AIAA 93-4784


B. Kent Joosten
NASA Lyndon B. Johnson Space Center, Houston, Texas

Lisa A. Guerra
Science Applications International Corporation, Washington, DC


Attempts to reestablish the capability for humans to explore the moon have suffered from high projected costs of spaceflight hardware development and production. An avenue for reducing these costs is proposed: the utilization of propellant produced on the moon at program inception. This technique permits drastic down-sizing of the required transportation systems and should enable lunar missions using a single STS-derived launch vehicle. Preliminary space transportation and surface system design efforts have validated the technical feasibility of this concept. The required technologies emphasize electric surface vehicles, automation and robotics, and highly reliable chemical processing; areas that have considerable relevance to the U.S. commercial sector. To evaluate the effectiveness of this strategy, cost assessments were performed and compared to estimates based on a more conventional lunar exploration scheme. The results indicate that an architecture emphasizing early production and utilization of lunar propellant has lower hardware development costs, lower cost uncertainties, and a reduction in human transportation costs of approximately fifty percent.

I. Introduction

This year, the U.S. Congress reaffirmed that "...the core mission of the National Aeronautics and Space Administration is, and depends upon, the extension of human presence beyond Planet Earth ... by the acquisition and development of knowledge necessary for expanding human presence beyond low Earth orbit to other celestial bodies... ." In addition, NASA was encouraged to "...leverage its investment in existing programs in a manner that can help reduce the technical and cost uncertainties related to human exploration."1

The primary obstacle to renewed human exploration of the moon is the high design, development, test and evaluation (DDT&E) and production costs of the required transportation systems. Past efforts at controlling production costs have emphasized reusable lunar transportation system concepts,2,3,4 but this strategy can exact a heavy toll in required technology development, infrastructure investment, and operations. In addition, the amortization time associated with these investments is highly dependent upon long-term flight rates, which are difficult to estimate. Approaches to reduce DDT&E costs often center about low-technology options with a minimum number of hardware elements, but the production costs of multiple flight units may be very high.5

Previous studies have examined the role of propellant produced in situ, but these have usually been in the context of a mature lunar base, with the goal of reducing downstream transportation costs.6,7 This strategy has several disadvantages. First, program development costs will not decrease, but increase when lunar oxygen production is included. This is true because the transportation system must be designed to initially operate with all-terrestrial propellants, and therefore cannot take advantage of the lunar propellant leverage in vehicle design and development.8 Second, there is a risk that once lunar base operations are underway, the resources necessary to initiate the development of the lunar propellant production equipment would not be available, and the expected economies would be deferred or would never materialize.

The concept outlined here relies upon lunar-produced oxygen (LUNOX) for piloted missions from program inception. The performance leverage is invested in reducing the size and complexity of the transportation systems and therefore their development costs. The economies in transportation production costs will accrue starting with the first piloted flight. The goal is to offset the development and deployment costs associated with the LUNOX production equipment through reduced transportation development and production costs after a "reasonable" number of flights.

To test these assumptions, two alternative lunar exploration scenarios have been developed. They both result in comparable lunar surface capabilities and space transportation requirements, but one utilizes the early LUNOX concept, the other does not. The development and production costs of the hardware associated with each scenario were estimated using the NASA Johnson Space Center's Advanced Missions Cost Model9, along with a statistical measure of confidence in the results.

II. Technical Strategy

Many writers have described the scientific, political, economic and social rationale for renewed human exploration of the moon,10-13 and it is not the intent of the authors to reiterate them here. Rather, we have tried to outline a set of lunar exploration capabilities which we feel represents a reasonable first step towards accomplishing a wide range of goals, regardless of the specifics of the lunar surface mission activities. These capabilities include habitation, surface mobility, extra-vehicular activity (EVA), and logistics/resupply. The levels defining the Initial Operational Capability (IOC) chosen for this analysis are:

Habitation - sufficient life support capabilities to sustain a crew of four on the lunar surface for a day-night-day cycle (42 Earth days). This provides two EVA teams two periods of high productivity.

Surface Mobility - capability to support a two-person EVA team for a lunar day and transport them over distances of at least 100 kilometers. This allows extensive geological traverses as well as servicing of remote instruments and installations.

