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Conceptual-level Human Lunar lander design concepts are presented for missions utilizing a single launch of the NASA Ares-V and a single launch of the Ares-I launch vehicles. Single stage, twostage and drop tank lander configurations are discussed, as are methodologies for off-loading cargo to the surface. Results show how propellant, staging location, delta-velocity, crew habitat mass and other choices affect lander payload delivery capability. The results serve as preliminary technical contexts for on-going evaluations. Landing systems for nuclear surface power generation stations are also presented.

This paper presents a partially reusable, two-stage-to-orbit launch vehicle concept for delivering 65 metric ton payloads to low Earth orbit. A winged booster and an expendable core stage each use identical main propulsion, thrust structure, and propellant feed systems. For both stages, four 745,000 lbf thrust, RS-68 oxygen/hydrogen rocket engines, presently used on the Delta-IV launch vehicle, are used. The booster, which does not require reentry thermal protection, ascends to aMach 2.4, 81,000 ft altitude condition, separates from the core stage, and glides back to a runway landing. Sixty-five metric tons in orbit would be enabling to a variety of exploration missions, including very large telescope emplacement missions, robust outer planet missions, and others requiring large payloadmasses or volumes. Use of a reusable booster, designed formaintainability and affordability, would, in combination with an expendable core using common propulsion, allow for significant reductions in recurring costs compared with a totally expendable vehicle with similar payload capabilities.

Conceptual in-space transfer stages, including those using solar electric propulsion and chemical propulsion with aerobraking or aerocapture assist at Mars, were evaluated. Roundtrip Mars sample return mission vehicles were analyzed to determine how technology selections influence payload delivery capability. Results show how specific engine, thruster, propellant, capture mode, trip time, and launch-vehicle technology choices would contribute to increasing payload delivered to Mars or decreasing the size of the required launch vehicles. Low-thrust trajectory analyses for solar-electric transfer were generated with the SEPTOP code.

Performance figures are presented for reusable, winged rocket stages launched from several large transport aircraft, including the Boeing 747, the Russian Anotonov An-226, a large supersonic aircraft comparable to the XB-70 aircraft, which achieved Mach 3.1 flight with conventional turbojet propulsion in 1964, and a Mach 6 conceptual aircraft based on the MA145-XAB ramjet demonstrated in 1968. Advantages of air launch include a reduced ascent-to-orbit delta-velocity, reduced drag, the capability to launch at any latitude, and simplified abort modes. Launch from the XB-70 flight condition would allow a 22% reduction in delta-velocity required of the rocket and a significant reduction in drag incurred by rocket flight beginning at an atmospheric density 1/20th that of sea level. Launch from a Mach 6, 85,000-ft altitude condition would allow a 33% reduction in delta-velocity; ar eusable rocket of 262,000 lb launched at this condition could deliver a 23,500-lb payload to orbit utilizing a 523-s vacuum specific impulse advanced rocket engine. Fluorine/lithium‐hydrogen engines achieved 523 s in U.S. Air Force and NASA development programs of the 1960s and 1970s. Comparisons to ground-launched, all-rocket vehicles delivering equivalent payloads to orbit are presented. Fluorine propellant reactivity and engine development history are also discussed.

Four approaches are presented to cone guring large, all-rocket, two-stage-to-orbit launch vehicles for delivering heavy payloads to low Earth orbit. The payload capability for each cone guration is 176,000 lb launched due east into a 220-n mile circular orbit. This capability was chosen so as to be comparable to the payload capability of the shuttle-derived Magnum expendable heavy lift launch vehicle concept characterized by NASA. In the e rst option presented, a reusable, winged stage is used to boost a parallel burn, expendable core stage to low Earth orbit. For the second option, both the booster and orbiter are winged, fully reusable stages. The third option differs from the second only in the fuel type used by the booster stage. The fourth cone guration differs in that the winged, reusable orbiter utilizes an expendablepayload shroud, rather than an internal payload bay. Its forward shroud is mounted to the orbiter’ s nose structure via a special structural attachment. Results indicate that a heavy-lift Earth-to-orbit strategy should consider using a shared reusable booster. The booster should be a common component of another two-stage-to-orbit launch system, to boost smaller core stages. This would allow recouping the boosters’ initial nonrecurring costs across a larger number of missions and programs.

