Cryogenic Propellant Depots Design Concepts and Risk Reduction Activities
Future InIn--SpaceSpace Operations (FISO( FISO))
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 Centaur Mission Mission Module Module Module NewLines ExistingLines LO2 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 use as the Simple -Broad area cooling Depot’s LOX module: -MMOD Centrifugal Force Centrifugal ─ Add valves to close existing Centaur purge lines LH2 vapor cool tank ─ New plumbing to allow for transfer of propellants. GO2 ─ Addition of TVS circumnavigating the Centaur to & He Purge shut-off valve reduce heating and maintain the LOX at the GH2 cool engine beam desired temperature Warm GH2 RCS ─ Cold gas nozzles added to the aft end of Centaur to vent the spent GH2 which is used to maintain Existing Lines New Lines the depot control ─ New wiring to support all of the new valves and instrumentation
Page_13 Simple Depot – Mission Module
Mission module resides between LOX and LH2 modules The mission module contains: ─ Flight computer, solar panels, batteries, fluid controls, avionics ─ Remote berthing arm and docking and fluid transfer ports Derived from CRYOTE Pup or Free Flyer with a variant of a Ball Commercial Platform (BCP) bus
ESPA structure Avionics Propellant control Transfer pumps Solar power Grapple arm Transfer Port(s)
Page_14 Simple Depot Extensibility to Larger Depots
5 m LH2 tank enables 70 mT depot capacity
5 m ACES and LH2 tank enables 120 mT capacity
5 m ACES and 6.5 m LH 2 tank enables 200 mT capacity
Page_15 TRL Increases with CRYOTE Experiments
Cryo Transfer Technology Current TRL TRL Post-CRYOTE TRL Post-CRYOTE Lite Pup, Free Flier 0-g Stld 0-g 10 -4 g 0-g 10 -4 g TransferSystemOperation 4 5 4 9 9* 9 PressureControl 4 9 6 9 9 9 LowAccelerationSettling N/A 9 N/A 9 N/A 9 Tankfilloperation 4 5 4 9 9* 9 ThermodynamicVentSystem 5 5 7 7 9 9 Multi-layerinsulation(MLI) 9 9 9 9 9 9 IntegratedMLI(MMOD) 6(2) 6(2) 9(7) 9(7) 9 9
Vapor Cooling (H 2 para-ortho)* 9(4) 9(4) 9 9 9 9 PassiveBroadAreaCooling(active) 9(4) 9(4) 9(4) 9(4) 9 9 Activecooling(20–80K)* 4 4 4 4 9 9 Ullageandliquidstratification 3 9 9 9 9 9 Propellantacquisition 2 9 9 9 9 9 Massgauging 3 9 9* 9 9* 9 Propellantexpulsionefficiency 3 9 9 9 9 9 Systemchilldown 4 5 4 9 9 9 SubcoolingP>1atm(P<1atm) 9(5) 9(5) 9(5) 9(5) 9(5) 9(5) Fluidcoupling 3 3 3 3 9 9
* Additional experiment to CRYOTE Core
Page_16 Technology for Fuel Depots
Zero-g cryogenic propellant transfer ─ The fluidic and heat transfer process associated with transfer of cryogenic propellants should be validated in zero-g prior to investments in FTD and operational depots ─ Optimization of transfer ConOps, including tank pre-chill, venting, transfer line conditioning, etc. can be validated in low-cost, medium risk programs such as CRYOTE prior to implementation of high cost FTD missions such as Simple Depot or CRYOSTAT Thermo-dynamic Vent System ─ Tank boil-off can be flash-boiled across orifice (FR-1 below), which results in a source of cooling to intercept system heat leaks ─ TVS also used to cool cryogenic propellants ─ For CRYOTE Core, PMD maintains wetted tank walls (heat exchanger can be located in tank)
Page_17 Technology for Fuel Depots (cont.)