Extra-Vehicular Activity - ability to service, clean, and repair space suits and portable life support equipment as well as routine pressurized element ingress/egress.

Logistics/Resupply - ability to transport logistics from Earth and provide storage volume on the lunar surface.


Propellant availability at the mission destination has extremely high flight performance advantages; not only can the space vehicle avoid carrying the return propellant, but also the propellant required to transport the return propellant. For this reason, oxygen extracted from the lunar regolith has long been recognized as potentially one of the most valuable resources on the moon. Oxygen is plentiful, being the major constituent of the regolith (Fig. 1) and many possible extraction techniques have been identified14 (Table 1). However, design challenges exist. Since the space transfer vehicles in this concept are designed to require LUNOX from the outset, human transport is not possible until a sufficient cache of oxygen has been produced. This implies that the delivery, deployment, checkout, and operation of the LUNOX production plant and regolith transport equipment must occur in an automated or teleoperated manner. (A rather interesting technique will be discussed later which could allow piloted flights to occur prior to LUNOX production.)

It is difficult at this point to identify the "best" extraction technique for early LUNOX production. Trades involving system mass, power requirements, process complexity, oxygen yields, reagent resupply, and equipment maintenance must be made in light of the automated, program critical nature of the LUNOX production. However, it was decided to adopt the ilmenite reduction process as a reference since a design study has been done for a processing plant consistent with the production quantities required by this analysis.15 This concept uses the combination of a small teleoperated front-end loader and regolith hauler to mine and transport feedstock and dispose of process tailings. Estimates of power requirements for the oxygen extraction and liquefaction equipment (40 - 60 kWe) indicate the desirability of a nuclear power source, especially since it would allow some operations during the two-week lunar night without the massive energy storage demands associated with photovoltaic systems.

Figure 1 - Typical Regolith Components

Table 1 - Candidate Processes

Space Transportation Because oxygen/hydrogen (LOX/LH2) rocket engines exhibit high performance and because the propellant is 85% LOX by mass, this combination is usually considered the best choice for LUNOX-based transportation systems. To derive the maximum benefit of LUNOX, the strategy proposed here conducts round-trip piloted missions using a lunar landing vehicle which is designed to land with "dry" LOX tanks (other than flight performance reserves). The lander must carry enough LH2 for the round trip, but the LOX tanks are refilled on the lunar surface for the return to Earth. The resulting gain in performance alters most of the rules of lunar transportation: lunar-orbit rendezvous is no longer advantageous since the penalty of carrying the Earth-return propellant to the lunar surface is essentially eliminated, and even staging on the moon offers little advantage. This allows development of a single crew module (essentially a scaled-up Apollo Command Module), and a single service module/lander combination. Most importantly, this mission concept decreases the trans-lunar injected spacecraft mass to the point that a piloted flight can be accomplished with a single launch of a Space Transportation System (STS)-derived launch vehicle.

"STS-derived" launch vehicle concepts have been proposed for numerous applications, usually due to the anticipated low development costs.2,4,8,16 The launch vehicle considered here consists of a core stage derived from the Space Shuttle External Tank (ET), a propulsion module incorporating an aft skirt and three Space Shuttle Main Engines (SSMEs), and two Redesigned Solid Rocket Motors (RSRMs) which are the only elements recovered for reuse. Mounted atop the launch vehicle is a LOX/LH2 trans-lunar injection (TLI) stage utilizing three Pratt & Whitney RL10-A4 engines, and the lunar spacecraft. This launch vehicle configuration was recently evaluated by NASA as an alternate Space Station launcher17, although the more familiar "Shuttle-C" side-mount arrangement may be preferable, since it allows processing and launch operations with minimal impact to Kennedy Space Center facilities.

"Man-rating" either launch vehicle design option should be straightforward since they consist primarily of STS components. In addition, robust "abort-to-orbit" capability would exist in the event of an SSME failure since the equivalent of 6 km/sec propulsive delta-V is available in the TLI stage and spacecraft. If such a failure occurred, the launch vehicle would simply continue thrusting on the remaining engines until its propellant was depleted, after which the TLI stage would separate and complete the injection into Earth orbit. The spacecraft would then wait for proper phasing with the desired emergency landing site, perform a deorbit maneuver, and land. An Apollo or Soyuz style tractor-rocket could be employed for crew escape during early launch phases, if deemed necessary. The spaceflight elements required by the LUNOX architecture are illustrated in Figure 3.

