Sprint 2: Return to the Moon with ISRU
Mission to Moon w/ ISRU
Sprint 2 Derived Goal Statements from Stakeholder Needs
- Government Needs:
- Continue developing a lunar-only architecture with ISRU
- Include sufficient depth in the lunar refueling elements to benefit the BAA with a focus on extensibility and reusability
- Commercial Needs:
- Identify opportunities in architecture where competition might yield benefits to both the government and the private parties. Note that for this to be valid, there would have to be a benefit to the government in the form of lower cost or higher reliability (due to multiple suppliers), and benefit to the private parties in the form of opportunity for profit, arising from innovation or process improvements.
- Include features in architecture that could have dual use for a private party such that private companies receive an anchor client, which might help to raise investment capital, and the government gets the service possibly at a reduced cost.
System Form Goals
- Lander
- Rover
- Fueling Station
System Function Goals
- Landing Equipment
- Long Duration Surface Suit
- Living Habitat
- Work Habitat
- Health Habitat
- Transportation and Handling Systems
- Communication and Navigation Systems
- Logistics Resupply
- ISRU Production
- Emergency Egress Systems
Goals Needed for Developing System
- Representative of Stakeholder Needs
- Ensures Opportunities for Commercialization
System Check on Goals
- Representative of Stakeholder Needs
- Consistent with System Function and Form Goals
Updated Morph Matrix
This Morph Matrix is an updated version I created of the matrix from Sprint 1 that incorporates lunar surface architecture and overall system considerations. New decisions that were added specifically related to the lunar surface architecture are:
- Moon ISRU: How much ISRU is expected to be done
- Moon Surface Stay Duration: How long sorties/missions are expected to stay on the surface
- Moon Mission Architecture: Whether most missions will be individual sorties or a whole campaign involving an established outpost on the lunar surface
- Moon Surface Mobility: Whether astronauts on a mission will stay in a local area once they have landed or will use rovers to travel to different locations and be able to access anywhere on the surface.
New decisions that were added specifically related to the overall system architecture are:
- Degree of Transportation Reuse: How much the same transportation elements (and/or equipment/structures related to it) are expected to be used repeatedly
- Degree of Commercial Involvement: How much commercial companies, efforts, and resources are expected to be involved in the operation of this architecture
Muramoto (talk) 21:19, 5 March 2019 (UTC)
| Decision | Option 1 | Option 2 | Option 3 | Option 4 | Option 5 | Option 6 |
|---|---|---|---|---|---|---|
| Gateway | Yes | No | ||||
| Earth Launch | LEO | L1 | L2 | Lunar Orbit | Lunar Surface | Gateway |
| Earth Orbit Rendezvous | Yes | No | ||||
| Lunar Orbit Rendezvous | Yes | No | ||||
| Lagrange Point Rendezvous | L1 Rendezvous | L1 Docking | L2 Rendezvous | L2 Docking | None | |
| Moon Surface Arrival | Direct from Earth | Direct from LEO | Arrival from L1/L2 | Arrival from LLO | Arrival from Gateway | |
| Lunar Surface Location | Near-side equator | Far-side equator | South pole | |||
| Moon ISRU | None | Medium | Extensive | |||
| Moon Surface Stay Duration | <14 Days | 14-45 Days | 45+ Days | |||
| Moon Mission Architecture | Sorties | Outpost Campaign | ||||
| Moon Surface Ops | Human EVA only | Sample Return | Unmanned Rover | Manned rover | ||
| Moon Surface Mobility | Local | Global | ||||
| Moon Departure | Gateway | LLO | L1/L2 | LEO | Earth | |
| Crew Selection | 1 | 2 | 3 | 4 | 5 | |
| Fuel Type | Solid | Liquid | Hybrid | Hypergolics | ||
| Degree of Transportation Element Reuse | None | Low | High | |||
| Degree of Commerical Involvement | None | Low | High |
Lunar Gateway and the Near-Rectalinear Halo Orbit
While the Apollo missions used a free return trajectory to low lunar orbit (LLO), improvements in orbital mathematics have created new opportunities for a lunar space station. The most popular location for a permanent station is the Lunar Gateway, which is located near the Earth-Moon Lagrange point (EML-2). The proposed orbit for the Lunar Gateway is a Non-Rectalinear Halo Orbit (NRHO) around EML-2. This orbit maintains a consistent line of sight with Earth while still positioned further from the Earth than the Moon, allowing it to relay signals to the back side of the moon. While the delta V required to rendezvous with a station in this orbit is larger than a conventional circular orbit around LLO, the Lunar Gateway allows for low-energy burns to escape the Earth-Moon system, as well as a convenient step-off point for a lunar excursion at any latitude. Andy
Delta-V Budget
| Earth | LEO | L1 | L2 | Gateway | Lunar Orbit | Lunar Surface | |
|---|---|---|---|---|---|---|---|
| Earth | 0 | 9.4 | 13.427 | 13.352 | 14.053 | 13.34 | 13.36 |
| LEO | 9.4 | 0 | 4.27 | 3.952 | 0.289 | 4.04 | 5.93 |
| L1 | 13.67 | 4.27 | 0 | 0.161 | 0 | 0.64 | 2.52 |
| L2 | 13.352 | 3.952 | 0.161 | 0 | 0 | 0.64 | 2.52 |
| Gateway | 9.689 | 0.289 | 0 | 0 | 0 | 0.82 | 2.9 |
| Lunar Orbit | 13.44 | 4.04 | 0.64 | 0.64 | 0.82 | 0 | 1.87 |
| Lunar Surface | 15.33 | 5.93 | 2.52 | 2.52 | 2.9 | 1.87 | 0 |
(Note: All values are in km/s)
Ranking Methodology for Lunar Mission Architectures
One of the tasks I set out to complete was to enumerate the various architectures that come out of our morphological matrix. After running the code, the total number of architectures, without applying any constraints, is equal to 25,920,000. That is a lot of architectures!
After applying a set of very simplistic logical constraints that focused primarily on transport to the Lunar Surface, I was able to reduce the number of feasible architectures to 691,200. That is roughly 2.6% of the entire architecture pool. As we learn more about the mission requirements, we will have a better idea of other constraints. Ericmagliarditi (talk) 01:35, 1 March 2019 (UTC)
Starting with all of the architectures created by full enumeration of the morph matrix we ranked each one by adding up the rankings of each of its components. Within each category of the morph matrix, each of the choices was given a weight with the weights from all the choices in that category adding up to one-hundred. Next, we are working to convert these rankings into our mission architecture figures of merit, so we will have several different ways of looking at the merit of each mission architecture.
| Decision | Option 1 | Option 2 | Option 3 | Option 4 | Option 5 | Option 6 | ||||||
|---|---|---|---|---|---|---|---|---|---|---|---|---|
| Gateway | Yes | 80 | No | 20 | ||||||||
| Earth Launch | LEO | 50 | L1 | 5 | L2 | 5 | Lunar Orbit | 10 | Lunar Surface | 10 | Gateway | 20 |
| Earth Orbit Rendezvous | Yes | 20 | No | 80 | ||||||||
| Lunar Orbit Rendezvous | Yes | 70 | No | 30 | ||||||||
| Lagrange Point Rendezvous | L1 Rendezvous | 20 | L1 Docking | 20 | L2 Rendezvous | 25 | L2 Docking | 25 | None | 10 | ||
| Moon Surface Arrival | Direct from Earth | 5 | Direct from LEO | 5 | Arrival from L1/L2 | 20 | Arrival from LLO | 30 | Arrival from Gateway | 40 | ||
| Lunar Surface Location | Near-side equator | 25 | Far-side equator | 5 | South pole | 70 | ||||||
| Moon ISRU | None | 15 | Medium | 60 | Extensive | 25 | ||||||
| Moon Surface Stay Duration | <14 Days | 20 | 14-45 Days | 70 | 45+ Days | 10 | ||||||
| Moon Surface Ops | Human EVA only | 15 | Sample Return | 40 | Unmanned Rover | 20 | Manned rover | 25 | ||||
| Moon Surface Mobility | Local | 45 | Global | 55 | ||||||||
| Moon Departure | Gateway | 35 | LLO | 40 | L1/L2 | 15 | LEO | 5 | Earth | 5 | ||
| Crew Selection | 1 | 5 | 2 | 10 | 3 | 25 | 4 | 40 | 5 | 20 | ||
| Fuel Type | Solid | 20 | Liquid | 25 | Hybrid | 25 | Hypergolics | 30 |
Ranking Breakdown
Gateway: Utilizing the gateway was given a heavy weight since NASA plans on utilizing the gateway in the future.
