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A lunar exploration architecture using lunar libration point one
Aerospace Science and Technology 12 (2008) 231–240 www.elsevier.com/locate/aescte
A lunar exploration architecture using lunar libration point one Kian Yazdi a,∗ , Ernst Messerschmid b,1 a EADS Astrium Ltd, Gunnels Wood Road, Stevenage, SG1 2AS, United Kingdom b Institut fuer Raumfahrtsysteme, Universitaet Stuttgart, Pfaffenwaldring 31, D-70550 Stuttgart, Germany
Received 22 June 2006; received in revised form 10 April 2007; accepted 1 June 2007 Available online 7 June 2007
Abstract Future space exploration activities very likely will incorporate a large number of human missions to various locations on the Moon surface and other destinations near the Earth–Moon system. A wide range of space infrastructure elements will be required in low Earth orbit and waypoints to these destinations. It is reasonable to start assembly and mission logistics from today’s frequently used near-Earth orbits, such as the orbit of the International Space Station (ISS). This allows for using launch systems and other infrastructure elements currently available in order to reduce costs. Besides new transfer, ascent and re-entry vehicles as well as planetary surface installations, space stations beyond low Earth orbit, namely in the Earth–Moon system, can be crucial elements for forthcoming exploration missions. The lunar Lagrange points have unique advantages in this context and offer promising options for the humans’ next steps into space. As shown in this paper, a lunar space station (LSS) is not too farfetched; and it is definitely not science fiction but rather the next logical step for solar system exploration. Like a stepping-stone, it could enable the immediate start of sustained development of outer space with technologies available today. The proposed lunar exploration architecture targets at an LSS at the lunar Lagrange point 1 to be used as a gateway for mid-term lunar surface exploration missions or servicing missions to other destinations. For study purposes the challenging constraint of only using existing and tailored infrastructure elements and technology of European/Russian heritage is assumed. The results manifest the feasibility of such a space station between 2015 and 2020. Presented are station configuration and modules, transfer vehicles for assembly, crew and cargo transport, an enhanced life support system and a logistics concept. The concept outlines enhancements of the current transportation and station infrastructure and shows that the ISS as a transportation node can beneficially support lunar scenarios. © 2007 Elsevier Masson SAS. All rights reserved. Keywords: Lagrange point; Lunar exploration; Mission architecture; Space stations; Human spaceflight Schlüsselwörter: Astronautik; Lagrangepunkt; Lunare Exploration; Missionsarchitektur; Mondraumstation
1. Introduction For more than three decades, no human space exploration effort beyond Low Earth Orbit (LEO) has existed since Apollo. Now, after the dawn of the 21st century, it seems that NASA and other space agencies have arrived at the conclusion that the time has come to leave Earth orbit again for new human spaceflight challenges. Recently, the US government has decided on NASA’s long-term strategic space programme leading back to * Corresponding author. Tel.: +44(0)143877-3749; fax: +44(0)143877-8910.
E-mail addresses: [email protected], [email protected] (K. Yazdi). 1 Member IAA. 1270-9638/$ – see front matter © 2007 Elsevier Masson SAS. All rights reserved. doi:10.1016/j.ast.2007.06.001
lunar activities and more to come. The European Space Agency (ESA) formulated its Aurora programme in 2001, which also includes the preparation and execution of the long-term human exploration of the solar system. In connection with this, an international human mission to Mars is envisioned with the possibility of using lunar locations as a way station. A lunar space station for orbital support of lunar surface activities is documented here as an example for a near-term start of a sustained approach to lunar exploration and exploitation. As a point of departure, a minimum configuration station, serving as a stepping-stone and gateway to further lunar and nearlunar activities was investigated. This addresses an option for ESA’s preparatory programme Aurora [1] indicating Europe’s long-term intentions for the next 30 years, and the US Space
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Exploration Policy of NASA focusing on human exploration of the solar system and primarily targeting the Moon and Mars [2–4]. The visions foresee a human return to the Moon surface by 2020 and the first human expedition to Mars by 2030. China and other countries such as India and Japan engage in space exploration as well. Activities in the Earth–Moon system are seen as the “next logical step” for the development of space and to prepare successive human space operations beyond. The actual range of possible utilizations of a lunar space station is manifold and shall be addressed in the beginning of this paper. This is followed by a discussion of the associated challenges and mission architecture of a scenario demonstrating the application of methodology and tools to a sample conceptual design problem. Conceptual design of space stations deals with systems of relatively large scale, long life time and with complex logistical and operational requirements. The applied conceptual design process needs the availability of sophisticated design and simulation tools, coping with the involved extensive calculations combined with high-level interactivity [5–7]. If properly implemented, this enables analysis and comparison of specific characteristics for evaluating the pros and cons of system and mission design concepts with rapid turn-around times. This paper focuses on the conceptual design of a lunar space station and supporting transfer vehicles. Relevant mission and system parameters are discussed and a summary of results is presented. Details may be found in [5]. 2. Lunar exploration and development rationales The Moon, as the celestial neighbour to Earth, is more accessible than all other bodies in the solar system. Thus human spaceflight infrastructures on or close to the Moon will be far
better maintainable and usable with access of shorter period and higher frequency. Furthermore, the Moon offers a long-term perspective for economical exploitation of space (in terms of propellant and raw material production and – possibly one day – tourism). A permanent human installation (a “gateway”) in the Earth– Moon system would enable efficient transportation to other destinations than LEO based scenarios. Simultaneously, the Moon and a lunar gateway offer excellent early utilization possibilities in scientific research and technological developments for successive steps of mankind into the universe, be it exploration or exploitation. 2.1. Science rationales By extending human activities to cis-lunar, science will benefit significantly. Building up of an infrastructure fosters research activities that a purely scientific budget could not afford and would not otherwise take place [8]. Beside its good accessibility, the most compelling arguments supporting the scientific case for a return to the Moon are based on geosciences, life sciences, astronomy and solar physics [9]. As summarized in Fig. 1 the causes for this are manifold: The Moon features no considerable atmosphere, has no tectonics and rotates slowly (33 arcsec per minute). Its surface is extremely old, preserving a record of the early evolution of terrestrial planets and the near-Earth cosmic environment [10]. This record includes the time period before 4.5 to 3.8 billion years that is not present on any other terrestrial body, except Mercury [11]. Our current understanding of planetary and solar system history and evolution bases primarily on lunar surface observation and samples brought back by the Apollo missions from only six landing sites
Fig. 1. Science and programmatic rationales for lunar exploration and development.
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in the vicinity of Moon’s equator on the near-side. Therefore, any approach for a better understanding ultimately will require thorough surface analysis including sample identification, collection and analysis on a far larger scale than before [12]. This will demand a considerable amount of fieldwork for collecting more and more diverse rock and soil and subsurface samples and returning the most promising to Earth. With an established access to the lunar surface, scientists will be interested in operating other instruments there e.g. seismic profiling devices and interferometers. 2.2. Resources and geopolitical issues of the Moon The Moon is not only of scientific interest; it could also be the “most valuable real estate in the solar system” [13] because of its location, properties and resources. The main arguments are based on the utilization and exploitation of lunar resources and the use of the lunar environment as a demonstration platform for a number of technologies required for robotic and human exploration of Mars and the solar system. Here, the task will be determining what lunar resources in which quantities indeed exist, proving their accessibility and verifying whether and how they can be economically utilized and exploited. For instance, hydrogen in some form must be present, especially at the poles [13]. This includes the possibility of over 10 billion tons of water ice [9]. Both, hydrogen and oxygen (or water), are assets for future space programmes due to their application in life support systems and for in-situ propellant production. In long perspective, the Moon could possibly offer energy. Besides supply by regenerative means, the long-term energy supply for Earth will potentially reside on thermonuclear fusion reactors using helium-3 as fuel. This volatile mass fraction on lunar regolith is only about 13 mg/ton [10], thus, a large area must be processed to obtain a considerable mass. However, only one kilogram of helium-3 is necessary to produce the same amount of energy as 10 billion kg of fossil fuel [9]. These examples emphasize the geopolitical aspect the Moon will possibly have in the second half of this century, in a world of scarce fossil fuels. This alone might be the driving cause for the US government to justify their renewed space initiative by arguments including national economy and security [2]. 2.3. SEL2, solar system exploration and the Moon Besides science, the “IAA Cosmic Study” [14] names the “exploration imperative – the innate drive of human beings to extend their boundaries and explore the unknown” as the rationale for human spaceflight. It proposes a staged approach into the solar system and selects following 4 steps and destinations: 1. 2. 3. 4.
