Key Technologies,Systems,and Infrastructure Enabling the Commercialization and Human Settlement of Low Earth Orbit,the Moon,and Cislunar Space

Introduction

More than 50 years have passed since Kubrick and Clarke's movie 2001: A Space Odyssey debuted in April 1968. ¹ For many of us this film brought to life the exciting possibilities awaiting humankind beyond the Apollo program—images of commercial space planes, large artificial gravity stations (AGS) orbiting Earth, and commuter flights to sprawling settlements on the Moon. The Moon is again a destination of great interest to the United States and the worldwide space community. Located just 3 days from Earth, the Moon has abundant resources and is an ideal location to demonstrate and test key technologies and systems (surface habitation, long-range pressurized rovers, surface power and resource extraction systems) that will allow people to explore, work, and live self-sufficiently both there, and afterward, on Mars.

Abroad, plans for human surface missions and settlements on the Moon in the 2025–2030 timeframe are openly being discussed in Europe, China, and Russia. In the United States, a number of private companies—Axiom Space, ² SpaceX, ³ United Launch Alliance (ULA), ⁴ and Blue Origin ⁵ —are discussing commercial ventures in low Earth orbit (LEO), the Moon, and Cislunar space along with public–private partnerships with NASA.

This renewed interest in the Moon by U.S. industry and other international space agencies prompted the Trump Administration to implement Space Policy Directive-1 in December 2017⁶ directing NASA to lead an innovative space exploration program to send American astronauts back to the Moon for long-term exploration and use. The schedule for returning American astronauts to the Moon was subsequently accelerated in March 2019, when Vice President Pence, speaking on behalf of the President, directed NASA to accomplish this goal by 2024. He also suggested the lunar south pole as the destination—a region thought to contain abundant water ice inside the permanently shadowed craters (PSCs) that exist there. ⁷

Lunar-derived propellant (LDP) production—specifically lunar-derived liquid oxygen (LLO₂) and lunar-derived liquid hydrogen (LLH₂)—has been identified as a key technology offering significant mission leverage ⁸ and it figures prominently in ULA's plan for developing a Cislunar space economy. ⁴ Similarly, Jeff Bezos' Blue Origin is developing a large lunar lander, called the Blue Moon, capable of delivering up to ∼4,536 kg of payload (PL) to the lunar surface (LS). ⁹ Powered by liquid oxygen (LO₂)/liquid hydrogen (LH₂) engines, it would also be able to use LDPs once they become available.

Samples returned from different sites on the Moon during the Apollo missions have shown that the lunar regolith has significant oxygen content. The iron oxide-rich volcanic glass beads returned on the final Apollo (17) mission have turned out to be a particularly attractive source material for oxygen extraction based on hydrogen reduction experiments conducted by Allen et al. ¹⁰ Post-Apollo lunar probe missions have also provided orbital data indicating the existence of large quantities of water ice trapped in deep PSCs located at the Moon's poles. ¹¹

So the stage is now set. NASA has been directed to return to the Moon for long-term exploration and use, and companies, large and small, both within and outside the United States, are anxious to start doing business on the Moon. ¹² What will the outcome of this next chapter in humankind's exploration of the Moon be? Today, 52 years after the Apollo 11 lunar landing, the images of Dr. Floyd's commuter flight to the Moon remain a source of inspiration to many, including the author (S.K.B.).

In this article, we examine some of the key technologies, systems, and supporting infrastructure that could be developed over the next 30 years by industry for NASA initially, and then for private commercial ventures, and that could allow the operational capabilities portrayed in 2001 to be achieved, although on a more Spartan scale. The intent of this article is not to provide a “business case” for the technologies, systems, and infrastructure presented here but rather to show a future vision of possibilities when they are integrated together. It also attempts to quantify the performance requirements of these systems, something that interested government and industry technologists and system engineers may find useful when considering development growth paths for these systems going forward.

Included in this article are the following topics. First, the benefits and options for using LDPs are discussed. Then, the two primary feedstock materials under consideration—icy regolith obtained from polar PSCs, and volcanic glass beads from vast pyroclastic deposits on the lunar nearside—are reviewed, and proposed concepts for their mining are discussed. Next, system descriptions of two candidate propulsion options—the LO₂/LH₂ RL10B-2 engine and a variant of the nuclear thermal rocket (NTR), the LO₂-Augmented NTR (LANTR)—are provided, including their currently existing and projected performance characteristics.

The important roles that fission power systems (FPSs) and space transportation nodes (STNs) are expected to play in supporting a reusable, space-based lunar transportation system (LTS) are discussed next. Concepts for crewed cargo transports (CCTs) and commuter shuttles are then presented along with the refueling requirements needed to support missions with varying trip times to polar (LPO) and equatorial lunar orbits (LLO). A comparison of the LDP production and mining requirements using icy regolith and volcanic glass follows, including a discussion of the synergy of LUNOX (another name for LLO₂) production with an evolving helium (He)-3 mining industry. The article ends with a summary of our findings and concluding remarks.

Benefits of Using and Options for Producing LDPs

Past NASA studies ⁸ have indicated a substantial benefit from using LDPs in a LTS using LO₂/LH₂ chemical rockets. Of the ∼6 kg of mass in LEO required to place 1 kg of PL on the LS, ∼70% (4.2 kg) is propellant assuming the engines operate with an oxygen-to-hydrogen mixture ratio (O/H MR) of ∼6:1. Also, since the cost of placing a kg of mass on the LS is approximately six times the cost of delivering it to LEO, ⁸ the ability to produce and utilize LLO₂ from processed lunar volcanic glass, or LLO₂ and LLH₂ from the electrolysis of lunar water (LH₂O), derived from lunar polar ice (LPI), can provide a significant mission benefit. By providing a local source of oxygen and hydrogen for use in life support systems, fuel cells, and the chemical rocket engines used on lunar landing vehicles (LLVs), the Initial Mass in Low Earth Orbit (IMLEO), launch costs, and LTS size and complexity can all be reduced. Greater quantities of “higher value” cargo (e.g., people, mining and propellant processing equipment, and scientific instruments) can also be transported to LEO and on to the Moon instead of bulk propellant mass, further reducing LTS costs.