Mechanical supports and thermal isolation ─ Mechanical attachment of cryogenic propellant tank to ‘ambient’ structure significant contributor to overall system heat leak ─ Various heritage launch vehicle technologies include both struts and conic sections ─ Designs concurrently optimized for structural and thermal performance ─ Separating strut technology can provide complete thermal isolation (LM PODS) ─ Cooling of struts to improve thermal performance demonstrated on-orbit ─ Launch of empty tanks reduces required cross section, reducing heat leak
CRYOTE Core implements CRYOTE ground test Ball Aerospace heritage Ball Aerospace cryogenic article conic section cryogenic strut technologies struts on Spitzer
Page_18 Technology for Fuel Depots (cont.)
Propellant Management and Acquisition technology ─ The ability to have guaranteed access to liquid or gaseous phases of stored cryogenic propellants is required for multiple fuel depot concepts ─ Access to gas phase allows to pressure cycle tanks with minimum propellant consumption ─ Access to liquid phase provides single phase during transfer, liquid source for TVS and RCS ─ Spin settling of the Simple Depot provides phase separation: transfer methodologies need further evaluation CRYOTE Core Propellant Management Device ─ Wets walls of tank to providing heat exchange interface to TVS, access to liquid for TVS operation, liquid has higher heat capacity ─ Maintains gas bubble in tank center for tank pressure cycling, cold gas thruster operation
Page_19 Technology for Fuel Depots (cont.)
Low acceleration settling ─ Rotation of fuel depots allows for gas/liquid phase separation using acceleration ─ Flight heritage approach for pressure control ─ Reducing liquid heating – much of the heating is absorbed by the ullage ─ Reducing ullage mass – provides warmer, less dense ullage
Liquid
He & LO2 Storage LH2 Storage (48 m 3; 53 mT) GO2 (110 m 3; 7.6 mT) Ullage
Page_20 Technology for Fuel Depots (cont.)
Integrated Multilayer Insulation (IMLI) Advanced MLI developed Insulation by Ball Aerospace and Quest Product Development Successfully completing a Phase II SBIR, TRL 6 achieved Uses polymer spacers instead of netting ─ Enables precise control over layer spacing ─ Not affect by gravity IMLI is a replacement for conventional MLI ─ More robust, layers bonded together ─ Lower heat leak or fewer layers and less mass Aluminized Mylar Layers
Polymer Spacers
Integrated MLI Concept Page_21 Technology for Fuel Depots (cont.)
Load Responsive MLI (LRMLI) Uses polymer spacers to support a thin vacuum shell Spacers elastically compress with atmospheric load and disconnect with low pressure Being developed by Ball and Quest Product Development Successfully demonstrated in a Phase I SBIR, TRL 4 achieved. Development continuing on a Phase II SBIR.
Page_22 Technology for Fuel Depots (cont.)
Micrometeroid/Orbital Debris Protection • MMOD-IMLI uses engineered polymer spacers to provide precise layer spacing and combinations of layer materials to provide excellent MMOD protection combined with thermal insulation • MMOD-IMLI with 0.070” inter-layer spacing has a modeled energy dissipation that predict 9-fold fewer layers than conventional MLI to stop particle penetration
Page_23 Technology for Fuel Depots (cont.)
Enhanced Area Cooling Vapor Cooled Shield (VCS) ─ Cryogenic boil-off is used to reduce tank area heat leak contributions by inserting a cooled shield within the MLI blanket layup ─ Shield is temperature matched within the blanket to provide highest performance ─ Part of baseline thermal system in CRYOTE Core ─ Ball Aerospace flight heritage approach Broad Area Cooling (BAC) ─ Similar to VCS, however employs a cryocooler to intercept significantly more heat than VCS ─ Ground demonstrated by Ball/ NASA GRC
Page_24 Technology for Fuel Depots (cont.)