Step: Beyond LEO, i.e. Moon and SEL2 Step: Deep Space, i.e. NEA Step: On to Mars, i.e. Phobos/Deimos Step: Down to Mars, i.e. Mars surface
For each step, the study identifies one new capability: a new transportation element, named geospace exploration vehicle to
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access the Moon and Sun–Earth Libration Point 2 (SEL2) for maintenance and servicing astronomical facilities, and an interplanetary exploration vehicle for roundtrips to Near Earth Asteroids (NEA) for studying these objects, which pose a potential impact threat to Earth. It is assumed that each step bases on the results of the previous steps and uses infrastructure elements built-up programmatically (technology development) or physically (staging). 2.4. Summary: Goals of the next steps Lunar scenarios are considered as first step for sustaining activities beyond LEO. The conclusion is that the destination Moon is of high priority, be it for cultural, political or scientific reasons. And whatever emphasis the future scenario will have: global access to Moon’s surface is an important requirement with the necessity of reaching each point repeatedly. To meet the long-term objectives, transfers must be possible to the Lunar Lagrange points, Sun–Earth Lagrange points, NEA, Mars, and destinations beyond. Implied by the requirement of affordability, risk reduction and sustainability, the design of the forthcoming interplanetary human spaceflight missions suggest the following general requirements: • Easy access via available or derived transportation systems (efficient utilization of launchers and their already existing launch sites) • Safe mission operations (i.e. large departure and arriving windows, robust free-return and/or abort options) [15] • Frequent departure and arrival windows (once per revolution, daily, weekly, monthly, etc.) • Minimum delta-v for transfers, station keeping and orbit maintenance • Minimum transfer time • A human factors supporting concept; especially enhancing ergonomics, habitability and reduction of the radiation hazard (i.e. accumulated radiation dose) Furthermore, constraints arise due to superior issues typically involved in future projects: • International cooperation: access from launch sites of all potential project partners • Utilization of the (already existing) International Space Station (ISS) • Sustained approach based on currently available or upgraded technology would enable development and use of common transportation vehicles. A space station or a human-tended space platform installed in the Earth–Moon system could play a significant role in nearterm space activities and would offer particular utilization objectives summarized in Fig. 2. As a demonstration example, a fictitious mission statement is assumed here, including following objectives, constraints and requirements.
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Fig. 2. Potential utilization objectives of a Lunar Space Station at the Lunar Lagrange point One (LL1).
2.5. Sample mission statement The project’s task is conceptualizing a Lunar Space Station (LSS) as a gateway to the Moon and beyond. This station shall prepare and support successive steps of lunar surface exploration by meeting following objectives: • Staging platform with permanent access to and from Earth and the surface of the Moon • Space-research beyond LEO (e.g. radiation) • Test and verification of technologies and processes (e.g. propulsion and enhanced life support systems) The station’s development, deployment and operation shall allocate minimum costs, risk and time. First level requirements and constraints are as follows: • • • • • •
European–Russian lead cooperation Minimum configuration but extendable Affordable, low risk and near-term Use of ISS hardware heritage Permanent crew of 3, temporarily 6 Start of operation between 2015 and 2018 with 10 years lifetime
The station utilization characteristics would allow for a robust contribution of Europe together with Russian partners, with own objectives within a larger international programme including the USA, Japan, China and other partners. But the mission emphasizes the intensifying collaboration between Europe and Russia in space by e.g. implementing the Soyuz or follow-on man-rated launchers at Europe’s spaceport at Kourou, French Guyana, and their agreement on joint future space activities. Recently initiated US developments following ESAS [3] with respect to project Constellation [4] are therefore not detailed
here, although they would be synergistically complemented by an LSS. 3. Lunar exploration architecture using a space station Most past and recent studies done on lunar architectures concentrate on surface bases and only very few studies have included non-surface installations or elements such as lunar space stations or platforms. Examples are a large lunar orbiting station named Space Operation Centre (SOC) [16], the rotating Clarke Station conceptually designed in a student project, and a gateway-like station in a high eccentric Earth–Moon cycling orbit drafted in another ISU student project. However, little can be directly applied for the mission objectives identified above. 3.1. Mission architectural decisions The goal is the creation of an initial, generic and affordable waypoint station in a short time frame to be used by succeeding crewed missions. The emphasis lies primarily on lunar exploration, utilization and exploitation. Thus the preparation and support of missions to the Moon and other destinations is the task, not providing the exploration capabilities itself. The main decisions to be made driving the project cost, complexity and risks are: • • • • • •
Launch site and vehicles LEO support (in/outbound) Station location Surface accessibility Surface stay time Transport elements/degree of re-usability
The transportation infrastructure includes launch systems to LEO or direct Lunar Transfer Orbit LTO-injection and in-space transportation systems (transfer vehicles). These systems will
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define, or rather limit, the size and mass of the station elements and the size and performance of the transfer vehicles. Besides systems available and already used today (Kourou based Ariane 5-ESV/ECA, and Baikonur based Proton M, Soyuz ST/FG/U) enhanced versions of systems planned to date are considered. This includes Ariane 5-ECB and Onega, which is an enhanced version of the Soyuz planned to carry a new Russian based crew transportation vehicle, called Clipper. Not yet planned but considered here, is an enhanced version of Ariane 5-ECB, called A5-ECB+ or A5-27 with an increased LEO payload performance of about 27 tons. The modifications necessary were baselined by industrial specialists in a study performed at ESA and were assumed to be feasible and available by 2015 [17]. The ESA study indicates an increase of delta-v between 8% (for Ariane 5 ESV) and 5% for a A5-27 rocket, thus the resulting benefits using the ISS should more than outweigh the reduction of payload. 3.1.1. LEO support Because no appropriate heavy lift launch vehicle of Saturn or Energia class exists today, the station elements and transfer vehicles cannot be launched directly to the destination location similar to the Apollo missions. This means transfer vehicles, i.e. injection stages, and some kinds of LEO operations are necessary for mating the elements for the individual missions. For the chosen example, two alternatives exist: One option is using a freely chosen parking orbit, where the elements automatically mate and are checked-out remotely before lunar transfer injection. The other is integrating the ISS into the concept as a lunar transfer preparation infrastructure (LTPI), i.e. a staging terminal, on which the vehicle and its elements could be docked, mated, checked out and crewed. Latter includes further advantages: • No demand for extended free-flying and autonomous operation of multiple spacecraft at the same time, thus reduced operational complexity and risks • The ISS provides safe haven at LEO reducing operational risks
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Furthermore, the ISS orbit imposes no major disadvantage for lunar transfer in terms of delta-v and transfer windows [18], thus, this solution appears attractive and is selected here as the baseline. 3.1.2. Station locations and properties The gateway station location must have good access characteristics to and from both the Earth and the Moon. Such “lunar locations” include different orbits, especially a Low Lunar Orbit (LLO), the cis-lunar Lagrange one (LL1) and the equilateral points LL4 and LL5 leading and trailing the Moon respectively. Other locations such as circular or high eccentric Earth orbits or halo orbits around the trans-lunar LL2 offer no significant advantages for the task chosen here, therefore are not further addressed. Table 1 summarizes the properties of the three considered locations. With respect to the mission statement, a location is considered more advantageous if it has: • • • •
Low delta-v for in/outbound Earth-transfer Low station-keeping delta-v Low delta-v to lunar surface locations Highest accessibility to the relevant locations (i.e. lunar surface latitudes, LEO and Earth return) • Low transfer time, especially for crewed flights • Best visibility and communication links As summarized in Table 1 all three locations have their respective pros and cons, but LL1 is selected for further analysis because it offers most advantageous characteristics, including lowest Earth return delta-v and transfer time as well as reasonable station keeping delta-v and lunar accessibility. 3.2. Transfer vehicle infrastructure For designing the in-space transportation systems, the following mission phases have to be considered: 1. Preparation, i.e. launch and in-orbit assembly of transfer vehicles with payload
Table 1 Locations properties for a Lunar Space Station (TWP: Transfer Window Period, LSB: Lunar Surface Base) Property
LLO (100 km)
LL1
LL4/5
v LEO/ISS → Lunar TWP v Earth landing ← TE Earth transfer time Earth TWP v Lunar landing ← v Lunar surface → TL Lunar transfer time v Station-keeping per year
4.2 km/s 2 h (equatorial) – 14 d (polar) 1.12 km/s 5.4 d 2 h (equatorial) – 14 d (polar) 1.87 km/s 1.87 km/s hours 80 m/s
3.89–4.58 km/s
Lunar surface accessibility Lunar surface visibility
Latitudes inclination Latitudes ∼inclination (high-res, successively) Periodical, variable 118 min 3 days (twice per lunar month)
3.81–3.93 km/s 7.7–11.3 d (mean: 9.9 d) 0.65 km/s 3.8 d Permanent 2.52 km/s 2.52 km/s 1–3 days 0.5–36 m/s (Lissajous orbit) Global Approx. half globe (med-res, near side) Permanent, fixed 27.2 d Permanent
Communication link from a LSB Orbital period Occultation free period
0.86 km/s 4.6 d 2.58 km/s 2.58 km/s 11–22/7–15 days N/A Global Approx. half globe (low-res, leading/trailing at ±60◦ ) Permanently, fixed 27.2 d Permanent
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Fig. 3. Transfer vehicles conceptually designed (not to scale).