LPI: Estimated Quantities and Locations

In the post-Apollo era, orbiting robotic spacecraft (SC)^13–15 have provided data indicating the existence of trapped water ice within a number of deep PSCs found near the Moon's poles. On the basis of the data provided by these SC, estimates of the water ice concentrations in the polar regolith varied from ∼0.7 to 8.5 wt % and the total quantity of LPI at both poles ranged from ∼600 million to ∼2 billion metric tons. Recently, the existence of surface water ice at the Moon's poles was confirmed and reported by Li et al., ¹⁶ whose team took a fresh look at data from NASA's Moon Mineralogy Mapper instrument that flew on the Chandrayaan-1 SC. ¹⁴ Figure 1 shows the distribution of surface ice at the Moon's south (Fig. 1a) and north poles (Fig. 1b). The turquoise dots represent the ice locations, and the gray scale corresponds to the surface temperature, with the darker gray representing colder areas and the lighter gray indicating warmer locations. As is evident, the ice is present at the coldest and darkest spots on the LS within 20° of both poles, and is more abundant in the south, where it is principally found at the bottoms of PSCs. In the north, the ice appears more widely dispersed and less concentrated.

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Fig. 1.

Location of surface water ice at the Moon's polar regions. ¹⁶ (a) North poles. (b) South poles.

Although considerable enthusiasm has been expressed about mining and processing LPI for rocket propellant, and using it to create a space-faring Cislunar economy, ¹⁷ the “ground truth” about LPI will need to be established before this enthusiasm is warranted. Robotic surface missions will need to be sent to potential sites of interest to quantify the physical state of the water ice (e.g., its concentration within the regolith), its vertical thickness and areal extent, and the levels of soil contamination. The depth, slope, and interior thermal environment of the PSCs must also be assessed.

Conditions and Concepts for LPI Mining

The PSCs where LPI exists can be deep and extremely cold, posing major engineering challenges for mining and processing the ice-bearing regolith. To put the operating temperature conditions into perspective, the world's 10 coldest mines are located in Russia, and all but one of these are located in Russia's Sakha Republic—a region in the country's extreme north that contains vast diamond, coal, and gold resources. At the coldest of these mines, Sarylakh, the temperatures can drop to nearly −50°C (∼223°K). By contrast, the temperatures inside PSCs can vary from ∼30 to 50°K—significantly colder than the coldest mines on Earth!

In addition to working in dark, extremely cold surroundings, mining equipment must be designed to operate in a hard vacuum, on electricity rather than petrol, and in gravity that is 1/6th that of Earth. It must also tolerate an increased radiation environment, and the abrasive nature of the lunar dust, which can cause increased rubbing friction, wreak havoc on machinery, and adheres to everything it touches.

Surface mining is the most common approach to mineral extraction here on Earth and different ice-regolith mixtures can impact the excavation process. Gertsch et al. ¹⁸ conducted load-penetration tests on samples of lunar regolith simulant (JSC-1) containing varying levels of water ice content (from 0 to ∼12 wt %). The samples were compacted and cooled to 77°K by using LN₂ to simulate conditions expected in lunar cold traps. From the test measurements, Gertsch et al. matched the different ice-regolith mixtures to the following types of terrestrial mined rocks: (1) at ∼0.6 to 1.5 wt % ice, the mixture behaves like weak shale or mudstone and is readily excavatable; (2) at ∼8.5 wt % ice, the mixture behaves like moderate-strength limestone and sandstone and is excavated by using mechanical excavators; and (3) at ∼10 to 12 wt % ice, the mixture behaves like strong limestone, sandstone, and high-strength concrete, and requires massive excavators.

An innovative technology for regolith excavation and transport known as the Regolith Advanced Surface Systems Operations Robot, or RASSOR for short, ¹⁹ has been developed by NASA's Swamp Works Laboratory at the KSC and is available for licensing through NASA's Technology Transfer Program. The RASSOR prototype (shown in Fig. 2) uses counter-rotating bucket drums positioned at the front and rear of its central mobility chassis to provide near-zero horizontal and minimal vertical net reaction force allowing it to load, haul, and dump regolith under extremely low gravity conditions with high reliability. When RASSOR's bucket drums are filled, it raises its arms (Fig. 2) allowing the central mobility chassis to drive to the processing facility where it unloads the collected regolith by spinning its drums in reverse, allowing the regolith to pour out into the collection bin.

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Fig. 2.

RASSOR—with raised rotating bucket drums. RASSOR, Regolith Advanced Surface Systems Operations Robot.

An alternative option to excavating and transporting icy regolith has been proposed and analyzed by the Colorado School of Mines. ²⁰ Known as “thermal mining,” this in situ approach uses directed sunlight from the crater rim to heat the surface of the icy regolith, producing sublimated water vapor that is then captured within a dome-shaped tent enclosure (Fig. 3) covering the heated surface. When the regolith's ice content is depleted, the tent would be moved to another location.

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Fig. 3.

Concept for LPI Thermal Mining Plant. LPI, lunar polar ice.

The collected water vapor is then vented into “cold trap” ice haulers for transport to a central processing plant for water purification and subsequent electrolysis to produce the LLO₂ and LLH₂ propellants used by surface-based LLVs. The purified LH₂O can also be shipped to orbiting STNs for conversion to propellant there.