Figure 3 - "LUNOX" Flight Elements

Figure 4 - "LUNOX" Flight Profile

Mission Profile The "direct mode" lunar transportation profile, along with the abort options afforded by the LUNOX concept are described next. A normal launch would be followed by a suborbital TLI stage ignition at a point consistent with acceptable performance and core stage disposal. After a short Earth-orbit coast to allow for proper phasing, the stage is re-ignited to complete the trans-lunar injection. Since the loiter duration in Earth orbit is on the order of a few hours, no debris shielding should be required. TLI stage disposal can be accomplished by targeting for lunar impact or for ejection from the Earth-moon system, thereby eliminating the risk of Earth reentry or debris generation. Abort options (either reinsertion into Earth orbit followed by deorbit and landing, or circumlunar return) exist during all phases of the TLI burn through utilization of the lander's propellant.

After a four-day coast to the moon, the spacecraft performs the lunar orbit insertion burn, waits in orbit for proper landing site phasing, and proceeds with the deorbit maneuver. At any time up to this point, the crew can return to Earth without conducting the lunar landing. (In fact, sufficient on-board propellant exist to permit an Apollo 8 style lunar orbital mission, should this prove desirable during early program phases.) Soon after initiating the final powered descent, the vehicle reaches an abort "mode-boundary", where continuation to the lunar surface is mandatory. Formulation of abort strategies and lander design must take this into consideration.

Following landing, the vehicle is powered-down for the duration of the surface mission, and the crew transfers to the surface habitation facilities. Prior to return to Earth, the crew refills the lander's tanks with ten tons of LUNOX. The lander's engines are restarted, performing an ascent into a temporary lunar phasing orbit, followed by trans-Earth injection. Since no staging is performed on the lunar surface, there will be no accumulation of useless hardware associated with the piloted flights. The lander body is jettisoned a few hours before Earth entry, after being targeted for safe disposal. The crew module reenters, using parachutes and retrorockets for recovery on land. Studies have indicated significant operational cost savings can be achieved by avoiding the need for mid-ocean recovery forces, and current GN&C systems should easily provide acceptable landing accuracy.18 Figure 4 illustrates the mission profile.

This "direct mode" offers several programmatic and operational advantages. It requires only one crew module; historically a very expensive development item. It requires no rendezvous, proximity operations, or docking avionics or hardware. And since no assets remain in lunar orbit, the Earth return "window" is continuously open, allowing freedom in selecting the lunar outpost location while preserving the option to return to Earth at any time.

Cargo The LUNOX production equipment, habitation elements, and exploration systems must be transported by dedicated cargo flights, as the piloted lander can provide only a few tons of cargo and remain within the lift capability of an STS-derived launcher. If the same launch vehicle and TLI stage are used for both piloted and cargo missions, a delivery capability of approximately 12 metric tons to the lunar surface results. The cargo lander departs somewhat in design from the piloted vehicle since there is no attempt to refuel it and because it must be designed for easy, automated payload unloading. However, there should be significant system and subsystem commonality between the two landers. The mass breakdown for the piloted and cargo flights is shown in Table 2.

As Table 2 suggests, it is possible to perform piloted flights with these vehicles prior to the establishment of LUNOX production. Since a payload capability in excess of 12 tons exists, a cargo flight could precede the human mission, delivering only a tank of LOX. After lunar landing, the crew could refill their lander with this terrestrial LOX in the same manner as they would with LUNOX. While this "lunar surface rendezvous" is probably not the optimum way to conduct human missions, it does provide program options and illustrates the high performance leverage of oxygen produced on the moon (one launch versus two).