2. Lunar transfer, i.e. lunar transfer injection (LTI), coast on lunar transfer orbit (LTO), mid-course corrections (MCC), destination orbit insertion (OI) 3. Rendezvous and docking including station acquisition and arrival 4. Docked phase (passive) 5. Station departure, i.e. undocking and leaving station vicinity 6. Earth transfer, i.e. Earth transfer injection (ETI), coast on Earth transfer orbit (ETO) and ETO-MCC 7. Earth atmospheric re-entry and landing, if applicable A set of transfer vehicles was conceptually designed and is shown in Fig. 3. The high-thrust transfer stage for transfer/orbit injection and corrections is the ATV-HD. This is a “heavy duty” version of the European Automated Transfer Vehicle (ATV) with the cryogenic Vinci engine and weighing 27 t at launch including 22.1 t liquid oxygen and hydrogen. The External Tank Module (ETM) is an optional propellant module with up to 17.8 t cryogenic propellants to be docked to an ATV-HD for heavy payloads. ATV-L is a small logistics version of 4.1 t comprising the ATV spacecraft subassembly (SCSA) for delivering payload and ETM modules of up to 20 t to the ISS. Three principle classes of missions are incorporated: assembly, cargo missions using the Logistics Transfer Vehicle (LTV), and crewed missions using the Crew Transfer Vehicle (CTV). Payload size of assembly missions can be varied making 2 or 3 launches (2L or 3L) for each mission necessary. Logistics missions can be performed with or without payload recovery.
Fig. 4. Transfer mission modes (top: heavy payload assembly missions, bottom: light payload/crewed mission to LSS). Table 2 Maximum payload masses of assembly missions (with 2 and 3 launches, including 1 or 2 ATV-HD vehicles) Tons
3L
2L
@LSS @LL1 @LTO
17.9 19.5 32.0
6.5 7.0 13.5
3.3. Transfer mission modes Fig. 4 shows the mission modes for heavy payloads. Expended elements shall not be discarded to lunar surface to prevent large-scale contamination due to regular lunar spaceflight activities. Instead, all expended elements are disposed by reentry into the atmosphere of Earth. Resulting payload performance of assembly missions is stated in Table 2. The values for cargo and crewed missions using conventional and advanced cryogenic vehicles are given in Table 3. The cryogenic vehicles were designed to increase the
Table 3 Nominal payload masses delivered of cargo and crewed vehicles (conventional and cryogenic) Tons
LTV
CTV
Cryo-LTV
Cryo-CTV
Up Down
2.7 1.5
0.120 0.050
3.8 2.15
0.41 0.60
performance to or from the LSS, but limit their operations lifetime due to propellant boil-off of 5 to 6 kg (ATV-HD) or rather 2.2 kg (Cryo-CTV/LTV) per day. In this design 30 days mission
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Fig. 5. Configuration of Lunar Space Station Eve.
duration are assumed. The table states nominal payload, which can be altered for each individual mission. The outbound mass can be increased in cost of the inbound mass and vice versa. Thus, with Cryo-LTV a maximum pressurized payload of 3.9 t up (cargo) and 6.5 t down (waste disposal) is possible. 3.4. LSS conceptual design Using the infrastructure described above, ISS derived space station elements of up to 19.5 t can be transported from LEO to LL1. The baseline configuration selected is shown in Fig. 5 and comprises 6 modules with following key elements. An improved Zvezda service module (SM) serves as first element providing initial orbital capability. An improved SPP truss with a centralized solar power generation and thermal radiator system is located at the docking compartment (DC) with power and robotic arm controls. The truss is used as an external platform supported by the European Robotic Arm (ERA). The Autonomous Habitation Module (AHM) is based on Columbus and ATV and is equipped with improved life support system components and personal crew compartments for three astronauts. To mitigate the risks of solar flare radiation the concept envisions the crew compartments to be surrounded by parts of the water tanks and storage space for water-rich food and wastes in order to reduce the astronauts total dose. Altogether, the module is self-sustaining, thus, it can be un-docked from the station to provide a contingency Safe Haven for the crew. If station control cannot be re-established, the crew can leave LL1 and return to Earth with the CRV docked to the aft port of the AHM. The added docking port is connected to the main compartment by a pressurized access tunnel. A dedicated Life Support Module (LSM) provides advanced water recovery. It hosts a hybrid system comprising both, physicochemical and biological systems for increased food provision quality and synergistic oxygen regeneration. The Airlock and Docking Module (ADM) provides a dedicated EVA capability relieving SM docking port as airlock. It consists of two segments that can be sealed to each other. The
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Fig. 6. Launch systems and on-orbit mass evolution of LSS (IMLEO: Initial Mass to LEO; DMLSS: Delivered Mass to LSS; IOC: Initial Orbital Capability, CRV: Crew Return Vehicle).
first segment is based on the ISS-Docking Compartment (DC) and is used for EVA preparation and provides space for equipment and gas consumables. The second is a SM derived docking node and serves as the actual airlock. In addition, the ADM provides multiple docking/berthing ports meeting the requirement for the station’s further growth. The Cargo and Stowage Module (CSM) is used to transport some of the LSM equipments and is then used as stowage space for spare parts and solid waste. 3.4.1. On-orbit mass evolution Fig. 6 shows the LSS mass evolution during assembly and initial operations. The delivered mass to the station (DMLSS) mass is about 90 t at assembly complete including one LTV and one CTV with crew. To achieve this, a total of 416 t must be launched into LEO (IMLEO) including all vehicles and propellant for ISS rendezvous and transfer. 3.4.2. Environmental control and life support Initial ECLSS capabilities rely on conventional functionalities of SM. Because neither water, nor oxygen regeneration is available at this stage, crew stay times are limited to resources on-board the station and the CTV. The electrolyzer Electron and traditional oxygen candles using lithium-perchlorate (LiClO4) secure oxygen generation and CO2 removal is provided by lithium-hydroxide (LiOH) cartridges. The ECLSS at assembly complete is shown in Fig. 7 and comprises various advanced physiochemical and biological systems allowing for significant reduction of resupply and disposal mass as well as enhancing food variety. The individual components are listed in Table 4 with stating their place of installation, approximate mass and Technological Readiness Level (TRL). With delivery and activation of LSM the station water recovery system becomes available, which includes VPCAR, AES and MilliQ. The station is capable of generating potable water for a full crew of three plus an additional three persons during subsequent crew rotations. Thus, transportation of water
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Fig. 7. The Environment Control and Life Support System (ECLSS) of the Lunar Space Station Eve.
decreases to an almost negligible level and resupply mass reduces significantly. AHM provides advanced physicochemical devices for regenerative CO2 removal and O2 generation and hence completes the recycling capability. It includes EDC, SFWE and the Sabatier reactor. Except due to leakage no transport of atmospheric gases is necessary. Finally, CSM delivers SWIS, which heavily reduces the need for transportation of combustible wastes back to Earth. The generated water is fed into the waste water loop and the filtered CO2-rich exhaust gas undergoes the same treatment as cabin air. CSM also delivers add-ons to the biological ECLSS. By providing on-board produced edibles including algae-based proteins and vegetables the photo-bio-reactor (PBR) and the “Salad Machine” [19] reduce the yearly food resupply mass by 285 kg (nominal operation, i.e. three crew members, two crew rotations per year).
