. Space Transportation Infrastructure Supported by Propellant Depots
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WHAT IS IN SITU RESOURCE UTILIZATION (ISRU)?
ISRU involves any hardware or operation that harnesses and utilizes ‘in-situ’ resources to create products and services for robotic and human exploration.
(1) Resource Assessment (Prospecting)/Assessment and mapping of physical, mineral, chemical, and volatile/water resources, terrain, geology, and environment.
(2) Resource Acquisition /Extraction, excavation, transfer, and preparation/ beneficiation before Processing.
(3) Resource Processing- Consumable Production /Processing resources into products with immediate use or as feedstock for construction and/or manufacturing Propellants, life support gases, fuel cell reactants, etc.
(4) In Situ Manufacturing / Production of replacement parts, complex products, machines, and integrated systems from feedstock derived from one or more processed resources.
(5) In Situ Construction /Civil engineering, infrastructure emplacement and structure construction using materials produced from in situ resources Radiation shields, landing pads, roads, berms, habitats, etc.
(6) In Situ Energy /Generation and storage of electrical, thermal, and chemical energy with in situ derived materials Solar arrays, thermal storage and energy, chemical batteries, etc.
CREDIT: Johnson Space Center Engineering Directorate / L-8: In-Situ Resource Utilization Capabilities by Jerry Sanders, November 2016.
Space transportation infrastructure supported by propellant depots
A space transportation infrastructure utilizes expendable launch vehicles as Delta IV Heavy, Atlas V and Falcon 9 for all crew, cargo and propellant launches to orbit.
First, propellant is launched to a Low-Earth-Orbit (LEO) Depot and at a Earth-Moon Lagrange Point 1 (L1) Depot to support new reusable in-space transportation vehicles. The LEO Depot supports missions to Geosynchronous Earth Orbit (GEO) for satellite servicing and to L1 for L1 Depot missions. The L1 Depot supports Lunar, Earth-Sun Lagrange 2 (ESL2), Asteroids and Mars missions. A Mars Orbital Depot is also described to support ongoing Mars missions.
Figure & Text Credit: David Smitherman, NASA Marshall Space Flight Center & Gordon Woodcock, Gray Research.
The new concepts of vehicle designs can be launched on current 5-meter diameter Expendable Launch Vehicles. These vehicles are based on International Space Station (ISS) heritage hardware. The high-energy depots at L1 and Mars orbit are compatible with, but do not require, electric propulsion Tug use for propellant and/or cargo delivery. New reusable in-space crew transportation vehicles include a Crew Transfer Vehicle (CTV) for crew transportation between the LEO Depot and the L1 Depot, a new reusable Lunar Lander for crew transportation between the L1 Depot and the lunar surface, and a Deep Space Habitat (DSH) to support crew missions from the L1 Depot to ESL2, Asteroid, and Mars destinations.
Do you remember? From 1967 to 1973, NASA utilized the Saturn V as a heavy-lift launch vehicle for human missions to the Moon. The gross lift-off mass was ~3,039,000 kg with a total payload mass of ~45,018 kg and a propulsion and propellant mass of ~2,993,982 kg. The payloads consisted of a 3.9 m diameter Command Module at ~5,809 kg, a 3.9 m diameter Service Module at ~24,523 kg, and a 4.3 m diameter Lunar Module in its folded configuration at 14,696 kg including propellants and consumables in each. So approximately 98.5% of the Saturn V mass was propellant and propulsion systems.
Today, the remaining 1.5% payload mass could be broken down and launched on current commercially available ELV systems. This means that space transportation is not a payload delivery problem, but a propellant and propulsion management problem. Here enters the idea for propellant depots in an infrastructure that can provide the same capabilities as the Saturn V for lunar missions with more flexibility to accommodate a variety of other missions simultaneously. In addition to propellant supply, the depots are servicing platforms where propellant is stored and transferred to the reusable vehicles as needed, vehicle maintenance and upgrades can be accomplished. More, crews with their logistics and payloads can be transferred between vehicles.
Development of this infrastructure begins with the LEO Depot and grows incrementally for access to L1, the Moon, and then Mars. The LEO depot is used repeatedly for all human missions, and the L1 Depot is used repeatedly for all human Lunar, ESL2, Asteroid, and Mars missions. This operating scenario makes the depots and reusable vehicles part of a permanent infrastructure that could eventually support dozens of missions simultaneously, both commercial and exploration oriented, for decades into the future. First thing first: LEO Infrastructure Buildup Satellite servicing could be the initial capability using a reusable CTV and a Reusable Upper Stage (RUS) operating out of a LEO Depot.
