Cargo Missions Research Articles

Air cargo transportation, loading, and phase-based maintenance service scheduling in demand channel routes

The article proposes an approach for modeling and solving the air cargo transportation, loading, and phase-based maintenance scheduling problems within demand channels. Demand channel transportation presents unique challenges to meeting the earliest pickup and delivery requirements during emergencies. The complexities in transportation involve issues with airport-aircraft compatibility, fleet selection, task assignments, route construction, and effective loading–unloading procedures. The challenges include coping with flight connections due to airport-aircraft compatibility, dealing with limited refueling and maintenance facilities, and swiftly adapting to unforeseen maintenance events by identifying and utilizing available fleets. In this context, the article delineates a demand channel route planning strategy that considers airport-aircraft compatibility, task integrity, priority, partial refueling strategy, and maintenance schedules spanning short to long-haul phases. Formulated as a Mixed Integer Linear Programming (MILP) problem, the solution employs a novel acyclic route construction procedure within a fuel-conscious network. We propose an insertion heuristic with multi-faceted algorithmic interconnections for producing a near-optimal solution suitable for real-world applications. Our insertion procedure has been shown to have very high computational performance, with an average 1.32 percent optimality gap. It can be scaled up to the most critical cargo mission, with up to 190 ports, 95 aircraft, and 380 tasks.

PEMODELAN SISTEM DINAMIK DAN IMPLEMENTASI SIMULINK PENGENDALIAN KESTABILAN UNMANNED AERIAL VEHICLE MULTIROTOR HEXACOPTER CARGO

The rapid growth of Unmanned Aerial Vehicle (UAV) technology, or drone, has shown its popularity and has been significantly applied to various purposes today. Nevertheless, with all the sophistication of drones, many related topics are still attractive, especially when a drone is designed to carry out a cargo mission. Thus, in this research, the dynamic model of a Hexacopter drone to deliver goods belongs to PT Aero Terra Scan is being developed. This dynamic modeling aims to further the drone's development by modeling it in 2 cases: no-payload and with a payload of 5 kg cases. The dynamic model of this Hexacopter is based on flight dynamics, a field of science studied in Aeronautical Engineering, and is implemented using Simulink. Through the results of this research, several conclusions have been withdrawn: (1) The drone's unstable nature characteristic inherently, even though it is analyzed from the initial hover condition. Thus, the drone and its system as a whole can never be separated with the feedback control that made it can maneuver adequately. (2) Several technical parameters of this Hexacopter, including the geometry, mass, the moment of inertia, until the estimation of motor throttle is required to achieve its hover conditions, both in the no-payload case and with-payload of 5 kg case. (3) The Hexacopter basic dynamic system model is based on the flight dynamics until its motion system control tuning through root locus map analysis using Simulink.

Phasing with near rectilinear Halo orbits: Design and comparison

In the next future, space agencies are planning to return to the Moon. The objective is to assemble an orbiting space station, called Gateway, on a Near Rectilinear Halo Orbit around the Moon as a base for future Moon and deep space missions. Within this framework, multiple side missions will be planned to sustain the Gateway (Artemis mission). The proposed work is thought in framework of the preliminary design of future cargo missions, in particular on the design of an efficient phasing trajectory, under the Circular Restricted Three body problem hypotheses, to bring a cargo vehicle from the end of the Earth-Moon transfer to the beginning of the proximity operations such as rendezvous and docking with the space station. The work aims covering the lack of literature in phasing trajectories with the NRHO by proposing three different strategies to connect the Earth-Moon transfer trajectory with the proximity operations. The three strategies are classified based on the choice of the parking orbits or the choice of the manifolds. Two strategies use butterfly and Halo orbits to park the vehicle before transferring to the target orbit. The third strategy, instead, uses manifolds to allow a direct phasing. In the paper, the three innovative strategies are designed and compare in a specific scenario.