tivities, namely its primary role as gateway and safe haven for lunar surface exploration missions. The delta-v for a round trip mission to the surface is around 5 km/s. This offers the possibility of operating a fully reusable lunar shuttle stationed at the LSS at LL1. Fig. 8 shows the proposed mission mode for such a Lunar Exploration Vehicle (LEV) transporting 2 crewmembers to the Moon and back to the station. Assuming advanced cryogenic propulsion with a specific impulse of 460 s and a mission length of 15 days, a LEV of 4.75 t dry mass including cargo (400 kg down, 200 kg up) would need less than 10 t propellants for a roundtrip mission. Two Ariane 5-27 based launches are sufficient for refuelling the LEV at the LSS. In contrast to Apollo all locations on the Moon can be reached and the surface stay time is not restricted by orbit mechanical reasons but only by the LEV resources in terms of life support and propellant. 3.6. LSS based missions to SEL2 and NEA
3.4.3. Logistics: LSS vs. ISS Table 5 lists resupply of ISS and LSS. The LSS system and subsystem consumables are conservatively estimated being equal to the ISS, although mass and volume of the LSS is smaller. Nevertheless, the resupply of consumables is 1726 kg less than compared to scaled ISS values. Three conventional LTV or rather two Cryo-LTV missions per year are sufficient for permanently crewed operation for resupply and waste disposal (Table 6). 3.5. LSS based surface exploration missions The space station was to be investigated as an intermediate step towards subsequent space exploration and exploitation ac-
The station could analogously be used as starting point for flights to SEL2 and Near Earth Asteroids (NEA). The deep space escape delta-v from LL1 is about 140 m/s – the lowest value for the entire Earth–Moon system [20]. Compared to 3220 m/s necessary from LEO this shows how much more efficient and capable round-trip missions to these destinations can be performed using the LSS. A 30 day roundtrip mission to SEL2 would require a delta-v about of 5 km/s when starting from LEO. Starting from LL1 this is reduced to below 1.3 km/s. Although this number alone does not account for the transportation effort of mission specific equipment and propellants to the LL1 gateway, it demonstrates the increase of efficiency when reusing crew and transfer vehicles and, thus, minimizing
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Table 6 Annual disposal mass of LSS
Table 4 ECLSS components
Atmosphere
Water
Waste
Food
Miscellaneous
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Component
Module
Static Feed Water Electrolysis (SFWE) Electro-chemical Depolarized CO2 Concentration (EDC) Trace Contaminant Control System (TCCS) Condensing Heat Exchanger (CHX) Photo-Bio-Reactor (PBR) Lithium hydroxide cartridges (LiOH) Oxygen candles (LiClO4)
AHM
Vapor Phase Catalytic Ammonia Removal System (VPCAR) Air Evaporation System (AES) MilliQ Sabatier Reactor (SR) Solid Waste Incineration System (SWIS) Algae Processor Salad Machine Storage Galley Solar collector Fire protection Hygiene facility Casualty wage (sick bay) Piping, ventilation Various Tanks
Mass [kg]
TRL
Transportation mass [kg]
Total LSS
200.00
8
AHM
89.00
5
ECLSS items System equipment Science instruments
3298.0 700.0+ 100.0+
SM, AHM SM, AHM LSM All
85.00
9
Summary Vehicle cargo capacity No. of vehicles Average cargo mass
4098.0 1500 3 1433
–
9 5 9
LSM
160.00 148 up 201 down* 243 up 171 down* 283.00
LSM
75.00
5
LSM AHM LSM+
100.00 43.00 150.00
5 7 5
LSM+ LSM+ All SM, LSM LSM All SM, AHM SM All SM, AHM, LSM
10.00 200.00 91.00 120.00
N/A 5/6 9
42.00 70.00 120.00
7 9 9
All
2800 2 2150
9 3/4
Fig. 8. Crewed missions to the Moon using LSS gateway. 60.00 672.00 1690.00
– 9 7–9
ECLSS elements and servicing capabilities including robotics, airlock, and power and thermal control system. 4. Conclusion
+ : components delivered by CSM, ∗ : 3 men/30 days.
Table 5 Annual resupply of ISS vs. LSS (ISS data based on Progress missions from 02/2003 to 08/2004; crew of 2) Transportation mass [kg]
Total ISS (scaled to 3)
Total LSS
ECLSS items System equipment Propellant Re-boost Science instruments
6325.6 1535.2 2181.6 1000.0 284.8
4599.2 1500.0 1140.0 0.0 300.0
Summary Vehicle cargo capacity
11327.2 2500.0
No. of vehicles Average cargo mass
5 2265.4
7540.0 2740 LTV 3 2200
3800 Cryo-LTV 2 3770
IMLEO or in-turn increasing the mass ratio of equipment and payload. A reusable SEL2 servicing vehicle of about 19 t based at the LSS at LL1 could be feasible requiring less than 10 t storable propellants. One module could be designed as servicing vehicle to synergistically sharing with the station its habitable volume,
The presented scenario of a Lunar Space Station at LL1 demonstrates a lunar exploration architecture with a gateway station in lunar vicinity rather than departing from Earth or LEO directly. The results reveal that such a LSS at LL1 is feasible and programmatically attractive. In the current situation of human spaceflight it offers various benefits for Europe and Russia: • Generating experience with human spaceflight beyond LEO. • Enabling demonstration and verification of key technologies in (semi) deep space as preparation for lunar surface, Mars and other missions. • Offering exploration of the Moon globally and systematically, with a Safe Haven system. • Enabling crewed missions to other near Earth destinations such as SEL2 for servicing and maintenance operations. • Allowing for new models for international cooperation, i.e. robust ESA–Russian elements embedded to an international exploration program led by the USA. • Finally, it is feasible and affordable in near-term by using existing knowledge and technology and even hardware.
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When thinking of human space exploration in general and systematic exploration of the Moon in specific, it is absolutely reasonable to take mission scenarios into account that involve crewed cis-lunar space infrastructures. Because the Moon is (visually) present in the every day life of all humans, public attention and support is provided, which is crucial for every future space program. This may have the potential to also accelerate the exploration of Mars, especially when we follow an integrated Moon/Mars exploration approach with taking first steps early. Acknowledgement The authors wish to thank Robert Lainé, Jean-François Clervoy and the ESA ATV team in Les Mureaux for the generous support provided. References [1] Aurora Programme, Presented and approved at the Ministerial Conference in Edinburgh, 11/2001, URL: www.esa.int/aurora. [2] G.W. Bush, New Vision for Space Exploration Program – Remarks by the President on U.S. Space Policy, Washington, D.C., NASA Special Publication NP-2004-01-334-HQ, January 2004. [3] D. Stanley et al., NASA’s Exploration Systems Architecture Study (ESAS01), Final Report, NASA-TM-2005-214062, November 2005. [4] Project Constellation, NASA, 2007, URL: http://www.nasa.gov/mission_ pages/constellation. [5] K. Yazdi, Conceptual design and flight simulation of space station mission beyond low earth orbit, Dissertation, Universität Stuttgart, 2006.