Piloted Flight
Crew Module 6743 kg
Dry 5935
Fluids 199
Crew and Support 609
Lander 26941 kg
Dry 5505
Cargo 2000
Outbound Prop. 16944
Return H2 2492
TLI Stage 50060 kg
Dry 6130
Propellant 43930
TOTAL 83744 kg
LUNOX Required 10165 kg
Cargo Flight
Payload 12454 kg
Lander 21295 kg
Dry 4717
Propellant 16578
TLI Stage 50060 kg
Dry 6130
Propellant 43930
TOTAL 83809 kg
Table 2 - "LUNOX" Vehicle Mass Breakdown

Surface Systems A typical lunar outpost buildup scenario begins with the delivery of a habitat module which may weigh from 17 to 30 tons, depending upon the level of outfitting19,20. Since this exceeds the cargo capability of the LUNOX transportation system, a strategy utilizing a combination of pressurized rovers and small, specialized modules to establish surface habitation has been developed.

Previous rover conceptual designs21,22 have been based upon requirements to support a crew of two for a lunar day, and transport them hundreds of kilometers. These requirements have been adopted for the present work. Since the vehicles must provide power, communications, thermal control, life support, and habitation volume as part of their prime mission, it is reasonable to take advantage of these features when the rovers are parked at the outpost. Typically, rover design concepts utilize hydrogen/oxygen fuel cells for mobile power, which are either regenerated or resupplied with reactants between excursions. If the rovers are employed for habitation during these times, external power must be supplied, and a more than adequate power supply will already be available on the lunar surface: the one associated with the LUNOX production. Therefore, if production is curtailed during the human surface mission, the associated power can be made available to the rovers through a power distribution system.

While the rovers can provide or augment many of the surface capabilities needed, they probably cannot supply the kind of robust EVA support that lunar missions will require. To maintain reasonable size and mass constraints, pressurized rover designs depend upon small "man-locks" or depressurization of the habitable volume for EVA transfer. This is acceptable during limited, remote EVAs, but more routine operations will require provision for transfer of materiel into and out of pressurized volumes, stowage and maintenance of space suits and portable life support equipment, dust control, and hyperbaric treatment. Therefore, a combination EVA support and airlock module will be required at the outpost location. This module will also allow pressurized connections to the rovers, logistics containers, and future expansions of the habitable volume.

In addition to the requirements for LUNOX production, surface mobility and habitation, the ability to transport and store logistics will be needed. Items such as crew consumables, system resupplies, and maintenance articles may require a pressurized environment, and a simple module which can be transported and docked to one of the pressurized volumes should suffice. These logistics modules can also be used to store trash and waste and transport them away from the outpost site. Finally, a mobile power source has been shown to be essential for several functions on the lunar surface21. This small vehicle, basically a mobile fuel cell/solar array combination, is used to supply "stay-alive" power to the piloted lander, and to augment the rovers' onboard power.

All of the components identified above have masses consistent with the delivery capabilities of the LUNOX launch vehicle and cargo lander. The technique used to integrate them on the lunar surface is outlined next.

Outpost Buildup Table 3 lists the hardware and associated masses required to achieve IOC. The initial cargo flight delivers the LUNOX production plant and the nuclear power supply. The plant utilizes the lander frame for structural support and the empty propellant tanks for LUNOX storage. The reactor, mounted on a small teleoperated cart, is deployed from the lander and is translated away from the plant, trailing a power cable with attachments for multiple connections. The reactor deploys its thermal radiators and begins power production. The second cargo flight delivers the teleoperated surface vehicles required for regolith and liquid oxygen transport, and LUNOX production begins at this time. One of the regolith loaders can also be used to establish a radiation shielding berm around the reactor. Technically, human round-trip lunar transportation is enabled as soon as a sufficient supply of LUNOX has been accumulated. However, surface habitation facilities have not yet been established.