Two and three impulses phasing strategy with a spacecraft orbiting on an Earth–Moon NRHO

The increasing interests in Moon exploration have led in recent years to international collaborations between space agencies aimed to assemble and operate the Gateway, an orbiting spaceport located on a Near Rectilinear Halo Orbit (NRHO) about the second libration point of the Earth–Moon system (EML-2). In this context, transfer strategies between the Earth and the spaceport cover a key role for both assembly and resupply missions. The presented work is thus focused on one leg of the transfer: the phasing maneuver. This work is conducted under the hypotheses of the Circular Restricted Three-Body problem (CR3BP) in which the Earth and the Moon are the only bodies that influence the spacecraft motion. The primaries are assumed to have a circular motion around their common barycenter. Under these hypotheses, is demonstrated the existence of periodic orbits, such as the Halo orbit family, that are exploited to design a proposed phasing maneuver. Two strategies are investigated: Halo parking orbit to NRHO two-impulses transfer and direct phasing with manifold exploitation. The first strategy is an optimal two-impulses transfer departing an EML-2 southern Halo and targeting the baseline NRHO. The second strategy considers the chaser already injected on the Gateway’s NRHO. Poincare maps are employed to identify unstable/stable manifolds intersections in search of low-energy phasing trajectories that leave the reference orbit along the unstable branch before re-entering it via the stable one. Three-impulses transfers with similar costs are found patching together these arcs. The two strategies are thus compared to highlight their advantages and disadvantages with the perspective of a real autonomous cargo mission around the Moon to refurnish the orbiting station.

Nuclear Power Concepts and Development Strategies for High-Power Electric Propulsion Missions to Mars

Under the Mars Transportation Assessment Study, NASA and the U.S. Department of Energy are performing analyses and generating concepts for crewed nuclear electric propulsion (NEP) missions to Mars. This paper presents the results of trade studies and concept development for the nuclear electric power system, consisting of the fission reactor, radiation shielding, power conversion, heat rejection, and power management and distribution (PMAD). The nuclear power team completed trade studies to evaluate different reactor and power conversion technologies and developed preliminary concepts for the crew shielding, waste heat radiators, and PMAD. The initial results suggest that a modified terrestrial microreactor combined with supercritical CO2 Brayton conversion could be used to perform the crew and cargo missions with satisfactory performance and modest risk. The paper includes preliminary development strategies that could bring the NEP technology to fruition for Mars missions in the late 2030s or early 2040s.

A new deployable pantographic lunar habitat

Deployable lunar structures have been and are being developed as an appropriate solution to the volumetric limitations of cargo mission flights. A deployable lunar habitat consisting of linear and planar elements connected through simple movable joints is introduced. This new structure is made of some linear units consisting of two types of scissor-like elements (SLE): 1) a modified polar scissor-like element and 2) a closed-loop of angulated scissor-like elements. After conducting theoretical research on lunar habitats, the evaluation criteria of transformable lunar structures are derived and then many new concepts are evaluated. The design considerations of such structures are discussed, and a proper form for these habitat types is selected. An experimental study is run to assess the best deployment strategy for the whole structure. By introducing the dimensional properties of the lunar habitats through physical and digital modeling process per crewmember, the derived alternatives from the previous stage are analyzed and the best transformation strategy that meets the aforementioned considerations is developed. Providing relatively high degrees of protection against environmental hazards through the combination of rigid plates with membrane, this proposed structure reduces the required extravehicular activity (EVA) time for assembly because of its quick and simple deployment mechanism. The folded state of the structure provides a low volume for the integration of subsystems.

Incidence and challenges of helicopter emergency medical service (HEMS) rescue missions with helicopter hoist operations: analysis of 11,228 daytime and nighttime missions in Switzerland

ObjectiveWe aimed to investigate the medical characteristics of helicopter hoist operations (HHO) in HEMS missions.MethodsWe designed a retrospective study evaluating all HHO and other human external cargo (HEC) missions performed by Swiss Air-Rescue (Rega) between January 1, 2010, and December 31, 2019.ResultsDuring the study period, 9,963 (88.7 %) HEMS missions with HHO and HEC were conducted during the day, and 1,265 (11.3 %) at night. Of the victims with time-critical injuries (NACA ≥ 4), 21.1 % (n = 400) reached the hospital within 60 min during the day, and 9.1 % (n = 18) at night. Nighttime missions, a trauma diagnosis, intubation on-site, and NACA Score ≥ 4 were independently and highly significantly associated with longer mission times (p < 0.001). The greatest proportion of patients who needed hoist or HEC operations in the course of the HEMS mission during the daytime sustained moderate injuries (NACA 3, n = 3,731, 37.5 %) while practicing recreational activities (n = 5,492, 55.1 %). In daytime HHO missions, the most common medical interventions performed were insertion of a peripheral intravenous access (n = 3,857, 38.7 %) and administration of analgesia (n = 3,121, 31.3 %).ConclusionsNearly 20 % of patients who needed to be evacuated by a hoist were severely injured, and complex and lifesaving medical interventions were necessary before the HHO procedure. Therefore, only adequately trained and experienced medical crew members should accompany HHO missions.