LUNOX: Extraction Efficiency and Siting Locations

As discussed earlier, the information regarding LPI is based on the analysis and interpretation of orbital data from past robotic science missions. By contrast, samples brought back on the Apollo missions have shown that nearly half the mass (∼43%) of the Moon's surface material is oxygen ⁸ and at least 20 different techniques ²¹ have been identified for its extraction. Hydrogen reduction of iron oxide (FeO) in the mineral “ilmenite” (FeTiO₃), or in FeO-rich volcanic glass, is among the simplest and best studied. The technique involves a two-step process in which the FeO is first reduced to metal, liberating oxygen and forming water. The water is then electrolyzed to produce oxygen, and the hydrogen is recycled back to the processing plant to react with more feedstock material. From an extensive set of hydrogen reduction experiments by Allen et al., ²² “ground truth” for oxygen release was established by using samples of lunar soil and volcanic glass beads returned by the Apollo missions. The results indicated that oxygen can be produced from a wide range of lunar soils and is strongly correlated with the Fe abundance in the soil, as shown in Figure 4. Iron-rich highland soils produced the smallest amount of oxygen, ∼1 to 2 wt %, whereas iron-rich mare soil samples produced ∼3.6 wt %. The highest yields—in the range of 4–5 wt %—were obtained from the pyroclastic (volcanic) glass collected at the Apollo 17 Taurus-Littrow landing site. The glass is extremely iron-rich with an Fe content of ∼17.8 wt %. The orange and black beads shown in Figure 4a have identical elemental compositions, but the black beads are largely crystalline whereas the orange beads are largely glass. Reduction of the orange glass beads produced an oxygen yield of ∼4.3 wt % whereas the black crystalline beads produced ∼4.7 wt %, the highest for any of the samples (Fig. 4b). ²² Assuming the hydrogen reduction process, volcanic glass feedstock, and a conservative oxygen yield of 4 wt %, 1 t of LUNOX can be produced from 25 t of volcanic glass—a significant improvement in oxygen yield over that using ilmenite-bearing feedstock material. ²¹

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Fig. 4.

Volcanic glass beads and oxygen yields from full range of Apollo samples. ²²

Besides its higher oxygen yield, volcanic glass is an attractive feedstock material because it is uniformly fine grained, reacts rapidly, and can be fed directly into the LUNOX production plant with little or no processing before reduction. More importantly, a significant number of large pyroclastic deposits, thought to be the result of continuous, Hawaiian-style, fire-fountain eruptions from large vents, have been identified on the lunar nearside. ²³ These dark mantle deposits (DMDs) are of regional extent and are composed largely of crystallized black beads, orange glass beads, or a mixture of the two. Large DMDs located just north of the lunar equator include: (1) the Aristarchus Plateau (∼49,015 km²); (2) Sinus Aestuum (∼10,360 km²); (3) Rima Bode (∼6,620 km²); (4) Sulpicius Gallus (∼4,320 km²); (5) Mare Vaporum (∼4,130 km²); and (6) Taurus Littrow (∼2,940 km²).

An attractive site for a possible commercial LUNOX production facility is the Taurus-Littrow DMD located at the southeastern edge of Mare Serenitatis (∼21°N, ∼29.5°E) ∼30 km west of the Apollo 17 landing site. This deposit of largely black crystalline beads covers ∼3,000 km², is thought to be tens of meters thick, and could yield well in excess of a billion metric tons of LUNOX by using the hydrogen reduction process, a 4.5 wt % oxygen yield, and a 5-m mining depth.

Figure 5 depicts a conceptual LUNOX production facility developed and first presented by the author (S.K.B.) in 1997. ²⁴ Shown in the lower left foreground are two lunar industrialists discussing planned expansions at the LUNOX facility. Toward the top, modular production units, supplied with volcanic glass feedstock by autonomously operating RASSOR-type excavator haulers, generate copious amounts of LUNOX stored in well-insulated tanks adjacent to the facility. At the top right, a bottom-loaded “Sikorsky-style” LLV lifts off from the surface carrying a tank of LUNOX to a propellant depot in LLO, while at the adjacent landing pad, a second LLV awaits servicing before its next mission. In the right foreground, increased numbers of government and industry personnel have taxed the capacity of several previously landed habitat modules necessitating construction of an inflatable dome for added living space. The dome's exterior is covered by bagged regolith to provide shielding against solar flares and galactic cosmic radiation. Lastly, FPSs—positioned within nearby craters with overhead surface radiators—will be critical to providing a good return to investors in the LUNOX enterprise. They provide abundant power at low mass to support continuous operation of the surface mining vehicles, production units, and habitat modules even during the 2-week lunar night.

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Fig. 5.

LUNOX facility near Taurus-Littrow DMD. ²⁴ DMD, dark mantle deposit.

As the production capacity of the LUNOX enterprise increases, additional supporting commercial activities are expected to emerge, including metals processing (e.g., iron and titanium from the H₂ reduction process), power generation, maintenance and operations of surface-based LLVs and LLO STNs, and, eventually, even a lunar tourism industry complete with routine commuter flights to and from the Moon.

Propulsion System Options: Chemical RL10B-2 and the LANTR

Two propulsion technologies are examined in this article. The first option is “now technology” and is represented by the LO₂/LH₂ RL10B-2 engine. ²⁵ Derived from the long line of proven RL10 engines, the RL10B-2 has been the “workhorse” of the commercial launch industry powering the upper stages of the medium and heavy-lift versions of the ULA's Delta IV launch vehicle, as well as the upper stage of the Delta III. It features the world's largest extendible carbon–carbon nozzle allowing the RL10B-2 to achieve the highest specific impulse (I_sp) of any cryogenic engine—465.5 s. The RL10B-2's thrust level, O/H MR, and thrust-to-weight ratio are 24.75 klb_f, 5.88:1, and 37.3, respectively. Pictures of the RL10B-2 engine, with its nozzle retracted and deployed, are shown in Figure 6. In this stowed configuration (Fig. 6a), the engine length is ∼2.2 m (86.5″). When deployed (Fig. 6b), the engine length is ∼4.15 m (163.5″), and the nozzle exit diameter is ∼2.15 m (84.5″). From an operational standpoint, the current service life and total number of engine starts for the RL10B-2 are reported ²⁶ to be 3,500 s and 15 starts. These values will have to increase significantly if tomorrow's RL10-derivative engines are to be used in a reusable LTS.

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Fig. 6.

RL10B-2 with extendible carbon-carbon nozzle.

The second propulsion option considered is the NTR—an important propulsion technology for Mars missions that is receiving considerable attention and funding from NASA at present. The NTR uses a compact fission reactor core containing uranium (U)-235 fuel used to generate 100s of megawatts of thermal power (MW_t) required to heat the LH₂ propellant to high exhaust temperatures for rocket thrust. ²⁷ Key features of an “expander cycle” NTR are shown in Figure 7.

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Fig. 7.

