Logistics and Operations versus Heavy Lift: Examining ...spacecraft.ssl.umd.edu/publications/2011/AIAA11-7221...the moon into low lunar orbit produces a maximum lunar crew module size

Logistics and Operations versus Heavy Lift:

Examining Approaches to Human Exploration in a

Cost-Constrained Era

David L. Akin,�

University Of Maryland, College Park, MD, 20742, USA

Although recent review panels have called into question the economic viability of ad-vanced heavy-lift vehicles, the conventional wisdom still demands some form of shuttle-derived heavy-lift launch vehicle prior to initiating human exploration beyond low Earthorbit. Recent publications by the author have demonstrated that existing evolved expend-able launch vehicles, speci�cally the current version of the Delta IV Heavy, along with asmaller human spacecraft and in-orbit modular propulsion stages, are capable of supportinga robust and extensible program of human lunar exploration, starting from single-vehiclelunar orbital missions to �ve-launch scenarios for lunar landing and return. This systemprovides routine lunar surface access for both humans and cargo, based on a architectureutilizing a low lunar orbit logistics site for stockpiling propulsion stages and supporting theassembly of lunar landing vehicles via autonomous rendezvous and docking. These priorpublications have also documented probabilistic risk analyses which demonstrate that themodular approach is capable of equal or higher reliability than a monolithic heavy-lift mis-sion, due to redundancy in propulsive options from active spare propulsion vehicles basedin low lunar orbit.

This paper continues and extends the analysis of a cost-constrained modular approachto exploration by examining the potential of such a system to provide access to otherFlexible Path sites, including human missions to near-Earth objects and Mars orbit. Insome cases, in-space technology additions such as inatable habitats for longer-durationhuman missions will su�ce to support these extended-range objectives. For the di�cultgoal of human Mars missions, the analysis will examine the feasibility of human missionssupported solely by current EELV launch vehicles, and will perform trade studies againstmissions with smaller numbers of launches by investigating the impact of larger modularpropulsion stages and larger vehicle currently under private development. Even with thesefar more ambitious mission objectives, the analyses documented in this paper still supportsthe basic concept of \spend the money ying", rather than postponing human explorationto await the more elegant solution of heavy-lift launch vehicles.

Acronyms

CPM Cryogenic Propulsion ModuleDIVH Delta IV HeavyEELV Evolved Expendable Launch VehicleEVA ExtraVehicular ActivityFH Falcon HeavyGEO Geostationary OrbitLEO Low Earth OrbitLLM Lunar Landing ModuleLLO Low Lunar OrbitLMO Low Martian OrbitLOI Lunar Orbit Insertion

�Director, Space Systems Laboratory. Associate Professor, Department of Aerospace Engineering. Senior Member, AIAA

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American Institute of Aeronautics and Astronautics

AIAA SPACE 2011 Conference & Exposition27 - 29 September 2011, Long Beach, California

AIAA 2011-7221

Copyright © 2011 by University of Maryland . Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

MMH Monomethyl HydrazineN2O4 Nitrogen TetroxideNASA National Aeronautics and Space AdministrationOPM Orbit Propulsion ModuleSSL Space Systems LaboratoryTEI Trans-Earth InsertionTLI Trans-Lunar InsertionUMd University of Maryland

I. Introduction

There would be little argument to the assertion that, as of the time of writing this paper, the U.S. humanspace ight program is in a state of limbo. With the 135th and �nal ight of the space shuttle program,it is unclear whether future U.S. human access to orbit will depend on NASA-developed or commercialspacecraft, own on existing or future launch vehicles. There is a directive that, over and above continuingthe International Space Station program, human space ight will be focused on visiting a near-Earth asteroid,although many in and out of NASA treat this more as a temporary aberration than an meaningful long-termchange in destination. Congressional support of NASA seems to be focused on its value as a high-tech jobsprogram, with the Space Launch System (or, perhaps more accurately, \Senate Launch System") mandatinga vehicle which will consume most of the available operations funds for more than a decade, and resultingin a vehicle which may well be impossible to y economically.1

In counterpoint to the \conventional wisdom" emphasizing the \Apollo on steroids" approach to humanspace exploration, the University of Maryland Space Systems Laboratory has been independently pursuingpossible alternate approaches. The core axiom of this e�ort can be simply stated: spend the money ying.Rather than pursue architectures which will require a decade or more of full NASA funding to create a heavy-lift launch vehicle, this research e�ort is focused on using existing vehicles and technologies to facilitate earlyand continuous human space ights in support of an ambitious program of space exploration.

This study is not limited to avoiding new launch vehicle development, but indeed to systems which canreduce the size and cost of architecture elements, even if capabilities are proportionately reduced from theambitious visions for the Constellation program. Is support of this e�ort, the following assumptions havebeen adopted a priori :

� Only existing (or currently funded) launch vehicles will be considered.