Changes in Exosome Release in Thyroid Cancer Cells after Prolonged Exposure to Real Microgravity in Space

Space travel has always been the man’s ultimate destination. With the ability of spaceflight though, came the realization that exposure to microgravity has lasting effects on the human body. To counteract these, many studies were and are undertaken, on multiple levels. Changes in cell growth, gene, and protein expression have been described in different models on Earth and in space. Extracellular vesicles, and in particular exosomes, are important cell-cell communicators, being secreted from almost all the cells and therefore, are a perfect target to further investigate the underlying reasons of the organism’s adaptations to microgravity. Here, we studied supernatants harvested from the CellBox-1 experiment, which featured human thyroid cancer cells flown to the International Space Station during the SpaceX CRS-3 cargo mission. The initial results show differences in the number of secreted exosomes, as well as in the distribution of subpopulations in regards to their surface protein expression. Notably, alteration of their population regarding the tetraspanin surface expression was observed. This is a promising step into a new area of microgravity research and will potentially lead to the discovery of new biomarkers and pathways of cellular cross-talk.

Families of transfers from the Moon to Distant Retrograde Orbits in cislunar space

This paper presents different families of transfers from Low Lunar Orbit to Distant Retrograde Orbits in the Earth-Moon system and analyzes their characteristics from the perspective of velocity increment requirement, transfer time, and sensitivity to small perturbations. Initially, families of transfer trajectories in the Circular Restricted Three-Body Problem are found by the predictor-corrector method and numerical continuation. Each transfer uses two impulsive maneuvers. These families are computed for different parking orbit altitudes, constants of motion, and insertion points. Then, to investigate the sensitivity characteristics of each family, a metric based on Local Lyapunov Exponent is introduced and computed along each trajectory. This work provides various choices for future cargo and human-crewed missions in the cislunar space.

Efficacy and efficiency of indoor nighttime human external cargo mission simulation in a high-fidelity training Centre

BackgroundThe human external cargo (HEC) operations conducted by Helicopter Emergency Medical Services (HEMS) rarely take place at night, making it difficult for crew members to attain and maintain the level of expertise needed to perform winch operations in the dark. As EASA requirements for training cannot currently be met, we evaluated whether simulation training could be an option.MethodsThis paper reports on a training concept using indoor simulation for the training of nighttime HEC operations. Participants’ experience and perceptions were evaluated with a survey and the procedural and economic advantages of the simulation approach were compared with those of the usual outdoor HEC training.ResultsMost participants had limited exposure to real-life nighttime HEC missions before undergoing the simulation-based training. The frequency of training cycles in simulation was much higher compared to conventional training (60 cycles indoors vs. 20 outdoors for HEMS-TC, 20 cycles indoors vs. 4 outdoors for MCM). Trainees perceived that their technical and non-technical skills (NTS) improved with the training. The estimated costs of standard outdoor-based nighttime HEC training (138€ per cycle) are at least 6.5 times higher than the costs of indoor simulated training (approximately 21€ per cycle). With a change to simulation, carbon dioxide emissions could potentially be reduced by more than 35 tons.ConclusionsIndoor simulation training of night HEC operations has advantages with regard to cost-effectiveness, environmental friendliness, and self-reported improvements in skills and knowledge. Its use is feasible and could improve crew and patient safety and fulfill regulatory demands for training intensity.

Low thrust Earth–Moon transfer trajectories via lunar capture set

A cislunar cargo spacecraft with low-thrust propulsion traveling between the Earth and the Moon is essential for sustainable, long-term manned lunar exploration. In low-thrust Earth–Moon transfer (LTEMT), lunar capture is the primary prerequisite for spacecraft subject to the circular restricted three-body model. Therefore, this study identifies sufficient conditions for lunar capture, which are determined by the Jacobi integral and Hill’s region. This paper proposes a guidance scheme that includes thrust direction, thrust efficiency, and a five-stage flight control sequence based on the variation of the Jacobi integral. The LTEMT problem is then converted to an initial value problem of a differential equation with three parameters. Lunar capture set theories (LCSTs), which are convenient for identifying lunar capture sets, are presented and proved according to the continuous properties of the ordinary differential equation. Finally, the solutions of the LTEMT trajectories departing from a geosynchronous orbit with an altitude of approximately 35,827 km are discussed for different thrust accelerations and cut-off values of the thrust efficiency. The robustness is analyzed assuming that navigation and switching time errors are present to demonstrate the adaptability of this method. The results reveal that the proposed guidance scheme and LCSTs can provide technical support for future cislunar cargo missions.