Mass gauging Accelerated mass gauging, with settled cryogenic propellants, allows level measurements to be obtained with temperature sensors or capacitance probes Measurements of propellant quantities in zero-g are more difficult to obtain due to the uncertainty of the gaseous phase propellant location within the tank The CRYOTE Core can host an array of zero-g Early CRYOTE Core concept cryogenic mass gauging technologies with CryoTracker ® ─ Sierra-Lobo CryoTracker ® ─ RF-mass gauging ─ Optical sensing ─ Others Unlike storable propellants, PVT measurements are difficult to implement due to propellant loss from boil- off and constantly varying tank temperature CRYOTE Core diode rack for level measurement during settled transfer
Page_25 Technology for Fuel Depots (cont.)
Sierra-Lobo CryoTracker ® Accurately measures fluid level and temperature profiles Calculates fluid mass and uncertainty in fluid mass Commercially available system Probe, feed-through, electronics, software Patented technology ─ U.S. Pat. No. 6,431,750 ─ U.S. Pat. No. 7,043,925 Lightweight, one-piece, flexible at cryogenic temperatures Easily customized to any length and width Designed for flight, easy to install, and clean Fluids used in: liquid hydrogen, nitrogen, oxygen, methane, kerosene, and water Applications include: launch vehicle, spacecraft, R&D facilities, and process industries
Page_26 Technology for Fuel Depots (cont.)
Sun Shields ─ Propellant depot located in LEO gets substantial radiation and reflection from the Earth ─ Conic sun shield that can protect the depots cryogenic tanks from both the Sun and Earth ─ Provides a view to space to reject heat ─ JWST uses a mechanical boom to deploy its sun shield ─ A pneumatic boom, inflated with waste GH2, can be used to deploy and support the sun shield (below)
Conic sun shield shades the ULA, ILC-Dover and NASA have developed a cryogenic upper stage to pneumatically deployed sun shield to minimize boil-off support long duration cryo-storage
Page_27 Technology for Fuel Depots (cont.)
Subcooling Propellant Subcooling represents a simple technique that can extend the operational life of a spacecraft, upper stage, or an in-space cryogenic depot for months with minimal mass penalty The heat capacity of the chilled hydrogen allows it to absorb the large quantities of energy that leak into the tank over time without the need to vent the cryogen, thus extending its in-space vent-free ‘hold-time’ Subcooling hydrogen to 16 K at 1 atmosphere pressure prior to launch triples its vent-free hold-time over hydrogen loaded into tanks at its normal boiling point of 20 K at 1 atmosphere The TCS concept for pressure subcooling cryogens on the launch pad. Courtesy NASA
Page_28 Technology for Fuel Depots (cont.)
Integrated vehicle fluids ─ Today’s DCSS missions are limited to three burns over the course of an eight hour mission due to the limited supply of hydrazine, helium and electrical power ─ To satisfy the existing range of missions from LEO to Medium Earth Orbit (MEO), GTO, and geosynchronous orbit (GSO), a variety of mission kits are required ─ To improve mission flexibility, an Integrated Vehicle Fluids (IVF) system to allow the use of hydrogen and oxygen from the upper stage primary tanks to satisfy the settling, attitude control, pressurization, and power requirements ─ The IVF will allow the elimination of hydrazine and helium from the vehicle while replacing the existing large capacity batteries with small rechargeable batteries ─ Development of the hydrogen/oxygen thruster, engine, pump, alternator, rechargeable battery and cryogenic compatible composite bottles are progressing well with concept testing under way ─ IVF will be especially valuable in a depot based transportation economy by eliminating the need to store and transfer additional commodities such as hydrazine and helium
The integrated vehicle fluids Testing has demonstrated module is designed to the functionality of this low support RCS, pneumatic cost hydrogen/oxygen and power requirements thruster (Credit Innovative (Courtesy ULA) Engineering Solutions)
Page_29 Summary
A significant paradigm shift in LEO departure capability is possible with on-orbit fueling
NASA and industry have recognized the importance of developing technologies to improve the capability to store and transfer cryogenic propellants
The use of low cost, high payoff precursor technology demonstration platforms such as CRYOTE Core minimize the risk to FTD class missions
FTD fuel depot missions can be developed using existing flight heritage designs, manufacturing and integration processes such as the Simple Depot based on heritage Centaur derived technologies
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