A LEO Depot is shown based on ISS heritage hardware consisting of a truss section with docking ports, solar arrays, radiators, sunshield protecting a propellant storage tank, and a pressurized node module. Attached vehicles include a Reusable Upper Stage, reusable Crew Transfer Vehicle, and a Crew Return Vehicle. It is similar in size to the ISS when it was under construction and had only one set of solar arrays in place.
The ISS orbits the Earth at a ~400 km altitude and a 51.6-degree inclination. The LEO Depot is envisioned to orbit the Earth under the ISS at a ~350 km altitude and a ~28.5-degree inclination. An approximate mass for the LEO Depot was estimated based on ISS hardware at ~43.5 MT. This infrastructure element is established in four launches, which includes: the depot truss section with attachments, the depot pressurized Node module, an RUS, and a CTV. An additional CRV crew mission can be included if human assembly operations are needed as done for the ISS.
GEO Satellite Servicing
With a LEO Depot, a Reusable Upper Stage and a reusable Crew Transfer Vehicle, human satellite servicing missions can be done as needed with 3 propellant launches and 1 crew launch per mission.
See Appendix A-1 for GEO Satellite Servicing Reference Profile details. The truss section is designed to handle all power and propellant requirements such that a variety of vehicles can dock to the truss for propellant transfer. The pressurized Node element is designed to handle transfers between all crew vehicles and to support servicing utilizing an airlock for extra-vehicular activity (EVA) and/or a pressurized human Free-Flyer vehicle. Each end of the platform is expandable by adding a truss section for additional power and propellant transfer capabilities at one end, and additional pressurized modules added to the Node at the other end.
The RUS has a dry-mass of ~4,300 kg and can hold ~32 MT of propellant. It is equipped with a docking adapter for propellant transfer and a structural truss adapter for attachment to the aft end of the CTV. Future RUS missions from the L1 Depot indicate that variations of the RUS could use structural truss adapters at both ends for multiple RUS stages and the addition of aero capture features with a spherical nose and tail flare for return from higher energy orbits. An aero capture system was estimated to add ~4,500 kg mass.
The CTV and the RUS hold ~32 MT of propellant each and are filled at the LEO Depot from ongoing propellant flights from a variety of vehicles with the largest capacity being ~22 MT delivered by the Delta IV Heavy and perhaps larger propellant loads by a future Falcon Heavy. This flexibility in propellant delivery means propellants can be taken out of the critical path for vehicle sizing of each mission.
Propellant can be collected on-orbit by contract from several sources, moving the largest part of the mission mass, the propellants and launch systems, into a competitive bid environment. This should have great potential for lowering overall mission cost and should help promote the commercial development of reusable launch systems for propellant delivery and other services.
The initial vehicles described here for propellant delivery include at least one reusable ~32 MT storage tanker with an active cooling system that stays at the depot with the remaining tankers being expendable. The long-term storage tanker(s) would be designed for cryogenic liquid oxygen (LOX), cryogenic liquid hydrogen (LH2), and a variety of storable propellants. In general, all the expendable tankers are used only once for propellant transfer to the LEO Depot. Once the expendable propellant tankers transfer their propellant to the reusable long-term storage tanker they are then disposed of by a re-entry maneuver to burn up the tanker in the atmosphere.
A long-term goal for this propellant delivery service is to transition to a reusable launch system, as mentioned above, that includes built-in reusable propellant tanks. Crew Transfer Vehicle. This configuration for the reusable Crew Transfer Vehicle shows a crew cabin and docking mechanism in the nose behind an open spherical heat shield, the remainder of the crew cabin behind payload bay doors, and an open bay for payloads that includes a human Free-Flyer servicing vehicle.
The CTV has a dry-mass of ~15,200 kg. A mass budget was calculated to include a 6,000 kg propulsion system, 4,200 kg for the human systems including crew cabin and servicing equipment, and 5,000 kg for the aero capture system including a nose cap and tail flare.
