Stakeholder Analysis

Stakeholder Value Network in Current State

File:3Current StakeholderValueNetwork.png

Stakeholder Value Network in a Mature Newspace Economy

Based on Stakeholder Analysis in Sprint 2 by John Hernandez

Stakeholder	Stakeholder Type	Need
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.
		National Pride
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
		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 advantageous to both NASA and the commercial sector
NASA	Sponsor	Provide a cloud of architectures considered during your analysis
Independent Investors	Funding	Fund viable business models with 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 or technical research

Based on Stakeholder Value Network in Cameron et al., 2011, Goals for space exploration based on stakeholder value network considerations

Also Based on Stakeholder Studies by Stella Tkatchova, 2018, Emerging Space Markets

Determining the optimal architecture for value producing exploration

Architecture Generation

The architecture generation process is driven by the morph matrix that was created at the beginning of sprint 1. However, as we progressed from sprint 1 to now sprint 3, the morph matrix has been updated to reflect the key decision points within the architecture. The morph matrix was thus simplified for this sprint, and focuses on trades such as where to launch from Earth to, whether or not to conduct Earth orbit rendezvous, whether or not to conduct Lunar Orbit Rendezvous, and other parameters as seen in the figure. The majority of the decisions will impact the overall trajectory taken to the lunar surface, and thus will impact the propellant requirements for various architectures.

File:Updated morph matrix.JPG

Outside of trajectory decisions, the morph matrix also examines key trades that will take place on the lunar surface. These include the number of crew to bring to the surface, the number of surface days, and how much mobility the crew will have. The number of crew and surface days were assumed to range between the values listed in the figure, but moon mobility was driven by the type of rover used on the surface. For example, if no rover was brought to the lunar surface, then the max mobility would be about a mile. The larger the rover, the more range, but more mass delivered to the lunar surface.

From all of the key decision points, we are able to apply logical constraints from a trajectory point of view in order to minimize the architectures needed for evaluation. Utilizing full factorial enumeration, the total number of architectures generated from our morph matrix is 2,211,840. However, after applying the trajectory constraints, the number of "feasible" architectures reduces to 34,560. This is roughly 1.5% of all the architectures, and a much more reasonable set of architectures to work with.

Delta V Metric

One key component of the morph matrix is the Delta V requirement metric. Previously, the Delta V parameter has focused on an Apollo 11 style trajectory architecture. In this iteration we introduce trajectories established and implemented during the later Apollo missions. We deviate from a traditional free return trajectory in favor of a hybrid free return trajectory. This hybrid trajectory was implemented for Apollo 13 and later, and allows for a status check and initial free return trajectory before performing a midcourse correction that reduces time of flight and energy requirements. While this midcourse correction inherently eliminates a free return abort, we have observed in Apollo 13 that it is possible to burn back onto a free return trajectory.

Gateway has been more thoroughly explored in this iteration, and we believe that the best insertion into the NRHO is from the L2 point. We also explore using the NRHO to navigate into low lunar orbit either at an equatorial inclination, or at polar latitudes.

File:Deltavr3.png

In this iteration we also explored the options for inserting spacecraft into a polar lunar orbit. A polar orbit would allow for exploration of Shackleton crater and other polar areas of interest. Polar exploration requires significantly more energy than a descent from an equatorial latitude. The polar orbit insertion trajectory is initiated as part of the transfer burn from Earth, rather than a change in inclination burn from low lunar orbit. The maneuver conducted is a hybrid transfer burn with vectoring that changes inclination simultaneously. Circularization occurs at perilune, and follows the same procedure as equatorial circularization seen in the Apollo 16 flight log.

While an equatorial Low Earth Orbit (LEO EQ) was included in the figure, it is unrealistic to utilize this orbit. Both an inclination change maneuver and a launch j-hook maneuver are expensive in delta V. Rather, a translunar injection from LEO at the Kennedy Space Center latitude (~28 degrees) allows for a midcourse correction that is less costly and has been thoroughly tested on all later Apollo missions.

Finally, some additional assumptions need to be touched on. We assume that a geosynchronous orbit, as well as L3, and L4/L5 orbits are irrelevant given the scope of the mission. We assume that all burns use high impulse engines, with the goal of minimizing time in flight to the best of our ability.