Earth Launch: We determined that launching to LEO is the most advantageous because it acts as a critical safety check prior to lunar or gateway insertion. Stopping in LEO gives planners time to check all systems before moving along in the mission.
Earth Orbit Rendezvous: We determined that rendezvous in earth orbit provided very little advantage, and just added system risk and complexity.
Lunar Orbit Rendezvous: Similar to Earth launch, stopping in lunar orbit provides a critical check before heading to the lunar surface.
Lagrange Point Rendezvous: We gave more weight to L2 compared to L1 primarily due to the fact that L2 provides exposure to the far side of the moon. However, the differences between L1 and L2 are small, and thus the rankings are quite uniform.
Moon Surface Arrival: Arriving from gateway requires a little bit more energy, but the gateway provides good exposure to the poles as well as station-keeping parameters, and thus was ranked higher than the other options.
Moon ISRU: Obviously no ISRU is not great, but extensive ISRU was deemed extremely costly and thus not a great alternative. Rather, medium ISRU capability can provide resources but hopefully within a reasonable cost range.
Moon Surface Duration: Too little time on the Moon and not enough can get accomplished, but too much time adds mass and risk to a mission. Thus the 14 - 45 day mission duration was the best choice.
Moon Surface Operations: All of these alternatives have solid rationale, but sample return, which can be in the form of ISRU resources is the goal and thus was heavily weighted.
Moon Surface Mobility: It would be great to have complete global coverage of the moon, but this could be a drag on resources and cost a great sum. Thus, both options were considered equivalent in weight, with global getting a few extra points.
Moon Departure: Similar to Earth launch, stopping in LLO is another critical step before any final and long term decisions can be made.
Crew Selection: 4 is an ideal crew size. Provides enough extra scientific return without adding extra cost.
Fuel Types: All fuel types were weighed similarly, but hypergolic were given a slightly better score since they have been proven on flight and used in the Apollo missions.
Based upon our analysis, the top 5 architectures in terms of raw score are shown below.
SSE-48: Mission Architecture Figures of Merit
- Exploration Capability: kg * crew-days-on-surface
- Commercial Merit: cost / crew-days-on-surface
- Launch Efficiency: Mass_to_lunar_surface / launch_mass
- Mass Level: initial mass in low earth orbit
- Overall Metric: mass_to_lunar_surface * number_of_crew / (delta_v * cost)
SSE-13: Moon Technologies Which Can Be Applied To Mars
Reuse from moon mission:
- Surface Habitat
- Food growing techniques
- Spacesuit
- Command and Service Modules
- Solar arrays
- Life-support systems
- ISRU techniques
- Surface rover
- Astronaut helper robots
- Launch vehicle
Develop new:
- Zero-g mitigation
- Deep space radiation protection
- Entry, Descent and Landing System
ISRU General Considerations:
ISRU is defined to include natural lunar resources and discarded materials such as a lunar descent module. Lunar reconnaissance orbiter indicates the presence of water ice buried under lunar regolith at certain locations 5% of regolith composed of water 5% of regolith contained methane, ammonia, hydrogen, carbon dioxide, and carbon monoxide Sunlight always strikes poles from low elevation above horizon Hills and craters produce persistent and even permanently shadowed regions Mountaintops may be bathed in persistent sunlight Polar shadowed regions are cold traps - impossible to warm up Apollo Rock samples revealed no water bearing minerals No ore bodies like those on earth have been discovered on moon, but some rock types have concentrations of minerals Oxygen composes roughly of lunar common soil by weight
Mission Requirements for ISRU Oxygen Outpost:
-1 Metric Ton per year for initial outpost 10 Metric ton per year for advanced outpost
ISRU Processes:
Regolith Extraction Regolith Treatment Electrolysis Oxygen cryogenic cooling Oxygen delivered to LLO, Gateway, etc
Note: ISRU requires a lot of power, thus understanding eclipse times at various landing locations is extremely important
Summary: On lunar surface, oxygen extraction seems to be the most viable due to the water that could be trapped in the cold regions within various craters. Oxygen can be used for a multitude of crew activities and lunar refueling. Unfortunately, we will not know if any of this is possible until we have ground truth of the resources. From an architecture selection standpoint, ISRU only impacts fuel requirements because you can assume a dry ascent vehicle that will fill on the lunar surface. This means you can increase mass of the descent stage, thus enabling more scientific and commercial payloads.
Ericmagliarditi (talk) 21:27, 3 March 2019 (UTC)
Failures and Contingencies
In order to conduct a thorough assessment of the potential failures and subsequent contingency plans for our architecture selection, it is clear that the architecture must have flexibility in the following four areas:
- 1. Programmatic: Our chosen Architecture shall adapt to changes in NASA priorities and budgets over several fiscal year cycles
- 2. Commercial: The Architecture shall adapt to changes in Commercial participation and changes to these companies priorities (to avoid the fate of programs like DARPA’s RSGS)
- 3. Technology: The Architecture shall adapt to changes in technological priorities and advances
- 4. Exploration: The Architecture shall adapt to changes in exploration priorities and changes in exploration methods.
We propose the following backup plan for NASA and the Commercial Partners to consider. This backup plan would see the launch vehicle lift off without astronauts on board. It would expend its first stage and second stage in turn, then its third stage would place itself plus the unmanned Command and Service Module (CSM) and Lunar Module (LM) spacecraft into parking orbit about the Earth. Since it would carry no crew, the CSM would need no Launch Escape System (LES) tower on its nose. The astronauts would reach Earth orbit separately in a ferry CSM on top of a two-stage rocket of choice. The ferry CSM would carry a special drogue - as was used for Apollo's backup - docking unit on its nose for linking up with the unmanned CSM's nose-mounted probe docking unit. The special drogue would need about one year and several million dollars to develop. Additionally, we consider the following 11 functional needs for our architecture elements in our contingency planning.
I also looked into the likelihood of the technologies being in existence particularly in relation to when our Commercial Partner would need them.
Lunar Surface Architecture
Although there are architectures that do not require landing on the lunar surface, there are many critical overall system decision points that require a surface architecture. After getting to the Moon, we have to decide what we want to accomplish there. What we can do on the Moon is dependent on what we can get there. Additionally, what we choose to bring will depend on what we hope to achieve. Therefore, the lunar architecture on a whole is both capability and mission-driven.
There are several elements that must be addressed in order to develop a full lunar surface architecture:
High Level Decisions
Complexity can also increase depending on the decisions being made, as shown below:
One of the key initial decisions that must be made to create a more defined lunar surface architecture is to choose between lunar missions that are either designed as individual sorties or more campaign-like and involve setting up semi-permanent outposts on the lunar surface. Sorties have the advantages of repeatability and flexibility, but outposts would enable more exploration and in-depth scientific research and technological development. ISRU technologies would be able to develop and mature. Permanent radio and telescope observatories would enable astronauts to study the Earth and deep space.
Landing sites are also important architectural decisions. Based on current scientific research, the lunar south pole is an area of interest because it is:
- Relatively safe due to its thermally moderate conditions
- Cost effective to land at due to a relatively small delta V and abundance of sunlight for solar power
- Rich in hydrogen, oxygen, and potentially other volatile chemicals from the polar ice, which can be used for ISRU
- Allows for flexibility with the incremental buildup of solar power, enhanced surface daylight ops, and more opportunities to launch
- Exciting because the area is unexplored and offers opportunities to learn about cold, dark craters
Architecture flexibility is essential for addressing long-term interests that may not currently be a high priority. Future lunar missions are expected to be highly capability-driven, so as incremental capability increases (due to new or incremental technology insertion), flexibility provides opportunities for relief to capability limitations. Technologies such as ISRU are unproven, and a flexible architecture allows for incremental steps towards a desired end state.
Below are some further infrastructure considerations to make:
- Schedule/flight rate
- Crew return opportunity frequency
- Overall minimum mission duration
- Cost/available budget
- Technological limitations
- System mass and performance
- Abort practicality
- Lunar Robotic Exploration
- Landing site recon
- Reduce risk for human missions
- Characterize critical environmental parameters
- Test technical capabilities as needed
- Demonstrate technologies
- ISRU
- Tribology (wear and tear)
- Power
- Communications
- Thermal systems
- Scientific exploration, obtain scientific data
- Identify lunar resources
- Characterize environment in permanently shadowed craters
Lunar Exploration Objectives:
Once on the Moon, astronauts will be able to conduct many different experiments in the unique lunar environment that will hopefully provide results to help inform the scientific community and further develop humanity's technological prowess and understanding of the universe. Furthermore, a lunar surface architecture would provide many economic and global opportunities that would help benefit society as a whole.
