A STUDY OF CRYOGENIC PROPULSIVE STAGES FOR HUMAN EXPLORATION BEYOND LOW EARTH ORBIT

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1 A STUDY OF CRYOGENIC PROPULSIVE STAGES FOR HUMAN EXPLORATION BEYOND LOW EARTH ORBIT Mark Schaffer SpaceWorks Enterprises, Inc., United States, Chauncey Wenner United Launch Alliance, LLC, United States, This paper summarizes a study to investigate the feasibility of an all-chemical, common-element solution for a human exploration architecture of the Moon, Near Earth Asteroids (NEAs), and Mars using proven, near-term, lowrisk propulsion systems and passive boil-off technologies. First, a parametric study was performed to understand the influence of different design variables on the initial mass of in-space Cryogenic Propulsive Stages (CPSs). The parametric study used a simplified sizing model and generic mission architectures. It quantified the degree to which propellant mass fraction is the driving design variable for all missions, boil-off rate is only important for long duration missions, and engine Isp is only a minor factor over the range available to LOX/LH2 engines. Second, using these insights, a set of design approaches for the CPSs was established. To improve propellant mass fraction, primary structures mass can be reduced by launching the CPS to Low Earth Orbit (LEO) as a non-loadbearing payload inside a fairing, and subsystems mass can be reduced by using an innovative Integrated Vehicle Fluids (IVF) technology. To reduce boil-off rates, the tanks can be intelligently designed and shielded with Multi-Layer Insulation (MLI), the number of thermal connection points to the tanks can be minimized, and a deployable sun shield can be included for long duration missions. With these characteristics in mind, conceptual level designs for two CPSs were developed: a ~100t gross mass CPS Alpha and a ~220t gross mass CPS Beta, both with low boil-off block upgrades available using deployable sun shields. An optional ~480t gross mass CPS Gamma was also designed as an alternative Earth departure stage for Mars missions. Using these stages, concepts of operations were developed for missions to the Moon, NEAs, and Mars. Missions departing from LEO and the Earth-Moon L1 point were also compared. Using L1 as a departure point reduces the number of elements required at the start of the mission, but increases the mission complexity and adds additional constraints for missions beyond cis-lunar space. Results presented include details on the masses and design of the CPSs, specific combinations of stages required for each mission, and in-space trajectories required to reach the selected destinations. Launch manifests to deliver the CPSs to LEO were not considered in this study, though efforts were made to constrain the CPS designs such that they would be capable of being launched empty or partially full on existing or proposed near-term launch vehicles. I. INTRODUCTION The Moon, Near-Earth Asteroids (NEAs), and Mars are natural destinations for near-term human exploration missions. Many of the architectures envisioned for these missions rely on advanced technologies for propulsion such as nuclear thermal rockets, or active cryogenic propellant management such as cryocoolers. These technologies often require significant time and resource investment to reach the maturity level required to achieve these missions. As an alternative, the authors performed a study to investigate the feasibility of an architecture solution 1 using only proven, near-term, and low-risk technologies. Only chemical propulsion was considered, with existing or near-term liquid rocket engines. To meet the demanding performance requirements of these missions, cryogenic propellants were selected. Furthermore, only passive, low-tech solutions were considered for the thermal management of the cryogenic propellants, as opposed to the use of active systems. This study was performed in three phases. In the first phase, the sensitivity of the initial mass in orbit of the in-space Cryogenic Propulsive Stages (CPSs) to different design parameters was analyzed. By

2 determining these sensitivities, the importance of each design variable was established quantitatively. In the second phase, using the insights gained from the design parameter sensitivities, innovative approaches and design practices were identified to maximize the performance of the CPSs. With a set of design approaches and philosophies having been established, an architecture solution was developed in the final phase of the study. The number of unique CPS designs was minimized by using common CPSs in series to achieve the range of trajectory performance required for the different missions of interest. When possible, common technologies and systems were selected for each unique CPS design to minimize the development investment