Key features of an expander cycle NTR engine. NTR, nuclear thermal rocket.

A three-engine cluster of Small Nuclear Rocket Engines (SNREs) ²⁸ is used in this article, with each engine producing ∼16.5 klb_f of thrust with an I_sp of ∼900 s. The total engine length is ∼5.8 m with its radiation-cooled, retractable nozzle section fully extended. The nozzle area ratio, exit diameter, and engine thrust-to-weight ratio are 300:1, ∼1.53 m, and ∼3.02, respectively.

LO₂-Augmented NTR

To take advantage of the mission benefits of refueling with LLO₂ and LLH₂ for Earth return, each SNRE is outfitted with an O₂ “afterburner” nozzle containing O₂ injectors and an O₂ feed system. The oxygen is stored as a cryogenic liquid at low pressure and is pressurized and gasified before its injection into the nozzle. This is accomplished by diverting a small fraction of the engine's hydrogen flow to an oxidizer-rich gas generator that drives an LO₂ turbopump assembly used to deliver the gasified O₂ (GO₂) to injectors positioned inside the divergent nozzle section downstream of the throat. Here, it mixes with the hot H₂ and undergoes supersonic combustion, adding both mass and chemical energy to the rocket exhaust—essentially “scramjet propulsion in reverse.” A simplified schematic of LANTR engine operation is illustrated in Figure 8.

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Fig. 8.

Simplified LANTR schematic. LANTR, LO₂-Augmented NTR.

By varying the O/H MR, LANTR engines can operate over a range of thrust and I_sp values while the reactor core power level remains relatively constant. As the MR varies from 0 (only LH₂) to 5 (LO₂-rich operation), the engine thrust level increases from 16.5 to ∼56.8 klb_f (over 344%) whereas the I_sp decreases from 900 to 516 s (∼57%), which is still 54 s higher than the RL10B-2. Although the LANTR engine can operate at higher I_sp than the RL10B-2, the LANTR is ∼9.5 times heavier and requires extra shielding mass to reduce crew radiation exposure. Additional performance characteristics of the SNRE-class LANTR are provided in Borowski et al. ²⁸

THE Importance of STNs and FPSs

Commercial STNs, providing propellant and cargo transfer services in LEO and lunar orbit, will be key to realizing a robust, reusable LTS in the second half of the 21st century. Supplied with LO₂ and LH₂ propellants from Earth, delivered by a new generation of low-cost, reusable heavy lift vehicles, and LDPs from the Moon, strategically located STNs will become transportation hubs for a variety of lunar transfer vehicles (LTVs) operating in Cislunar space. A concept design for a LEO STN, called Oasis, is shown in Figure 9 along with the key features it needs to perform its propellant and cargo transfer functions. It is here that LTVs such as the CCT (shown docked to Oasis in Fig. 9) will be resupplied with propellant and cargo for their next scheduled delivery to the Moon. An FPS is used to supply the high electrical power Oasis needs (approximately 0.5–1 MW_e) for cargo and propellant transfer operations, on-board cryofluid management, and habitat module life support. It uses twin liquid metal-cooled, fast spectrum reactors with uranium nitride fuel, Brayton power conversion, and a deployable, fold-out radiator system, which can be collapsed, allowing the entire FPS to be delivered on a single Space Launch System (SLS) launch. ²⁹

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Fig. 9.

Oasis commercial LEO STN—key features and activities. LEO, low Earth orbit; STN, Space Transportation Node.

As LDP production levels increase and operation of surface-based LLVs become routine, development of commercial STNs would be expected in both polar and equatorial lunar orbits. Because abundant deposits of volcanic glass are located at a number of sites just north of the lunar equator, a STN established in equatorial LLO could be routinely supplied with LUNOX by tanker LLVs operating from LUNOX production facilities. Similarly, LH₂O, derived from processing icy regolith at the poles, could be transported to a STN in LPO by water tanker LLVs. Here, the water would be electrolyzed and the LDPs stored for subsequent use. Besides providing their propellant and cargo transfer function, lunar STNs will also provide convenient staging locations where CCTs and commuter shuttles can drop off cargo and passengers that would then be picked up by LLVs for transport to the LS.

The LLO STN, Serenity Shores, shown in Figure 10, derives its name from the FeO-rich volcanic glass DMD located at the southeastern edge of the Sea of Serenity. It is a clone of Oasis and has all of the same features needed to unload cargo from arriving CCTs destined for the LS. While cargo is being unloaded, the CCT would be refueled with LLO₂, delivered by LUNOX tanker LLVs, for its return to Earth. Both activities are shown in Figure 10. Periodic shipments of Earth-supplied LH₂ (ELH₂) would supply the STN with the LH₂ needed by the LLVs. For the LPO STN, higher electrical power levels from the FPS will be required to support onboard water electrolysis and propellant production.

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Fig. 10.

Serenity Shores commercial LLO STN.

One-way transit times to and from the Moon on the order of ∼72 h would be the norm initially. Eventually, however, as lunar outposts grow into permanent settlements staffed by visiting scientists, engineers, and administrative personnel representing both government and private ventures, more frequent flights of shorter duration could become commonplace. As Table 1 shows, decreasing the Earth-to-Moon transit time from 72 to 36 h increases the outbound ΔV (velocity change increment) requirement from ∼4.0 to 4.9 km/s and the total roundtrip ΔV requirement by ∼1.8 km/s. Decreasing 1-way flight times from 72 to 24 h increases the round trip ΔV requirement by ∼4.9 km/s, to ∼12.9 km/s! As a result, long lifetime engines and LDP for refueling will be needed for LTVs of reasonable size.

To access a STN in a 300-km circular LPO, a three-burn lunar orbit capture (LOC) maneuver is assumed. The first burn captures the LTV into an elliptical orbit around the Moon with a perigee altitude of 300 km—the same as the final parking orbit. A second burn is then performed at apogee to change the plane of the orbit to match the inclination of the desired parking orbit—in this case 90° for LPO. The third and final burn is performed near perigee to lower the orbit's apogee, resulting in the final 300-km circular LPO. The duration of the LOC maneuver used in this study is 2.5 h, and a larger total capture ΔV is required compared with the single LOC burn ΔV values shown in Table 1.