� New systems will be minimized in number, size, and complexity to minimize nonrecurring costs anddevelopment time.

� Crew and spacecraft sizes will minimized to save mass and cost.

� New technologies will be restricted to only those which are enabling, and which are demonstrated toproduce clear and immediate improvement in system e�ectiveness.

� Cost mitigation will focus on learning curve e�ects from economy of scale in larger production runs,rather than reusability.

� Cost analysis will include opportunity costs of money, which favors higher operational costs if so doingminimizes initial nonrecurring costs.

These assumptions were used to de�ne a series of speci�c design choices which have been examinedthrough trade studies and propagated throughout the study. These speci�c architecture decisions include

� 3 person nominal crew size

� Propellants limited (at least at �rst) to standard storable hypergolic systems

� No propellant transfer or depots

� Multi-launch vehicle on-orbit integration limited to conventional docking technologies

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In the �rst phase of this study, analysis2 showed that small spacecraft supporting 2-3 person crewswere suitable for Delta IV Heavy (DIVH) missions, up to and including single-launch human lunar orbitmissions. This required the development of a crew launch and entry vehicle below 6500 kg (including theaccommodation of a launch escape system for human launch), and a standard orbital propulsion module(OPM) for lunar orbit insertion (LOI) and trans-Earth insertion (TEI). Trade studies showed that on-orbitlogistics and operations for lunar missions was best performed in lunar orbit, to take maximum advantageof the high speci�c impulse of cryogenic upper stages of the launch vehicles. This resulted in a 9980 kg masslimit for payloads delivered to trans-lunar insertion (TLI). Using standard vehicle performance parametersbased on regression analysis of past vehicles, it was shown that the favored method for payload transportto LLO is to create a standard orbital propulsion module which matches the maximum payload delivered.Rather than design a smaller module to provide the 837 m/sec lunar orbit insertion �V, the architectureadopted used two OPMs per launch: one fully loaded with propellant to be delivered to LLO, and one witha partial propellant load su�cient to deliver the fully loaded one into lunar orbit. This stage of the analysisidenti�ed a program plan which consisted of four three-crew ISS rotation missions per year, two human lunarmissions per year each with a pre-emplaced cargo module of 1870 kg at the landing site, and a Flexible Pathmission in alternate years. By using existing launch vehicles and limiting new vehicle development to thepropulsion module, a moderate OPM variant incorporating landing gear and avionics for terminal landingon the moon, and a small (4700 kg) human spacecraft, this scenario allowed reaching operational missionstatus in less than a decade and accommodated steady-state ight operations at this scale for less than $3Bper year total.

One of the obvious concerns about a system with multiple launches and in-orbit docking operations is theoverall reliability of a mission with so many critical elements. In the second phase of this ongoing research,analysis focused on probabilistic risk analysis and mitigating strategies. Since most of the Earth launchescarry standard interchangeable OPMs, it was shown3 that the ability to store additional mission elements atthe LLO staging location produced the ability to increase mission reliability to levels equalling or exceedingmore monolithic architectures. This phase also examined the use of human-rated Atlas 402 launch vehiclesfor the ISS crew rotation missions, as the more expensive DIVH has substantially more payload performancethan required for a simple LEO mission.

The current (third) phase of this research project is focused on two parallel e�orts: increasing the �delityof the vehicle design parameters through more sophisticated models and sensitivity analyses, and lookingin detail at evolutionary missions including human missions to near-Earth objects (NEOs) and to Phobosand low Martian orbit (LMO). Assumptions made in the earlier phases as to vehicle design parameters aretested against more detailed models, and the revised baseline architecture components will be analyzed forapplication to lunar, NEO, and Mars missions.

II. Revisiting Vehicle Design Estimates

Before going further into analysis on more ambitious missions, it would make sense to revisit past as-sumptions to verify that best practices are being followed, and that data to date �ts known trends wherepossible. To this end, several of the more speculative assumptions from past publications will be analyzedhere for correspondence to known data, and used to reassess past studies based on more accurate (andinvolved) predictive models.

A. Mass-Dependent Inert Mass Fraction

As summarized above. past analyses demonstrated the bene�ts of adopting a single propulsion vehicle design,which is sized to be delivered fully fueled to LLO by a second propulsion module launched partially fueled.Based on a regression analysis of similar vehicles, a stage inert mass fraction

� =minert

minert +mprop= 0:10 (1)

was assumed.4 This value was further assumed to be invariant based on total propulsion module size. For the9980 kg TLI injected payload of the Delta IV Heavy, the optimum worked out to be a fully fueled propulsionmodule mass of 6950 kg, with 695 kg of inert mass and 6255 kg of propellants. Since one of these propulsionmodules is used for crew module ascent from the moon, the ascent �V of 2334 m/sec from the surface of

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the moon into low lunar orbit produces a maximum lunar crew module size of 4966 kg. A survey of pasthuman spacecraft in reference2 showed that this was entirely feasible, falling slightly behind Apollo in entryvehicle mass but considerably larger than Gemini.