required for the architecture. This study was only concerned with the design of inspace propulsive elements. Other architecture elements were assumed to be of fixed mass based on current publically available development objectives for these elements. These elements serve as the payloads carried by the CPSs. II. PAYLOAD DEFINITION Four architecture elements were identified as representative payloads for future human exploration missions and are shown in Figure 1. Masses for each Figure 1. Elements for Human Exploration Architecture (not shown to scale) element were based on a literature search [1-17, 37, 38]. The estimates for each mass are shown in Table 1. Orion / Multi-Purpose Crew Vehicle (MPCV) The MPCV is currently being developed by NASA to support missions beyond Low Earth Orbit (LEO). This vehicle serves as the primary exploration vehicle for local Earth-Moon system missions and provides abort capability and Earth re-entry capability for deep space missions. It can sustain a crew of 4 for 21 days on a deep space mission. Lunar Lander Payload The lunar lander provides crew transportation from Low Lunar Orbit (LLO) to the lunar surface and from the lunar surface back to LLO. It is also capable of supporting a crew of 4 for a 7 day sortie on the lunar surface. Space Exploration Vehicle The Space Exploration Vehicle (SEV) is a small, 2- crew in-space vehicle proposed for close-range exploration of NEA. This vehicle provides mobility and safety in vicinity of unpredictable asteroid surfaces, and can sustain a crew of 2 for up to 28 days. Deep Space Habitat Wet Mass (t) Orion Multi-Purpose Crew Vehicle 21.2 Lunar Lander 30.0 Space Exploration Vehicle 6.5 Deep Space Habitat 1 (180 day) 23.5 Deep Space Habitat 2 (1,000 day) 30.0 Table 1. Architecture Element Masses The Deep Space Habitat (DSH) provides in-space habitat for the crew for long term missions outside of the Earth-Moon system. Two versions of the DSH were considered: DSH-1 and DSH-2. DSH-1 is capable of sustaining a crew of 4 for 180 days, suitable for NEA missions. DSH-2 can support a crew of 4 for 1,000 days as required for Mars missions. 2

3 III. MISSION DEFINITION Three missions were identified as representative future human exploration missions. These missions are as follows: 1. A round trip mission to the lunar surface using a single CPS to perform Trans-Lunar Insertion (TLI) and Lunar Orbit Insertion (LOI), a lunar lander for access to the lunar surface from LLO, and the MPCV for in-space crew habitation and Earth entry and recovery. The MPCV performs Trans-Earth Injection (TEI). 2. A round trip mission to a NEA using multiple CPSs to perform the Earth departure, NEA arrival, and NEA departure maneuvers. This mission also includes a DSH-1 for in-space crew habitation, SEV for near-asteroid operations, and MPCV for Earth entry and recovery. 3. A round trip mission to Mars orbit using multiple CPSs. This mission includes a DSH-2 for inspace crew habitation and MPCV for Earth entry and recovery. Mars surface access can be achieved through a second cargo mission using the same propulsive elements. The cargo mission was not considered in this study. All missions are assumed to end with direct Earth entry using the MPCV. For each mission, a concept of operations was established and required ideal ΔVs for each maneuver were determined. An outline of these concepts of operations is shown in Figure 2. Departures from Low Earth Orbit (LEO) and the Earth-Moon Langrangian point 1 (E-M L1) were considered. For the case of the deep space missions to a NEA or Mars, windows of opportunity for launching to these targets in the timeframe were also determined. The required ideal ΔVs for each mission were analyzed using internal tools and programs. These ΔVs were then cross-referenced with published values. This process is detailed in this section for each mission. A summary of the required ideal ΔVs for each maneuver for each mission are shown in Table 2. These ΔV values represent ideal ΔVs applied instantaneously. Lunar Mission Time of flight from LEO to LLO was assumed to be Figure 2. Concept of Operations for Moon, NEA, and Mars Missions 3