Similar to the capture maneuver, trans-Earth injection (TEI) requires three burns and 2.5 h to complete as well. The first burn raises the apogee of the orbit, resulting in a highly elliptical orbit around the Moon. The second burn is a plane change burn performed near apogee that adjusts and aligns the plane of the elliptical orbit from 90° to that needed for departure. The third and final burn is again performed near perigee and after it is completed, the LTV has escaped the Moon and is on its trajectory back to Earth.

Conestoga: A Reusable, Space-Based CCT

The Conestoga wagons of old, developed in Pennsylvania in the early 1700s, were called the “Ships of Inland Commerce” and were used to transport settlers, farm produce, and freight across Pennsylvania and the nearby states of Maryland, Ohio, and Virginia for more than 150 years. ³⁰ Named after its ancestor, the Conestoga CCT can deliver varying amounts of cargo (from 10 to 36 t) to lunar orbit depending on the desired transit times out to the Moon and back. Conestoga's four key elements shown in Figure 11 include: (1) a common three-engine propulsion stage (PS) carrying ∼40 t of LH₂ propellant; (2) an in-line tank assembly with a conical adaptor that carries ∼111 t of LO₂ propellant; (3) a 16-m long, four-sided star truss with attached PL; and (4) a forward habitat module that supports a crew of 4 and has a mass of ∼10 t. Refueling ports and twin photovoltaic arrays (PVAs) are also located at the forward ends of the PS and in-line tank assembly. Additional information on the missions, PLs, and transportation system ground rules and assumptions used in this article can be found in Borowski et al. ²⁸

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Fig. 11.

Key features and dimensions for the Conestoga CCT. CCT, crewed cargo transport.

Once loaded with cargo and propellant, Conestoga separates from the LEO STN (Fig. 12), performs the trans-lunar injection (TLI) burn, and departs for the Moon. After capture into either a lunar equatorial or polar orbit, it rendezvouses and docks with the lunar STN where its PL is removed, and its propellant tank(s) are refueled for the trip back to Earth (shown in Fig. 10).

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Fig. 12.

Conestoga in LEO with attached payloads. LEO, low Earth orbit.

Conestoga-class missions carrying 20 t of PL to LPO were compared by using RL10B-2 and LANTR propulsion. The mission objective was to determine the minimum amount of LO₂ and LH₂ propellant required in both LEO and LPO. The assumed LO₂-to-LH₂ refueling ratio was 8:1, the same O/H mass ratio produced during water electrolysis. For 72-h one-way transit times, the LANTR engines were run H₂-rich on the outbound mission leg and O₂-rich inbound (for the TEI and Earth orbit capture burns), lowering the CCT's refueling requirements in LEO and LPO to ∼102 and ∼61 t, respectively. The IMLEO, mission ΔV, and total engine burn time are ∼186 t, ∼8.378 km/s, and ∼30 min, respectively.

When using RL10B-2 engines, with their higher O/H MR (5.88:1) and lower I_sp (465.5 s), the CCT's refueling requirements are larger in LEO (∼124 t) and LPO (∼62.6 t). The larger LEO refueling requirement is offset, however, by the RL10B-2's lower PS mass (∼53% that of the LANTR CCT). The LANTR CCT uses heavier engines (∼9.5 times that of its RL10B-2 counterpart), and additional shield mass is required on each engine to reduce crew radiation exposure during the mission. The result is only a slightly larger IMLEO (∼192.5 t) for the RL10B-2 option. However, because of the larger propellant loading out and back, and lower I_sp, the total engine burn time for the RL10B-2 CCT is longer at just more than 42 min.

For minimum transit time CCT missions (on the order of ∼40 h), the LANTR CCT refueling requirements are ∼151 t in LEO and ∼123 t in LPO. For the RL10B-2 option, the LEO and LPO refueling requirements are ∼135 and ∼118 t, respectively. The LANTR engines also run O₂-rich out and back, lowering their I_sp advantage over the RL10B-2. This plus the heavier PS mass increases the IMLEO for the LANTR CCT option to ∼235 t compared with ∼203 t for a CCT using RL10B-2 engines. With shorter transit times out and back, the total mission ΔV and engine burn time also increase to ∼10.865 km/s and ∼34 min for the LANTR CCT and ∼10.550 km/s and ∼58 min for RL10B-2 CCT option.

For missions to LLO, the CCT operates between LEO and equatorial LLO, uses only ELH₂, and refuels with only LLO₂ before returning to Earth. Without the additional ΔV required to access and depart from LPO, the total mission ΔVs are noticeably lower for both propulsion options, and especially so for the short transit time missions. The minimum transit time achievable by a LANTR CCT delivering 20 t of PL to LLO is ∼46 h with corresponding IMLEO, total mission ΔV, and engine burn time values of ∼232 t, ∼8.883 km/s, and ∼25 min. For the RL10B-2 CCT, the minimum transit time is ∼51.5 h and the corresponding values are ∼206 t, ∼8.710 km/s, and ∼46 min. Additional details on the missions just cited can be found in Borowski et al. ²⁸

With improvements in engine service lifetime (to 10 h or more), and the availability of LDPs at strategically positioned STNs in LPO and LLO, Conestoga-class CCTs can provide the basis for a robust and flexible LTS that offers a wide range of cargo delivery capability and transit times. Today, “time is money” for the long-distance freight haulers traveling our highways, oceans, and skies. In the future, Conestoga-class vehicles could play the same important role in establishing Cislunar trade and commerce as the Conestoga wagons of old did for more than a century throughout Pennsylvania and its neighboring states.

Feasibility of Commuter Shuttle Missions to the Moon

In 2001: A Space Odyssey, Dr. Floyd must attend an important meeting—a meeting on the Moon. He departs from a large AGS orbiting Earth and arrives there 25 h later ³¹ aboard a large spherical-shaped LTV, which touches down on a landing pad that subsequently descends into a large sprawling lunar settlement located underground. Today, 53 years after its debut, the images portrayed in Kubrick and Clarke's film remain well beyond our capabilities and 2100: A Space Odyssey seems a more appropriate title for the movie. In this section, the feasibility and requirements for commuter flights to the Moon using RL10B-2 and LANTR propulsion, along with LDPs, are evaluated to see whether the operational capabilities presented in 2001 can be achieved although on a more Spartan scale.