However, there has long been a item of \conventional wisdom" that there is a physical economy of scalein larger vehicle systems. From a survey of published sources of mass estimating relations, the followingequations for stage inert mass fractions were derived by regression from past vehicle data:

�storables = 1:6062 (Mstagehkgi)�0:275 (2)

�LOX=LH2 = 0:987 (Mstagehkgi)�0:183

(3)

The trend of these equations as a function of stage gross mass is seen in Figure 1. Smaller stages are ata clear disadvantage due to higher stage inert mass fractions. It is also clear that the prior assumption of�=0.10 was particularly optimistic, given the focus on small stage sizes in this architecture.

Figure 1. Trend of stage inert mass fraction with stage gross mass for high-density (e.g., storable) and low-density (LOX/LH2) propellants

B. Lunar Orbit Delivery Architecture

In the �rst phase of this study, the conclusion was reached that delivering loaded OPMs to lunar orbit wasbest done by launching two OPM systems at a time: one fully loaded with propellant to perform requiredmissions to and from the moon, and one partially fueled to perform the lunar orbit insertion (LOI) maneuver.Given the conversion to mass-dependent inert mass fraction estimation, it seems reasonable to revisit thisidea as compared to the creation and use of a ideally-sized LOI stage.

Figure 2 shows the e�ect of launch vehicle size on the relative sizes of payload delivered. Although thereis a penalty associated with the use of a partially-loaded OPM stage for LOI, the relative size of the masspenalty is small. Figure 3 shows a much more critical variation in these approaches. With two separatevehicles to be designed and developed, the nonrecurring costs for the optimized LOI stage approach is almosttwice as large as for the use of a single module design for both LOI and OPM functions. Based on theseresults, the original decision for a single OPM design stands.

It should be noted that Figures 2 and 3 illustrate the rationale behind the basic assumption of this studythat focus should be placed on existing upper-end launch vehicles, rather than waiting to develop super-heavy launch vehicles. As the size of the launch vehicle grows, the resultant size of the in-orbit spacecraftgrows, along with the cost. To minimize \up-front" costs in a program, keeping all of the vehicles (andtheir non-recurring costs) as small as practical is key to cost containment. As will be discussed later, theexception to this paradigm is the size of the human spacecraft, which tends to be more invariant than thepropulsion stages.

Prior analysis also assumed the availability of storable propellant propulsion systems (N2O4/MMH) ata speci�c impulse of 320 seconds; this is feasible, but may be optimistic for this type of application. Atrade study was performed to investigate the sensitivity of OPM module size achievable as a function ofpropulsion system exhaust velocity. Figure 4 shows that the vehicle size is not stronly sensitive to speci�c

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Figure 2. Mass trends of OPMs with and without ded-icated LOI stages

Figure 3. Total cost for OPM architecture with andwithout dedicated OPM stages

impulse. Nonetheless, the decision was made to reduce the assumed speci�c impulse to 310 seconds to beconservative.

Figure 4. E�ect of speci�c impulse on OPM performance

III. Mission Applications

As a result of the trade studies, it was clear that the assumptions of the previous phases of this analysiswere essentially correct, but optimistic. The logical next step was clearly to revisit the baseline design forhuman lunar exploration in light of these revised baseline parameters.

A. Lunar Exploration Program

The basic concept of operations for a logistics-based human lunar mission is shown in Figure 5. A seriesof Delta IV Heavy launches deliver three Orbital Propulsion Modules, one Lunar Landing Module, and ahuman spacecraft. The assembly is docked together in LLO and used for a lunar landing. The landedcon�guration is shown in Figure 6. A key sizing factor was the desire to use a standard fully-fueled OPM tolaunch the crew module back into low lunar orbit, where it would rendezvous and re-dock to the same OPMwhich brought it into LLO, which has su�cient propellant to perform the TEI burn for return to Earth anda direct entry and landing.

The size of the human spacecraft is bounded by two constraints: keeping within the lunar launch payloadcapability of an OPM, and keeping it spacecraft mass low enough that it can be launched with su�cientpropellant in its OPM to perform both a lunar orbit insertion and a trans-Earth insertion. This constraintwas selected for crew safety: while a number of \spare" OPMs will be based in LLO, this assumption ensures

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Figure 5. Original baseline CONOPS for human lunar landing2

Figure 6. Original vehicle con�guration post-lunar landing2

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that the crew has the capability to abort to Earth at any point in the mission (up until lunar landing) withoutthe requirement for precursor orbital operations.

1. Revised DIVH Mission Baseline

Table 1 compares the critical design parameters for the orbital propulsion modules of the prior study, andafter the revised baseline assumptions documented above. As can be seen, the revised assumptions result inan OPM with reduced gross mass, but increased inert mass. This should allow su�cient margin for detaileddesign without extraordinary e�orts to maintain minimum mass. Table 2 shows the same data for the lunarlanding module. In both cases, the LLM was essentially an OPM with an additional 50% inert mass marginto account for landing gear, descent avionics, and other systems required for the actual lunar landing phase.