Exploration Preparation
- Evaluate and employ dust mitigation techniques to protect crews, materials, and instruments during extended surface stays
- Demonstrate robust EVA capability
- Determine the internal structure and dynamics of the Moon
- Monitor space weather to determine risks to explorers
Scientific Knowledge
- Characterize impact cratering over the Moon’s geologic history to understand early solar system history
- Characterize the Moon’s atmosphere to better understand surface boundary exospheres
- Determine the origin and distribution of lunar volatiles to support studies of the origin, composition, and structure of the Moon and other planetary bodies
- Study the lunar regolith to understand the nature and history of solar emissions, galactic cosmic rays, and the local interstellar medium
Human Civilization
- Develop and validate tools, technologies, and systems that excavate lunar material
- Provide position, navigation, and timing capabilities to support lunar operations and evolve to support Mars operations
- Understand the effects of the integrated lunar environment on human performance
- Characterize the radiation bombardment on the lunar surface and subsurface
- Research to inform Mars mission
- Galactic Cosmic Radiation
- Long-duration microgravity exposure (countermeasures, artificial-g)
- Cryogenic propellant storage
- Autonomous rendezvous and capture
- Simulation of Mars missions
- EVA systems
- Habitation
- Power
- Operations
- ISRU
Economic Expansion
- Use private sector to deliver payloads to the Moon
- Utilize innovative commercial entertainment and media outlets to broadcast to the public high bandwidth video, imagery, and other information
- Deploy effective on-site remove health care systems to ensure crew health on the Moon
- Utilize commercial sector to perform lunar resource extraction
Global Partnership
- Establish an appropriate global framework, able to encompass both commercial and governmental involvement, to coordinate the lunar activities of all interested parties
- Establish standards (ex. common interface designs) to enable interoperability of systems developed by a global community
Outpost Location(s), Number(s), Size(s), and Permanence
If the lunar surface architecture consisted of outposts, one of the primary sites would be at the rim of the Shackleton Crater. Located near the south pole of the Moon, it has the benefits of being thermally moderate, potentially having large deposits of lunar ice, offering lots of sunlight in some spots, and containing a diverse set of terrestrial features including a deep crater which is permanently shadowed.
An outpost at this location would most likely rely on a solar power architecture that has the potential for nuclear power augmentation later. The U.S would most likely build the transportation infrastructure, initial communication and navigation infrastructure, and initial surface EVA capability. Making the architecture open would allow for modular functions and operations later in development, along with opportunities for commercial and international collaboration.
Due to transportation and resource constraints, outposts will take time and several iterations to become fully developed. Below are base concepts that NASA roughly conceptualized in the Exploration Strategy and Architecture presentation at the 2nd Space Exploration Conference:
If the lunar surface architecture depends on outposts, utilizing rovers will allow for the additional potential of having smaller outposts in more locations around the lunar surface. Quick and long-range transport will allow for astronauts to have a presence all over the lunar surface while still receiving the necessary resources to sufficiently live on the Moon. Additionally, multiple outposts will significantly increase exploration capabilities and allow for more astronauts, more EVAs, and more time dedicated to scientific research and discovery.
As mentioned previously, lunar surface architectures are significantly limited by technological capabilities. At the current state of technology and resources, only a few hardware systems have been developed by the U.S. Therefore, NASA is actively trying to embrace an open architecture and open opportunities for international cooperation so joint efforts can increase with time.
Surface Habitation
At the early stages of lunar surface exploration, habitats will be expected to be mostly pre-assembled and pre-integrated with the launch/transportation vehicle. This reduces the necessity for astronauts to assemble their habitats and increase safety by ensuring that safe habitats are already set up and functional before the astronauts arrive on the surface. At the absolute basic level, they must provide power, life support, and communication capabilities.
In order to reduce assembly and increase mission operation flexibility, modular mobile habitations are being considered. These will be able to be used in "super sorties” which would provide shelter, life support, and communication for astronauts who are even hundreds of kilometers away from the nearest permanent outposts. Utilizing mobile habitats would facilitate greater lunar access to capture exploration and science objectives. As with all habitats, emergency egress and redundancies must be considered.
Surface Mobility
The Lunar Surface Mobility Systems Comparison and Evolution (MOBEV) Final Report Vol. 2 Book 3 on the systems engineering of lunar roving vehicles (Nov 1966) details the design requirements that were set for future lunar rovers:
- Along with the primary Design Point Vehicle (DPV), one additional heavy-duty vehicle was required for surface grading and leveling
- Trailers exist primarily for material handling and transport functions
- All two and three-man cabin vehicles shall contain a seven-day contingency supply of expendables. Supply shall be based on reduced food and oxygen requirements, and only emergency communications
- All manned vehicles whose failure during lunar storage would cause an abort of the entire manned portion of the mission shall be designed for six-month lunar storage
- Unmanned remote-controlled vehicles shall not have a storage requirement, but shall commence their mission upon arrival on the lunar surface
- Those vehicles delivered by the LM-Shelter will obtain power for thermal control during lunar storage from the LM-Shelter.
For base support, two discrete kinds of vehicles were required: prime movers (exploration vehicle with scientific payload removed) and trailers. A backhoe would be attached to the front of the vehicle for soil excavation, loading soil onto trailers, or for burying equipment such as nuclear reactors. The backhoe would have required mechanical and electrical connections. To utilize a trailer, the prime mover would need a trailer hitch and electrical power for towing.
In the design process, four rover cabin concepts were detailed:
- One-man, open cockpit
- One-man pressurized cabin, no airlock (cabin atmosphere dumped)
- Two-man pressurized cabin, airlock plus pump
- Three-man pressurized cabin, airlock plus pump (long-duration)
Ultimately, the R3DE (a 90-day, three-man, isotope powered vehicle) was designated as the primary Design Point Vehicle (DPV). The R3BE (42-day, three-man, fuel cell powered vehicle was also included in the primary DPV category due to earlier availability and lower development risk potential.
Below are two of the many proposed lunar rover designs in the report:
For future lunar surface missions, NASA has stated that 4 basic mobility elements are necessary for its envisioned architecture:
- Unpressurized rover (15-20km range)
- Pressurized rovers (50km range)
- Large payload mover vehicle (move payload or whole lander)
- Crew Aids
Small pressurized rovers augment EVA operations by allowing astronauts to explore in ‘shirt sleeve’ environment using EVA judiciously. Although trades must be done between pressurized rovers and mobile landers/mobile habitats, many new rovers with modern and future technology have been envisioned. Below is information on a small pressurized rover designed by NASA.
(Information and pictures regarding a lunar surface architecture were found on this NASA Lunar Architecture Update at the 3rd Space Exploration Conference & Exhibit in Feb 2008)
Muramoto (talk) 06:59, 6 March 2019 (UTC)
SSE-17 The Benefits of Future Science Missions to the Moon
The following is a brief summary of the state of the geology and astrobiology of the Earth-Moon system.
Apollo 14 found the evidence of the early Earth
Few geological traces of the early Earth have survived tectonic recycling, but a sample brought back from the lunar surface by Apollo Apollo 14 harbors zircons crystals formed after the Late Heavy Bombardment. The scientists found that one rock contained a 0.08-ounce (2 grams) fragment composed of quartz, feldspar, and zircon, all of which are rare on the moon but common in volcanic processes here on Earth. Geochemical analyses indicated that the fragment crystallized in an oxidized environment, at temperatures consistent with those found in the near subsurface of the early Earth. This suggests the moon surface could be the repository of rocks formed on Earth during the Hadean.
Later impacts deposited terrestrial rocks on the lunar surface. Astronauts of Apollo 14 found the rock in the middle sized of a crater on the moon, which aids this hypothesis.
The Moon as a Biomarker
The origin of life on Earth remains an ongoing area of investigation. Rival theories include abiogenesis at deep-ocean hydrothermal vents and the panspermia model, where extremophiles are thought to have seeded the early Earth via meteor impacts. A further study of the Moon might find biomarkers and hydrocarbon aromatics not detectable with 1970's technology. Additionally, radiocarbon dating of the Moon can further constrain the evolution of the Earth-Moon system and offer insight into the dynamical evolution of the system.
Reference
"Terrestrial-like zircon in a clast from an Apollo 14 breccia" on "Earth and Planetary Science Letters Volume 510, 15 March 2019, Pages 173-185" Terrestrial-like zircon in a clast from an Apollo 14 breccia
SSE-18: What is the benefit of human exploration of the moon?