The Space Launch System (SLS) Program is making progress toward delivering a new capability for Beyond Earth Orbit (BEO) exploration. Developed with the goals of safety, affordability and sustainability in mind, SLS will start with 10% more thrust than the Saturn V that launched astronauts to the Moon 40 years ago. From there it will evolve into the most powerful launch vehicle ever flown, via an upgrade approach that will provide building blocks for future space exploration. The SLS provides a critical heavy-lift launch capability enabling diverse deep space missions. In this paper SLS capabilities are described.

The human Mars Mission architectures recently under study are compared. The evolution of mission modes utilizing indigenous resources for propellant production is described. A historical summary of human Mars architecture design is given as a way of introducing some of the challenges inherent to mission planning. A Mars architecture option based on minimum development costs and risks, rather than minimized mass in low Earth orbit, is presented. A multiyearMars vehicle set, characterized by chemical propulsion transfer stages (with aerocapture forMars orbit capture) is described. It is maintained that, for the current Mars Design Reference Mission, the development of more costly advanced propulsion technologies (such as nuclear thermal propulsion, nuclear electric propulsion, or solar electric propulsion) can be considered as later, evolutionarysteps from this initial, less costly, all-chemical approach. Emphasis is given to the minimization of new technology development programs and the associated impacts on total Mars mission costs. Analysis is based on the availability of an 80-metric-tonpayload-class launch vehicle. An important ground rule is the preclusion of on-orbit assembly, allowing only for the joining of complete and fully fueled transfer stages in orbit.

The majority of NASA manned Mars exploration studies undertaken over the past three decades have advocated high-performance nuclear thermal propulsion systems as a means of reducing spacecraft mass and ameliorating the high Earth-to-orbit launch costs associated with these missions. In these studies the use of nuclear propulsion has been uniformly limited to the Earth-Mars interplanetary transfer maneuvers, however. Recent interest in smaller, higher-thrust-to-weight engines coupled with advances in nuclear fuel material capability have prompted consideration of these engines for other applications. This investigation presents an evaluation of the utilization of a small nuclear propulsion system for the Mars ascent burn. The performance of a nuclear ascent-stage concept is compared with several chemical propulsion concepts, including those benefiting from Mars atmospheric in situ propellant production. Propellant requirements, boiloff losses, mission delta velocities, and other parameters relevant to a piloted Mars mission are discussed. Also, a unique low-initial-mass mission trajectory suited for Mars base crew rotation, especially benefited by ascent systems utilizing high-specific-impulse nuclear systems or in situ propellant production, is described.