[6] R. Bertrand, Conceptual design and flight simulation of space stations, Aerosp. Sci. Technol. 5 (2001) 147–163. [7] A. Hinüber, Attitude and orbit analysis for conceptual space station design, Dissertation, Universität Stuttgart, 2002. [8] I.A. Crawford, The scientific case for human spaceflight, Astronomy & Geophysics 39 (6) (December 1998) 14–17. [9] Human Spaceflight Vision Group (HSVG), Moon: the 8th continent – Human Spaceflight 2025, Final Report, ESA, December 2003. [10] Mission to the Moon, ESA SP-1150, ESA 1992. [11] I.A. Crawford, Human exploration of the Moon and Mars: Implications for aurora, Astronomy and Geophysics 45 (2). [12] I.A. Crawford, Back to the Moon?, Astronomy and Geophysics 44 (2) (2003) 15–17. [13] P. Lucey, Lunar Poles – Status of Understanding a Potential Resource, Presentation, LPI 03/2004. [14] W.T. Huntress et al., Final Report on the Next Steps in Exploring Deep Space – Cosmic Study, IAA, 9 July 2004, URL: http://iaaweb.org/iaa/ Studies/nextsteps.pdf. [15] C. Brown, Spacecraft Mission Design, 2nd ed., AIAA, 1998. [16] H.-H. Koelle, Concept and Simulation of a Lunar Orbit Service Centre, ILR Mitt. 258, TU Berlin 2002. [17] S. Hovland, ESA Study Report CDF-23(A), ESA 2004. [18] K. Yazdi, E. Messerschmid, Analysis of parking orbits and transfer trajectories for mission design of cis-lunar space stations, IAC-03-IAA. 10.1.07, in: New Opportunities for Space, Acta Astronautica 55 (3–9) (2004) 759– 771. [19] M. Kliss, R.D. MacElroy, Salad Machine – A vegetable production unit for long duration space missions, SAE Transcriptions 99 (1) (1999) 722– 729. [20] D.G. Boden, S.J. Hoffmann, Orbit selection and astrodynamics, in: P. Larson (Ed.), Human Spaceflight – Mission and Analysis, McGraw-Hill, 2000.
A lunar exploration architecture using lunar libration point one Kian Yazdi a,∗ , Ernst Messerschmid b,1 a EADS Astrium Ltd, Gunnels Wood Road, Stevenage, SG1 2AS, United Kingdom b Institut fuer Raumfahrtsysteme, Universitaet Stuttgart, Pfaffenwaldring 31, D-70550 Stuttgart, Germany
Received 22 June 2006; received in revised form 10 April 2007; accepted 1 June 2007 Available online 7 June 2007
Abstract Future space exploration activities very likely will incorporate a large number of human missions to various locations on the Moon surface and other destinations near the Earth–Moon system. A wide range of space infrastructure elements will be required in low Earth orbit and waypoints to these destinations. It is reasonable to start assembly and mission logistics from today’s frequently used near-Earth orbits, such as the orbit of the International Space Station (ISS). This allows for using launch systems and other infrastructure elements currently available in order to reduce costs. Besides new transfer, ascent and re-entry vehicles as well as planetary surface installations, space stations beyond low Earth orbit, namely in the Earth–Moon system, can be crucial elements for forthcoming exploration missions. The lunar Lagrange points have unique advantages in this context and offer promising options for the humans’ next steps into space. As shown in this paper, a lunar space station (LSS) is not too farfetched; and it is definitely not science fiction but rather the next logical step for solar system exploration. Like a stepping-stone, it could enable the immediate start of sustained development of outer space with technologies available today. The proposed lunar exploration architecture targets at an LSS at the lunar Lagrange point 1 to be used as a gateway for mid-term lunar surface exploration missions or servicing missions to other destinations. For study purposes the challenging constraint of only using existing and tailored infrastructure elements and technology of European/Russian heritage is assumed. The results manifest the feasibility of such a space station between 2015 and 2020. Presented are station configuration and modules, transfer vehicles for assembly, crew and cargo transport, an enhanced life support system and a logistics concept. The concept outlines enhancements of the current transportation and station infrastructure and shows that the ISS as a transportation node can beneficially support lunar scenarios. © 2007 Elsevier Masson SAS. All rights reserved. Keywords: Lagrange point; Lunar exploration; Mission architecture; Space stations; Human spaceflight Schlüsselwörter: Astronautik; Lagrangepunkt; Lunare Exploration; Missionsarchitektur; Mondraumstation
1. Introduction For more than three decades, no human space exploration effort beyond Low Earth Orbit (LEO) has existed since Apollo. Now, after the dawn of the 21st century, it seems that NASA and other space agencies have arrived at the conclusion that the time has come to leave Earth orbit again for new human spaceflight challenges. Recently, the US government has decided on NASA’s long-term strategic space programme leading back to * Corresponding author. Tel.: +44(0)143877-3749; fax: +44(0)143877-8910.
E-mail addresses: [email protected], [email protected] (K. Yazdi). 1 Member IAA. 1270-9638/$ – see front matter © 2007 Elsevier Masson SAS. All rights reserved. doi:10.1016/j.ast.2007.06.001
lunar activities and more to come. The European Space Agency (ESA) formulated its Aurora programme in 2001, which also includes the preparation and execution of the long-term human exploration of the solar system. In connection with this, an international human mission to Mars is envisioned with the possibility of using lunar locations as a way station. A lunar space station for orbital support of lunar surface activities is documented here as an example for a near-term start of a sustained approach to lunar exploration and exploitation. As a point of departure, a minimum configuration station, serving as a stepping-stone and gateway to further lunar and nearlunar activities was investigated. This addresses an option for ESA’s preparatory programme Aurora [1] indicating Europe’s long-term intentions for the next 30 years, and the US Space
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Exploration Policy of NASA focusing on human exploration of the solar system and primarily targeting the Moon and Mars [2–4]. The visions foresee a human return to the Moon surface by 2020 and the first human expedition to Mars by 2030. China and other countries such as India and Japan engage in space exploration as well. Activities in the Earth–Moon system are seen as the “next logical step” for the development of space and to prepare successive human space operations beyond. The actual range of possible utilizations of a lunar space station is manifold and shall be addressed in the beginning of this paper. This is followed by a discussion of the associated challenges and mission architecture of a scenario demonstrating the application of methodology and tools to a sample conceptual design problem. Conceptual design of space stations deals with systems of relatively large scale, long life time and with complex logistical and operational requirements. The applied conceptual design process needs the availability of sophisticated design and simulation tools, coping with the involved extensive calculations combined with high-level interactivity [5–7]. If properly implemented, this enables analysis and comparison of specific characteristics for evaluating the pros and cons of system and mission design concepts with rapid turn-around times. This paper focuses on the conceptual design of a lunar space station and supporting transfer vehicles. Relevant mission and system parameters are discussed and a summary of results is presented. Details may be found in [5]. 2. Lunar exploration and development rationales The Moon, as the celestial neighbour to Earth, is more accessible than all other bodies in the solar system. Thus human spaceflight infrastructures on or close to the Moon will be far
better maintainable and usable with access of shorter period and higher frequency. Furthermore, the Moon offers a long-term perspective for economical exploitation of space (in terms of propellant and raw material production and – possibly one day – tourism). A permanent human installation (a “gateway”) in the Earth– Moon system would enable efficient transportation to other destinations than LEO based scenarios. Simultaneously, the Moon and a lunar gateway offer excellent early utilization possibilities in scientific research and technological developments for successive steps of mankind into the universe, be it exploration or exploitation. 2.1. Science rationales By extending human activities to cis-lunar, science will benefit significantly. Building up of an infrastructure fosters research activities that a purely scientific budget could not afford and would not otherwise take place [8]. Beside its good accessibility, the most compelling arguments supporting the scientific case for a return to the Moon are based on geosciences, life sciences, astronomy and solar physics [9]. As summarized in Fig. 1 the causes for this are manifold: The Moon features no considerable atmosphere, has no tectonics and rotates slowly (33 arcsec per minute). Its surface is extremely old, preserving a record of the early evolution of terrestrial planets and the near-Earth cosmic environment [10]. This record includes the time period before 4.5 to 3.8 billion years that is not present on any other terrestrial body, except Mercury [11]. Our current understanding of planetary and solar system history and evolution bases primarily on lunar surface observation and samples brought back by the Apollo missions from only six landing sites
Fig. 1. Science and programmatic rationales for lunar exploration and development.