The five primary objectives of human space exploration as identified by the International Space Exploration Coordination Group (ISECG) in the Global Exploration roadmap are:
- Expand human presence into the solar system
- Understand humanity’s place in the universe
- Engage the public
- Stimulate economic prosperity
- Foster international cooperation
In pursuit of these objectives through the development of hardware, systems, and and technology, as well as through engagement and outreach, the benefits of space exploration can be organized into several categories: (i) innovation; (ii) culture and inspiration; and (iii) new means to address global challenges. These can be further sorted into the “direct” and “indirect” benefits shown below.
More concretely, humans bring unprecedented efficiency to exploration when compared to robotic explorers. As noted by Crawford, human exploration occurs for both strictly science reasons and for socio-political reasons. While science would be the “major beneficiary” of human exploration, there are cultural contributions and inspirational derivatives that are much greater than those from robots. Here, the Moon serves as a ‘proving ground’ to demonstrate both technical concepts on smaller scales and timeframes, and show that human exploration will inspire humanity and bring about the the expected cultural benefits.
Business Considerations
Potential Stakeholders
| Stakeholder | Stakeholder Type | Need | Priority |
|---|---|---|---|
| MIT | Support | Acquire BAA Funding | |
| U.S. Government | Support/Funding | Develop a plan viable to both the U.S. Government, international entities and the commercial sector | |
| Foreign Governments | Support/Funding | Contribute to the efforts of achieving sustainable lunar surface operations. | |
| U.S. Public | Support | Develop a cost effective and economically beneficial plan to achieve lunar surface operations. | |
| International Public | Support | Develop a cost effective and economically beneficial plan to achieve lunar surface operations. | |
| Congress | Support, Funding | Develop a plan viable to both the U.S. Government, international entities and the commercial sector | |
| Commercial Entities (e.g. Blue Origin, SpaceX, Draper) - Jeff and Javier | Partner, Investor, Support | Build the architecture to accommodate both NASA and the commercial space industry | |
| How many dollars per gallon of fuel would entities be willing to pay for? | |||
| Make money using lunar natural resources | |||
| Build a profitable cargo transport system | |||
| Build an architecture which demonstrates a private company owning the Gateway will be adventageous to both NASA and the commercial sector | |||
| Provide the lifecycle costs associated with the proposed architecture | |||
| Provide an NPV for your prposed architecture demonstrating a good investment | |||
| Provide a delta-v analysis of your architecture orbits | |||
| Split the lunar surface into different landing site nodes and validate your selection with orbital analysis | |||
| Provide abort capabilities during descent/ascent lunar phases | |||
| Provide the fuel type and burn rate for each abort option | |||
| Provide a solution for landing a pressurized rover on the lunar surface (too big for a crew to take down to the surface) | |||
| Provide a scalable architecture, which accomodates different lunar landing sites | |||
| Provide a list of lessons learned from the moon missions which can be scaled to a Mars mission | |||
| NASA - Oli | Sponsor | Provide a cloud of architectures considered during your analysis | |
| Build a paretto frontier populated with your architecture cloud | |||
| Calculate the dollar value per hour of astronaut time on the surface of the moon for each location considered | |||
| Independent Investors | Funding | Propose viable business models including breakeven point analysis and partnership opportunities | |
| Scientific Community (U.S./International) | User | Provide both lunar samples and data | |
| Astronauts | User | Develop safe systems used to travel to the lunar surface and establish sustained operations | |
| Universities (U.S./International) | User | Provide both lunar samples and data | |
| Contractors (Production) | User | Provide space system specifications to build to |
Stakeholder Interaction Matrix
Moon with ISRU Commercial Business Opportunites
| Architectural Decision Points | Commercial Business Opportunities | Dual Use Opportunities | Cost Reduction Opportunities |
|---|---|---|---|
| Launch from Kennedy | Tourism | Tourist/cargo transport | Cargo transport cost reduction (tourists) |
| Trans-lunar Injection | Cargo transport (e.g. supplies for Gateway, lunar surface compounds) | Fuel station cargo storage (e.g. fuel), tourists, docking | Partial capital investment from private sector and Gov for infrastructure development |
| L2 orbit | LEO fuel depot operations | Fuel Depots/Tourist/Cargo transport/Repair Station/Manufacturing/Communication?/Science R&D | Cargo transport cost reduction (tourists) |
| Dock with Fuel Storage Station | NHRO fuel depot operations | Refuel/Repair | Partial capital investment from private sector and Gov for infrastructure development |
| Spacecraft split (CM remains with Gateway) | Material transport from Gateway to lunar surface | Cargo transport / Lunar Telecommunication | Economies of scale |
| LM descends to lunar surface | Material transport from lunar surface to Gateway | Science / Mining & Maunfacturing for habitat | Learning Curve/Cumulative Experience |
| Load LM with unprocessed resources (fuel) | Lunar surface natural resource mining | Science/ Manufacturing & Refueling | |
| Launch from lunar surface | Lunar surface natural resource processing | ||
| Dock with fuel station | Lunar surface natural resource storing | ||
| Deposit fuel | Lunar surface infrastructure construction | ||
| Repeat | Lunar surface power production |
NHRO Fuel Depot System Dynamics Model
File:NHRO Fuel Depot Model v1.1.png
SSE-46 Create lifecycle costs to include cash-flow analysis and revenue
This shows break-even price and cost analysis of Lunar Surface ISRU for propellant use.
Summary
- The price of propellant on the Lunar Surface would be $26,900/kg. But if it’s delivered to LLO, the price would be $134,000/kg because of the high delivery cost from the surface of the moon to LLO. Lunar ISRU could be used only on the surface of the moon.
- If we consider projects until 2026 where commercial or governmental activities on the moon would be limited, we don’t have to use Lunar ISRU.
Break-even price
- Deterministic Prices per kilogram for LOX/LH2 propellant produced via Lunar ISRU are as follows (at 22.7% Weighted Average Cost of Capital):
- Sale to a government or commercial customer on the Lunar Surface: $26,900/kg
- Delivery to LLO to fuel two government customer LSAM descent stages: $134,000/kg
- Price per kilogram for propellant delivered to LLO is roughly five times the price of propellant purchased on the Lunar surface
- This difference in price is a direct result of costs for delivery of propellants to LLO
- Development costs for the case of delivery to LLO, including the development of a Lunar Transfer Vehicle derived from an ESAS LSAM Descent Stage, are more than twice the development costs for the case of propellant on the Lunar surface
- Transportation costs from the Earth to the Moon are double that of the Lunar surface case due to the need to transport the Lunar Transfer Vehicle as well as the ISRU production plant
- The Lunar Transfer Vehicle must use 25 MT of propellant to deliver 21 MT of propellant for sale in LLO
- Sale of excess oxygen extracted from water during propellant production results in a modest reduction of propellant price
ISRU Plant system design
- ISRU plant system design, specifications, and capability provided by the Shimizu Corporation Space Project Office of Tokyo, Japan
- Elements shown are not to scale, but represent those that are included in the plant landed by the lunar cargo lander
Cost structure and its probability (containing uncertainty)
| Deterministic / Most Likely | Minimum | Maximum | ||
|---|---|---|---|---|
| Parameter | Case 1: Lunar Surface | Case 2: LLO | All Cases | All Cases |
| DDT&E Cost [$M, FY2006] | $957 M | $ 2,157 M | -25% | +75% |
| Nuclear Power Plant* | $200 M | $200 M | ||
| Excavation/Processing/Storage Facility Cost* | $595 M | $595 M | ||
| Mass of Excavation/Processing/Storage Facility* | $162 M | $162 M | ||
| Lunar Tanker Vehicle | - | $1,200 M | ||
| Acquisition Cost [$M, FY2006] | $319 M | $ 1,019 M | -25% | +75% |
| Nuclear Power Plant* | $67 M | $67 M | ||
| Excavation/Processing/Storage Facility Cost* | $198 M | $198 M | ||
| Mass of Excavation/Processing/Storage Facility* | $54 M | $54 M | ||
| Lunar Tanker Vehicle** | - | $700 M | ||
| Transportation Cost to Lunar Surface [$M, FY2006] | $1,445 M | $2,220 M | -10% | +25% |
| Cargo Launch Vehicle (CaLV)*** | $560 M | $1,120 M | ||
| Earth Departure Stage (EDS)**** | $215 M | $430 M | ||
| Lunar Surface Access Module (LSAM)**** | $670 M | $670 M | ||
| Mission Operations Cost [$M/year, FY2006] | $35 M | $35 M | -10% | +50% |
Notes: United States Dollars FY2006 unless otherwise noted ‘* - Source: Shimizu Corporation (75% development cost, 25% acquisition cost) ‘** - Source: SEI internal cost estimates derived from previous work; development cost to the commercial company is for modification of existing stages, not for complete development of a new vehicle ‘*** - Source: Charania, A., "The Trillion Dollar Question: Anatomy of the Vision for Space Exploration Cost," AIAA- 2005-6637, Space 2005, Long Beach, California, August 30 - September 1, 2005. ‘**** - Source: Exploration Systems Architecture Study (ESAS) Draft Report, Section 12.