A 24-h commuter flight to the Moon is a daunting challenge. This is about the time it now takes to fly from Washington, DC to Melbourne, Australia with a 3-h layover in San Francisco. As Table 1 shows, decreasing the Earth-to-Moon transit time from 72 to 24 h increases the round trip ΔV requirement by ∼4.9 km/s, to ∼12.9 km/s. At these higher velocities, free return trajectories are no longer available, so multiple engines will be required to maximize shuttle reliability and ensure passenger safety.

What might a typical commuter flight to the Moon involve? It might originate from a future commercial AGS like Ad-Venture shown in Figure 13. Operating in LEO, Ad-Venture is powered by a 2.5 MW_e FPS ²⁹ and has facilities supporting a variety of activities, including zero-gravity R&D, manufacturing, developing Cislunar industries, and an emerging space tourism market. The rotation rate (ω = 2 rpm) and radius (∼37.5 m) of the eight habitation modules located at Ad-Venture's mid-section produce a 1/6thg AG level, providing Earth tourists the opportunity to experience what it would be like to live on the Moon. Similarly, long-time lunar colonists and individuals born on the Moon in the future could travel to LEO to experience the Earth's beauty “up close and personal” while being exposed to a comfortable lunar gravity environment.

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Fig. 13.

Features, characteristics, and activities at commercial AGS Ad-Venture. AGS, artificial gravity station.

Ad-Venture also functions as a transportation hub for flights to and from the Moon. Its forward transportation element has multiple docking ports to accommodate a variety of SC and PLs. A possible scenario for a commuter flight to the Moon might start with passengers boarding an Earth-to-orbit mini-shuttle for a flight to Ad-Venture (Fig. 14a). There they would enter a passenger transport module (PTM) that contains its own life support, power, instrumentation and control, and reaction control system. The PTM provides the “brains” for the commuter shuttle and is home to the 18 passengers and 2 crewmembers operating it while on route to the Moon. After undocking from Ad-Venture (Fig. 14b), the PTM rendezvouses and docks with the refueled shuttle, awaiting it a short distance away. After system checkout, the shuttle fires its engines to depart LEO and the commuter flight to the Moon begins (Fig. 14c).

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Fig. 14.

Commuter shuttle mission to the Moon—key transportation system elements. (a) Earth-to-orbit mini-shuttle for a flight to Ad-Venture. (b) Undocking from Ad-Venture. (c) Shuttle fires its engines to depart LEO and the commuter flight to the Moon begins. (d) PTM detaches and docks with a waiting Sikorsky-style LLV. (e) PTM is lowered to a “flatbed” surface vehicle and electronically engaged, providing the PTM with surface mobility. LEO, low Earth orbit; LLV, lunar landing vehicles.

After the ∼1 to 1.5-day transfer to the Moon, the shuttle captures into the lunar orbit where the PTM detaches and docks with a waiting Sikorsky-style LLV (Fig. 14d) that delivers it to the LS. The “orbit to surface” transfer time is ∼1 h. From here the PTM is lowered to a “flatbed” surface vehicle (Fig. 14e) and electronically engaged, providing the PTM with surface mobility. The PTM then drives itself to the lunar base airlock for docking and passenger unloading (shown in the lower right corner of Fig. 5). This scenario is reversed on the return trip to Earth. During the PTM transfer to the LS and back again, the Serenity Shores STN (Fig. 10) refuels the shuttle with the LDPs needed for its return to Earth.

The commercial commuter shuttle we envision utilizes the same PS, engine types, and in-line LO₂ tank assembly used on the Conestoga CCT. However, for the commuter shuttle application, the CCT's habitat module, star truss, and PL pallets are removed and replaced with a 20-person PTM (shown in Fig. 15). The fully loaded PTM has an estimated mass of ∼15 t and its outer diameter and length are ∼4.6 m by ∼8 m, respectively.

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Fig. 15.

Relative size of a CCT and commuter shuttle. CCT, crewed cargo transports.

RL10B-2 and LANTR shuttle missions to LPO and LLO were analyzed to determine the fastest transit times possible, along with the associated LEO and LDP refueling requirements, engine burn times, and IMLEO required for the mission.

The fastest one-way transit time a LANTR-propelled commuter shuttle can deliver a PTM to and from LPO is ∼33.5 h, which includes the additional 2.5 h required for the LPO insertion and departure maneuvers. The shuttle's PS and in-line LO₂ tank are refueled to their maximum capacities at the LEO STN, and its engine's run O₂-rich out to the Moon. At the LPO STN, it refuels (at an 8:1 ratio) with ∼109 t of LLO₂ and ∼13.63 t of excess LLH₂ (available after water electrolysis) to supplement the ∼15.3 t of ELH₂ remaining in the shuttle's PS for Earth return. The associated IMLEO, mission ΔV, and total engine burn time are ∼207.5 t, ∼12.326 km/s, and ∼34.1 min, respectively.

The RL10B-2 commuter shuttle operates with a fixed engine thrust level (24.75 klb_f), O/H MR (5.88:1), and I_sp (465.5 s) out and back so its one-way transit times are slightly longer at ∼36.6 h. At the start of the mission, the shuttle's PS and in-line LO₂ tank contain ∼19.4 t of LH₂ and ∼82.8 t of LO₂ (∼49% and ∼74.5%, respectively, of their maximum capacities), and during the outbound mission leg, the shuttle uses ∼80.6 t of LO₂ and ∼13.7 t of LH₂. At the LPO STN, the shuttle again refuels with ∼109 t of LLO₂ and ∼13.63 t of excess LLH₂ to supplement the ∼4.9 t of ELH₂ remaining in the PS for Earth return. The IMLEO, mission ΔV, and engine burn time are ∼141.4 t, ∼11.688 km/s, and ∼51.1 min (∼50% longer than the LANTR shuttle). The significantly lower IMLEO is again attributed to the lower dry mass of the RL10B-2's PS (∼52.5% that LANTR) and the lower initial propellant loading required in LEO.