Cargo Flight 112379 kg
LUNOX Plant7269
Nuclear Reactor5110
Cargo Flight 28322 kg
Tanker (2)2942
Loader (2)3456
Hauler (2)1924
Cargo Flight 37694 kg
Pressurized Rover5150
Mobile Power Unit1544
Science Payload1000
Cargo Flight 47694 kg
Pressurized Rover5150
Mobile Power Unit1544
Science Payload1000
Cargo Flight 511010 kg
Cargo Flight 612454 kg
Table 3 - Cargo Manifest for "LUNOX" IOC

The habitation infrastructure emplacement begins with the third mission and the delivery of the first pressurized rover. Once unloaded onto the lunar surface, it is remotely driven to the location desired for the outpost. This process is repeated with the fourth flight. The airlock/EVA support module, transported on Flight 5 is mounted on a wheeled chassis and is also remotely driven and attached to the power supply to serve as the "hub" of the outpost assembly. The components can now be configured in the arrangement shown in Figure 5, or if this is deemed too complex an operation to be accomplished through teleoperation, it can be performed by the first lunar crew. The final cargo flight prior to IOC delivers consumables, spares, and any equipment required for additional outfitting. The facilities are now in place to support six-week, four-person missions.

Figure 5 - "LUNOX" Surface Systems

From this description of the LUNOX strategy, it is apparent that its feasibility relies heavily upon the development of key technologies: high duty-cycle electric surface vehicles for regolith, LUNOX, and cargo transport, automation and robotics for mining and surface system integration, and extraterrestrial chemical processing for unattended lunar oxygen production. This is somewhat of a departure from the technology expertise of the aerospace industry where the focus tends to be in the area of vehicle and propulsion systems. However, this can be viewed as an opportunity to broaden the relevancy of government-sponsored space technology to the commercial sector. In addition, the programmatic risk associated with the oxygen extraction techniques can be mitigated by proof-of-concept precursor missions involving small lunar landers equipped with process demonstration packages.

"First Lunar Outpost"

The "First Lunar Outpost" (FLO) was an effort initiated by NASA's Office of Exploration aimed at understanding the technical, programmatic, schedule, and budgetary implications of restoring U.S. lunar exploration capability. It was intended to be a benchmark against which competing strategies could be measured, which is how it is treated in this analysis. The approach taken in establishing the lunar surface capabilities was much different than that taken in the LUNOX architecture. Integration of elements and complex operations on the lunar surface were minimized, and the number of items to be developed was also reduced. The systems required had a high reliance on inheritance from past and current programs in the anticipation of lowering hardware development costs. This strategy resulted in a few large, pre-integrated systems, a distinct contrast from the LUNOX approach.

Figure 6 - "FLO" Lunar Habitat

Cargo Flight 130785 kg
Hab Module30785
Cargo Flight 224764 kg
Pressurized Rover (2)10310
Table 4 - Cargo Manifest for "FLO" Outpost Buildup

Figure 7 - "FLO" Launch Vehicle

Surface Systems The lunar surface habitat was designed to be delivered to the moon with a single launch and to automatically activate all needed systems upon landing. It was derived from a Space Station Freedom Program habitat module design and featured an airlock, a photovoltaic/regenerative fuel cell power system, thermal control system, logistics storage volume, and a spacesuit maintenance facility integrated into a single unit. To eliminate the need for massive unloading equipment and to avoid the associated surface operations, it was decided to leave the hab integrated with the lunar lander, as shown in Figure 6.

To accommodate the requirement for long-range surface roving, pressurized rovers with the same characteristics as those used in the LUNOX scenario were proposed. While the vehicles were not integrated into the habitat structure, they did depend upon the habitat power system for recharge. Table 4 gives a mass breakdown of the two required cargo flights

Figure 8 - "FLO" Piloted Vehicle

Space Transportation The pre-integrated nature of the lunar hab, combined with the desire to avoid on-orbit operations and multiple launches resulted in the requirement for a large new launch vehicle design. A concept derived from the Apollo-Saturn vehicle, shown in Figure 7, used combinations of re-engineered Rocketdyne F-1A engines on the boosters and first stage, and J-2Ss on the second and TLI stages. A single lander design was utilized for both cargo and piloted flights. When used in the piloted mode, the payload was replaced by an Earth-return stage and the crew module (Fig. 8). Like the LUNOX concept, the "direct mode" was used to avoid developing two piloted vehicles, but because all terrestrial propellants were used, a much larger vehicle design resulted. Table 5 shows the mass breakdown for these vehicles.