Mass on Lunar Surface

One of the most important variables to consider for each architecture is the mass delivered to the lunar surface. This mass is a key variable, and an excellent one to trade on, because each decision point will greatly impact the mass delivered to the lunar surface. For this pre-phase A analysis, we made some key assumptions that allowed the analysis to be more straightforward and less dependent on other key design trades such as what type of lunar module to utilize, etc. We assumed that we only had to care about a few variables that make up the mass on the lunar surface, all of which are derivative of some of the decisions within the morph matrix.

The mass on the lunar surface was the summation of the consumables brought to the lunar surface, the mass of the rover brought to the lunar surface, and any additional ISRU infrastructure that needs to be deployed on the lunar surface. Obviously, there are other variables that must be brought to the lunar surface such as power systems, communication systems, etc., however these were not considered because as stated above, are highly dependent on the lunar architecture, as in lander design, that is chosen. Consumables brought to the lunar surface is directly related to the number of crew days on the lunar surface, or number of crew * number of days on surface. The mass of the rover is directly related to the lunar mobility defined in the morph matrix. ISRU capability and associated mass was analyzed for two different architectures as described in the next section.

Once the mass on the lunar surface is obtained, we can utilize the rocket equation to determine the mass of the system at Earth launch. This is an extremely important variable and will help us understand the differences in the architectures. The rocket equation was used for each staging point in the trajectory, and a wet mass was determined for each stage along the route. This does not look into more nuanced design trades such as the number of modules, etc., but simply assumes that if there is a change in trajectory, i.e. you go from LEO to Gateway to the Lunar surface, a different propulsive module will be used. This will give us a large final mass value, but this is a good thing at this point along the process.

Trade Studies

Figures of Merit

The following Figures of Merit (FOMs) were selected to evaluate the lunar missions.

Cargo delivered to lunar surface [kg]
Exploration Capability [kg * crew days]
Relative Exploration Capability [vs. Apollo 17]
Injected Mass into Low Earth Orbit [kg]
Return on Investment ROI or Net Present Value [%], [%]
Risk = Probability of Loss of Mission P(LOM), Loss of Crew P(LOC), Loss of Element P(LOE), Number of mission critical events (burns, docking etc)
Total Investment Cost [$]
Commercial Merit: cost / crew-days-on-surface
ISRU Production = Ultimate long term production mass capability (kg)
Overall Metric: mass_to_lunar_surface * number_of_crew / (delta_v * cost)

Each of these FOMs was weighted to deliver an overall preferred architecture. Exploration capability was the most weighted FOM it is the most representative of the stakeholder needs of both government and commercial interests.

ISRU on the Moon

Why ISRU?

In order to tackle the topic of lunar ISRU, we must first examine if the endeavor is worth pursuing. (All data and results in this section were excerpted from a publication written by a postdoc in Professor de Weck’s lab):

Ishimatsu, Takuto, Olivier L. de Weck, Jeffrey A. Hoffman, Yoshiaki Ohkami, and Robert Shishko. “Generalized Multicommodity Network Flow Model for the Earth–Moon–Mars Logistics System.” Journal of Spacecraft and Rockets 53, no. 1 (January 2016): 25–38.

Ishimatsu applied a Generalized Multi-Commodity Network Flow (GMCNF) Model to the Earth-Moon-Mars Logistics System to determine optimal logistical strategies for missions to Mars.

Production rates are described as the amount of resource produced (in kg) per year per kilogram of system mass. Anything above 1.0 signifies that the ISRU system is producing more resources than its own weight, which means it is productive. The GMCNF modeling found that lunar ISRU becomes beneficial above a production rate above 1.9 kg/year/kg. Beyond that, the production rate generally increases, although the system mass fluctuates up and down in a zigzag manner. Above 3.6 kg/year/kg, lunar ISRU is strongly favored over Mars IRU. Therefore, engineers should aim to create a lunar ISRU system with at least these capabilities.

File:ISRUprod.png

ISRU resource production rates Lunar and Martian missions

Key findings:

Liquid Oxygen/Liquid Hydrogen (LOX/LH₂) is much more compatible with ISRU water production than Nuclear Thermal Rocket (NTR) propulsion
Aerocapture significantly reduces Total Launch Mass to LEO (TLMLEO), development of lightweight aeroshell/thermal protection system should be encouraged
Lunar water ice is the most valuable resource for Mars missions. Production rates above 1.9kg/year/kg for lunar ISRU and 0.6kg/year/kg for Mars ISRU would make the lunar/Mars ISRU system benefit exceed the cost in terms of overall launch mass.