Table 1. Comparison of previous and revised orbital propulsion module design parameters

Parameter Old Baseline Current Baseline

Gross Mass (kg) 6950 5973

Stage Inert Mass Fraction � 0.10 0.160

Propellant Mass (kg) 6255 5105

Inert Mass (kg) 695 868

Speci�c Impulse (sec) 320 310

Table 2. Comparison of previous and revised lunar landing module design parameters

Parameter Old Baseline Current Baseline

Gross Mass (kg) 6950 5973

Stage Inert Mass Fraction � 0.15 0.24

Propellant Mass (kg) 5908 4543

Inert Mass (kg) 1042 1430

Speci�c Impulse (sec) 320 310

At this point, it became possible to derive the constraints on the human spacecraft. In the previous study,a spacecraft mass of 4966 kg was selected to allow a single OPM launch from the moon back into lunar orbit,which also met the constraint for single-OPM insertion and departure from LLO. The current baselineanalysis produced a human spacecraft mass limit of 4886 kg for the LLO insertion-departure requirement,but the new OPM design will only deliver 3548 kg to LLO with a single module. The original report2

discussed required spacecraft mass in some detail, with reference to past human spacecraft; while 3548 kg iscomparable to a Gemini spacecraft, it was felt at that stage that a spacecraft mass of approximately 5000kg would provide for a three-person crew and direct entry return to Earth from a lunar trajectory.

To meet this shortfall, the spacecraft mass was assumed to be 4886 kg from the LLO constraint. Tosuccessfully launch this spacecraft from the lunar surface back into LLO, two OPMs will be landed with itfor lunar launch; one fully fueled, and one loaded with 1670 kg of propellant. While it seems wasteful touse an OPM with only 28.3% of its propellant loaded, the only other choices would be to either design aspecialized smaller version of the OPM, or to develop a larger lunar-launch stage and use a larger launchvehicle to deliver it.

The revised baseline lunar landing system can now be de�ned. Where the original baseline requiredone human spacecraft, one LLM, and four OPMs for six DIVH launches per human lunar landing, the newbaseline requires six OPMs (one partially loaded), a LLM, and a human spacecraft for eight DIVH launchestotal. Assuming the current DIVH costs of approximately $300M per launch, the total launch costs for thisapproach would be $2.4B per mission, which is still competitive with expected launch costs of the SpaceLaunch System.

In an active, robust program of lunar exploration, there will be a signi�cant need for routine cargo deliveryto the moon. Whether for landing exploration robots or pre-emplacing logistics and equipment for future

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human landings, the components developed for human missions can also provide one-way cargo deliveries tothe lunar surface at a variety of payload sizes, as detailed in Table 3.

Table 3. One-way cargo capabilities to the lunar surface

Vehicle Con�guration Cargo Landed (kg)

LLM only 1732

LLM + 1 OPM 5025

LLM + 2 OPMs 8292

LLM + 3 OPMs 11,553

LLM + 4 OPMs 14,811

2. Mixed Fleet Options

Since the prior publications of this study, SpaceX has announced the development of Falcon Heavy, a three-core heavy-lift version of the Falcon 9. Recognizing that all SpaceX vehicles are at an early stage of maturitycompared to EELVs, Falcon Heavy is nonetheless a heavy-lift vehicle which is both funded and in activedevelopment. For that reason, it will be used to examine the e�ects of a heavy-lift launch vehicle on thislunar exploration architecture.

Falcon Heavy is projected to carry 53,000 kg to LLO and 16,000 kg to TLI. Using the same paradigmwith two identical OPMs, using one partially fueled to deliver the fully-fueled OPM into LLO, producesa Falcon Heavy-based orbital propulsion module (OPM-F) design with a gross module mass of 9739 kg,propellant mass of 8483 kg, and inert mass of 1256 kg (�=0.144). This module, used alone, will be capableof carrying 6082 kg from the lunar surface to LLO, which represents the baseline human spacecraft and 1196kg of optional up-cargo. This scenario reduces the total launches to seven, replacing two DIVH launcheswith a single Falcon Heavy. Since OPMs come in integer quantities, four fully-fueled OPMs and the LLMprovide su�cient performance to land the spacecraft, OPM-F for lunar launch, and 186 kg of down-cargo.

Two other cases in the mixed-eet model were also examined. In case 3, all OPMs were replaced byOPM-F stages, with only the DIVH version of the LLM still used. For the last case, an LLM-F was designedbased (as in the case of the DIVH LLM) on a 50% increase in inert mass for the OPM-F. The LLM-F massparameters are therefore a gross mass of 9739 kg, a propellant mass of 7855 kg, and an inert mass of 1884kg.