Piloted Flight
Crew Module7259 kg
Crew and Support609
Return Stage24124 kg
Lander61654 kg
TLI Stage146057 kg
TOTAL239094 kg
Cargo Flight
Payload35894 kg
Lander57143 kg
TLI Stage146057 kg
TOTAL239094 kg
Table 5 - "FLO" Vehicle Mass Breakdown

III. Cost Analysis

In examining alternative approaches to human exploration to the moon, cost becomes a convenient metric. In this analysis of comparing the LUNOX and FLO approaches, only the prime contract cost was evaluated. The prime contract cost includes the DDT&E phase as well as the hardware production costs through the initial operational capability. The costs not considered for this study include spares, flight and ground software, launch operations, mission control operations, communications, recovery systems, program support, and program management. Therefore, the LUNOX and FLO cost assessments emphasize the hardware investment differences incurred early in a lunar program. Given inevitable constraints on exploration expenditure, this analysis poses the question: should resources be spent on usable surface infrastructure or expendable launch transportation systems? In particular, the investment issue focuses on technology development (automation, chemical processing techniques, surface mobility) versus the resurrection of a large Apollo-era booster.

Groundrules and Assumptions

Certain groundrules and assumptions were used to prepare the LUNOX and FLO cost assessments. Economic assumptions included: 1) all cost estimates are in constant fiscal year 1993 dollars (FY93$), and 2) a contractor's fee is added at 10%. The following mission assumptions also applied to the study:


The LUNOX and FLO cost assessments used the Advanced Missions Cost Model (AMCM) to estimate all flight hardware prime costs (except for costs noted in the groundrules and assumptions). The AMCM was developed at the NASA Johnson Space Center as an alternative to parametric, weight-based models. The AMCM particularly tries to identify cultural differences as a cost driver (i.e., planetary spacecraft cost more than submarines due to the NASA culture they are developed in). The AMCM performs cost estimates at the system level, such as the crew module; thus, providing the analyst with a means of assessing design/mission variations from a top-down approach.

The parametric inputs required by the AMCM for each system include the dry mass, the number of development articles, the number of production units, the system block number, the year of initial operational capability, the specification value (mission classification), and the difficulty factor (measure of technical and programmatic difficulty). For the difficulty factor, rather than the mission designer selecting a value assessed relative to other similar historical systems, a separate survey was issued to capture the individual manager's technique for project and engineering management. The difficulty factor survey appears as Table 6.

Programmatic Difficulty
1) Type of requirements (level of specs)Detailed build specsSome build-to specsProduct performance
2) Management style/systemsCustomer imposed/duplicateCustomer imposed/streamlinedSupplier's systems only
3) Number of changesThousands/yearModerateVery few; tight control
4) Budget strategyIncremental funding/yearlyIncremental funding/5 yearsFull commitment w/guarantees
5) Schedule stabilityAnnual evaluationFirm within 5 yearsFull commitment for duration
6) Competition through Start of detailed designEnd of preliminary design (C/D start)Commitment to fixed price by supplier
7) Contract typeCost plus fixed, award feeFixed or award Fixed price and incentives
8) Complexity of buyer/seller interactionHighly interactiveInteractive at discipline levelInteractive through intermediary
9) Complexity of external/internal interfacesNumerous; multinational projectModerate; multiple locationsSmall; single location
Technical Difficulty
1) Technology status at full scale developmentProof of conceptUsually provenTotally proven technology
2) Length of full scale development6-15 years3-6 years2-3 years
3) Design complexityNew design at advanced SOTANew design at current SOTAMods to an existing design
4) Fabrication complexityMachined plate; high precision Machined plate; heavy millingMachine castings; low precision
5) Integration complexityDiff. interface precision alignNew, but familiar interfacesRoutine interfaces
6) Design marginsLarge (>50%)Moderate (10-50%)Small (<10%)
7) Type of tooling technologyMin. automation; hand-craftedModerate automationHighly automated; SOTA
8) SR&QACustomer specifiedCustomer imposedIndustry and supplier standard
Table 6. - Difficulty Factor Survey

The column on the left of Table 6 designates the categories of interest. The three following columns express "ways of doing business", from business as usual (approximately like the Space Shuttle Orbiter/calibrated at 0) to aggressive business practices (like the X-24 and Project Mercury/calibrated at -1.5). The LUNOX and FLO design leads treated the survey like a shopping list, selecting one response for each line under programmatic and technical difficulty. The values for each selection were averaged to arrive at a programmatic difficulty factor and a technical difficulty factor.