It is important to note that these results can vary greatly depending on the parameters and assumptions used in the model, as well as the figures of merit used for optimization. So far, the paper proves that the GMCNF model is useful as a preliminary guide for the logistics of lunar a lunar architecture with ISRU infrastructure.

(Data and results in this section excerpted from this Ph.D thesis written by a student in Professor de Weck’s lab:)

Ho, K. (June 2015). Dynamic Network Modeling for Spaceflight Logistics with Time-Expanded Networks (Doctoral dissertation). Retrieved from DSpace@MIT. (Identifier No. 920684579)

In his dissertation, Dr. Koki Ho developed a new dynamic time-expanded GMCNF model and applied it to several cases with lunar ISRU capability: an architecture for human exploration of Mars, an architecture for human exploration of Near Earth Objects (NEO), and an architecture for human exploration of both Mars and a NEO.

In his case study for human exploration of Mars, Ho found that as lunar ISRU capabilities increase, the required initial mass to LEO (IMLEO) decreases. When ISRU capability is 5 kg/kg plant/year, IMLEO is about 15.7% lower than the case without ISRU (511.5 MT to 431.3MT IMLEO reduction).

File:LunarISRU.png

Lunar ISRU provides decreased IMLEO requirement

Sensitivity analyses for all cases studies revealed that IMLEO values all decrease as lunar ISRU capability increases. The cost of ISRU pays off if its capability is large enough. The more capable the ISRU system is, the more efficient it will be to use a common space infrastructure for multiple missions, even those with different destinations. Developing lunar ISRU capability is critical for a lunar architecture, and Ho suggests that it would be very beneficial to design Mars and NEO missions together if ISRU technology can achieve and possibly exceed the range of 1-7 kg/kg plant/year. Ishimatsu showed that the target for resource generation should be above 3.6 kg/kg plant/year, and this value falls within Ho’s range.

How to do Lunar ISRU?

Previous research does suggest that lunar ISRU does bring plenty of benefits to human space exploration missions, especially in terms of IMLEO estimates. Now, it is important to examine if the technology is capable of achieving the capability goals set by Ishimatsu and Ho.

(Data and results in this section excerpted from a S.M thesis of a graduate student in Professor de Weck’s lab:)

Schreiner, S.S. (May 2015). Molten Regolith Electrolysis Reactor Modeling and Optimization of In-Situ Resource Utilization Systems (S.M. thesis). Retrieved from DSpace@MIT. (Identifier No. 921147148)

Schreiner focused on an oxygen production technique called Molten Regolith Electrolysis (MRE), in which molten lunar regolith is directly electrolyzed to produce oxygen gas and metals (iron and silicon).

File:RegElec.png

Schematic diagram of Molten Regolith Electrolysis Reactor producing oxygen gas at the anode and molten metals at the cathode.

The MRE reactor model used in the thesis predicts that an MRE reactor is able to produce oxygen with a specific mass on the order of ≈10 (kg O₂/year)/(kg reactor) and ≈5 (W)/(kg O₂/year). These values provide initial evidence that an MRE reactor can be a viable option for producing oxygen from lunar regolith to resupply ECLSS consumables and provide oxygen for chemical propulsion systems.

The more oxygen the ISRU system produces annually, the more the total system mass is expected to increase. For a system that produces 10,000kg oxygen/year processing Highlands regolith, the largest system mass breakdowns are given below (in relation to total system mass):

Oxygen liquefaction/storage system (Designed to hold 6 months of oxygen production - can be relaxed depending on system needs): 30%
Reactor: 30%
Power System: 25%

File:Annualox.png

Resource requirement breakdown for 10t peak O₂ production

Optimized models predict that:

400kg, 14kW MRE-based ISRU system can produce 1,000kg oxygen per year from lunar Highlands regolith.
1593kg, 56.5kW system can produce 10,000kg oxygen per year

Note that the optimal design of an MRE-based ISRU system does not vary significantly with regolith type, which gives flexibility in landing site designation.

As seen in this paper, Ishimatsu’s 3.6 kg/kg plant/year capability and Ho’s 1-7 kg/kg plant/year capability range can easily be met using an MRE-based ISRU system.

What is the optimum lunar ISRU architecture?

Now that lunar ISRU seems beneficial and is a feasible endeavor worth pursuing, we need to look into the details of an ISRU architecture and conduct a preliminary trade study to determine preliminary IMLEO estimates.