Table 4 summarizes the four cases considered with the revised baseline architecture. It should be notedthat every OPM used for lunar landing or launch requires a second OPM for the lunar orbit insertionmaneuver, which is then expended. Similarly, LLM and human spacecraft each require an OPM for LOI.The table also lists the available up-cargo and down-cargo masses, which represent the payload mass marginsfor landing and launch maneuvers based on fully-fueled OPMs, as well as the total requirement for DIVHand FH launch vehicles.

It should be noted that, even for mission architectures which are otherwise exclusively launched on FalconHeavy vehicles, it is assumed that the human spacecraft still ies on a Delta IV Heavy with an OPM asmaneuvering stage. This is justi�ed on the basis that DIVH is currently ying, and human-rating the DeltaIV family is already under consideration. While SpaceX intends to human-rate all of its launch vehicles, thelarger size of FH is ill-matched to the relatively small size of the baseline human spacecraft.

Further analysis is important to examine the various lunar options on the basis of cost and reliability.Even the most capable of the options considered (all FH-sized modules) still uses six ights as comparedto �ve for the old baseline; this is indicative of the reduced performance (in terms of both � and Isp) ofarchitecture components in the more recent analysis. Future e�orts will seek to �nd optimum program-level plans which accommodate upgrades in capability through phased development of new architecturecomponents throughout an ongoing ight program.

B. Selecting Extended Exploration Parameters

With the revisions to the parametric description of the in-space logistics architecture, the next issue to exam-ine is the applicability of these components to extended space exploration. Two next-generation evolutionary

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Table 4. Summary of mission logistics for mixed-eet human lunar surface missions

All DIVH OPM-F for All OPM-Fs All OPM-Fs

modules launch only and DIVH LLM and LLM-F

Mission Con�guration Spacecraft Spacecraft Spacecraft Spacecraft

2 OPM (ascent) 1 OPM-F (ascent) 1 OPM-F (ascent) 1 OPM-F (ascent)

1 LLM 1 LLM 1 LLM 1 LLM-F

4 OPM (descent) 4 OPM (descent) 3 OPM-F (descent) 3 OPM-F (descent)

OPMs 14 10 2 1

LLM 1 1 1 0

OPM-Fs 0 2 8 9

LLM-F 0 0 0 1

Up-cargo (kg) 0 1196 1196 1196

Down-cargo (kg) 1414 186 3381 0

DIVH launches 8 6 2 1

FH launches 0 1 4 5

mission goals will be considered: a six-month trip to a near-Earth object (NEO), and a human mission toPhobos, the inner moon of Mars.

Prior to the detailed analysis of deep-space missions, it is important to consider the choice of a construc-tion and check-out site for the mission preparation. This could be anywhere in cis-lunar space, excludingonly Earth orbits above approximately 1000 km altitude up to 65,000 km altitude, due to the heightenedradiation environment due to the Van Allen radiation belts. Other speci�c locations include the �ve librationpoints of the Earth-Moon system.

The original trade study of optimum staging location showed that the most advantageous staging sitefor lunar exploration is low lunar orbit. Given the signi�cantly higher performance of the LOX/LH2 upperstage of the Delta IV Heavy, it was highly advantageous in that analysis to get the staging site as farout of Earth’s gravity well as practical via direct injection of payloads upon launch. Either of the stableEarth-Moon librations points, L4 or L5, require nearly equal performance for orbit insertion as the selected100 km low lunar orbit used for lunar exploration staging. For this reason, it was decided to consider theEarth-Moon L4 libration point as one possible option for space-based logistics caching and vehicle assembly.The obvious alternative, as in the original study of logistics site for the lunar program, is low Earth orbit.Both of these potential sites will be considered in this study.

C. Near Earth Object Missions

Earlier this year ENAE 484, the University of Maryland senior capstone class in spacecraft design, performeda detailed study of a human mission to four NEO candidates between 2020 and 2030. The critical elementsof the mission design chosen were a round-trip travel time of six months or less, and a total mission �V notto exceed 7 km/sec from the LEO staging site assumed for that study.

Two of the four candidate NEO targets from the ENAE 484 study were adopted for this analysis:2007XB23 and 2001QJ142. These two targets were the greatest and least �Vs from a LEO departure forthe four considered. The calculated mission �Vs were recalculated for Earth departure from L4 rather thanLEO, which brought all of the L4-based mission �Vs into the range of 4.0-4.5 km/sec. The �V requirementsby mission increments are detailed in Table 5.