Cargo logistics for a Notional Mars Base using Solar Electric Propulsion

The current aim of NASA's Journey to Mars is a stepwise approach towards landing humans on the Red Planet, culminating in a sustained presence. There are many recent studies on how this could be achieved in an evolvable and affordable manner. Most architectures begin with crewed missions to Phobos or Mars orbit in the mid-2030's, progress toward short-stay missions on the surface, and then culminate with regular, long-stay missions at a permanent outpost in the 2040's. A common factor of these architectures is that many robotic launches would be required in order to support the crew by prepositioning mission elements and other needed supplies. In this paper, the use of 150 kW reusable Solar Electric Propulsion (SEP) tugs as a means to deliver elements both to orbit and the surface is studied. The conceptual SEP tugs make use of continued technology developments that were initialized through the Asteroid Redirect Robotic Mission (ARRM). They would also be used to deliver food and supplies to sustain the crews similar to resupply missions for the International Space Station. These SEP tugs would cycle (with loitering) between staging orbits in cislunar space and Mars orbit. The concepts presented here focus on the use of SEP for purposes of delivering cargo, and could potentially be complimentary to many human Mars architectures presented in the literature.In order to characterize mission design parameters such as dates, masses, and durations, thousands of optimized trajectories were run using low-thrust optimization software. Solutions are found for all launch/arrival date pairs for the years 2038–2053. They can be displayed as contour plots called Bacon plots – the SEP equivalent of porkchop plots. Possible mission architecture concepts for a steady-state human presence on Mars along with the cargo missions needed to keep it functioning are described and the relevant mission parameters such as launch dates, masses, arrival dates, etc., are given. It was found that the reusable SEP tug architecture would be highly beneficial to the logistics of a sustainable Mars outpost.

Optimization of In-Space Supply Chain Design Using High-Thrust and Low-Thrust Propulsion Technologies

In-space propellant supply chain can be effectively established for manned missions if high-thrust crew vehicles and cargo tugs can be used in conjunction with low-thrust cargo tugs. Optimizing the campaign logistics network by considering all involved missions together, rather than focusing on mission-level planning alone, can provide an efficient method to support top-level architectural decisions. In this work, a framework is presented to help mission designers quickly obtain the Pareto front of initial mass in low Earth orbit versus campaign length by optimizing propellant supply locations for crew missions concurrently with propulsion technology options (e.g., high-thrust or low-thrust) for corresponding cargo deliveries. This methodology is demonstrated for a return-to-the-moon campaign that uses the cislunar space as a staging area. Discrete propulsion system sizes are used for cargo delivery vehicles, and reuse of these tugs is permitted in the model. Both high-thrust and low-thrust trajectory options are considered for each cargo mission. Although the costs for high-thrust trajectories are part of the inputs to the formulation, the costs for low-thrust trajectories are calculated internally because of their dependency on the thrust-to-payload-mass ratio. For an efficient evaluation of low-energy, low-thrust transfers in the Earth–moon system, an approximation method is implemented based on a feedback control law, Q-law, and invariant manifolds. With the developed methods, optimal predeployment strategies are calculated for a campaign of one, two, or three human lunar missions, thus quantifying the game-changing impact of low-thrust propulsion for logistics supply planning in the selected demonstration cases.

Integrated Targeting and Guidance for Powered Planetary Descent

This paper presents an on-board guidance and targeting design that enables explicit state and thrust vector control and on-board targeting for planetary descent and landing. These capabilities are developed utilizing a new closed-form solution for the constant thrust arc of the braking phase of the powered descent trajectory. The key elements of proven targeting and guidance architectures, including braking and approach phase quartics, are employed. It is demonstrated that implementation of the proposed solution avoids numerical simulation iterations, thereby facilitating on-board execution of targeting procedures during the descent. It is shown that the shape of the braking phase constant thrust arc is highly dependent on initial mass and propulsion system parameters. The analytic solution process is explicit in terms of targeting and guidance parameters, while remaining generic with respect to planetary body and descent trajectory design. These features increase the feasibility of extending the proposed integrated targeting and guidance design to future cargo and robotic landing missions.