In addition to providing a baseline cost estimate for LUNOX and FLO, this study included a range for both scenario costs based on mission design uncertainty. The cost uncertainty analysis involved applying a risk model to the AMCM.25 Four uncertain input variables were modeled with probability distributions - specification value, difficulty factor, dry mass, and development quantity. A normal distribution function was applied to the specification value and difficulty factor, adhering to their derived, modeled form within the AMCM. A triangular distribution function was applied to the system mass and development quantity. A range of -10% and +25% was applied to the throughput cost values. A Latin Hypercube simulation was used to arrive at the probability distribution for the total LUNOX and FLO costs.


Table 7 - Prime Cost Estimate Summary
(FY93$ billions through IOC)
Launch Vehicle4.529%12.650%
Space Transportation6.935%7.329%
Habitation Systems0.00%2.19%
LUNOX Systems3.417%0.00%
Surface Systems4.021%1.98%
Science Payloads0.84%1.14%

The baseline cost estimates for LUNOX and FLO appear in Table 7. This summary separates the cost results in terms of system categories - launch vehicle, space transportation, habitation systems (for FLO), surface systems, oxygen production systems (for LUNOX), and science payloads. The prime cost for FLO is approximately $25 billion, $6 billion dollars more than the LUNOX scenario for a comparable initial operational capability in the year 2005. Note that the FLO approach applies almost 80% of the total prime cost on transportation, while the LUNOX approach maintains a 60%-40% split between transportation costs and lunar infrastructure costs.

The disparity is obvious in the two launch vehicle estimates, with the FLO launch vehicle consuming 50% of the resources as compared to 24% for LUNOX. Table 8 highlights the specific launch vehicle element costs to emphasize the difference between an STS-derived, 80 metric ton-to-LEO vehicle and a new, 240 metric ton vehicle. The values displayed in this table include the development costs and the flight production costs for three FLO launches (two cargo and one crew) and seven LUNOX launches (six cargo and one crew).

Table 8 - Prime Cost Estimate for Launch Vehicle Systems
(FY93$ billions through IOC)
Main Engines0.91.2
Second Stage0.02.3
TLI Stage1.21.5
Shroud & Integration0.30.7

Figure 9 - Prime Costs through IOC

Figure 9 further highlights the distribution of costs between LUNOX and FLO through the initial operational capability. The development costs for all the transportation systems and the surface systems are distinguished from the production/flight costs. Again, the LUNOX scenario has the larger share for surface system (cargo) development and production; FLO maintains the larger share for transportation.

Figures 10 and 11 display the costs associated with piloted transportation and cargo transportation to the moon, respectively. In both approaches, the piloted transportation graph assumes the development costs for the launch vehicle. The first piloted mission to the moon in FLO costs approximately $4 billion, while the same crew of four to the moon in LUNOX costs approximately $1.5 billion. Since the same launch vehicle and lunar descent stage are required for FLO piloted missions as FLO cargo missions, there are no development costs attributed to FLO in Figure 11. The LUNOX development cost refers to the cargo lander. The first cargo mission to the moon in FLO costs approximately $3.2 billion (delivering 32 mt), while the first cargo mission in LUNOX costs almost $1 billion (delivering 12 mt).

Figure 11- Prime Costs for Cargo Mission

Figure 10- Prime Costs for Piloted Mission

Figure 12 projects the prime cost for follow-on piloted missions to the moon for both approaches. The cost of an additional piloted flight includes the launch vehicle, TLI stage, and lunar lander. The FLO piloted lander has the capacity to transport the crew's logistics for their 42 Earth-day stay. In contrast, the LUNOX approach requires an additional cargo landing for every two piloted missions in order to deliver the essential logistics. In Figure 12, flight numbers 3 and 5 for LUNOX include the cost for two additional cargo flights. As previously noted, the cost delta after the first piloted flight is approximately $6 billion. After six piloted flights the cost delta between FLO and LUNOX is approximately $18.5 billion. Additional oxygen production infrastructure, beyond IOC capability, was not included in this estimate for LUNOX. Further analysis regarding the compatibility between the oxygen production rate and the LUNOX mission flight rate is required.