4 Destination Moon NEA Mars Origin Origin Departure ΔV (m/s) Destination Arrival ΔV (m/s) Destination Departure ΔV (m/s) LEO 3, ,050 E-M L ,050 LEO 3,350 2,000 2,150 E-M L ,000 2,150 LEO 4,400 2,200 2,550 E-M L1 1,770 2,200 2,550 Outbound Time of Flight (days) Destination Stay Time (days) Return Time of Flight (days) Table 2. Mission Required ΔVs and Times of Flight in the range of 3-4 days. To achieve this range, the ΔVs required for a 3 day transfer time were assumed as a conservative estimate. Similarly, boil-off propellants lost in transit were assumed to be lost over a 4 day time period, again representing a conservative estimate. An internal Earth-Moon patched conic trajectory tool was used to determine ideal required Trans-Lunar Injection (TLI), Lunar Orbit Insertion (LOI), and Trans-Earth Insertion (TEI) ΔVs as a function of time of flight. A literature search was then performed to find published values for these ΔVs for comparison [7, 18, 19, 24, 25]. The ΔVs from the Apollo missions were also collected [30-35]. The results of this analysis and research are shown in Table 2. These ΔVs represent the ΔV required for a 72 hour time of flight from LEO to LLO. A plane change maneuver to adjust for the inclination of the Moon s orbit with respect to the Earth was not considered in this study. NEA Mission The time of flight from the Earth to a NEA was assumed to be in the range of days for the different NEA mission opportunities available, with stay times of 30 days and total mission times not exceeding 180 days. Boil-off propellants lost in transit were assumed to be lost over a 120 day time period regardless of NEA, representing a conservative estimate. To determine the required mission times and ΔVs for this mission, 27 candidate NEAs for human missions were identified that have been considered in previous studies [2,26,27,28]. For expediency, this list was narrowed to 6 candidate NEAs based on launch date and ΔV requirements that have launch windows in the timeframe. The down-selected list of NEAs is shown in Table 3. The ΔVs found in literature were independently verified using Bullseye, a SpaceWorks Software commercially available interplanetary trajectory tool, NEO Launch Date ΔV From Literature (m/s) Earth Post Escape Depart 4 LEO Depart Calculated ΔV (m/s) E-M L1 Depart NEA Arrival NEA Depart 2007 UN12 5/22/2020 3,300 1,450 3, QJ142 4/24/2024 3,490 3,400 3, ,220 1, AO10 9/19/2025 3,320 3,740 3, ,020 2, LN6 12/21/2025 3,330 3,690 3, ,230 2, SG344 4/27/2028 3,340 3,220 3, , UQ216 8/15/2028 3,710 3,550 3, ,800 1,790 Table 3. NEA Mission Opportunities Summary and Selected Mission ΔVs

5 based on ephemeris data from the JPL HORIZONS [36] online database. The NEA ephemeris used in Bullseye to generate these results was the latest available and may be different than the ephemeris used in the literature references. The results of this independent analysis are shown in Table 3. Based on this analysis, a required design ΔV of 3,350 m/s was selected for the Earth departure, while ΔVs of 2,000 m/s and 2,150 m/s were selected for the NEA arrival and departure ΔVs respectively. Based on these assumptions, five of the six selected NEAs would be accessible with the mission architecture described in this study; a mission to 2003 LN6 would require significantly more energy. The largest value possible was selected for the final maneuver of the mission, the NEA departure maneuver. This represents a conservative assumption for the purposes of calculating boil-off losses. selected ΔVs shown in Table 2 allow for all four 30 day launch windows for Earth departure and Mars departure in the timeframe. Though the split between Mars arrival and Mars departure ΔVs changes with each mission, the sum of these two ΔVs stays relatively constant. Both Mars ΔVs are performed by a single CPS in this study. The time of flight from the Earth to Mars was assumed to be in the range of days for the different conjunction-class Mars mission opportunities available, with stay times up to 550 days. The 550 day stay time coincides with a 200 day outbound travel time, for a total of 750 days combined outbound and stay time; this represents the longest total mission time for the launch windows considered. For boil-off calculations, a 200 days outbound and 550 days Mars stay time were assumed for conservatism. Mars Mission Mars missions are typically classified as short duration (one year or less), opposition, or conjunction class missions. Because they represent the lowest total mission ΔVs with chemical propulsion systems, a conjunction class mission was chosen for this study. A sparse sweep of conjunction class Earth-to-Mars trajectories in the timeframe was performed using SpaceWorks Software s Bullseye. This sweep was used to identify four mission opportunities in this timeframe for more detailed analysis. These missions were subject to a maximum time-of-flight constraint of 250 days. Within each opportunity, a 30 day launch window was identified. These 30 day windows were used to determine the required mission ΔVs. A summary of the launch windows and required ΔVs for each opportunity are shown in Table 4. The Data Point Nominal Launch Date LEO Departure ΔV (m/s) E-M L1 Departure ΔV (m/s) 5 Earth re-entry velocities vary from 11.5 to 12.4 km/s. This represents at maximum a 10% increase over entry velocities seen during a typical Apollo mission [32]. Design Variables IV. SENSITIVITY ANALYSIS Three design variables were selected for this analysis: Propellant Mass Fraction (PMF), engine specific impulse (Isp), and propellant boil-off rate. These variables represent the major top-level vehicle design drivers that will affect the performance of the CPSs. An outline of these variables and the values considered for each is shown in Table 5. Engineering experience was used to select ranges of interest for these variables. PMF is defined as the total usable propellant mass plus boil-off propellants lost during the mission Mars Arrival ΔV (m/s) Mars Departure ΔV(m/s) Total Mars ΔV (m/s) Opportunity 1 7/9/2020 4,000 1,320 2,200 2,550 4,750 Opportunity 2 9/2/2022 4,400 1,770 2,500 2,300 4,700 Opportunity 3 10/11/2024 4,400 1,770 2,550 2,150 4,650 Opportunity 4 11/10/2026 4,250 1,570 2,500 2,050 4,550 Table 4. Mars Mission Opportunities Summary and Selected Mission ΔVs

6 divided by the total stage wet mass. Boil-off rates are measured as a percent of fully fueled propellant capacity lost per day. A fixed mass loss rate is calculated at the beginning of the mission and held constant throughout the mission. It is assumed that the boil-off propellant mass generated before any particular maneuver is vented before that maneuver occurs, and that this is a combined loss of fuel and oxidizer. Mission Scenario and Analysis Tool Mission Scenario and Analysis Tool (MSAT) modeling was developed in 2004 by SpaceWorks as an architecture-level modeling methodology that combines various disciplinary design models into one overarching simulation. MSAT models can be used to measure architecture-level impacts of design variables on multi-disciplinary figures of merit including weights, life cycle costs, and reliability metrics. Using optimization software, the architecture can be optimized to any of the figures of merit, for example the lowest cost mission design. The MSAT model created for the parametric study was built in Microsoft Excel to determine the masses of the CPSs required. Stage inert mass, usable propellant mass, and boil-off propellant mass are generated with the model using the three design variables and the rocket equation in an iterative process. Analysis The parametric sizing model was executed and the CPS stage fully sized for all combinations of PMF, engine Isp, and propellant boil-off rate for each candidate mission. The sheer number of combinations of design variable values considered produces more results than can be reasonably addressed in this paper. Instead, a reference point was selected for each mission and the sensitivity of the design to each design variable is shown based on that reference point. In this sensitivity analysis, ΔV and payload are held constant, and the CPS is rescaled along each trend line. Inert mass represents the sum of the dry mass and any unused propellants (residuals and reserves) as well as other fluids and consumables. Value Description Propellant Mass Fraction Human Exploration Framework Team 0.75 (HEFT) assumed low-end CPS mass fraction 0.80 Conservative value Ares V EDS-like (Earth Departure 0.85 Stage) mass fraction 0.90 Centaur-like mass fraction 0.95 High-end possible mass fraction Engine Specific Impulse (Isp) 448 sec J-2X 451 sec RL10-A sec RL10-B2 or Next Generation Engine Boil-off Rate 0.001%/day 0.01%/day 0.05%/day 0.1%/day Though sensitivities were performed using both LEO and E-M L1 as departure points, only the LEO departure solutions are shown here. The sensitivities are comparable for E-M L1 departure. Lunar Mission Results Requires active cooling Aggressive boil-off rate with passive thermal protection Reasonable near-term boil-off rate with passive thermal protection Centaur boil-off rate achievable via already reviewed modifications Table 5. Parametric Study Design Variables For the lunar mission, a single CPS was sized to perform TLI and LOI. The MPCV is assumed to perform TEI. The impact of all three design variables is shown in Figure 3. Each line represents a departure from the reference point by varying the considered design variable while holding the other design variables constant at the reference point value. The vehicle mass breakdown for the lunar mission reference case is shown in Table 6 based on a PMF of 0.90, an engine Isp of 451s, and a boil-off rate of 0.05%/day. Because the assumed transit time for boil-off calculations is only 4 days, the boil-off propellant is small regardless of boil-off rate. 6