Economic optimization of cargo airships

Historical strengths and weaknesses of airships were investigated to determine a mission suited for airships. The transatlantic cargo mission was selected to take advantage of the high payload and endurance qualities of airships while minimizing the frequency of ground handling. An optimization was performed to minimize the cost per ton mile of the airship with maximum velocity as the variable. Other design parameters were held constant and based on historical airship studies. The cost per ton mile and von Karman efficiency were used to compare the optimized airship designs with other modes of cargo transportation. An airship with a volume of 200,000 m3, which was the volume of the Hindenburg, would achieve a cost per ton mile of $1.03. This value equates to about 85 % the cost of an airplane, and five times the cost of a truck. A graph of von Karman efficiency showed that the airships proposed by this study could occupy a niche market between airplanes and trucks in terms of both efficiency and velocity.

High-Current Lanthanum Hexaboride Hollow Cathode for 10-to-50-kW Hall Thrusters

Space missions continue to demand higher power Hall thrusters that provide high thrust and long life. For in-space propulsion applications such as high-power orbit raising and cargo missions, Hall thrusters capable of operating in the range of 10-50 kW are in development. The hollow cathodes for these thrusters will be required to produce discharge currents of 20-100 A with lifetimes well in excess of 10 kh. A lanthanum hexaboride (LaB <sub xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">6</sub> ) hollow cathode with these capabilities has been developed that features graphite tubes or sleeves to provide a diffusion boundary that protects the LaB <sub xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">6</sub> electron emitter from chemical reactions with the refractory metal support structure and includes a long-life heater capable of igniting the high-temperature electron emitter. Cathode operation has been fully characterized at currents of 20-100 A including probe measurements of the plasma parameters inside the LaB <sub xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">6</sub> insert and energetic ion measurements in the cathode plume. The cathode was also tested at up to 200 A of discharge current to explore the cathode failure mechanisms. While the LaB <sub xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">6</sub> cathode operates at higher temperatures than conventional barium oxide-impregnated dispenser cathode, LaB <sub xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">6</sub> provides very high discharge currents, long life, and significantly less sensitivity to propellant impurities and air exposure than conventional dispenser cathodes.

Solid Thoriated Tungsten Cathode Arc Discharges for Electrically Propelled Spacecraft

Since the early 1960s to mid-1960s, laboratory studies have demonstrated the unique ability of magnetoplasmadynamic (MPD) thrusters to impart an exceptionally high level of specific impulse and thrust at large power processing densities. These intrinsic advantages are why MPD thrusters have been identified as a prime candidate for future long-duration space missions, including piloted Mars, Mars cargo, lunar cargo, and other missions beyond the low Earth orbit. The large total impulse requirements inherent of the long-duration space missions demand the thruster to operate for a significant fraction of the mission burn time while requiring the cathodes to operate at 1-10 000 kW for 2-10 000 h. The high current levels lead to high operational temperatures and a corresponding steady depletion of the cathode material by evaporation. This mechanism has been identified as the life-limiting component of MPD thrusters. Continuing work in the Laboratory of Astronautical Plasma Dynamics at the University of Southern California (USC) is focused on extending the operational life of the cathode. In this paper, axial temperature profiles obtained under varying current levels (20-60 A) and argon gas mass flow rates (450-640 sccm) of a solid 2% thoriated tungsten cathode are presented. The electron temperatures and electron densities near the “active zone” (the surface area of the cathode responsible for approximately 70% of the emitted current) were measured using Langmuir probing. The results will be set as boundary conditions for solid cathode models in development at USC.

VX-200 Magnetoplasma Thruster Performance Results Exceeding Fifty-Percent Thruster Efficiency