Price calculation of Propellant
- The price is calculated at the point of Net Present Value (NPV) of zero
- The baseline Weighted Average Cost of Capital (WACC) is 21.7 % based on a debt to equity ratio of three, equity beta of comparable industries (Aerospace, Air Transport, E-Commerce), tax rate of 30%, average nominal interest rate of 7.5%, inflation of 2.1%, and risk-free rate of 4%
- Probabilistic simulation of each case involved 1000 Monte Carlo runs with triangular distributions on the cost variables as previously defined
- Probabilistic simulation in all cases resulted in higher mean price per kilogram than deterministic analysis due to distributions on cost variables skewed toward higher cost
Propellant on Lunar Surface
Propellant in LLO
Reference
"ECONOMIC ANALYSIS OF A LUNAR IN-SITU RESOURCE UTILIZATION (ISRU) PROPELLANT SERVICES MARKET 58th International Astronautical Congress (IAC) IAC-07-A5.1.03" http://www.sei.aero/eng/papers/uploads/archive/IAC-07-A5.1.03_present.pdf
SSE-84 Business Case for a Commercial Fuel Depot in L2
ULA Depot-Based Space Transportation Architecture
- 70% of the mass we move to LEO orbit is simply propellant
- There must be a constant flux of mass through the depot to justify its presence
- 120T capacity with 300T of flow through the depot over the course of a year
- 75% of the total propellants lifted to LEO will be stored in a depot 25% is used to fuel vehicles that dock
- Since all or a fraction of the transfer stage propellant can be off-loaded, the separately launched spacecraft with payload and/or crew could have a larger mass or use a smaller launch vehicle
Development Costs
- Depending on launch vehicle, market pressures and delivery rate the cost of propellant at LEO should fall in the range of 5 to 10 $M/mT
- The total launcher and depot hardware cost and services are expected to be roughly $1B, including orbital checkout
- 2-4 crewed/robotic asteroid and cis-lunar missions per year requires 200T at an approximate cost of under $2B/year
- Propellant demand for 2 lunar landings in a year is approximately 300T/year
- Development cost of depot architecture: $100B including lunar surface systems
- Recurring launch cost is estimated at $3B/year with an IOC (Initial Operational Capability) between 2020-2025
- The cost of the payload elements with their unique low-rate designs is estimated to be approximately $2B/year
- ACES (Advanced Common Evolved Stage) upper stage estimated to be approximately $3.5B
- The depot evolution is estimated to cost, including the deployment of operational Centaur and ACES depots at LE and L2, approximately $3B
- The entire depot based architecture is estimated to cost less than $40B with an IOC for the initial elements in 2016
- The recurring lift cost for propellants is approximately $2B/year and this is sufficient to support two major crewed missions or multiple DoD and NASA planetary missions
- Developmental costs for many elements are diluted across both NASA and DoD budgets
- If the initial investment is $40B with an IOC of 2016 and there is a recurring propellant lift cost of $2B/year then the break even point is 2056 (40 years) if the cost of propellant is 10 $M/mT and there is a constant annual demand of 300T which is equivalent to 2 lunar landings
- If the initial investment is $40B with an IOC of 2016 and there is a recurring propellant lift cost of $2B/year then the break even point is 2036 (20 years) if the cost of propellant is 20 $M/mT and there is a constant annual demand of 200T which is equivalent to 2-4 crewed/robotic asteroid and cis-lunar missions
- Ex-NASA administrator Mike Griffin commented at the 52nd AAS Annual Meeting in Houston, November 2005, that "at a conservatively low government price of $10,000/kg in LEO, 250 MT of fuel for two missions per year is worth $2.5 B, at government rates
Amplification of Vehicle Capabilities with Depot Operations
Existing Atlas/Centaur performance is highly constrained and upgrading to a larger upper stage, while valuable, is only a small incremental benefit. Taking those same stages and departing from a LEO or L2 depot enables the same hardware to deliver far larger payloads to extremely high energies. These improvements can establish a long-duration human presence on the moon and Mars
LEO Depot (upper graph) and L2 Depot (lower graph) Supporting Various Planetary, Crew and Geostationary Orbit Missions
2-4 crewed/robotic asteroid and cis-lunar missions per year requires 200T at an approximate cost of under $2B/year
The Advanced Common Evolved Stage
ACES design is optimized with long-duration cryogenic applications in mind with a number of passive-thermal management features
SpaceX Interplanetary Transport System (ITS)
- ITS tanker is a propellant tanker delivery system where a single tanker performs a rendezvous and docking with an on-orbit spacecraft
- The tanker is designed to transport approximately 380 tonnes of propellant to low Earth orbit
SSE-20: Investigate aspects of program sustainability
For continued human space exploration to be feasible, aspects of cost reduction and sustainability must be considered. Sustainability in this case encompasses two primary objectives: programmatic sustainability and element reusability.
Programmatic sustainability means investing in elements and capabilities that can still deliver value to stakeholders even if program objectives change. There are 4 primary objectives to programmatic sustainability:
- Ensure benefits to all stakeholders
- Affordable in acquisition and operations
- Manage risks and communicate residual risk
- Policy robustness
Recommendations for meeting the 4 objectives include:
- Develop and invest in capabilities gradually and incrementally that can later enable more complex missions
- Minimize the development of large, mission-specific elements
- Reuse elements where possible
- Use commonly applied technologies
- Use international and commercial partnerships where possible
- Use analysis tools such as Multi-Attribute Tradespace Exploration principles to assess how to incorporate aspects of flexibility and sustainability into design choices. Two examples of factors to incorporate into such a tradespace analysis are:
- Value robustness: “the ability of the system to maintain value delivery in the context of changing system external and internal forces, including stakeholder expectations”
- Versatility: “the ability of a system to satisfy diverse expectations on the system without the need for changing form”
NASA BAA
Parametric Library of Launch Vehicles
Using data from NASA, NASA Launch Services Program, and commercial manufacturer websites, I created this updated parametric library of actual/estimated/claimed specifications of possible launch vehicles currently in service or in development. Note that the launch vehicles in this library vary in completeness; some are currently in service, and some are still being developed. These vehicles undergo several iterations and specifications change constantly. Some of the upmass figures for vehicles without publically published specifications include calculations and assumptions from NASA's Launch Services Program and found in NASA’s Plans for Human Exploration Beyond Low Earth Orbit.The numbers shown in the table are current as of March 2019.
| NASA | Commercial Currently in Service | Commercial Currently in Development | |||||||||
|---|---|---|---|---|---|---|---|---|---|---|---|
| SLS Block 1 | SLS Block 1B | SLS Block 2 | SpaceX Falcon 9 (Block 5) | SpaceX Falcon Heavy | ULA Atlas V | ULA Delta IV Heavy | NGIS Antares (232 Var) | Blue Origin New Glenn | ULA Vulcan | NGIS OmegA | |
| Scheduled completion date | Jun-20 | 2024 | 2029 | In Service | In Service | In Service | In Service | In Service | 2021 | Mid 2020s | 2021 |
| Cost/Launch ($) | 1.5B - 2.5B (Est.) | 50M-62M | 90M (reusable)
150M (expendable) |
110M | 350M | 80M-85M | - | 99M target (base, no solid boosters) | >100M | ||
| Cargo payload fairing diameter (m) | 5 | 5 or 8.4 | 8.4 or 10 | 5.2 | 5.2 | 5.4 | 5 | 3.9 | 7 | 5.4 | 5.25 |
| Upmass to LEO (Mt) | 95 | 95 | 130 | 22.8 | 63.8 | 18.85 | 28.37 | 4.4 | 45 | 34.9 | 5.3-7.8 (GEO) |
| Upmass to cislunar orbit (Mt) | >26 | 34-37 (crew)
37-40 (cargo) |
>45 | ~7.6 | ~21-23 | 2.1-6.3 | 10.5 | 1.5 | - | 14 | - |
| Upmass to Mars (Mt) | N/A | N/A | >45 | 4 | 16.8 | 1.4-4.8 | 8.1 | 1 | - | 10.5 | - |
(It is important to note that existing U.S. launchers such as ULA’s Atlas V and Delta IV Heavy could provide services but may be phased out in the 2020s by the Vulcan that is currently under development.)
