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Technology for Fuel Depots (Cont.) Subcooling Propellant

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Cryogenic Propellant Depots Design Concepts and Risk Reduction Activities Future InIn--SpaceSpace Operations ((FISOFISO)) March 2, 2011 Christopher McLean 303303--939939--71337133 [email protected] Introduction The capability to provide on-orbit cryogenic refueling for LEO departure stages represents a paradigm shift in the architecture required to support: ─ NASA’s Exploration program ─ Deep-space robot missions ─ National security missions ─ Commercial missions Fuel depots enables large, beyond LEO missions without super heavy lift vehicles This discussion covers an evolutionary approach to flight demonstrate key technologies required for operational fuel depots: ─ Low cost Missions of Opportunity ($50M – $100M) ─ Technology Demonstration Missions (TDM’s) ($150M – $250M) ─ Flagship Technology Demos (FTD’s) ($400M - $1B) Technology developed for these cryogenic fuel depots also increases robustness and capacity of existing launch platforms ─ Technologies to reduce cryogenic propellant boil-off also enhance long-term (>24 hours) storage of cryogenic propellants ─ Increases operational flexibility Page_2 State of the Art Cryogenic Propulsion Systems Current cryogenic propulsion stages rapidly lose residual propellant once on orbit Studies for the Exploration EDS resulted in changing ConOps ─ Initial goal was launch with 90 day on-orbit dwell in LEO ─ Final goal reduced to 4 days due to boil-off rates, desire not to employ active cooling ─ 4 day LEO dwell results in significant system level constraints Cryogenic boost vehicles employ Spray-on-Foam Insulation (SOFI) ─ Mitigates build up of liquid oxygen on tank external surfaces prior to launch ─ Once in orbit, provides no significant insulation capability and is a system mass penalty Long duration cryogenic storage for LOX and LH2 demonstrated on every STS flight ─ Power Reactant Storage and Distribution (PRSD) tanks employ vacuum shells for ground hold, in-flight thermal isolation ─ Vacuum shell not applicable for large propulsion stages Long duration, on-orbit passive storage has been demonstrated for space telescopes ─ Spitzer: super fluid helium, 66 months ─ Wide-Field Infrared Survey Explorer (WISE): solid hydrogen, 9 months Page_3 Exploration Technology Investments Improvements in cryogenic propellant storage technology represents the largest potential mass savings for NASA Space Technology investment Robert D. Braun, NASA Chief Technologist, May 5, 2010 “Investment in the Future: Overview of NASA’s Space Technology” Page_4 Mission Capability On-orbit cryogenic propellant refueling can dramatically impact Exploration architecture Topping ARES V EDS in LEO adds significant payload for TMI, LLO, L2, and TLI missions Page_5 Evolutionary Fuel Depot Technology Development Roadmap Flagship Technology Mission of Opportunity Technology Demonstration Mission Demonstration – $50M-$100M projected budget – $150M–$200M projected budget – $400M–$1B projected – 1.5–2-yr Development – 3-yr Development budget – 5-yr Development CRYOTE Lite CRYOTE Free Flyer – Full Stand Alone Mission with • 2 year development • 2 year development plus competitive a Dedicated Launch Vehicle • Experiment duration ~17 hr bid lead time • Baseline cryogenic experiments • Experiment duration ~1 yr CRYOSTAT (PTSD) Separate spacecraft from the • • Long term in-space cryogenic Centaur propellant storage CRYOTE Pup • BCP-100 with 3-axis control, solar • Automated rendezvous and power and communications • 2 year development docking • Accommodates multiple NASA • Experiment duration 3 mon - 1 yr • Multiple cryogenic propellant experiments • BCP-100 with 3-axis control, solar power transfer demonstrations and communications • Mission lasts at least 1 yr • Accommodates multiple NASA experiments Perceived NASA goals • Simple Depot potential CRYOSTAT platform All CRYOTE missions elevate multiple technologies to TRL 9 Page_6 CRYogenic Orbital Testbed (CRYOTE) CRYogenic Orbital Testbed (CRYOTE): experimental test platform utilizing residual launch vehicle propellants to validate cryogenic propellant storage and transfer technologies CRYOTE is comprised of two main subsystems ─ BCP-100 spacecraft bus ─ CRYOTE Core CRYOTE BCP-100 spacecraft bus ─ Provides power (600 – 800W) and RS-422 interfaces for up to three experimental payloads (in addition to CRYOTE Core) ─ Provides Attitude Determination and Control System (ADCS), Guidance Navigation and Control (GNC), Command and Data Handling (C&DH), Power, and Thermal Control ─ Attitude control provided by cold gas H2 thrusters or hydrazine RCS( mission dependent) CRYOTE Core ─ 1000 liter cryogenic propellant tank adaptable to multiple technology demonstration missions ─ With BCP-100 can remain