The same basic human spacecraft from the lunar analysis (mass=4886 kg) was kept as the launch andentry spacecraft. However, the much greater mission duration of a NEO mission will require substantialadditional human support infrastructure. A six-month, three-crew mission requires 540 crew-days of con-sumables, along with an expanded crew volume for long-term habitability. Typical rules of thumb for crewconsumables range from 10 kg/day for an open-loop system, down to 1 km/sec for an aggressive degree ofrecycling of water and air loops. A detailed design study for an extended-duration lunar habitat concluded

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Table 5. �Vs for candidate near-Earth object missions (all values in m/sec)

2007XB23 2001QJ142

LEO departure 4160 3460

L4 departure 3834 1687

NEO arrival 180 1260

NEO departure 310 2040

Total �V (LEO) 4650 6760

Total �V (L4) 4324 4987

that the installed life support system mass would be about 600 kg, with a consumables usage rate of 5crew-day, for a total life support mass of 3300 kg for three crew for 180 days.5 The crew would also needexpanded habitable volume for a longer-duration mission. Recent SSL experience in the development andtesting of inatable habitats resulted in the demonstration of a space-quali�able inatable habitat with aninated volume of 60 m3 and a total mass of 350 kg.6

With these parameters in mind, it was assumed that the human payload for the NEO missions wouldconsist of the 4886 kg launch and entry vehicle, along with a 6083 kg supplemental human payload. Thiswould consist of long-term life support systems and power generation, crew consumables, and an inatablehabitat module for additional living volume. The 6083 kg limit is based on the maximum payload deliverableby the DIVH-based logistics infrastructure, where the payload is unmanned and does not require an onboardreturn capability for crew safety. This brings the mission payload to 10,859 kg.

Results of the mission analysis are summarized in Table 6.

Table 6. Summary of mission logistics for NEO missions from L4

2007XB23 2007XB23 2001QJ142 2001QJ142

DIVH Modules FH Modules DIVH Modules FH Modules

Mission Con�guration Spacecraft Spacecraft Spacecraft Spacecraft

Hab Module Hab Module Hab Module Hab Module

9 OPMs 5 OPM-Fs 11 OPMs 7 OPM-Fs

OPMs 20 2 24 2

OPM-Fs 0 10 0 14

Payload Margin (kg) 1208 228 0 637


FH launches 0 5 0 7

Clearly, these architectures require a large number of assembled modules for a mission, with a corre-sponding number of heavy-lift launch vehicle ights. Unlike the lunar landing case, where a module failureduring descent will result in an immediate (and redundant) abort back to LLO and safety, a module failureduring a NEO mission could strand the crew a long way from home.

To address both the cost and safety concerns, the NEO mission seemed to be a good point to considerthe incorporation of advanced technology, speci�cally a long-duration cryogenic propulsion module for theFalcon Heavy launch vehicle (CPM-F). Given LOX/LH2 technology for a deep-space stage, it is also notunreasonable to assume the development of a cryogenic upper stage for the Falcon Heavy (currently underearly development as the \Raptor" stage). A brief calculation based on comparison to other launch vehicleswith cryogenic upper stages produced an estimate of 24,000 kg delivered to C3=0 (lunar orbit/L4).

Rather than continue the approach of ying two identical modules and using one partially fueled todeliver the other, it would make sense to examine the pre-existing designs (OPM and OPM-F) as deliverystages to insert the CPM-F into the L4 logistics site. Analysis shows that the OPM is too small, but thestorable OPM-F with a 68% propellant load will deliver a CPM-F with a gross mass of 16,964 kg, propellant

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mass of 15,093 kg, and inert mass of 1872 kg.

Table 7. Summary of mission logistics for cryogenic FH NEO missions from L4

2007XB23 2001QJ142

Mission Con�guration Spacecraft Spacecraft

Hab Module Hab Module

2 CPM-Fs 3 CPM-Fs

OPMs 2 2

OPM-Fs 2 3

CPM-Fs 2 3

Payload Margin (kg) 3175 5893

DIVH launches 2 2

FH launches 2 3

Although the cryogenic propulsion module based on a Falcon Heavy launch brings the mission scenariodown to 4-5 launches, the large number of launches required for the storable systems intuitively would seemto suggest that the propulsion modules are too small for the larger �Vs of the NEO mission. In order to getthe largest propulsion modules possible, it makes sense to compare the L4 logistics site to a more traditionalLEO site.

For completeness, Table 8 list the nominal mission characteristics of the three stages designed and ana-lyzed so far.

Table 8. Orbital Propulsion Stage parameters for L4 staging

OPM OPM-F CPM-F

Gross mass (kg) 5973 9739 16,964

Propellant mass (kg) 5105 8483 14,150

Inert mass (kg) 868 1256 2814

Speci�c Impulse (sec) 310 310 460

Stage inert mass fraction � 0.1453 0.1290 0.1659

For LEO staging, components such as OPMs can be designed to take up the entire LEO payload massof the launch vehicle: 23,000 kg for the DIVH, and 53,000 kg for the Falcon Heavy. For completeness, bothlaunch vehicles will be considered with propulsion stages using storable and cryogenic propellants. (DIVHCPMs were not considered for L4 staging, as it was assumed the very high number of OPMs required wouldnot be signi�cantly reduced by the higher performance of the small DIVH CPM stage allowed after transportto the L4 staging site.) The design characteristics for all four of these new propulsion module designs arelisted in Table 9. Note that the pre�x \L" is added for a LEO-speci�c propulsion module, so an \LOPM-F" isa Falcon Heavy-sized orbital propulsion module with storable propellants designed for use following deliveryto low Earth orbit.