A slide on NASA Science Mission Directorate (SMD) Cislunar Activities Overview presentation from August 2018 shows that the Lunar Gateway Power and Propulsion Element (PPE) is expected to be launched on either of the four vehicles shown below, the Falcon 9, Blue Origin New Glenn, ULA Vulcan, and Northrop Grumman's OmegA. Deep space resupplies are also expected to be launched on these vehicles.
Muramoto (talk) 22:11, 5 March 2019 (UTC)
Refueling Element
If a lunar gateway were to be created and utilized, as NASA believes it should, then it must be refueled during its lifespan. The European Space Agency (ESA) has conceptualized and been actively developing a refueling module, currently named the European System Providing Refuelling, Infrastructure and Telecommunications (ESPRIT). This module will be multipurpose: its primary function is to refuel, but can also resupply and provide additional communication capabilities to the gateway. It provides augmentation for additional fuel capacity (designed to be able to refill the xenon and hydrazine propellant tanks of the Power and Propulsion Element (PPE)) and adds a scientific airlock (ability to move payloads internal to external). The module can serve many purposes, even as a platform for science and tech demos. It is designed to “support payloads inside, affixed outside, free-flying nearby, or on the lunar surface” according to the NASA HQ Cislunar and Gateway Overview Presentation by Gerstenmaier and Crusan.
Known specifications of the ESPRIT module as of mid-2017:
| In-orbit mass | Approximately 4 tons |
| Liftoff mass on Ariane-64/HTV-X service module | 7,400 kg |
| Liftoff mass on Atlas-5/commercial service module | 6,000 kg |
| Developer | European Space Agency (ESA) |
| Module's length along the axial direction of the station | 3.91 m (As of Apr 2018) |
ESPRIT is expected to be launched together with an American-provided Utilization module on Orion and crew on Exploration Mission 3 (EM-3) in 2024 on the Space Launch System (SLS), the rocket currently being developed by NASA. Orion would then dock to the modules that rode underneath it inside the SLS upper stage, then rendezvous with the PPE that was launched first into lunar orbit and dock together, building the first piece of the Gateway.
ESPRIT would carry an antenna designed to maintain communication with assets on the moon’s surface and a separate S-band radio for spacecraft communications. A high-data flow communication channel with the Earth will also be installed, which will allow the ESPRIT module to act as a relay station between ground control and lunar surface hardware.
Depending on final configuration, the liftoff mass of the entire stack is estimated to be between 6,000 - 7,400 kilograms, which would require either the Atlas-5 or the Ariane-64 rocket. By mid-2018, the Falcon Heavy became a potential carrier for delivering the module. Launching ESPRIT on the more capable SpaceX rocket would enable the module to carry more propellant, reaching a total module mass of between 5-6 tons.
The European Space Agency (ESA) planned in 2018 that further development of the ESPRIT module would be conducted under competitive contracts with industry. Between May and July 2018, two parallel contracts began. One contract was given to the team comprised of Thales Alenia Space and OHB and another to Airbus Defense and Space. Muramoto (talk) 22:41, 5 March 2019 (UTC)
(Detailed technical information regarding ESPRIT was found on this website)
Resource Prospector and RESOLVE History and Relevance
The BAA is released on the heels of the cancellation of NASA's Resource Prospector mission. This mission was 10 years in development and began with the RESOLVE payload, which was ultimately woven into the Resource Prospector Mission. Publicly available mission data and highlights include the following requirements and partnership status. It was interesting to find that most partners were international government and not Commercial partners. And while the ISRU potential is also of interest, the unmanned nature of this mission means our architectures have significantly different considerations.
Resource Prospector Level 1 Requirements
1.1 RESOURCE PROSPECTOR SHALL LAND AT A LUNAR POLAR REGION TO ENABLE PROSPECTING FOR VOLATILES
- Full Success Criteria: Land at a polar location that maximizes the combined potential for obtaining a high volatile (hydrogen) concentration signature and mission duration within traverse capabilities
- Minimum Success Criteria: Land at a polar location that maximizes the potential for obtaining a high volatile (hydrogen) concentration signature
1.2 RESOURCE PROSPECTOR SHALL BE CAPABLE OF OBTAINING KNOWLEDGE ABOUT THE LUNAR SURFACE AND SUBSURFACE VOLATILES AND MATERIALS
- Full Success Criteria: Take both sub-surface measurements of volatile constituents via excavation and processing and surface measurements, at multiple locations
- Minimum Success Criteria: Take either sub-surface measurements of volatile constituents via excavation and processing or surface measurements, at multiple locations
Past Cases of In-space Refueling
- Mir: Frequently refueled by Progress spacecraft[1].
- Salyut 6 and 7[2]: refueled on orbit.
- International Space Station (ISS): Frequently refueled
- Unmanned spacecraft to demonstrate refueling:
- ISS (Robotic Refueling Mission)[3]: RRM is a NASA technology demonstration mission with equipment launches in both 2011 and 2013 to increase the technological maturity of in-space rocket propellant transfer technology by testing a wide variety of potential propellant transfer hardware, of both new and existing satellite designs.
- Trends: Heading to the International Space Station aboard December’s SpaceX Commercial Resupply Services Mission-16 (CRS-16), the Robotic Refueling Mission 3 (RRM3) will test new methods for transferring and storing cryogenic fluids like liquid methane in space.[4]
- Trends:Beyond the Moon, carbon dioxide in the Martian atmosphere also has the potential of being converted to liquid methane, a cryogenic fluid. RRM3 techniques could then be applied to refuel departure rockets from Mars.[5]
- Orbital Express mission[6]: a 2007 U.S. government-sponsored mission to test in-space satellite servicing technologies with two vehicles designed from the start for on-orbit refueling and subsystem replacement.
- Space Infrastructure Servicing (SIS)[7]: SIS is a spacecraft being developed by Canadian aerospace firm MacDonald, Dettwiler and Associates to operate as a small-scale in-space refueling depot for communication satellites in geosynchronous orbit.
- Customer: In June 2017, SSL announced its first commercial customer, Luxembourg-based satellite owner/operator SES S.A.
- Mission Extension Vehicle (MEV)[8]: proposed by ViviSat, a 50/50 joint venture of aerospace firms U.S. Spaceand ATK, to operate as a small-scale in-space satellite-refueling spacecraft. ViviSat would connect to the target satellite in the same way as MDA SIS, but not transfer fuel. It would rather use "its own thrusters to supply attitude control for the target. In December 2017, the US telecommunications regulator approved a plan submitted by OrbitalATK to use an MEV to service an Intelsat satellite, Intelsat 901, that was originally launched to geostationary orbit in June 2001 for a planned in-service life of 13 years. That satellite has already been replaced in orbit. The first MEV, MEV-1, is planned to launch with Eutelsat's 5WB commsat, no earlier than 2019.[5][6] MEV-1 will also need a licence from National Oceanographic and Atmospheric Administration (NOAA).
- Business Model: Extend life/ take-over maneuver/ even repair for out-of-fuel satellites or fueled spacecraft, sell service to satellite operators.