For commuter shuttle flights to and from LLO, the LANTR shuttle uses only ELH₂ and refuels with only LUNOX. By fully loading the shuttle's LH₂ and LO₂ propellant tanks to their maximum capacities (∼39.8 and ∼111.2 t, respectively) before TLI, refueling with ∼80.3 t of LUNOX, and operating the LANTR engines O₂-rich (O/H MR = 5, I_sp ∼516 s) out and back, the LANTR shuttle can achieve one-way transit times of ∼33 h. The IMLEO, mission ΔV, and engine burn time are ∼204 t, ∼10.419 km/s, and ∼25.3 min, respectively.

For the same commuter shuttle mission to LLO using RL10B-2 engines, shorter one-way transit times of ∼31 h are possible. Before TLI, the shuttle's LH₂ and LO₂ tanks are filled with ∼32.2 and 111.2 t of propellant. At the LLO STN, the shuttle refuels with ∼76.5 t of LUNOX, which it then burns with the ∼13 t of ELH₂ remaining in the shuttle's PS on its way back to Earth. The corresponding IMLEO, mission ΔV, and total burn time are ∼181.9 t, ∼10.966 km/s, and ∼50 min—twice that of the LANTR shuttle.

A variant of the commuter shuttle mission that focuses on delivering high priority cargo to the Moon is shown in Figure 16. The priority cargo container (PCC) envisioned has a gross mass of ∼7.5 t, the same outer mold line as the PTM, and carries ∼5 t of cargo within a pressurized volume. The cargo shuttle can deliver the 7.5 t PCC to and from LLO, with one-way transit times of ∼27 h by using RL10B-2 engines, LUNOX refueling, and only ELH₂ for the round trip mission. The IMLEO for this priority cargo mission is ∼173.6 t, and the associated LUNOX refueling requirement, mission ΔV, and engine burn time are ∼73.2 t, ∼12.050 km/s, and ∼49 min, respectively.

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Fig. 16.

Cargo shuttle departing LEO for the Moon. LEO, low Earth orbit.

Estimated Total LDP Mission Needs, Mining Area, and Processing Requirements

In the previous two sections, Conestoga: A Reusable, Space-Based CCT and Feasibility of Commuter Shuttle Missions to the Moon sections, CCT and commuter shuttle missions to LPO and LLO were examined and compared. For LPO missions we assumed that LPI deposits are mined and processed to produce LH₂O that is then transported to an LPO STN where it is electrolyzed to produce the LLO₂ and LLH₂ needed to refuel the CCT and commuter shuttle flights. Two types of surface-based LLVs are used in this study. The first is a single-stage lunar descent and ascent vehicle (LDAV) that can carry a crew of four and deliver 5 t of surface PL stored in two 2.5-t PL pallets mounted on each side of the LDAV's crew cab. The second type is a Sikorsky-style LLV that has side-mounted engines and propellant tanks, and carries its PL attached to the underside of the LLV structure. Operated either autonomously or semi-autonomously, the Sikorsky-style LLVs are used for transporting tanks of LDP or LH₂O to orbiting STNs, as well as PTMs and PCCs from orbit to settlements on the LS, and back again. The dry masses, propellant loadings, and PLs carried by the LDAV and Sikorsky-style LLVs are discussed in Borowski et al. ²⁸

Because of water's composition (8:1 O/H mass ratio), ∼1.125 t of LH₂O must be produced and electrolyzed for every ton of LLO₂ required for LTV refueling. Additional water must also be produced to supply the LDP that the tanker LLVs need to deliver water to the lunar STN. In this article, the LLVs use throttled LO₂/LH₂ chemical rockets operating at an O/H MR of 5.5:1 and I_sp of 450 s. As a result, it will be necessary to overproduce on water (∼9 t of LH₂O for every ton of LH₂ needed by the LLVs) unless additional ELH₂ is supplied to the STN for their use.

To determine the range of LDP needed at both the orbiting STN and surface ice mining and processing facility, it is necessary to look at the different mission types, their transit times, and frequency of occurrence. The needs of the various LLVs supporting each mission type must also be taken into account. To illustrate the point, we examine the 72-h LANTR CCT mission discussed in Conestoga: A Reusable, Space-Based CCT section and assume the CCT delivers 20 t of cargo to LPO six times a year. Supporting six CCT flights per year between LEO and LPO will require annual LDP and LH₂O production rates of ∼1,509 and ∼1,951 t/year, respectively. Approximately 367 t of LLO₂ and LLH₂ propellant is required by the LANTR CCT, ∼543 t by 24 LDAV flights transporting 120 t of cargo from LPO to the LS, and ∼599 t by 15 tanker LLVs delivering 25 t of LH₂O per flight to the STN. For an electrolysis rate of ∼1 t/day, the electrical power at the STN needed just for electrolysis is estimated to be ∼0.205 MW_e with the electrolysis power (in MW_e) equal to ∼0.2042 × (H₂O electrolysis rate, t/day). The power level and electrolysis rate supporting LDAV and LLV tanker operations at the polar propellant facility is ∼0.886 MW_e and ∼4.34 t/day.

In a recent commercial lunar propellant study ³² assuming in situ thermal mining and an annual LO₂/LH₂ propellant production rate of ∼1,640 t/year, the estimated electrical power required was ∼2.2 MW_e (∼1.4 MW_e for electrolysis plus ∼800 kW_t used for the thermal mining process). Assuming a 4% ice content in the regolith and a 25 kg/m² yield, ³² the estimated mining area was ∼98,400 m². For the CCT mission scenario outlined earlier, the corresponding area to be processed via thermal mining and the total electrical power demand at the propellant plant are estimated to be ∼78,040 m² and ∼1.4 MW_e.