7 Item Mass (t) Payload 51.2 Inert 10.4 Usable Propellant 93.0 Boil-Off Propellant 0.2 Table 6. Mass Breakdown for Lunar Mission Reference Point Figure 3. Impact of Design Variables on Initial Mass for Lunar Mission PMF has the largest impact on the total stage size; small improvements in propellant mass fraction can yield significant reductions in initial mass for the ranges considered. By comparison, even significant improvements in Isp will only yield modest reductions in initial mass. Boil-off rate has very little impact on the overall vehicle size over the range considered due to the short mission transit time. NEA Mission Results For the NEA mission, two CPSs were sized and operated in series. CPS 1 was assumed to perform the Earth departure maneuver; CPS 2 was assumed to perform the NEA arrival and departure maneuvers. The impact of all three design variables is shown in Figure 4. Each line represents a departure from the reference point by varying the considered design variable while holding the other design variables constant at the reference point value. The solid line represents the impact of CPS 1, while the dashed line represents the impact of CPS 2. The vehicle mass breakdown for the NEA mission reference case is shown in Table 7 based on a PMF of 0.90, an engine Isp of 451s, and a boil-off rate of 0.05%/day for both stages. With an assumed transit time of 120 days and NEA stay time of 30 days for boil-off calculations, the boil-off propellants lost in transit become a significant fraction of the total mass of CPS 2. Because CPS 1 is completely expended in the first maneuver, there is no boil-off propellant on this stage. Even with two CPS stages and longer mission times, PMF still has the largest impact on the total stage size. Boil-off rate has only a moderate impact on the Figure 4. Impact of Design Variables on Initial Mass for NEA Mission 7

8 overall vehicle size. As expected from the rocket equation, the sensitivity of the design to changes on the second stage is stronger than the sensitivity to changes on the first stage. Mars Mission Results For the Mars mission, two CPSs were sized and expended in series. CPS 1 was assumed to perform the Earth departure maneuver; CPS 2 was assumed to perform the Mars arrival and departure maneuvers. A representative example of a Mars mission vehicle mass breakdown is shown in Table 8 based on a PMF of 0.90, an engine Isp of 451s, and a boil-off rate of 0.05%/day for both stages. With assumed transit time of 200 days and Mars stay time of 550 days for boiloff calculations, the boil-off propellants lost in transit become a significant fraction of the total mass. Because CPS 1 is completely expended in the first maneuver, there are no propellants lost to boil-off on this stage. The impact of all three design variables is shown in Figure 5. Each line represents a departure from the reference point by varying the considered design variable while holding the other design variables constant at the reference point value. The solid line represents the impact of CPS 1, while the dashed line represents the impact of CPS 2. Unlike the other missions, the propellant mass fraction and boil-off rate both have a significant impact on the total stage size, because of the long mission times for the Mars mission. Conservative estimates for both of these variables can cause the vehicle to be infeasible. Again, as expected from the rocket equation, the sensitivity of the design to changes on the second stage is stronger than the sensitivity to changes on the first stage. Summary Item Mass (t) Payload 51.2 CPS 1 Inert 24.5 CPS 1 Usable Propellant CPS 1 Boil-Off Propellant 0.0 CPS 2 Inert 11.9 CPS 2 Usable Propellant 99.2 CPS 2 Boil-Off Propellant 8.0 Table 7. Mass Breakdown for NEA Mission Reference Point For the ranges considered, PMF is the more important design variable. Improvements in PMF yield to significant reductions in initial mass. Boil-off rates only become significant for the Mars mission for the ranges considered. Engine Isp is not a significant driver over the range available to likely Figure 5. Impact of Design Variables on Initial Mass for Mars Mission 8

9 LOX/LH2 engines. V. IMPLEMENTATION As seen in the parametric study, an all-chemical architecture solution is possible for missions to the Moon, NEA, and Mars with boil-off rates achievable through passive technologies, as long as the CPS PMF is sufficiently large. Based on this philosophy, a set of design approaches and innovation solutions were proposed for use in the full architecture design. These approaches to structures, subsystem design, equipment layout, and thermal control design meet the goal of improving PMF while achieving low boil-off rates passively. In addition, propulsion systems were selected to satisfy the goal of using proven, near-term, low-risk propulsion systems. Structures Item Mass (t) Payload 51.2 CPS 1 Inert 62.3 CPS 1 Usable Propellant CPS 1 Boil-Off Propellant 0.0 CPS 2 Inert 21.5 CPS 2 Usable Propellant CPS 2 Boil-Off Propellant 35.3 Table 8. Mass Breakdown for Mars Mission Reference Point Structures masses can be reduced by reducing launch loads. One approach is to bypass launch loads by housing the CPS inside a fairing. Additionally, launching the CPS partially full (with minimal LOX) further reduces sizing structures