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in the vicinity of Moon’s equator on the near-side. Therefore, any approach for a better understanding ultimately will require thorough surface analysis including sample identification, collection and analysis on a far larger scale than before [12]. This will demand a considerable amount of fieldwork for collecting more and more diverse rock and soil and subsurface samples and returning the most promising to Earth. With an established access to the lunar surface, scientists will be interested in operating other instruments there e.g. seismic profiling devices and interferometers. 2.2. Resources and geopolitical issues of the Moon The Moon is not only of scientific interest; it could also be the “most valuable real estate in the solar system” [13] because of its location, properties and resources. The main arguments are based on the utilization and exploitation of lunar resources and the use of the lunar environment as a demonstration platform for a number of technologies required for robotic and human exploration of Mars and the solar system. Here, the task will be determining what lunar resources in which quantities indeed exist, proving their accessibility and verifying whether and how they can be economically utilized and exploited. For instance, hydrogen in some form must be present, especially at the poles [13]. This includes the possibility of over 10 billion tons of water ice [9]. Both, hydrogen and oxygen (or water), are assets for future space programmes due to their application in life support systems and for in-situ propellant production. In long perspective, the Moon could possibly offer energy. Besides supply by regenerative means, the long-term energy supply for Earth will potentially reside on thermonuclear fusion reactors using helium-3 as fuel. This volatile mass fraction on lunar regolith is only about 13 mg/ton [10], thus, a large area must be processed to obtain a considerable mass. However, only one kilogram of helium-3 is necessary to produce the same amount of energy as 10 billion kg of fossil fuel [9]. These examples emphasize the geopolitical aspect the Moon will possibly have in the second half of this century, in a world of scarce fossil fuels. This alone might be the driving cause for the US government to justify their renewed space initiative by arguments including national economy and security [2]. 2.3. SEL2, solar system exploration and the Moon Besides science, the “IAA Cosmic Study” [14] names the “exploration imperative – the innate drive of human beings to extend their boundaries and explore the unknown” as the rationale for human spaceflight. It proposes a staged approach into the solar system and selects following 4 steps and destinations: 1. 2. 3. 4.
Step: Beyond LEO, i.e. Moon and SEL2 Step: Deep Space, i.e. NEA Step: On to Mars, i.e. Phobos/Deimos Step: Down to Mars, i.e. Mars surface
For each step, the study identifies one new capability: a new transportation element, named geospace exploration vehicle to
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access the Moon and Sun–Earth Libration Point 2 (SEL2) for maintenance and servicing astronomical facilities, and an interplanetary exploration vehicle for roundtrips to Near Earth Asteroids (NEA) for studying these objects, which pose a potential impact threat to Earth. It is assumed that each step bases on the results of the previous steps and uses infrastructure elements built-up programmatically (technology development) or physically (staging). 2.4. Summary: Goals of the next steps Lunar scenarios are considered as first step for sustaining activities beyond LEO. The conclusion is that the destination Moon is of high priority, be it for cultural, political or scientific reasons. And whatever emphasis the future scenario will have: global access to Moon’s surface is an important requirement with the necessity of reaching each point repeatedly. To meet the long-term objectives, transfers must be possible to the Lunar Lagrange points, Sun–Earth Lagrange points, NEA, Mars, and destinations beyond. Implied by the requirement of affordability, risk reduction and sustainability, the design of the forthcoming interplanetary human spaceflight missions suggest the following general requirements: • Easy access via available or derived transportation systems (efficient utilization of launchers and their already existing launch sites) • Safe mission operations (i.e. large departure and arriving windows, robust free-return and/or abort options) [15] • Frequent departure and arrival windows (once per revolution, daily, weekly, monthly, etc.) • Minimum delta-v for transfers, station keeping and orbit maintenance • Minimum transfer time • A human factors supporting concept; especially enhancing ergonomics, habitability and reduction of the radiation hazard (i.e. accumulated radiation dose) Furthermore, constraints arise due to superior issues typically involved in future projects: • International cooperation: access from launch sites of all potential project partners • Utilization of the (already existing) International Space Station (ISS) • Sustained approach based on currently available or upgraded technology would enable development and use of common transportation vehicles. A space station or a human-tended space platform installed in the Earth–Moon system could play a significant role in nearterm space activities and would offer particular utilization objectives summarized in Fig. 2. As a demonstration example, a fictitious mission statement is assumed here, including following objectives, constraints and requirements.
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Fig. 2. Potential utilization objectives of a Lunar Space Station at the Lunar Lagrange point One (LL1).
2.5. Sample mission statement The project’s task is conceptualizing a Lunar Space Station (LSS) as a gateway to the Moon and beyond. This station shall prepare and support successive steps of lunar surface exploration by meeting following objectives: • Staging platform with permanent access to and from Earth and the surface of the Moon • Space-research beyond LEO (e.g. radiation) • Test and verification of technologies and processes (e.g. propulsion and enhanced life support systems) The station’s development, deployment and operation shall allocate minimum costs, risk and time. First level requirements and constraints are as follows: • • • • • •
European–Russian lead cooperation Minimum configuration but extendable Affordable, low risk and near-term Use of ISS hardware heritage Permanent crew of 3, temporarily 6 Start of operation between 2015 and 2018 with 10 years lifetime
The station utilization characteristics would allow for a robust contribution of Europe together with Russian partners, with own objectives within a larger international programme including the USA, Japan, China and other partners. But the mission emphasizes the intensifying collaboration between Europe and Russia in space by e.g. implementing the Soyuz or follow-on man-rated launchers at Europe’s spaceport at Kourou, French Guyana, and their agreement on joint future space activities. Recently initiated US developments following ESAS [3] with respect to project Constellation [4] are therefore not detailed
here, although they would be synergistically complemented by an LSS. 3. Lunar exploration architecture using a space station Most past and recent studies done on lunar architectures concentrate on surface bases and only very few studies have included non-surface installations or elements such as lunar space stations or platforms. Examples are a large lunar orbiting station named Space Operation Centre (SOC) [16], the rotating Clarke Station conceptually designed in a student project, and a gateway-like station in a high eccentric Earth–Moon cycling orbit drafted in another ISU student project. However, little can be directly applied for the mission objectives identified above. 3.1. Mission architectural decisions The goal is the creation of an initial, generic and affordable waypoint station in a short time frame to be used by succeeding crewed missions. The emphasis lies primarily on lunar exploration, utilization and exploitation. Thus the preparation and support of missions to the Moon and other destinations is the task, not providing the exploration capabilities itself. The main decisions to be made driving the project cost, complexity and risks are: • • • • • •
Launch site and vehicles LEO support (in/outbound) Station location Surface accessibility Surface stay time Transport elements/degree of re-usability
The transportation infrastructure includes launch systems to LEO or direct Lunar Transfer Orbit LTO-injection and in-space transportation systems (transfer vehicles). These systems will
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define, or rather limit, the size and mass of the station elements and the size and performance of the transfer vehicles. Besides systems available and already used today (Kourou based Ariane 5-ESV/ECA, and Baikonur based Proton M, Soyuz ST/FG/U) enhanced versions of systems planned to date are considered. This includes Ariane 5-ECB and Onega, which is an enhanced version of the Soyuz planned to carry a new Russian based crew transportation vehicle, called Clipper. Not yet planned but considered here, is an enhanced version of Ariane 5-ECB, called A5-ECB+ or A5-27 with an increased LEO payload performance of about 27 tons. The modifications necessary were baselined by industrial specialists in a study performed at ESA and were assumed to be feasible and available by 2015 [17]. The ESA study indicates an increase of delta-v between 8% (for Ariane 5 ESV) and 5% for a A5-27 rocket, thus the resulting benefits using the ISS should more than outweigh the reduction of payload. 3.1.1. LEO support Because no appropriate heavy lift launch vehicle of Saturn or Energia class exists today, the station elements and transfer vehicles cannot be launched directly to the destination location similar to the Apollo missions. This means transfer vehicles, i.e. injection stages, and some kinds of LEO operations are necessary for mating the elements for the individual missions. For the chosen example, two alternatives exist: One option is using a freely chosen parking orbit, where the elements automatically mate and are checked-out remotely before lunar transfer injection. The other is integrating the ISS into the concept as a lunar transfer preparation infrastructure (LTPI), i.e. a staging terminal, on which the vehicle and its elements could be docked, mated, checked out and crewed. Latter includes further advantages: • No demand for extended free-flying and autonomous operation of multiple spacecraft at the same time, thus reduced operational complexity and risks • The ISS provides safe haven at LEO reducing operational risks
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Furthermore, the ISS orbit imposes no major disadvantage for lunar transfer in terms of delta-v and transfer windows [18], thus, this solution appears attractive and is selected here as the baseline. 3.1.2. Station locations and properties The gateway station location must have good access characteristics to and from both the Earth and the Moon. Such “lunar locations” include different orbits, especially a Low Lunar Orbit (LLO), the cis-lunar Lagrange one (LL1) and the equilateral points LL4 and LL5 leading and trailing the Moon respectively. Other locations such as circular or high eccentric Earth orbits or halo orbits around the trans-lunar LL2 offer no significant advantages for the task chosen here, therefore are not further addressed. Table 1 summarizes the properties of the three considered locations. With respect to the mission statement, a location is considered more advantageous if it has: • • • •
Low delta-v for in/outbound Earth-transfer Low station-keeping delta-v Low delta-v to lunar surface locations Highest accessibility to the relevant locations (i.e. lunar surface latitudes, LEO and Earth return) • Low transfer time, especially for crewed flights • Best visibility and communication links As summarized in Table 1 all three locations have their respective pros and cons, but LL1 is selected for further analysis because it offers most advantageous characteristics, including lowest Earth return delta-v and transfer time as well as reasonable station keeping delta-v and lunar accessibility. 3.2. Transfer vehicle infrastructure For designing the in-space transportation systems, the following mission phases have to be considered: 1. Preparation, i.e. launch and in-orbit assembly of transfer vehicles with payload
Table 1 Locations properties for a Lunar Space Station (TWP: Transfer Window Period, LSB: Lunar Surface Base) Property
LLO (100 km)
LL1
LL4/5
v LEO/ISS → Lunar TWP v Earth landing ← TE Earth transfer time Earth TWP v Lunar landing ← v Lunar surface → TL Lunar transfer time v Station-keeping per year
4.2 km/s 2 h (equatorial) – 14 d (polar) 1.12 km/s 5.4 d 2 h (equatorial) – 14 d (polar) 1.87 km/s 1.87 km/s hours 80 m/s
3.89–4.58 km/s
Lunar surface accessibility Lunar surface visibility
Latitudes inclination Latitudes ∼inclination (high-res, successively) Periodical, variable 118 min 3 days (twice per lunar month)
3.81–3.93 km/s 7.7–11.3 d (mean: 9.9 d) 0.65 km/s 3.8 d Permanent 2.52 km/s 2.52 km/s 1–3 days 0.5–36 m/s (Lissajous orbit) Global Approx. half globe (med-res, near side) Permanent, fixed 27.2 d Permanent
Communication link from a LSB Orbital period Occultation free period
0.86 km/s 4.6 d 2.58 km/s 2.58 km/s 11–22/7–15 days N/A Global Approx. half globe (low-res, leading/trailing at ±60◦ ) Permanently, fixed 27.2 d Permanent
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Fig. 3. Transfer vehicles conceptually designed (not to scale).