Figure above illustrates the conceptual layout envisioned for the CTV. Since it is designed for aero capture maneuvers, when in its closed configuration, it can be launched on an ELV without a payload fairing. The crew cabin can vary in size from that shown, and the docking port arrangement can be in the nose and/or the topside payload bay area. Both are shown in Figure 4 to illustrate the flexibility available with this configuration. The forward docking mechanism is exposed when the spherical heat shield is opened up, which facilitates docking maneuvers at the depot and facilitates multiple vehicle configurations for future deep space missions out of the L1 Depot. Large forward facing windows with sunshields are also illustrated. On the topside with the payload bay doors open is a docking port that is sized to match an ISS payload hatch. This configuration would allow the transfer of large ISS rack-sized payloads in an out of the CTV for transfer to other destinations at L1 as the infrastructure grows. Behind the crew cab is a human Free Flyer4 vehicle docked to a port for use in servicing other vehicles in space. This is an option proposed in lieu of, or in addition to, traditional EVA as done from the ISS.
The color-coding on the sides of the CTV indicates the approximant sizes of the large LH2 tank and smaller LOX tank supplying a reusable propulsion system at the aft end. Propellant transfer mechanisms to resupply these tanks with cryogenic propellants are not shown, and are a point for further research and development of this design. Conceptually, propellant transfer could be done through feed lines at the forward docking port, separate feed lines between the LH2 and LOX tanks, or a separate docking port at the aft end. These options will need further trades for safety and on-orbit operations considerations.
Lunar Lander. This configuration for a reusable Lunar Lander shows a crew cabin at the forward end with four landing engines behind each of the four deployed landing legs. Two orbital transfer and descent engines are located at the aft end.
Cargo Lander. This configuration for a reusable Cargo Lander utilizes two standard crew Landers without the crew cabin at the forward end. The payload is suspended from a truss structure between the two vehicles.
Deep Space Habitat Cut-Away View. This configuration for a Deep Space Habitat is about the size of the ISS Lab and Node modules combined. It includes ISS type racks, an internal airlock, and docking ports at each end and on the sides at the airlock location.
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In-Situ Ressources Utilizations (ISRU)
Tips to Know
A Propellant need to meet many criteria to be considered as basis for transportation architecture. First, it must have thermal stability to operate in a liquid rocket engine, means the ability to cool engine throat critical heat flux, avoiding thermal decomposition and coking in engine coolant channels, as well as offers sufficiently high engine specific impulse (Isp).
From a vehicle system, it is the combined characteristic of propellant Isp and bulk density in meeting the vehicle impulsive velocity (DeltaV) mission requirement that offers either the lowest mass or lowest propellant tank volume that warrants the selection. So...
The cryogenic Liquid Hydrogen(LH2), with its Normal Boiling Point at 36.6 degree Rankin, has an excellent gravimetric heat of combustion (energy per mass) and can generates an High Engine Specific Impulse when combusted with Liquid Oxygen (LO2). Used in launch vehicles for first stage and upper stage applications, it is the fuel of choice for In-Space Propulsion Stage because that high Isp value. LEARN MORE
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Jim Keravala, the chief operating officer and co-founder of Shackleton Energy Company Inc., a firm intending to build in-space propellant depots discusses his companies plans.
Government Futures Lab, Reconstitutional Convention / Jim Keravala, Co-founder of Shackleton Energy Company.
NEAR EARTH ASTEROIDS: ~85% OF METEORITES ARE CHONDRITES
Ordinary Chondrites 87%
FeO:Si = 0.1 to 0.5 Pyroxene Fe:Si = 0.5 to 0.8 Olivine Plagioclase Diopside Metallic Fe-Ni alloy Trioilite -FeS
-» Source of metals(Carbonyl)
Carbonaceous Chondrites 8%
Highly oxidized w/ little or no free metal Abundant volatiles: up to 20% bound water and 6% organic material -» Source of Water/volatiles
Enstatite Chondrites 5%
Highly reduced; silicates contain almost no FeO 60 to 80% silicates; Enstatite & Na-rich plagioclase 20 to 25% Fe-Ni Cr, Mn, and Ti are found as minor constituents
-» Easy source of oxygen
SPACE RESOURCES AND PRODUCTS OF INTEREST
Three major resources
(1)Regolith: oxides and metals−Ilmenite 15% −Pyroxene 50% −Olivine 15−Anorthite 20% (2)Solar wind volatiles in regolith: −Hydrogen 50 –150 ppm −Helium 3 –50 ppm −Carbon100 –150 ppm ( 3)Water/iceand other volatiles in polar shadowed craters −1-10% (LCROSS) −Thick ice (SAR)
Three major resources
(1)Atmosphere: −95.5% Carbon dioxide, −2.7% Nitrogen, −1.6% Argon (2) Water in soil: concentration dependant on location −2% to dirty ice at poles (3) Oxides and metals in the soil