- The new systems are based on the Mission Extension Vehicle (MEV), a satellite life extension vehicle that Orbital ATK currently offers. The MEV docks with a satellite and takes over maneuvering of that satellite, including stationkeeping as well as relocation and disposal into graveyard orbits.Under the new approach, a Mission Robotic Vehicle, based on the MEV design, will carry 10 to 12 Mission Extension Pods. The Mission Robotic Vehicle would approach a customer’s satellite and use a robotic arm to attach a pod to that satellite. The pod would then take over stationkeeping, proving up to five years of additional life.[9]
- ConeXpress[10] — an ESA concept vehicle to perform the same mission
Business Model Options for Refueling Element in BAA
Business Morph Matrix based on a refueling element in the NHRO Orbit
| Function | Subfunction | Service | Product | Customers | Partners | Monetarization |
|---|---|---|---|---|---|---|
| Docking Platform | NA | Yes | No | Spacecraft operator | NA | By service time |
| Space Tourism | NA | By service time | ||||
| Science Airlock for Exterior Experiments[11] | NA | By service time | ||||
| Tug/Platform to transfer from Gateway to Low Lunar Orbit | NA | By service time | ||||
| Storage | ISRU Resource | Yes | No | ISRU operator | Gateway | Volumn/Time discounted for value |
| People | Yes | No | Space Agency | Gateway Operator | Number/Time | |
| Fuel/Resource from Earth | Yes | No | Space Agency/Commercial | Gateway Operator | Volumn/Time | |
| Refuel | Refeul for Ascend/Descend | Yes | Yes | Ascend/Descend Vehicle operator (Lockheed Martin)[12] | NA | By Fuel price |
| Refeul for Deep Space Missions | Yes | Yes | Deep Space Vehicle operator | NA | By Fuel Price | |
| Refeul for End-of-life Satellites/Spacecrafts | Yes | Yes | Satellite/Spacecraft operator | NA | By Fuel Price/ Fuel Price Plus Extended Revenue Share | |
| Repair / Assembly | Repair Satellites/Spacecraft | Yes | No | Satellite/Spacecraft operator | NA | Service Fee |
| Telecommunication Enabler | Telecom in Lunar Orbit | Yes | No | Lander/Spacecraft operator | NA | Service Fee |
Business Considerations
| Architectural Decision Points | Commercial Business Opportunities | Dual Use Opportunities | Cost Reduction Opportunities |
|---|---|---|---|
| Launch from Kennedy | Tourism | Tourist/cargo transport | Cargo transport cost reduction (tourists) |
| LEO insertion for fuel | Cargo transport (e.g. supplies for Gateway, lunar surface compounds) | Gateway cargo storage (e.g. fuel), tourists, docking | Partial capital investment from private sector and Gov for infrastructure development |
| Trans-lunar Injection | LEO fuel depot operations | Economies of scale | |
| NHRO injection | NHRO fuel depot operations | Refuel for Lunar Mission/ Deep Space Missions/ Satellite Life Extension / Lunar Telecommunication | Learning Curve/Cumulative Experience |
| Gateway Docking | Material transport from Gateway to lunar surface | Tourist/ Science Airlock for Exterior Experiments / Spacecraft stationing | |
| Spacecraft split (CM remains with Gateway) | Material transport from lunar surface to Gateway | Cargo transport / Lunar Telecommunication | |
| LM descends to lunar surface | Lunar surface natural resource mining | Science / Mining & Maunfacturing for habitat | |
| Load LM with processed/unprocessed resources | Lunar surface natural resource processing | Science/ Manufacturing & Refueling | |
| Launch from lunar surface | Lunar surface natural resource storing | ||
| Dock with Gateway | Lunar surface infrastructure construction | ||
| Transfer resources to Gateway for storage/use (e.g. fuel) | Lunar surface power production |
SSE-77: Evaluate merits of hydrogen vs. methane for refueling module
Evaluate Merits of Hydrogen as a Fuel
•LH2 = liquid hydrogen
•Requires special materials and processes
•Caused headaches/delays on STS
•Welds that work for other materials leak LH2
•Methane easier to store than hydrogen
•Low boiling point at ~23 K
•Harder to store because active cooling is required to keep in cryogenic; even with active cooling, it would still vent over time, making it bad for long-duration missions
•Venting required to prevent explosion
•Specific impulse: ~450 (Astronautix) (highest of any rocket fuel)
•Byproducts of heat and water are easy to manage
•Heat is useful to keep astronauts warm
•Water can be drank by astronauts
•Vehicles that use LH2: Atlas, Boeing’s Delta III and Delta IV, Space Shuttle, Russian Angara
•Boiling point far from oxygen’s means you must have a complex bulkhead design to manage both elements
•“In addition, liquid hydrogen causes hydrogen embrittlement, where hydrogen atoms alloy themselves into their metal containers, and so weaken the structure. At high pressures, this can be catastrophic.”
•ISRU on Moon: Ice at poles could be used to make hydrogen
•ISRU on Mars: CO2 atmosphere, so water would need to be present in order to make LH2
•The modifications required to handle hydrogen as a fuel negate the advantages of its high specific impulse
Evaluate Merits of Methane as a Fuel
•CH4
•Minimizes costs, surprises
•Maintain fast tempo for development, testing, operations
•High boiling point at ~112 K
•Easier to store because passive cooling can keep it cryogenic
•Less bulky/higher density so it requires less volume on launch vehicle
•“While lower in mass, carrying hydrogen (H2) from Earth to make O2 /methane (CH4 ) is volumetrically and technically difficult
•H2 is <1/3 the mass but 3 times the volume compared to CH4 brought from Earth” (NASA Sanders)
•Clean burn means there is no residue (“coking”) buildup in the rocket, which is ideal for reusability
•Coking: polymerization that occurs when a fuel is burned during a rocket launch; this creates residue inside the rocket
•Specific impulse: ~322 (DLR paper)
•Boiling point close to oxygen’s means you can have a simple bulkhead design to manage both elements
•Potentially manufacturable on Mars: would require import of hydrogen or use of water in-situ, but then CO2 could be converted to CH4
•ISRU on Moon: There is carbon monoxide on the moon that could be used to make methane
•ISRU on Mars: Bring hydrogen, then use CO2 in atmosphere to make methane
Compare Hydrogen and Methane
| Hydrogen | Methane | |
|---|---|---|
| Specific Impulse | 450 s | 322 s |
| Boiling Point | 23 K (uses active cooling) | 112 K (uses passive cooling) |
| Bulkhead Design (Compatibility with Oxygen) | Difficult | Simple |
| Reusability | Poor (embrittlement) | Good (no coking) |
| Storage | Difficult | Feasible |
| Welding | Difficult | Feasible |
| Cooling to Prevent Boiling | Difficult | Feasible |
| ISRU on Moon | Feasible | Feasible |
| ISRU on Mars | Feasible | Feasible |
Conclusion
While liquid hydrogen was used heavily by NASA in the past, methane appears to be a better fuel moving forward. This is reflected by choices made by SpaceX, Blue Origin, and even now, NASA. Bbrowder (talk) 15:39, 6 March 2019 (UTC)
SSE-27: Cryogenic Storage and Transfer Techniques
There is no evidence that methane can be found in any significant quantity on the Moon. Therefore, ISRU fuel production on the Moon will be primarily hydrogen and oxygen based. An orbiting refueling station will either need to maintain fuel stock with periodic dedicated fuel resupply missions from Earth of either LH2 or LNG, as well as LOX, or be resupplied with LH2 and LOX from the lunar surface produced with ISRU. Resupply will have to occur at a consistent rate, even if transfer and fuel use are minimized because of natural boiloff. The best designed systems for LH2 storage in space applications so far not been able to reduce losses below 2-20% per day. Any boiloff can create a dangerous pressure build up if not handled carefully. For example, hydrogen has an isobaric 1:851 liquid to gas volume ratio at room temperature. Oxygen is even greater at 1:860. Methane’s expansion ratio is slightly lower at 1:578, but still a definite safety hazard.
Storage tanks have traditionally been made of aluminum alloys, but new technological developments in carbon fiber composite tanks may reduce mass requirements for future installations. All storage techniques also require cooling and pressurization equipment. Traditionally, in-space cooling has been done with multi-layer insulation (MLI), a polyester film, metallized with aluminum, copper, or gold that greatly reduces radiative heat transfer. Active cooling methods like heat pumps may be applicable but would require additional power mass. Fuel transfer will likely require active cooling to keep in line temperatures low. Early research also shows that prechilling of all pipes and transfer lines before pumping is necessary to prevent to formation of vaporization bubbles.
An in-space refueling depot in lunar orbit will likely have a single docking interface through which cryogenic fuels can be deposited or withdrawn. This reduces system complexity as the same transfer line cooling system can be used for both operations. This system will also require the operation of reversible pumps to accommodate fluid flow into and out of the refueling element. A hydrogen cooling system will increase in size and complexity due to hydrogen’s lower density and the lower required temperature maintenance as compared to methane.
SSE-32: Investigate aspects of Refueling Element
Although reusability is not a preliminary requirement of the design of the refueling element, reusing the element in some way may potentially provide an additional avenue for cost savings, and therefore increase the competitiveness and sustainability of a proposal. By 2028, NASA would like to complete a demo that makes use of two refueling elements, one to refuel the Transfer Vehicle and one to refuel the Ascent Element.
Considerations for Reusability
While a reusable system may turn out to be more cost effective in the long run, there are additional costs, risks, and design considerations associated with reusability that must be weighed against the potential benefits of a reusable system. Successfully evaluating the viability of a reusable system means comparing the options for reusability and completing an analysis of each design option that takes into account:
- Understanding of system wear and degradation over time and expected operating conditions
- The useful life of the system
- The use of potentially more expensive manufacturing materials that will endure multiple uses
- Other stakeholder needs that may align with re-use options of the element (ie. using spent fuel tanks as a hab)
Selected Reusability Scenarios
Selected reusability scenarios and their associated pros and cons are listed below:
| Reusability Scenario | Constraints | Pros | Cons |
|---|---|---|---|
| No reusability | Base constraints: Wet mass at launch less than 16 metric tons.