Future space transportation architectures and designs must be affordable. Consequently, their Life Cycle Cost (LCC) must be controlled. For the LCC to be controlled, it is necessary to identify all the requirements and elements of the architecture at the beginning of the concept phase. Controlling LCC requires the establishment of the major operational cost drivers. Two of these major cost drivers are reliability and maintainability, in other words, the system's availability (responsiveness). Potential reasons that may drive the inherent availability requirement are the need to control the number of unique parts and the spare parts required to support the transportation system's operation. For more typical space transportation systems used to place satellites in space, the productivity of the system will drive the launch cost. This system productivity is the resultant output of the system availability. Availability is equal to the mean uptime divided by the sum of the mean uptime plus the mean downtime. Since many operational factors cannot be projected early in the definition phase, the focus will be on inherent availability which is equal to the mean time between a failure (MTBF) divided by the MTBF plus the mean time to repair (MTTR) the system. The MTBF is a function of reliability or the expected frequency of failures. When the system experiences failures the result is added operational flow time, parts consumption, and increased labor with an impact to responsiveness resulting in increased LCC. The other function of availability is the MTTR, or maintainability. In other words, how accessible is the failed hardware that requires replacement and what operational functions are required before and after change-out to make the system operable. This paper will describe how the MTTR can be equated to additional labor, additional operational flow time, and additional structural access capability, all of which drive up the LCC. A methodology will be presented that provides the decision makers with the understanding necessary to place constraints on the design definition. This methodology for the major drivers will determine the inherent availability, safety, reliability, maintainability, and the life cycle cost of the fielded system. This methodology will focus on the achievement of an affordable, responsive space transportation system. It is the intent of this paper to not only provide the visibility of the relationships of these major attribute drivers (variables) to each other and the resultant system inherent availability, but also to provide the capability to bound the variables, thus providing the insight required to control the system's engineering solution. An example of this visibility is the need to provide integration of similar discipline functions to allow control of the total parts count of the space transportation system. Also, selecting a reliability requirement will place a constraint on parts count to achieve a given inherent availability requirement, or require accepting a larger parts count with the resulting higher individual part reliability requirements. This paper will provide an understanding of the relationship of mean repair time (mean downtime) to maintainability (accessibility for repair), and both mean time between failure (reliability of hardware) and the system inherent availability.

Cost control must be implemented through the establishment of requirements and controlled continually by managing to these requirements. Cost control of the non-recurring side of life cycle cost has traditionally been implemented in both commercial and government programs. The government uses the budget process to implement this control. The commercial approach is to use a similar process of allocating the non-recurring cost to major elements of the program. This type of control generally manages through a work breakdown structure (WBS) by defining the major elements of the program. If the cost control is to be applied across the entire program life cycle cost (LCC), the approach must be addressed very differently. A functional breakdown structure (FBS) is defined and recommended. Use of a FBS provides the visibifity to allow the choice of an integrated solution reducing the cost of providing many different elements of like function. The different functional solutions that drive the hardware logistics, quantity of documentation, operational labor, reliability and maintainability balance, and total integration of the entire system from DDT&E through the life of the program must be fully defined, compared, and final decisions made among these competing solutions. The major drivers of recurring cost have been identified and are presented and discussed. The LCC requirements must be established and flowed down to provide control of LCC. This LCC control will require a structured rigid process similar to the one traditionally used to control weight/performance for space transportation systems throughout the entire program. It has been demonstrated over the last 30 years that without a firm requirement and methodically structured cost control, it is unlikely that affordable and sustainable space transportation system LCC will be achieved.

In 2008 the Boeing company conducted a six month design study for the NASA Johnson Space Center as part of the of the Altair Lunar lander concept desi gn effort. Information presented here is taken from Boeings final report presented to NASA in October 2008 in Houston to the Altair program office. This report contains a summary of one aspect of that stu dy; its objective is to present propulsion system t rade studies and configuration options for the Altair As cent Module. Vehicle subsystem concept design sought to maximize crew safety and vehicle reliability whi le staying within the mass and volume envelope provided by the Ares-V launch vehicle. Tradable ele ments in the Ascent module design included main engine location and number (single vs several, exte rnal vs embedded, bottom vs side placement); propellant type (hypergolic N2O4/MMH vs LO2/LCH4); Ascent crew cabin layout (horizontal or vertical cabin alignment) and several other items.