With all these parameters determined, the various components of the architecture can be examined todetermine the necessary arrangement of propulsion modules to accomplish each NEO mission with each ofthe four candidate approaches to LEO systems logistics. This is summarized in Tables 10 and 11.

One major di�erence between the architecture tables for the lunar surface missions as compared to theNEO missions is the absence of additional modules for transport to a remote staging site. All of the payloadof the launch vehicles goes directly to a LEO staging site, eliminating expended stages prior to the mission.Also, both the crew spacecraft and the additional habitat module �t easily within the LEO payload capabilityof a single DIVH launch.

In summary, it is obvious that the LEO logistics base is far superior to the L4 site for NEO missionsand, by extension, for all non-lunar missions. By staging from LEO, even storable DIVH architectures canreach NEO targets for 4-6 launches, and cryogenic Falcon Heavy systems will allow similar missions with 1

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Table 9. Orbital Propulsion Stage parameters for LEO staging

LOPM LCPM LOPM-F LCPM-F

Gross mass (kg) 23,000 23,000 53,000 53,000

Propellant mass (kg) 20,631 19,391 48,692 45881

Inert mass (kg) 2369 3609 4308 7119

Speci�c Impulse (sec) 310 460 310 460

Stage inert mass fraction � 0.1030 0.1569 0.0813 0.1343

Table 10. Systems architectures for 2007XB23 missions using LEO staging

DIVH DIVH Falcon Heavy Falcon Heavy

Storables Cryos Storables Cryos

Mission Spacecraft Spacecraft Spacecraft Spacecraft

Con�guration Hab Module Hab Module Hab Module Hab module

3 LOPM 2 LCPM 2 LOPM-F 1 LCPM-F

Payload Margin (kg) 2973 6124 10,995 7437


FH launches 0 0 2 1

Table 11. Systems architectures for 2007XB23 missions using LEO staging






Payload Margin (kg) 254 904 1606 24,531


FH launches 0 0 3 2

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manned DIVH launch and 1-2 FH launches of propulsion modules, depending on the speci�c NEO target.(It should be noted that the LEO case does not assume the development of the Raptor cryogenic upper stagefor Falcon Heavy, as it is unnecessary for this architecture.)

D. Phobos Mission

To further extend the applications of this modular architecture, the same vehicle components derived forNEO missions in the preceding section will be applied to a human mission to Phobos, the inner moon of Mars.Beyond the interest in Phobos in its own right and as a moon probably gravitationally captured by Marsin the past, Phobos could provide su�cient mass for highly e�ective radiation shielding for long-durationhuman missions, using the proximity to the Mars surface to teleoperate high-performance robots with verylow latencies in the control loop.

�Vs were calculated for a Phobos mission using standard patched conic techniques, and are summarizedin Table 12. This table also presents the �V calculation results for reaching a 100 km altitude orbit aboveMars rather than Phobos orbit; this lower orbit would be ideal for staging a Mars landing program, whichis beyond the scope of the current paper.

Table 12. �V requirements for Mars missions to Phobos and 1000 km orbits

Maneuver Phobos orbit 1000 km orbit

�V (m/sec) �V (m/sec)

LEO departure 3165 3165

L4 departure 2264 2264

Mars orbit arrival 1885 2030

Mars orbit departure 1885 2030

Total �V (LEO) 6.934 7.225

Total �V (L4) 6.034 6.342

While a Phobos mission is operationally just another trip to another asteroid, the fact that it is farfrom Earth introduces its own set of challenges. Foremost among them is the additional trip time; theincrease in mission duration from 6 to (approximately) 24 months drives the life support consumables massup considerably. Keeping the 5 kg/crew-day parameter used earlier, this would add 8250 kg of consumablesto the mission payload. Longer missions optimize for higher levels of closed-loop life support, which shouldreduce this parameter by half. For convenience, it was assumed that another DIVH-delivered cargo module inL4 would be devoted to life support consumables, making the total payload mass for a minimum three-personMars mission 16,832 kg.

Initially, a complete set of L4-staging scenarios was analyzed for feasibility in a Phobos mission, althoughthe NEO mission results cast doubt upon the e�cacy of this approach given a 50% larger payload andan additional 2000 m/sec �V. Results demonstrated that an L4-based mission using Falcon Heavy-sizedpropulsion modules would require 16 OPM-F modules to complete the mission, and a DIVH-based missionwould require 27 OPMs and 30 DIVH ights. Even though the Falcon Heavy system with cryogenic propul-sion modules could complete the Phobos mission with only four modules, the much greater e�ectiveness ofthe LEO staging site demonstrated in the NEO section led to the abandonment of L4-based scenarios fordeep-space missions.