For the same CCT mission scenario to LLO, ²⁸ but now using RL10B-2 instead of LANTR engines, an annual LUNOX production rate of ∼1,234 t/year is needed to supply the six CCT flights plus the LLV flights supporting it. The CCTs use only ELH₂ and refuel with only LUNOX before returning to Earth. Approximately 329 t of LUNOX is used by the CCTs, ∼459 t by the four LDAVs transporting the CCT cargo from the STN to the LS, and ∼446 t by three LUNOX tanker LLVs, each flying approximately four to five resupply missions to the LLO STN over the course of a year. Although the CCT is supplied at the LEO STN with all the LH₂ it needs for the mission, additional shipments of LH₂ to the LLO STN will be required to supply the LLV flights. The ∼165 t of ELH₂ required by the LLVs can be supplied by using two NTR tanker vehicles, ²⁸ each carrying ∼28 t of LH₂ and flying three flights to the LLO STN each year.

Synergy of LUNOX Production and He-3 Mining

In 1986, Wittenberg et al. ³³ estimated that a million metric tons of SWI He-3 is embedded in the near-surface lunar regolith. It is divided roughly equally between the mare and the highlands with the highest concentrations of He-3 found in mare regoliths that are rich in titanium-oxide (TiO₂), which is contained in the mineral ilmenite. The University of Wisconsin's Fusion Technology Institute has designed an automated lunar miner (shown in Fig. 17) that is capable of producing ∼33 kg of He-3 per year while operating during the lunar days to take advantage of beamed solar power (∼200 kW_e) used for its process heat and operation. ³⁴

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Fig. 17.

Automated Mark II Lunar Miner for extracting He-3 and SWI volatiles. ³⁴ He, helium; SWI, solar wind implanted.

During the He-3 extraction process, each lunar miner also recovers significant quantities of other important volatiles (shown in Table 2). Especially noteworthy are the large quantities of H₂ and H₂O produced as “by-products” for each kg of He-3 collected. As a result, the 1,440 t of LH₂ required to support weekly commuter shuttle flights to LLO can be supplied by eight Mark II miners producing ∼264 kg of He-3 annually.

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Table 2.

Gaseous Volatiles Released During Heating of Lunar Ilmenite to ∼700°C

Isotope, Molecule, or Compound	t of Volatile Released per kg of He-3
H2	6.1
H2O	3.3
He-4	3.1
CO	1.9
CO2	1.7
CH4	1.6
N2	0.5
Total volatiles = 18.2

He, helium.

Source: Kulcinski et al. ³⁴

Mare Tranquillitatis is an attractive potential site for He-3 mining. With its titanium-rich regolith and large surface area estimated at ∼190,000 km², this region could contain ∼7,100 t of He-3³⁵ along with ∼43 million tons of SWI H₂. To the northwest is Mare Serenitatis, another potential He-3 mining location and also a candidate site for LUNOX production using iron-rich volcanic glass. The Taurus-Littrow DMD, consisting largely of black crystalline beads, covers ∼3,000 km² and is thought to be tens of meters thick. Assuming an area of ∼2,000 km² (equivalent to a square ∼28 miles on each side), a mining depth of ∼5 m, a soil density for the volcanic glass of ∼1.8 g/cm³, and a mining mass ratio of 25 to 1 (equivalent to a 4% O₂ yield), Figure 18 shows that the Taurus-Littrow DMD could provide ∼720 million tons of LUNOX. The production of ∼12,000 t of LUNOX annually to support a weekly commuter shuttle service requires a glass throughput of ∼300,000 t/year. Assuming twelve 1,000 t/year LUNOX production plants, each using two scaled-up and autonomously operated RASSOR-type excavator/loaders and four haulers, a soil mining rate of ∼4 t/hour per excavator/loader is required at each plant. This rate assumes a 35% mining duty cycle that corresponds to mining operations during 70% of the lunar daylight hours (∼3,067 h/year). ²¹ The total electrical power required per production plant is ∼1.5 MW_e. ²⁸

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Fig. 18.

LUNOX mining areas and production rates.

Figure 18 also shows that the mining areas needed to support commuter flights to the Moon are not unrealistic at ∼0.033 and ∼0.167 km² for 1–5 flights/week, respectively. Even at five times the higher rate of ∼60,000 t/year, there are sufficient LUNOX resources at this one site to support 25 round trip commuter flights carrying 450 passengers each week for the next ∼2,400 years and more sites containing even larger quantities of iron-rich pyroclastic glass have been identified. ²³ For 25 flights per week, ∼36,000 t/year of LLH₂ would be needed to fuel the tanker and transport LLVs. This amount of LLH₂ is consistent with an He-3 production rate of ∼5.9 t/year.

Summary and Conclusions

The commercialization and human settlement of LEO, the Moon, and Cislunar space will be greatly aided by the development and utilization of LDPs, FPSs, STNs, and reusable propulsion systems with long operating lifetimes—10s of hours not 10s of minutes. Propellant derived from LPI is currently receiving a lot of attention. However, other source materials for LDPs should not be overlooked. Vast deposits of volcanic glass on the lunar nearside can supply well in excess of 25 billion tons of LUNOX, and in the longer term, ∼5 billion tons of SWI volatiles can be extracted for propellant and life support use, from processed TiO₂-rich mare regolith during He-3 mining.

Combining LDP with chemical and LANTR propulsion can allow a robust LTS with unique mission capabilities. However, LANTR engines are heavy, require radiation shielding, and the mission operations must deal with engine cooldown and management of the associated cooldown thrust that can last for hours to days. For many of the missions examined in this article, those using the RL10B-2 engine show performance comparable to or better than missions using LANTR.

Scalable, megawatt-class FPSs are another key technology needed on the Moon. Although nearly continuous solar power may be available at a few select sites at the lunar poles, only FPS can satisfy the requirements for abundant “24/7” electrical power, at low mass, needed for the continued growth of commercial activities in LEO, lunar orbit, and on the LS.

Strategically positioned STNs will also be important. Besides providing a propellant depot and cargo transfer function, orbiting STNs offer convenient staging locations where propellant, cargo, and passengers can be dropped off and/or picked up.

Having just celebrated the 50th anniversary year of the Apollo 11 mission in 2019, it is comforting to know that work is underway on many of these technologies. With industry interested in developing Cislunar commerce and competitive forces at work, the timeline to develop and implement the capabilities discussed here could well be accelerated so that future Dr. Floyds may have the opportunity to experience “for real”—a routine flight to the Moon.