(Data and results in this section excerpted from a S.M thesis of a visiting student that Professor de Weck sponsored:)

Schrenk, Lukas, “Development of an In-Situ Resource Utilization (ISRU) Module for the Analysis Environment HabNet”, Visiting Student from TU Munich, December 2015

For a soil-based ISRU on Moon and Mars, a soil processor for water extraction and a water electrolysis unit has been implemented. The figure below depicts the design configuration of a water extraction system enhanced with additional subsystems (heat exchanger, filter, condenser, heater, and auger heater)

File:Schemsoiloven.png

Interbartolo, M. A., Sanders, G. B., Oryshchyn, L. et al. (2013), ‘Prototype Development of an Integrated Mars Atmosphere and Soil-Processing System’, Journal of Aerospace Engineering, 26.

Soil processing requires water extraction and water electrolysis(split water into oxygen and hydrogen, which is then provided as fuel at EML fuel depot where it is stored as liquids to refuel spacecraft)

Water Extraction Model

Currently available data suggest adapting water extraction sizing models by applying a sizing uncertainty factor to undetermined mass, volume, and power results:

Mass: 1.8 x calculated mass
Volume: 11 x calculated volume
Power: 1 x calculated power

The maximum size of the water extraction oven is constrained by various factors:

Chemical process inefficiencies in drying process due to large oven volume can decrease productivity rate at a certain point
Module size and assembly strategy can limit the maximum possible volume of the water extraction plant

Analysis conducted for three different water abundances (Minimum [0.5%], Mean [2.7%], Maximum [6.5%]) from Lunar Environmental Description Model (EDM) water map within +60 degree and -60 degree latitude.

File:Waterproduction3.png

Productivity for the water extraction process, for 3 different water contents in martian soil. Specific system mass for the water extraction process, for 3 different water contents in Martian soil. Specific mass of a water electrolysis system, for different hydrogen production rates.

Performance of water extraction system strongly depends on water availability at the target site.

Water Electrolysis Model

Not enough data on electrolysis units for space applications to draw a final conclusion. Sizing uncertainty parameters for mass, volume, and power:

Mass: 4 x calculated mass (mean value)
Volume: 1.64 x calculated volume
Power: 1 x calculated power

The thesis looked at producing LOX/LH₂ propellant using lunar surface water ice found in shadowed craters. The production process is detailed in the figure below:

File:Propellantproductprocess.png

Overview of propellant production process from excavation to storage

Case study of two architecture ideas that provide in-situ fuel for a depot at EML:

Architecture A -(Decentralized Complexity) Fuel produced on orbit and shuttle transports water ice to fuel depot. Water electrolysis unit attached to EML fuel depot and a shuttle transports water ice from lunar surface to depot
- Advantages
  - Water can be provided at fuel depot in addition to LOX/LH₂
  - Water electrolysis unit at EML allows flexibility with regard to changing demand
- Disadvantages
  - Ice transport: Common LOX/LH2 ratios are 5:1, and water has a ratio of 8:1. Therefore, more oxygen than hydrogen is transported (fixed oxidizer to fuel ratios)
  - Two electrolysis systems needed since shuttle cannot be refueled at the depot (shuttle propellant budget analysis showed that it uses more propellant on the whole trip than it can transport), one at lunar surface is needed for shuttle propellant production

Architecture B -(Centralized Complexity) Water electrolysis unit deployed with water extraction oven at lunar surface and shuttle carries LOX and LH₂ to orbit. Produces propellant for both shuttle and fuel depot
- Advantages
  - Oxygen and hydrogen can be transported independently of each other
  - Single electrolysis system on the surface can supply both shuttle and depot
  - Maintenance of the system can be checked with a single mission
- Disadvantages
  - Delayed response to demand changes since production and transport are from lunar surface
  - Multiple cryogenic storage tanks: EML, Lunar surface, Shuttle
  - Cannot provide water without major changes
  - Everything is on Moon, which is more expensive to get to than Lagrange points

This conceptual design figure shows the major lunar surface systems necessary to produce LOX/LH₂

File:Fueldepot.png

Source: Charania, A. C., and DePasquale, D. (2007), ‘Economic Analysis of a Lunar In-Situ Resource Utilization (USRU) Propellant Services Market’, 58th International Astronautical Congress

Full factorial search was applied to find the combination with minimum initial mass in LEO (IMLEO). Feasibility check used a time step analysis of the fuel demand scenario, preventing fuel depot tanks from dropping below zero (which would regard configuration infeasible). Optimum IMLEO determined by the mass of systems and propellant required to transport them to the surface or EML

Assumptions:

Demand evaluation function considers boil off (depends on tank size and location) and different launch strategies
- EML tank exposed to space and solar radiation
- Lunar surface tank assumed to be in shadowed areas and only exposed to thermal radiation from the surface and thermal radiation to space
Transportation evaluation constrained by the maximum number of launches during the demand scenario. Assumed that a round-trip to/from the lunar surface and EML fuel depot takes 8 days.
Use shuttle for transport: Estimated Delta-V required for a shuttle from surface → EML fuel depot: 2.64km/s, 2km/s necessary to reach LLO

Mass Comparison of Architecture A and B for combined NEO and Mars propellant demand scenario
	A	B
IMLEO	19844 (100%)	23988 (121%)
ISRU Systems	2383 (100%)	2644 (111%)
ISRU Mass to EML	428 (100%)	0 (0%)
ISRU Mass to Moon	1955 (100%)	2644 (135%)
Shuttle, Structural Mass to Moon	884 (100%)	1065 (121%)
Maximum Payload	6031 (100%)	5290 (88%)
Maximum O₂	5361 (100%)	4513 (84%)
Maximum H₂	670 (100%)	777 (116%)

Architecture B has a significantly higher IMLEO than Architecture A. This is because:

Mass to lunar surface is larger and more propellant is required for transportation
Due to nonconstrained Hydrogen and Oxygen relation, Architecture B transports less oxygen due to demand scenario
Due to higher boil-off losses (during surface storage and transport), more hydrogen has to be transported than Architecture A
Shuttle mass is higher because cryogenic storage tanks are heavier than a water ice tank due to insulation and higher internal pressure
Heavier shuttle for B leads to higher propellant production demand and therefore heavier ISRU system

Conclusion: Architecture A is the better solution to the lunar ISRU supported fuel depot

Advantages:

Provides more flexible fuel depot that can also supply manned missions with water and excess oxygen

Disadvantages:

Development of an additional electrolysis system for zero gravity operations at the fuel depot
Heavier fuel depot, which requires more propellant for station keeping

Muramoto (talk) Ferrous (talk) 21:30, 19 March 2019 (UTC)

Business Perspective

Commercial, Government Partnership Considerations The fuel depot business case (excel file attached) provides the basis for a sustainable presence on the moon. We can’t start from the moon – doing so assumes an unrealistic time and monetary investment up front without proof of return on investment. Our team proposes two fuel depots established in LEO, which can serve to propel government(s) sanctioned lunar exploration and jump start the commercial space economy (e.g. tourism, in-space manufacturing). Moon Missions The fuel depots serve as a staging point for optimizing the mass transported to the lunar surface. This can be materials for surveying, materials for infrastructure buildup, etc. As infrastructure is built up on the lunar surface, ISRU processing and storage becomes a potential revenue stream. An L2 fuel depot could represent the next staging point for deep-space missions. Space Economy The fuel depot provides a life source from which to develop a space economy. Fuel Depot The cost to develop and operate the fuel depot(s) is almost prohibitively expensive for one entity to should the amount. Partners will be required to get this project completed. In our initial business case, we propose government entities provide 70% of the cost for DDTE up to deployment and commercial entities provide 30% of the cost. Long-term, government entities enjoy a heavy discount on LEO fuel costs and the depots help stimulate space commerce in their respective countries. Strategic Resource To prevent one country from holding the project hostage, we recommend each country (government) contribute a percentage of their GDP (TBD) up front. This significant up-front investment anchors the country to the project. If they elect to leave the project, they agree to (initial contract) abandon their contribution to date. This mechanism incentivizes government entities from threatening to leave the project, possibly hampering progress. The biggest economic strategic resource in this project will the fuel depot interface. If an agreement can be made to provide one country exclusive rights to manufacture and produce this interface, essentially setting the industry standard, this entity will control everything. In particular, costs of fuel for this specific project. Since the most significant barrier to entry is up front and sustained capital investment, a very long-term strategic horizon it afforded this entity. Our team brings this point up as a topic of discussion for ownership and control of project assets and rights for long-term bargaining.

Fuel Depot Analysis Media:Fuel_Depot_Business_Case.png

Key Assumptions and Requirements

File:ISRU Processing Requirements 2.png

Source: NASA (2019), ‘NASA In-Situ Resource Utilization (ISRU) Development & Incorporation Plans’

Tanker and Propellant Launch from Earth to Cis-lunar Aggregation Point
- Architecture 1: SLS
- Architecture 2: Commercial Launch Vehicles

File:Concepts of Operations 1.png