Table 13 details the analysis results for the four LEO-based approaches developed in the NEO section. Allthree of the payload modules can reach Earth orbit on a single DIVH launch. Although the DIVH approachwith storable propellants is somewhat cumbersome, six ights for a storable propellant solution with FalconHeavy launches are completely in line with baselined missions for the lunar surface program. Cryogenic FHpropulsion modules can bring a human Phobos trip down to a three-launch mission, although the practicalimplications of maintaining liquid hydrogen without boilo� for a two-year mission would require advancedtechnologies, un-front nonrecurring development funds, and additional mission risk.

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Table 13. Systems architectures for human Phobos missions using LEO staging





Logistics Module Logistics Module Logistics Module Logistics Module


Payload Margin (kg) 1358 3166 4463 0


FH launches 0 0 5 2

IV. Future Work

The \holy grail" of the human space exploration program is a human landing on the surface of Mars. Thispaper has demonstrated that Phobos is an entirely feasible objective for human space ight with minimalnew development required, but analyses of feasible logistics trains for Mars surface exploration will have towait until a future paper.

This paper also expanded the paradigm of this ongoing research project from absolute adherence to aphilosophy of \no new development" (or \spend the money ying") to show that some destinations willrequire advanced technologies to keep the overall architecture manageable. Future work should examine theoptimum choices for additional technology developments. Is cryogenic propellant maintenance more e�ectivethan aerobraking? When (if ever) do larger launch vehicles make sense, and does the answer change if launchvehicle development costs have to be borne by the human exploration program?

In the same way that this paper re-examined simple assumptions from earlier work and resulted in moreconservative (and hopefully accurate) estimation practices, there is a real need to perform a detailed designstudy of critical elements of the proposed architecture, such as human spacecraft and baseline propulsionmodules, to ensure that the parametric estimation used here results in feasible design estimates. This isespecially true in the case of Mars landing architectures, where published estimates of mission infrastructuremay vary in mass by an order of magnitude or more.

Lastly, while this paper has focused on parametric design and performance calculations, it is in peril offalling prey to the "conventional wisdom" of fewer ights is always better. While even the author would notargue in favor of a mission scenario requiring 30 launches, it should be kept �rmly in mind that the goal hereis to optimize for the maximum opportunity to y missions within the strict funding limitations apparentlybeing applied to NASA today. Costing of these advanced scenarios should proceed to inform decisions onwhich architectures to select, as well as to perform optimal sequencing of development programs so that asteady stream of new technologies and capabilities ow into the program at regular intervals without starvingan robust ight program from the funding required for sustenance.

V. Conclusions

As before, the basic conclusion of this paper is that human space exploration does not have to waita decade or more for the development of super-heavy lift launch vehicles. Existing, ight-proven DeltaIV Heavy vehicles, if modi�ed for human-rating, are fully capable of supporting near-term human lunarsurface missions without waiting a decade for super-heavy lift launch vehicles, long-term cryogenic storage,or propellant handling at orbital depots. While the low lunar orbit staging location, or any similar high orbitin cislunar space, appears to be disadvantageous for longer-range exploration, modular spacecraft assembledin low Earth orbit are fully capable of taking humans to near-Earth objects or to explore the moons of Mars.

References

1-, \ESD Integration: Budget Availability Scenarios" NASA Internal Brie�ng, August 19, 2011

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2David L. Akin, \In-Space Operations: Developing a Path to A�ordable, Evolutionary Space Exploration" AIAA 2010-2293, SpaceOps 2010 Conference, Huntsville, Alabama, April 25-30, 2010.

3David L. Akin, \Build a Little, Fly a Lot: An A�ordable Evolutionary Approach to Flexible Path, Lunar Surface, andBeyond" AIAA 2010-8610, AIAA Space 2010 Conference and Exposition, Anaheim, California, August 30-September 2, 2010.

4David L. Akin, \Mass Estimating Relations" ENAE 483/788D Principles of Spacecraft De-sign Lecture Notes, Department of Aerospace Engineering, University of Maryland, Fall Term 2011hhttp://spacecraft.ssl.umd.edu/academics/483F11/483F11L09.MERs/483F11L09.MERs.pdfi

5David L. Akin, Massimiliano Di Capua, Adam D. Mirvis, Omar W. Medina, William Cannan, and Kevin Davis, MinimumFunctionality Lunar Habitation Element, Final Report for NASA Grant NNJ08TA89C, University of Maryland Space SystemsLaboratory, July, 2009.

6Massimiliano Di Capua, David Akin, Kevin Davis, and Justin Brannan, "Design, Development, and Testing of an Inat-able Habitat Element for NASA Lunar Analogue Studies" AIAA-2011-5044, 41st International Conference on EnvironmentalSystems, Portland, Oregon, July 17-21, 2011.

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