attached to Centaur to simulate fuel depot or independent fuel transfer vehicle as a Free Flyer Page_7 CRYOTE Core (for Lite, Pup, Free Flyer) Propellant transfer methodologies ─ Residual cryogenic propellant transfer between from Centaur to CRYOTE S/C ─ Vented, non-vented, pump transfer methods can implemented Propellant Storage ─ Thermal design includes flight heritage Ball Aerospace long duration cryogenic dewar technologies ─ Testbed allows for accurate measurement of system heat leaks ─ Evaluation of active/passive technologies (pressure duty cycle, cryocoolers, disconnecting struts, MLI, internal tank heat exchangers etc.) possible by measurement of the impact to total system heat leak Advanced MLI ─ Flight evaluation of advanced MLI technologies such as LDMLI, IMLI, LRMLI Passive Thermodynamic Vent System (TVS) ─ Baseline includes TVS cooling using tank propellant expanded across expansion orifice with direct heat exchange to fluid through tank walls (walls maintained wetted by PMD) ─ Baseline includes TVS outlet flow coupled to Vapor Cooled Shield (VCS) ~ temperature matched in MLI Microgravity mass gauging ─ Core includes diode rake to verify fill level in low (10^-4 g) transfer environment ─ Spun tank design allows mission specific tank features to support multiple experiments ‘zero-g’ mass gauging technologies such as the Sierra Lobo CryoTracker, RF mass gauging HD video for correlation etc. Propellant Management Device ─ Baseline PMD provides gas-free and liquid-free access to the stored cryogen ─ Alternate PMD/ Liquid Acquisition Device (LAD) technologies can replace/augment Core PMD Page_8 CRYOTE Core Configurations CRYOTE Options CRYOTE Lite CRYOTE Pup CRYOTE Free Flier Duration on orbit <17 hours ~1 year ~1 year Lead Time to ILC 2 years 2 years ATP+2 yrs (minimum) Power Centaur batteries Solar Arrays Solar arrays Cold gas thrusters using Vented hydrogen, torque Attitude Control Centaur hydrazine vented hydrogen, storable rods, control gyros, propellants storable propellants Telemetry Centaur MDU CRYOTE antennas CRYOTE antennas Closed-loop Ball BCP Closed-loop Ball BCP Flight Computer Centaur - open-loop avionics avionics Upper Stage LV No seperation from No seperation from Seperation from Centaur Separation Centaur Centaur LH2 Source* Centaur residual LH2 Centaur residual LH2 Centaur residual LH2 Tank Size 1000L; 48" dia 1000L; 48" dia Can be >1000L; 48 dia * Potential LOX transfer with modifications to Centaur Page_9 CRYOTE Core Subsystem Chart TVS/vapor cooling (2X) -Z propulsive venting cryogenic flanges (2x) upper shell valving (dome mounted) ESPA access port cryogenic struts (6x) propellant management device girth weld GSO battery kit (2x) CRYOTE Lite only C-29 mounting interface C-29 lower shell valving (dome mounted) MLI not shown for clarity Page_10 Simple Depot as Primary Mission A Centaur derived fuel depot can be deployed in the middle of the decade ─ Demonstrates capabilities required for NASA’s CRYOSTAT program with heritage space transportation system hardware ─ Support of Orion missions to the Earth Moon Lagrange points, lunar fly-by missions possible Technologies and infrastructure saleable to larger depots He & LO2 Storage LH2 Storage (48 m 3; 53 mT) GO2 (110 m 3; 7.6 mT) L=16.3m L= 9.6m L= Centaur Mission Module Module Module New Lines New Lines Existing LO2 LO2 LH2 LH2 Page_11 Simple Depot – LH2 Module Modified Centaur tank Launched with helium to optimize design Tether counter weight for on-orbit thermal performance ─ No SOFI required LH2 Tank (launched empty) All plumbing connected to mission module side of LH2 tank ; 7.6 mT) Integrated MLI blanket 3 ─ Plumbing to enable liquid hydrogen to -Vapor cooled shield be transferred in and out of the tank -Broad area cooling LH2 Storage LH2 (110m -MMOD ─ Pneumatic tubes pressurize the tank to support transferring LH 2 out of the tank ─ Vent plumbing allows tank pressure Centrifugal Force Centrifugal Vapor cool tank control as well as a pressure sense line GH2 ullage to measure tank pressure Composite Strut ─ Vent and pressure control valves are Transfer Valve closely coupled with the top of the tank Pressurization line to vent the lines, reducing conduction Wiring Employs TVS, VCS, Ball Integrated MLI Page_12 Simple Depot – LOX module Modified Centaur stage Cryocooler Centaur LH2 tank used for LOX storage Transfer Port(s) ─ Thermally efficient upper stage completely encapsulated with MLI Composite Adapter ─ Centaur LH2 tank is efficient if there is minimal O2 pres/vent line LO2 Transfer Line thermal gradient across common bulkhead Centaur LOX tank used for thermal isolation of LH2 vapor cooling to condition LO2 ; mT) ; 54 LOX with helium and GO2 3 Several modifications to the Centaur are Integrated MLI blanket LO2 Storage -Vapor cooled shield (47mt required to enable its
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