “Afterburner” Nozzle Test “Revolutionary Capability in an Evolutionary Manner” 18

  19. Notional NTP Foundational Technology Development and System Technology Demonstration Schedule AES NCPS Project is Focused on Foundational Technology Development Lunar NTR Stage SOTA Reactor Core & Engine Modeling NERVA“Composite” Fuel Crew Return in MAV Small “Fuel-Rich”Engine Hot GasSource Fuel Element IrradiationTesting in ATR at INL “Cermet” Fuel Ground & Flight Technology Demonstrators Affordable SAFE Ground Testingat the Nevada Test Site (NTS) NTR Element Environmental Simulator (NTREES) 19

  20. Mission versatility increased with smaller (15-25 klbf) NTR engines; the time and cost to design, build, test and fly is also reduced Size Comparison of RL 10B-2 and Lower Thrust NTR Engine Designs NTR “Key Performance Parameters” (KPPs): Tex ~2700 K, pch ~1000 psia, Nozzle Area Ratio () ~300:1, Isp ~910 s DRA 5.0 Fvac: 25-klbf ~525 MWt RL10B-2 KPPs: Tex ~3167 K, pch ~620 psia, Nozzle AR () ~285:1, Isp ~463 s GTD / FTD Enginein 2020 / 2023 Fvac: 15-klbf ~315 MWt RL10B-2 Fvac: 24.75-klbf Used on Delta IV Fvac: 5 - 7.5-klbf ~105 MWt 6.23 m 20.5 ft 5.36 m 17.6 ft 4.19 m 13 ft 4.27 m 13.9 ft 1.87 m 6.13 ft 2.16 m 7 ft 0.84 m 2.74 ft 1.45 m 4.75 ft Ref: Russ Joyner, PWR 20

  21. NTP Stage Approach for Flight Demo Atlas 5 • NTP stage concept can be leveraged from Delta 4 DCSS of the same diameter and approx. length Delta 4 ~40-ft (12.2 m) ~16-ft (5 m) • Remove LO2 Tank, Lines, Valves • Remove RL10B-2 Use Elements ofLO2/LH2 Delta 4 Cryogenic Second Stage (DCSS) • Add small NTP with retractable nozzle skirt • Increase LH2 lines • Similar thrust structure • NTP Cryogenic Stage for FTD can Be Made Affordable via Delta 4 Cryogenic Second Stage Components 2012 GLEX Conference, Washington, DC, May 22 - 24 21

  22. Frequently Asked Questions about NTP • Launching Nuclear Systems: • Fission reactor systems (fission surface power or NTR engines) have negligible quantities of radioactive material within them prior to being operated (few 100 Curies vs 400,000 Curies in Cassini’s 3 RTGs) • Fission product buildup only becomes appreciable at the end of the TMI burn as the MTV is departing Earth orbit for heliocentric space • Fission systems designed to generate thermal power not to explode • Inadvertent criticality accidents prevented by design safety features (e.g., neutron poison wires, control drum interlocks) or reactor design (e.g., cermet fuel NTR operating on fast neutrons) • Improvements in fuel element CVD coatings and claddings expected to significantly reduce or eliminate fission product gas release within the engine’s hydrogen exhaust • Cost for Engine Development & Ground Testing will not “break the bank”: • Separate effects tests (non-nuclear, hot H2 testing under prototypic operating conditions -- pch, temperature & H2 flow -- in NTREES followed irradiation testing in ATR will validate fuel element design • Small engine (5 klbf) scalable to higher thrust levels will be developed, ground, then flight tested first using a common fuel element design • Lower thrust-class engines (up to ~25 klbf) can use / adapt existing RL10-derived engine components (per discussions with PWR) • Small engine size, SAFE ground test approach and use of Nevada Test Site (NTS) assets (e.g., Device Assembly Facility for 0-power critical tests), mobile control trailers, etc., rather than large fixed test structures, indicate lower costs. Recent estimates (Dec. 2011) from the NSTec and the NTS for the SAFE capital cost are ~45 M$ (site and all supporting equipment) with ~2 M$ recurring cost for each additional engine test SAFE: Subsurface Active Filtration of Exhaust; also know as “Borehole” Phoebus-2A – 5000 MWt / 250 klbf NTR engine being transported to Test Cell C at NTS in 1968. Note technicians riding at the front of the engine NTR Element Environmental Simulator (NTREES) Affordable SAFE Testing 22

  23. Summary of Results and Key “Take Away” Points on NTP • Nuclear Thermal Propulsion (NTP) is a proven technology; 20 NTR / reactors designed, built and tested at the Nevada Test Site (NTS) in the Rover / NERVA programs • “All the requirements for a human mission to Mars were demonstrated” – thrust level, hydrogen exhaust temperature, max burn duration, total burn time at power, #restarts • The smallest engine tested in the Rover program, the 25 klbf “Pewee” engine, is sufficient for human Mars missions when used in a clustered engine arrangement – No major scale ups are required as with other advanced propulsion / power systems • In less than 5 years, 4 different thrust engines tested (50, 75, 250, 25 klbf – in that order) using a common fuel element design – Pewee was the highest performing engine • “Common fuel element” approach used in the AISP / NCPS projects to design a small (~7.5 klbf), affordable engine for ground testing by 2020 followed by a flight technology demonstration mission in 2023. PWR sees strong synergy between NTP and chemical • SAFE (Subsurface Active Filtration of Exhaust) ground testing at NTS is baseline; capital cost for test HDW is ~45 M$ with ~ 2M$ for each additional engine test (NTS Dec. 2011) • Cost for engine development and ground testing will not “break the bank” & the system will have broad application ranging from robotic to human exploration missions 23

  24. Summary of Results and Key “Take Away” Points on NTP • NTP consistently identified as “preferred propulsion option” for human Mars missions: - NASA’s SEI – Stafford Report (1991) listed NTP as #2 priority after HLV - NASA’s Mars Design Reference Missions (DRMs) 1 (1993) – 4 (1999) - NASA’s Design Reference Architecture (DRA) 5.0 (2009) • Using NTP, the launch mass savings over “All Chemical” and “Chemical / Aerobrake” systems amounts to 400+ metric tons (~ISS mass) or ~4 or more HLVs. At ~1 B$ per HLV, the launch vehicle cost savings alone can pay for NTP development effort • The DRA 5.0 crewed MTV “Copernicus” has significant capability allowing reusable “1-yr” NEA missions & short (~1.5 yrs) Mars / Phobos orbital missions before a landing • JSC’s “NEA Accessibility Study” presented by Bret Drake to Doug Cooke (April 7, 2011). Findings: NTR outperforms chemical, SEP/Chemical & all SEP systems, allowing access to more NEAs over larger range of sizes and round trip times for fewer HLV launches. • With more LH2, faster “1-way” transit times to from Mars are possible if desired • Lastly, NTP has significant growth capability (other fuels, bimodal & LANTR operation) 24

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