Figure 12 - Post-IOC Costs for Continued Piloted Missions

Cost Uncertainty Analysis

One thousand iterations of a Latin Hypercube simulation were run to arrive at the LUNOX and FLO cost probability distributions. Table 9 displays the minimum, 10%, mean, 90%, and maximum values for both the LUNOX and FLO prime costs at the initial operational capability. The range of potential estimates for the FLO program spans a cost delta of $48 billion, while the LUNOX delta between the minimum and maximum costs is $31 billion. Such a large spread in potential LUNOX and FLO costs clearly reflects the immaturity (and uncertainty) in the mission and systems designs. However, FLO's larger cost spread is compounded by the uncertainties in the heavy lift launch vehicle, the highest cost system in the approach.

Table 9 - Prime Cost Probability Distribution
(Values in FY93$ billions)
10 Percent18.623.5
90 Percent30.341.6

Figure 13 highlights the uncertainty for each system classification by measuring the standard deviations for LUNOX and FLO. The contribution of the launch vehicle uncertainty for FLO to the total program uncertainty is evident. Because of the lack of definition of the LUNOX systems required to produce lunar oxygen, they have a higher standard deviation than the habitat systems of FLO. However, because their cost magnitude is smaller than that of the FLO launch vehicle, their impact on the uncertainty of the total LUNOX cost is not as significant.

Figure 13 - Measurement of Uncertainty

Probabilistic uncertainty analysis allows program managers to discuss confidence levels for cost estimates and ultimately the amount of reserve to carry for such a mission. The cost S-curve gives the probability of a project's cost not exceeding a given cost estimate. Figure 14, the cost S-curve or cumulative probability distribution, clearly demonstrates the confidence levels for: 1) the estimate for FLO as modeled by AMCM at about 17% and for LUNOX at about 20%, 2) the simulation means, and 3) the 90% confidence values. The estimated costs do not correspond to the mean values because the cost uncertainties for the individual systems were usually skewed toward overruns, not underruns. The steepness of the S-curve signifies how much the level of confidence improves when a small amount of reserves are added. From Figure 14, it is evident that the LUNOX curve is steeper than the FLO curve, thus emphasizing the decreased cost risk in the LUNOX mission approach. Furthermore, the 90 percentile confidence value for LUNOX is less than the mean value for FLO.


This study presents an exploration scenario keyed to the availability of lunar-produced oxygen and teleoperated surface vehicles prior to the return of humans to the moon. The cost metric used effectively demonstrates the difference in early investment between such a surface technology development strategy versus a more conventional approach. The results indicate that the lunar propellant approach has high leveraging potential with respect to prime costs due to accumulated savings in launch vehicle development and space transportation costs. In comparison to FLO, the LUNOX approach saves approximately $6 billion (from the modeled estimate) and demonstrates less uncertainty based on a tighter probability distribution.

To truly validate the argument, a life cycle cost analysis is required. The launch operations costs must be compared, i.e., fewer launches of a large vehicle versus multiple launches of a smaller vehicle. The early technology development/demonstration costs for the LUNOX systems also need to be assessed, as well as the long-term oxygen production capability and associated flight rate. Once a life cycle cost is generated, a cost uncertainty analysis of the "total" program can be generated. The resulting distributions would provide additional insight into the maturity of the concepts under consideration and allow discussion of reserve amounts.

Finally, it is recognized that the LUNOX scenario represents a more complex approach to human lunar exploration than some competitive concepts. A risk assessment must be performed to identify the segments with the highest contribution to overall mission risk, and investigate methods to minimize the impacts.

Figure 14 - LUNOX and FLO Prime Cost Cumulative Distribution


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Copyright 1993 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17, U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for Governmental purposes. All other rights are reserved by the copyright owner. Copyright 1993 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17, U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for Governmental purposes. All other rights are reserved by the copyright owner. 17 Copyright 1993 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States under Title 17, U.S. Code. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for Governmental purposes. All other rights are reserved by the copyright owner.