H IGH-POWER electric propulsion thrusters can reduce propellant mass for heavy-payload orbit-raising missions and cargo missions to the moon and near-Earth asteroids, and they can reduce the trip time of robotic and piloted planetary missions [1–4]. The Variable Specific Impulse Magnetoplasma Rocket (VASIMR®) VX-200 engine is an electric propulsion system capable of processing power densities on the order of 6 MW=m with a high specific impulse (4000 to 6000 s) and an inherent capability to vary the thrust and specific impulse at a constant power. The potential for a long lifetime is due primarily to the radial magnetic confinement of both ions and electrons in a quasi-neutral flowing plasma stream, which acts to significantly reduce the plasma impingement on the walls of the rocket core. High-temperature ceramic plasma-facing surfaces handle the thermal radiation: the principal heat transfer mechanism from the discharge. The rocket uses a helicon plasma source [5,6] for efficient plasma production in the first stage. This plasma is energized further by an ion cyclotron heating (ICH) RF stage that uses left-hand polarized slow-mode waves launched from the high field side of the ion cyclotron resonance. Useful thrust is produced as the plasma accelerates in an expanding magnetic field: a process described by conservation of the first adiabatic invariant as the magnetic field strength decreases in the exhaust region of the VASIMR [7–9]. End-to-end testing of the VX-200 engine has been undertaken with an optimum magnetic field and in a vacuum facility with sufficient volume and pumping to permit exhaust plumemeasurements at low background pressures. Experimental results are presented with the VX-200 engine installed in a 150 m vacuum chamber with an operating pressure below 1 10 2 Pa (1 10 4 torr), and with an exhaust plume diagnostic measurement range of 5 m in the axial direction and 1 m in the radial directions. Measurements of plasma flux, RF power, and neutral argon gas flow rate, combined with knowledge of the kinetic energy of the ions leaving the VX-200 engine, are used to determine the ionization cost of the argon plasma. A plasmamomentum flux sensor (PMFS)measures the force density as a function of radial and axial positions in the exhaust plume. New experimental data on ionization cost, exhaust plume expansion angle, thruster efficiency, and total force are presented that characterize the VX-200 engine performance above 100 kW. A semiempirical model of the thruster efficiency as a function of specific impulse has been developed to fit the experimental data, and an extrapolation to 200 kWdc input power yields a thruster efficiency of 61% at a specific impulse of 4800 s.

Single-Channel Hollow Cathodes in 5–20-eV Argon Discharge for Spacecraft Thruster Applications

High-power magnetoplasmadynamic (MPD) thrusters are an attractive option for providing primary propulsion on flagship-class spaceflight missions beyond the Low Earth Orbit (LEO), such as piloted Mars and Mars cargo, and more near-term cargo missions supporting a crewed lunar outpost due to their unique ability to process large amounts of power in a relatively small footprint size. This results in significant savings in system mass and volume over competing propulsion options such as ion or Hall-effect thrusters. This paper presents a partial summary of data from recent research at the University of Southern California, in collaboration with NASA's Jet Propulsion Laboratory, investigating several concerns of the mechanisms influencing the performance and lifetime of high-current single-channel hollow cathodes (SCHCs), the central electrode, and the primary life-limiting component in MPD thrusters. The overall objective of such research is to extend the operational lifetime of the cathode to mission-enabling lengths, generally considered to be in the range of 5-10 000 h. Specifically covered are the trends seen in the discharge efficiency and power, the size of the plasma attachment to the cathode (the active zone), the cathode exit plume plasma density and energy, and the plasma property distributions of the internal plasma column of an SCHC.

Designing from minimum to optimum functionality

This paper discusses a multifaceted strategy to link NASA Minimal Functionality Habitable Element (MFHE) requirements to a compatible growth plan; leading forward to evolutionary, deployable habitats including outpost development stages. The discussion begins by reviewing fundamental geometric features inherent in small scale, vertical and horizontal, pressurized module configuration options to characterize applicability to meet stringent MFHE constraints.A proposed scenario to incorporate a vertical core MFHE concept into an expanded architecture to provide continuity of structural form and a logical path from “minimum” to “optimum” design of a habitable module.The paper describes how habitation and logistics accommodations could be pre-integrated into a common Hab/Log Module that serves both habitation and logistics functions. This is offered as a means to reduce unnecessary redundant development costs and to avoid EVA-intensive on-site adaptation and retrofitting requirements for augmented crew capacity. An evolutionary version of the hard shell Hab/Log design would have an expandable middle section to afford larger living and working accommodations.In conclusion, the paper illustrates that a number of cargo missions referenced for NASA’s 4.0.0 Lunar Campaign Scenario could be eliminated altogether to expedite progress and reduce budgets. The plan concludes with a vertical growth geometry that provides versatile and efficient site development opportunities using a combination of hard Hab/Log modules and a hybrid expandable “CLAM” (Crew Lunar Accommodations Module) element.