2. Lunar transfer, i.e. lunar transfer injection (LTI), coast on lunar transfer orbit (LTO), mid-course corrections (MCC), destination orbit insertion (OI) 3. Rendezvous and docking including station acquisition and arrival 4. Docked phase (passive) 5. Station departure, i.e. undocking and leaving station vicinity 6. Earth transfer, i.e. Earth transfer injection (ETI), coast on Earth transfer orbit (ETO) and ETO-MCC 7. Earth atmospheric re-entry and landing, if applicable A set of transfer vehicles was conceptually designed and is shown in Fig. 3. The high-thrust transfer stage for transfer/orbit injection and corrections is the ATV-HD. This is a “heavy duty” version of the European Automated Transfer Vehicle (ATV) with the cryogenic Vinci engine and weighing 27 t at launch including 22.1 t liquid oxygen and hydrogen. The External Tank Module (ETM) is an optional propellant module with up to 17.8 t cryogenic propellants to be docked to an ATV-HD for heavy payloads. ATV-L is a small logistics version of 4.1 t comprising the ATV spacecraft subassembly (SCSA) for delivering payload and ETM modules of up to 20 t to the ISS. Three principle classes of missions are incorporated: assembly, cargo missions using the Logistics Transfer Vehicle (LTV), and crewed missions using the Crew Transfer Vehicle (CTV). Payload size of assembly missions can be varied making 2 or 3 launches (2L or 3L) for each mission necessary. Logistics missions can be performed with or without payload recovery.
Fig. 4. Transfer mission modes (top: heavy payload assembly missions, bottom: light payload/crewed mission to LSS). Table 2 Maximum payload masses of assembly missions (with 2 and 3 launches, including 1 or 2 ATV-HD vehicles) Tons
3L
2L
@LSS @LL1 @LTO
17.9 19.5 32.0
6.5 7.0 13.5
3.3. Transfer mission modes Fig. 4 shows the mission modes for heavy payloads. Expended elements shall not be discarded to lunar surface to prevent large-scale contamination due to regular lunar spaceflight activities. Instead, all expended elements are disposed by reentry into the atmosphere of Earth. Resulting payload performance of assembly missions is stated in Table 2. The values for cargo and crewed missions using conventional and advanced cryogenic vehicles are given in Table 3. The cryogenic vehicles were designed to increase the
Table 3 Nominal payload masses delivered of cargo and crewed vehicles (conventional and cryogenic) Tons
LTV
CTV
Cryo-LTV
Cryo-CTV
Up Down
2.7 1.5
0.120 0.050
3.8 2.15
0.41 0.60
performance to or from the LSS, but limit their operations lifetime due to propellant boil-off of 5 to 6 kg (ATV-HD) or rather 2.2 kg (Cryo-CTV/LTV) per day. In this design 30 days mission
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Fig. 5. Configuration of Lunar Space Station Eve.
duration are assumed. The table states nominal payload, which can be altered for each individual mission. The outbound mass can be increased in cost of the inbound mass and vice versa. Thus, with Cryo-LTV a maximum pressurized payload of 3.9 t up (cargo) and 6.5 t down (waste disposal) is possible. 3.4. LSS conceptual design Using the infrastructure described above, ISS derived space station elements of up to 19.5 t can be transported from LEO to LL1. The baseline configuration selected is shown in Fig. 5 and comprises 6 modules with following key elements. An improved Zvezda service module (SM) serves as first element providing initial orbital capability. An improved SPP truss with a centralized solar power generation and thermal radiator system is located at the docking compartment (DC) with power and robotic arm controls. The truss is used as an external platform supported by the European Robotic Arm (ERA). The Autonomous Habitation Module (AHM) is based on Columbus and ATV and is equipped with improved life support system components and personal crew compartments for three astronauts. To mitigate the risks of solar flare radiation the concept envisions the crew compartments to be surrounded by parts of the water tanks and storage space for water-rich food and wastes in order to reduce the astronauts total dose. Altogether, the module is self-sustaining, thus, it can be un-docked from the station to provide a contingency Safe Haven for the crew. If station control cannot be re-established, the crew can leave LL1 and return to Earth with the CRV docked to the aft port of the AHM. The added docking port is connected to the main compartment by a pressurized access tunnel. A dedicated Life Support Module (LSM) provides advanced water recovery. It hosts a hybrid system comprising both, physicochemical and biological systems for increased food provision quality and synergistic oxygen regeneration. The Airlock and Docking Module (ADM) provides a dedicated EVA capability relieving SM docking port as airlock. It consists of two segments that can be sealed to each other. The
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Fig. 6. Launch systems and on-orbit mass evolution of LSS (IMLEO: Initial Mass to LEO; DMLSS: Delivered Mass to LSS; IOC: Initial Orbital Capability, CRV: Crew Return Vehicle).
first segment is based on the ISS-Docking Compartment (DC) and is used for EVA preparation and provides space for equipment and gas consumables. The second is a SM derived docking node and serves as the actual airlock. In addition, the ADM provides multiple docking/berthing ports meeting the requirement for the station’s further growth. The Cargo and Stowage Module (CSM) is used to transport some of the LSM equipments and is then used as stowage space for spare parts and solid waste. 3.4.1. On-orbit mass evolution Fig. 6 shows the LSS mass evolution during assembly and initial operations. The delivered mass to the station (DMLSS) mass is about 90 t at assembly complete including one LTV and one CTV with crew. To achieve this, a total of 416 t must be launched into LEO (IMLEO) including all vehicles and propellant for ISS rendezvous and transfer. 3.4.2. Environmental control and life support Initial ECLSS capabilities rely on conventional functionalities of SM. Because neither water, nor oxygen regeneration is available at this stage, crew stay times are limited to resources on-board the station and the CTV. The electrolyzer Electron and traditional oxygen candles using lithium-perchlorate (LiClO4) secure oxygen generation and CO2 removal is provided by lithium-hydroxide (LiOH) cartridges. The ECLSS at assembly complete is shown in Fig. 7 and comprises various advanced physiochemical and biological systems allowing for significant reduction of resupply and disposal mass as well as enhancing food variety. The individual components are listed in Table 4 with stating their place of installation, approximate mass and Technological Readiness Level (TRL). With delivery and activation of LSM the station water recovery system becomes available, which includes VPCAR, AES and MilliQ. The station is capable of generating potable water for a full crew of three plus an additional three persons during subsequent crew rotations. Thus, transportation of water
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Fig. 7. The Environment Control and Life Support System (ECLSS) of the Lunar Space Station Eve.