[Abstract] Boeing performed lunar lander design trade studies during 2008 to support the NASA Altair concept definition effort. The goal of these trade studies was to optimize configuration layouts and figures of merit (FOMs) for the Ascent Module (AM), Descent Module (DM), structural concept, and the integrated vehicle. Multiple AM concepts were developed. Four alternate lander DM configuration concepts were developed, each having different basic structural concepts and propellant tank configurations. These trade studies, presented at the AIAA Space 2008 conference, produced a wealth of useful design data and provided significant support to NASA’s Altair lunar lander concept definition effort. They also pointed out the desirability of performing additional design optimization to enhance the ability of lander DM structure and systems to be scavenged on the lunar surface and reused by future crews, as potential sources of spare parts and hardware for construction of lunar outposts. Reusability of lander DM structure and systems has significant potential to reduce the amount of mass that must be transported from earth for this purpose. Additional work was performed during 2009 to optimize two of the alternate Altair DM configurations that have potential for reuse of a significant portion of structure and systems: The DM with a single LH2 tank and four LO2 tanks, the DM with single LH2 and LO2 tanks separated by a common bulkhead. A third DM design, having single, concentric toroidal LH2 and LO2 tanks, was identified as having the potential for significant parts reuse due to its geometry and placement of subsystems. The purpose of this paper is to present these three lander designs, with design assumptions, requirements, and configuration drawings, and compare their relative Figures of Merit (FOMs) for reusability of DM structure and systems with the Boeing Reference Space Truss configuration. It is anticipated that these results will provide significant support to NASA’s ongoing Altair lunar lander concept definition effort.

[Abstract] This paper describes lunar lander design trade studies performed by Boeing using the Altair concept mission definitions. Complex and multi-layered design interrelationships surfaced during this effort. Trade spaces were developed for the Ascent Module (AM) and Descent Module (DM) to address these interrelationships. Figures of merit (FOMs) were developed to rate the trade study candidates in various performance categories. The trade study discussion provides the rationale and methodology for the trade studies, including the set of possible trade candidates, the candidates that were excluded, and a description of the FOMs. The AM and DM trade studies were performed to optimize the layout and FOMs of candidate AM and DM designs. Promising AM and DM concepts were rated side-by-side using the FOMs. The DM structural concept trade study provided inputs to FOM ratings for candidate DM concepts and guidance and design rules for the layout of integrated lunar lander vehicles. Trade study results were incorporated into four integrated lunar lander vehicle concepts: A lander with four liquid oxygen (LO2) tanks and four liquid hydrogen (LH2) tanks, a lander with a single LO2 tank and four LH2 tanks, a lander with a single LH2 tank and four LO2 tanks, and a lander with single LH2 and LO2 tanks separated by a common bulkhead. Structural concepts for the four lander vehicles range from a full space truss for the eight-tank design to a full shell for the two-tank design, with a mixture of truss and shell elements for the two five-tank designs. The four lander vehicles incorporate two promising candidates from the AM trade study: An AM with a single, central, recessed engine, and an AM with six, smaller, side-mounted engines. Each lander design is presented with design assumptions, a configuration description, drawings, and a summary of design advantages and disadvantages. The Boeing lunar lander trade studies have produced a wealth of useful design data and four viable, alternative lunar lander vehicle concepts.

Conceptual in-space transfer stages, including those utilizing solar electric propulsion, chemical propulsion, and chemical propulsion with aerobraking or aerocapture assist at Mars, were evaluated. Roundtrip Mars sample return mission vehicles were analyzed to determine how specific system technology selections influence payload delivery capability. Results show how specific engine, thruster, propellant, capture mode, trip time and launch vehicle technology choices would contribute to increasing payload or decreasing the size of the required launch vehicles. Heliocentric low-thrust trajectory analyses for Solar Electric Transfer were generated with the SEPTOP code.

A high Mach number, high altitude aircraft, utilized as a launch platform for a reusable air-launched rocket, can provide several benefits for access to low Earth orbit. Performance figures are presented for reusable, winged rocket stages launched from a large supersonic aircraft comparable to the large XB-70 aircraft. The XB-70 achieved mach 3.1, 73,000 ft altitude flight with conventional turbojet propulsion in 1964. Advantages of air-launch from this condition include a 23,644 ft/s ascent-to-orbit ideal delta-velocity (a 22% reduction compared to a ground launch), reduced drag incurred by rocket flight beginning at an atmospheric density one-twentieth that of sea level, and the capability to launch at any latitude. A reusable rocket of 343,000 lbs launched from an appropriated sized aircraft could deliver a 23,500 lb externally mounted payload pod to low Earth orbit utilizing a 523 sec Isp advanced rocket engine. Fluorine/lithiumhydrogen engines achieved 523 sec in USAF and NASA development programs of the 1960’s and 1970’s, an increase of 70 sec Isp over that of the O2/H2 SSME. Together, the 22% delta-V and 70 sec Isp advantages would enable significant reductions in rocket weight compared to a ground-launched all rocket vehicle delivering the same payload to low Earth orbit. Also presented are data for a similarly configured 465 sec Isp O2/H2 air-launch rocket. Fluorine reactivity pasivation and engine development history is also discussed.