10 MT propellant. Must fit within 6.3 m diameter dynamic envelope. |
Do not need to develop landing capabilities on refueling element.
No costs of recovery or refurbishment. |
Long duration spaceflight or multiple missions would require multiple trips of a refueling element - cost of production of refueling elements would increase. |
| Reuse refueling element for additional missions - Earth refueling | Base constraints. Must include landing capabilities. | Potential cost savings if element is to be used frequently enough. | Need to develop landing/ recovery capabilities. Need to understand the useful life of the element to accurately calculate potential benefits. |
| Reuse refueling element for additional missions - lunar refueling | Base constraints. Must include landing capabilities and have launch component. Requires ISRU. | Smaller delta V required to reach the gateway. | Requires significantly more ISRU. |
| Reuse design of refueling element from Moon to Mars | Base constraints. Need to meet the mass, volume, and power targets necessary to for vehicles in Mars missions. | Potential cost savings from reduced design, development, and testing costs. | Need to fully understand Mars-related mission requirements ahead of time. Must also be able to accurately characterize power considerations and drag considerations from Martian atmosphere that are different than lunar. |
| Reuse refueling element as a hab on moon | Base constraints. Space requirements for human habitation - 25 m^3/ person required by standards. Considerations about toxicity of propellant. | Savings of at least the cost of a launch that would otherwise be required to transport a habitation element. | Will need to refuel many times - only need a few habs. |
| Reuse spare parts from refueling element | Base constraints. | Do not necessarily need requirements well ahead of time to be able to reuse parts. Reduces the need to spend fuel and money launching spare parts. | Must consider the safety and effectiveness of reusing parts. Must design complementary systems that can feasibly use these parts. |
SSE-59 Identify technology maturation areas and implement maturation plans for these areas for Refueling Element
Fluid Control
The low gravity environment of space results in unpredictable fluid behavior if there are no measures taken to ensure the fluid settles. The current highest TRL method for this is inertial settling, through the use of an applied thrust, which causes the fluid to settle at the back of the tank. Alternatively, the fluid storage can be spun about its axis, and this has been demonstrated on orbit successfully.
Other methods, which are all of lower TRL, are electrodynamic tether, gravity gradient, and electromagnetic settling. The first two involve large and complex systems to implement, while the latter has serious questions about how much mass superconducting magnets require. If not using superconducting magnets, the question exists of whether electromagnets are capable of settling large LH2 tanks
If a refueling element is to be implemented and developed, the quickest path would involve using inertial settling via a centrifugal force. This has already been demonstrated to work, does not require mass to induce its settling, and only requires advances in GNC which are already under development by NASA.
Thermal Control
Thermal control of cryogenic fluids has been flight proven, as there are plenty of space systems which have used LO2 and LH2. The methods used do not meet the mass efficiency requirements of a long-term refueling depot. To circumvent this, an active thermal control method has been proposed which uses the boil-off of LH2 also as station-keeping propellant, making this boil-off loss negligible in terms of the efficiency.
Rendezvous, Docking, and Transfer
Fluid transfer has been demonstrated on orbit, but not cryogenic fluid transfer. The primary change is in the mechanical hardware used to do the fluid transfer, which ULA proposed off of work on the Atlas Centaur upper stage. Additionally, if centrifugal fluid settling is used, which is the highest likelihood given current TRL’s, higher fidelity GNC may need to be developed than has been currently demonstrated. GNC improvements to rotation rate handling are already a part of NASA’s Technology Roadmap and should be available for use in any near-term developments.
The use of a space tug merits further and deeper development consideration, as it greatly increases mass efficiency of refueling systems and reduces the complexity of any tankers and depots.
References
1) Realistic Near-Term Propellant Depots: Implementation of a Critical Spacefaring Capability: https://arc.aiaa.org/doi/pdf/10.2514/6.2009-6756
2) ULA DMSP-18 Inertial Propellant Setling Results: https://www.ulalaunch.com/docs/default-source/supporting-technologies/successful-flight-demonstration-conducted-by-the-air-force-and-united-launch-alliance-will-enhance-space-transportation.pdf
3) A Practical, Affordable Cryogenic Propellant Depot Based on ULA’s Flight Experience https://arc.aiaa.org/doi/pdf/10.2514/6.2008-7644
4) 2 for 2 Tech demo spacecraft achieve objectives: http://www.boeing.com/news/frontiers/archive/2007/october/i_ids03.pdf
5) Cryogenic Propellant Storage and Transfer Technology Demonstration: Prephase A Government Point-of-Departure Concept Study: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20120016527.pdf
SSE-65: Assess overlap between NASA and Commercial Interests
General overview
Public-private partnerships between NASA and commercial enterprises have been successful at advancing research interests while simultaneously being advantageous for profit-driven companies. As the customer of both near-term and long-term space technologies, NASA actively seeks to invest in R&D of products which both have an existing market and those which need more funding and development to be commercially viable. Funding for the second category may be particularly interesting for commercial companies which would like to innovate and develop capabilities to get in on the ground level in emerging markets, but who would otherwise not have the capital or risk appetite to invest in basic research or emerging technology.
Cislunar Econosphere
The following chart (derived from United Launch Alliance) shows some examples of the “Cislunar Econosphere”, or areas of potential overlap between NASA and commercial interests of space technologies and capabilities:
| Capability | Location | Market readiness | One time or recurring service? |
|---|---|---|---|
| Remote Sensing | LEO | Existing market | Recurring |
| Communication | LEO | Existing market | Recurring |
| Observation | LEO, GEO | Existing market | Recurring |
| Commercial station | LEO | Emerging market | Recurring |
| Manufacturing | LEO, Lunar surface | Future market | Recurring |
| Propellant transfer | LEO | Future market | Recurring |
| Repair station | GEO, High Earth Orbit | Future market | Recurring |
| Human outpost | Lunar surface | Future market | One time |
| Mining | Lunar surface | Future market | Recurring |
Reference
Cislunar Econosphere source: United Launch Alliance, slide 6, http://sciences.ucf.edu/class/wp-content/uploads/sites/58/2017/04/ULA-UCF.pdf
SSE-76 Forward extensibility of lunar lander towards Mars
Overview
Extensibility, as defined in the BAA, is a system design principle where the initial implementation and deployment considers future intended use cases. It measures how easily a system can have its capability extended and the level of effort required to expand that capability.
Extensibility is a system design principle where the implementation takes into consideration future growth. It is a systemic measure of the ability to extend a system and the level of effort required to implement the extension. Extensions can be through the addition of new functionality or through modification of existing functionality. The central theme is to provide for change while minimizing impact to existing system functions. The design of an avionics suite to readily support software changes from one lander to the next, more capable lander is an example. NASA will work with partners to develop a strategy that will enable efficient transitioning of reusable and extensible aspects of the architecture.
Lander systems that can be used with no or minimal modification
- Command and Service Modules
- Life-support systems
- ISRU techniques
- Launch vehicle
Lander systems requiring significant modification
- Power Generation
- Entry, Descent and Landing
- Ascent / Orbit
- Communication
| System | Reason | Change | Implementation Difficulty |
|---|---|---|---|
| Power | Solar irradiance on Mars is ~ 590 W/m2 vs 1000 W/m2 on Earth surface. | Requires larger solar panels, alternate power generation capability, or significantly lower power usage | Moderate |
| EDL | Martian atmosphere ranges from 30 Pa to 1.1 kPa | Requires atmospheric interface hardware, such as heat shield, or significant ΔV margin to slow before entry to minimize heating | Significant |
| Ascent | ΔV to low Mars orbit 3.3 km/s vs ~1 km/s | Significant increase in ΔV required to reach stable orbit | Significant |
| Communication | Both systems will likely use the Deep Space Network, but Mars system requires far greater range and power | Increase communications system power and mass budgets, use antennas raged for the distance and pointing needs seen in a Mars mission profile | Minor |
Conclusion
Many of the lander systems can be reused in a Mars architecture with minimal modification, the driving requirements will be EDL and ascent capability. These are hard physical requirements dictated by fundamental differences between the Moon and Mars, and will require substantial design changes to enable capability. Ferrous (talk) 06:13, 6 March 2019 (UTC)