decreases to an almost negligible level and resupply mass reduces significantly. AHM provides advanced physicochemical devices for regenerative CO2 removal and O2 generation and hence completes the recycling capability. It includes EDC, SFWE and the Sabatier reactor. Except due to leakage no transport of atmospheric gases is necessary. Finally, CSM delivers SWIS, which heavily reduces the need for transportation of combustible wastes back to Earth. The generated water is fed into the waste water loop and the filtered CO2-rich exhaust gas undergoes the same treatment as cabin air. CSM also delivers add-ons to the biological ECLSS. By providing on-board produced edibles including algae-based proteins and vegetables the photo-bio-reactor (PBR) and the “Salad Machine” [19] reduce the yearly food resupply mass by 285 kg (nominal operation, i.e. three crew members, two crew rotations per year).
tivities, namely its primary role as gateway and safe haven for lunar surface exploration missions. The delta-v for a round trip mission to the surface is around 5 km/s. This offers the possibility of operating a fully reusable lunar shuttle stationed at the LSS at LL1. Fig. 8 shows the proposed mission mode for such a Lunar Exploration Vehicle (LEV) transporting 2 crewmembers to the Moon and back to the station. Assuming advanced cryogenic propulsion with a specific impulse of 460 s and a mission length of 15 days, a LEV of 4.75 t dry mass including cargo (400 kg down, 200 kg up) would need less than 10 t propellants for a roundtrip mission. Two Ariane 5-27 based launches are sufficient for refuelling the LEV at the LSS. In contrast to Apollo all locations on the Moon can be reached and the surface stay time is not restricted by orbit mechanical reasons but only by the LEV resources in terms of life support and propellant. 3.6. LSS based missions to SEL2 and NEA
3.4.3. Logistics: LSS vs. ISS Table 5 lists resupply of ISS and LSS. The LSS system and subsystem consumables are conservatively estimated being equal to the ISS, although mass and volume of the LSS is smaller. Nevertheless, the resupply of consumables is 1726 kg less than compared to scaled ISS values. Three conventional LTV or rather two Cryo-LTV missions per year are sufficient for permanently crewed operation for resupply and waste disposal (Table 6). 3.5. LSS based surface exploration missions The space station was to be investigated as an intermediate step towards subsequent space exploration and exploitation ac-
The station could analogously be used as starting point for flights to SEL2 and Near Earth Asteroids (NEA). The deep space escape delta-v from LL1 is about 140 m/s – the lowest value for the entire Earth–Moon system [20]. Compared to 3220 m/s necessary from LEO this shows how much more efficient and capable round-trip missions to these destinations can be performed using the LSS. A 30 day roundtrip mission to SEL2 would require a delta-v about of 5 km/s when starting from LEO. Starting from LL1 this is reduced to below 1.3 km/s. Although this number alone does not account for the transportation effort of mission specific equipment and propellants to the LL1 gateway, it demonstrates the increase of efficiency when reusing crew and transfer vehicles and, thus, minimizing
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Table 6 Annual disposal mass of LSS
Table 4 ECLSS components
Atmosphere
Water
Waste
Food
Miscellaneous
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Component
Module
Static Feed Water Electrolysis (SFWE) Electro-chemical Depolarized CO2 Concentration (EDC) Trace Contaminant Control System (TCCS) Condensing Heat Exchanger (CHX) Photo-Bio-Reactor (PBR) Lithium hydroxide cartridges (LiOH) Oxygen candles (LiClO4)
AHM
Vapor Phase Catalytic Ammonia Removal System (VPCAR) Air Evaporation System (AES) MilliQ Sabatier Reactor (SR) Solid Waste Incineration System (SWIS) Algae Processor Salad Machine Storage Galley Solar collector Fire protection Hygiene facility Casualty wage (sick bay) Piping, ventilation Various Tanks
Mass [kg]
TRL
Transportation mass [kg]
Total LSS
200.00
8
AHM
89.00
5
ECLSS items System equipment Science instruments
3298.0 700.0+ 100.0+
SM, AHM SM, AHM LSM All
85.00
9
Summary Vehicle cargo capacity No. of vehicles Average cargo mass
4098.0 1500 3 1433
–
9 5 9
LSM
160.00 148 up 201 down* 243 up 171 down* 283.00
LSM
75.00
5
LSM AHM LSM+
100.00 43.00 150.00
5 7 5
LSM+ LSM+ All SM, LSM LSM All SM, AHM SM All SM, AHM, LSM
10.00 200.00 91.00 120.00
N/A 5/6 9
42.00 70.00 120.00
7 9 9
All
2800 2 2150
9 3/4
Fig. 8. Crewed missions to the Moon using LSS gateway. 60.00 672.00 1690.00
– 9 7–9
ECLSS elements and servicing capabilities including robotics, airlock, and power and thermal control system. 4. Conclusion
+ : components delivered by CSM, ∗ : 3 men/30 days.
Table 5 Annual resupply of ISS vs. LSS (ISS data based on Progress missions from 02/2003 to 08/2004; crew of 2) Transportation mass [kg]
Total ISS (scaled to 3)
Total LSS
ECLSS items System equipment Propellant Re-boost Science instruments
6325.6 1535.2 2181.6 1000.0 284.8
4599.2 1500.0 1140.0 0.0 300.0
Summary Vehicle cargo capacity
11327.2 2500.0
No. of vehicles Average cargo mass
5 2265.4
7540.0 2740 LTV 3 2200
3800 Cryo-LTV 2 3770
IMLEO or in-turn increasing the mass ratio of equipment and payload. A reusable SEL2 servicing vehicle of about 19 t based at the LSS at LL1 could be feasible requiring less than 10 t storable propellants. One module could be designed as servicing vehicle to synergistically sharing with the station its habitable volume,
The presented scenario of a Lunar Space Station at LL1 demonstrates a lunar exploration architecture with a gateway station in lunar vicinity rather than departing from Earth or LEO directly. The results reveal that such a LSS at LL1 is feasible and programmatically attractive. In the current situation of human spaceflight it offers various benefits for Europe and Russia: • Generating experience with human spaceflight beyond LEO. • Enabling demonstration and verification of key technologies in (semi) deep space as preparation for lunar surface, Mars and other missions. • Offering exploration of the Moon globally and systematically, with a Safe Haven system. • Enabling crewed missions to other near Earth destinations such as SEL2 for servicing and maintenance operations. • Allowing for new models for international cooperation, i.e. robust ESA–Russian elements embedded to an international exploration program led by the USA. • Finally, it is feasible and affordable in near-term by using existing knowledge and technology and even hardware.
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When thinking of human space exploration in general and systematic exploration of the Moon in specific, it is absolutely reasonable to take mission scenarios into account that involve crewed cis-lunar space infrastructures. Because the Moon is (visually) present in the every day life of all humans, public attention and support is provided, which is crucial for every future space program. This may have the potential to also accelerate the exploration of Mars, especially when we follow an integrated Moon/Mars exploration approach with taking first steps early. Acknowledgement The authors wish to thank Robert Lainé, Jean-François Clervoy and the ESA ATV team in Les Mureaux for the generous support provided. References [1] Aurora Programme, Presented and approved at the Ministerial Conference in Edinburgh, 11/2001, URL: www.esa.int/aurora. [2] G.W. Bush, New Vision for Space Exploration Program – Remarks by the President on U.S. Space Policy, Washington, D.C., NASA Special Publication NP-2004-01-334-HQ, January 2004. [3] D. Stanley et al., NASA’s Exploration Systems Architecture Study (ESAS01), Final Report, NASA-TM-2005-214062, November 2005. [4] Project Constellation, NASA, 2007, URL: http://www.nasa.gov/mission_ pages/constellation. [5] K. Yazdi, Conceptual design and flight simulation of space station mission beyond low earth orbit, Dissertation, Universität Stuttgart, 2006.
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