Systems analyses were performed for a Titan Explorer Mission characterized by Earth‐Saturn transfer stages using solar electric power generation and propulsion systems for primary interplanetary propulsion, as well as chemical propulsion for capture at Titan. An examination of a range of system factors was performed to determine their effect on the payload delivery capability to Titan. The effect of varying launch vehicle type, solar array power level, ion thruster number, specific impulse, trip time, and Titan capture stage chemical propellant choice was investigated. The major purpose was to demonstrate the efficacy of applying advanced ion propulsion system technologies like NASA’s evolutionary xenon thruster (NEXT), coupled with state-of-the-art (SOA) and advanced chemical technologies to a Flagship class mission. This study demonstrated that a NASA design reference mission (DRM) payload of 406 kg could be successfully delivered to Titan using the baseline advanced ion propulsion system in conjunction with SOA chemical propulsion for Titan capture. In addition, the solar electric propulsion system (SEPS)/chemical system of this study is compared to an all-chemical NASA DRM. Results showed that the NEXT-based SEPS/chemical system was able to deliver the required payload to Titan in 5 years less transfer time and on a smaller launch vehicle than the SOA chemical option.

ElectroDynamic (ED) Tethers can generate continuous low thrust in a low Earth orbit. An induced current running through the length of the tether reacts with the geomagnetic field to produce thrust. The amount of thrust scales with tether lens!th and current. The International Space Station (ISS) requires periodic reboost to maintain an approximately circular orbit t above the Earth. The baseline reboost method is a traditional bi-propellant rocket thruster and tankage system which must to be refueled via Soyuz / Progress or other launch vehicle. The estimated propellant costs associated with keeping ISS in the designated orbit over a 10-year life have been extremely high. The ED Tether would draw energy from the renewable ISS Solar Array electrical power system. Propulsion requirements for ISS vary depending on solar wind and other conditions. It is projected that a ED Tether could provide the majority of the required reboost thrust for ISS for a nominal solar year. For above nominal solar wind years the ISS would have to use the rocket reboost system, but at a greatly reduced level. Thus resulting in substantial cost savings, via the reduction in the number of Earth-to-orbit launch vehicle flights to the ISS that must bring reboost propellant. However, the purposes of this paper is to further Previous research on an ISS ED Tether and examine the operational and technical issues working against using a ED Tether on ISS. Issues such as Shuttle rendezvous and flight path concerns raise serious safety concerns and restrictions on tether use. Tether issues such as tether librations and off angle thrust raise concerns about impacts to microgravity payloads and the long-term effect on ISS orbital path and inclination. Operational issues such as peak power available to an ED Tether and allowable duty cycle may impose severe restrictions on tether design and ultimately limit the practicality of an ED Tether on ISS. Thus while at first glance the cost numbers appear to be strongly in favor of an ED Tether the limitations imposed by safety, operations and technical concerns may severely undermine the economic model. Possible Solutions to these problems have been investigated and proposed, however some items like off angle thrust are still being actively investigated for an adequate solution.

Advanced solar electric power systems and electric propulsion technology constitute viable elements for conducting human Mars transfer missions that are roughly comparable in performance to similar missions utilizing alternative high thrust systems, with the one exception being their inability to achieve short Earth-Mars trip times. A modest solar electric propulsion human Mars scenario is presented that features the use of conjunction class trajectories in concert with pre-emplacement of surface assets that can be used in a series of visits to Mars. Major elements of the Mars solar electric transfer vehicle can be direct derivatives of present state-of-the-art Solar array and electric thruster systems. During the study, several elements affecting system performance were evaluated, including varying Earth orbit altitude for departure, recapturing the transfer stage at Earth for reuse, varying power system mass-to-power ratio, and assessing solar array degradation on performance induced by Van Allen belt passage. Comparisons are made to chemical propulsion and nuclear thermal propulsion Mars vehicles carrying similar payloads.

Space solar power satellites have the potential to provide abundant quantities of electricity for use on Earth. One concept, the Sun Tower, can be assembled in geostationary orbit from pieces transferred from Earth. The cost of transportation is one of the major hurdles to space solar power. This study found that autonomous solar-electric transfer is a good choice for the transportation from LEO to GEO.

This paper will briefly identify a promising fusion plasma power source, which when coupled with a promising electric thruster technology would provide for an efficient interplanetary transfer craft suitable to a 4 year round trip mission to the Jovian system. An advanced, nearly radiation free Inertial Electrostatic Confinement scheme for containing fusion plasma was judged as offering potential for delivering the performance and operational benefits needed for such high energy human expedition missions, without requiring heavy superconducting magnets for containment of the fusion plasma. Once the Jovian transfer stage has matched the heliocentric velocity of Jupiter, the energy requirements for excursions to its outer satellites (Callisto, Ganymede and Europa) by smaller excursion craft are not prohibitive. The overall propulsion, power and thruster system is briefly described and a preliminary vehicle mass statement is presented.

This report investigates various methods of reducing the cost in space transportation systems for human Mars missions. The reference mission for this task is a mission currently under study at NASA. called the Mars Design Reference Mission, characterized by In-Situ propellant production at Mars. This study mainly consists of comparative evaluations to the reference mission with a view to selecting strategies that would reduce the cost of the Mars program as a whole. One of the objectives is to understand the implications of certain Mars architectures, mission modes, vehicle configurations, and potentials for vehicle reusability. The evaluations start with year 2011-2014 conjunction missions which were characterized by their abort-to-the-surface mission abort philosophy. Variations within this mission architecture, as well as outside the set to other architectures (not predicated on an abort to surface philosophy) were evaluated. Specific emphasis has been placed on identifying and assessing overall mission risk. Impacts that Mars mission vehicles might place upon the Space Station, if it were to be used as an assembly or operations base, were also discussed. Because of the short duration of this study only on a few propulsion elements were addressed (nuclear thermal, cryogenic oxygen-hydrogen, cryogenic oxygen-methane, and aerocapture). Primary ground rules and assumptions were taken from NASA material used in Marshall Space Flight Center's own assessment done in 1997.

Future Lunar exploration plans will call for delivery of significant mounts or cargo to provide for crew habitation, surface tansportation, and scientific exploration activities. Minimization of costly surface based infrastructure is in large part directly related to the design of the cargo delivery/landing craft. This study focused on evaluating Lunar lander concepts from a logistics oriented perspective, and outlines the approach used in the development of a preferred configuration, sets forth the benefits derived from its utilization and describes the missions and system considered. Results indicate that only direct-to-surface downloading of payloads provides for unassisted cargo removal operations imperative to efficient and low risk site buildup, including the emplacement of Space Station derivative surface habitat modules, immediate cargo jettison for both descent abort and emergency surface ascent essential to piloted missions carrying cargo, and short habitat egress/ingress paths necessary to productive surface work tours for crew members carrying hand held experiments, tools and other bulky articles. Accommodating cargo in a position underneath the vehicles structural frame, landing craft described herein eliminate altogether the necessity for dedicated surface based off-loading vehicles, the operations and maintenance associated with their operation, and the precipitous ladder climbs to and from the surface that are inherent to traditional designs. Parametric evaluations illustrate performance and mass variation with respect to mission requirements.