loads. Oversizing propulsion for CPS drives unnecessary loads through the stage, also increasing tank mass. Tailoring the propulsion system to the mission needs can improve structures performance. Subsystems Vehicle subsystems, including power storage and generation, tank pressurization, and attitude control, constitute a significant percentage of vehicle inert mass. These systems often require their own set of specialized of tanks and fluids, driving system masses higher and adding complexity to the vehicle. To streamline vehicle design and reduce subsystem mass, United Launch Alliance is developing Integrated Vehicle Fluids (IVF), an innovative system that uses boil-off gases from cryogenic propellant tanks to provide power generation, tank pressurization, and attitude control and eliminates the need for separate attitude control reactants, pressurant tanks and gases, and electric power generation consumables. [42] A diagram of the IVF system is shown in Figure 6. With IVF, boil-off gases from the main propellant tanks are pressurized and stored in separate, smaller accumulator tanks. Electrical power is used to drive cryogenic positive displacement pumps to pressurize these gases. Tank pressurization, purges, pneumatic power, and the power generation O2 supply are fed from accumulator tanks. Power generation for peak loads and battery recharge is provided by a very small GOX/GH2 Internal Combustion Engine (ICE). The ICE burns at very low oxygen to hydrogen mixture ratio (0.5 to 2.0) and doubles as engine starter motor. It produces electrical power with a generator and charges a rechargeable Choice of tank design (e.g. monocoque, common bulkhead stainless steel construction vs. separated aluminum isogrid tanks) can have a factor of two impact on tank mass. Figure 6. Integrated Vehicle Fluids System 9

10 battery, analogous to a hybrid car. The ICE is cooled regeneratively by incoming cold hydrogen gas from main tank ullage. The IVF system also provides attitude control through 8 small thrusters that can be activated in bipropellant combustion or monopropellant cold gas mode. Bi-propellant combustion mode can be achieved by drawing GH2 and GOX from the accumulator tanks into a combustion chamber. Cold gas mode can be achieved by releasing gas directly from the accumulator tanks. Attitude control reactants and tank pressurants, and dedicated storage vessels, are wholly eliminated with IVF. Thermal Control To reduce boil-off rates passively, the tanks can be completely covered with Multi-Layer Insulation (MLI). MLI is composed of a series of thin layered sheets that significant reduce heat loss through radiation and has been used on many deep space missions for passive thermal control. To reduce heat losses to conduction, the number of structural connection points to the tanks can be minimized. The number of conduction paths into the tank can be reduced by placing all subsystems hardware on an equipment shelf that is integrated into a thermally isolated aft thrust structure. [43] Innovative propellant tank design can further reduce propellant boil-off. Vapor-cooling paths, where vented hydrogen is used to intercept the remaining high-load heat paths, can be integrated into the tank structure. Enhanced vacuum insulation panels can also be built directly into the tank structure. [43] Based on thermal analysis performed by United Launch Alliance, the combination of MLI and innovative tank and structure design can reduce the CPSs average boil-off rates to 0.03%/day of total propellant lost. This is sufficient for short term missions, but long term missions to NEAs and Mars will require further boil-off reductions. For long term missions, United Launch Alliance is developing a lightweight, deployable sun shield, shown in Figure 7. The sun shield is composed of multiple layers forming a concentric cone around the CPS. While in orbit around the Earth or Mars, the CPS maintains a northern ecliptic orientation, which enables the sun shield to shadow the tank from both solar and planetary heating throughout the orbit. The multiple, concentric conical shield layers are maintained at different angles and provide an open path for thermal energy to be directed out into deep space, and away from the cryogenic propellant tank. [44] With the sun shield, internal analysis by United Launch Alliance predicts that the average propellant boil-off losses will be reduced by a factor of two. Based on their estimate of 0.03%/day for the short term CPS design, the long term CPS with sun shield will have a boil-off rate of roughly 0.015%/day per day. It is assumed that these technologies are sufficient to reduce the LOX boil-off rate to essentially zero. In the common bulkhead vehicles, any excess heat in the LOX tank can be transferred directly to the LH2 tank. Therefore only LH2 is assumed to boil-off, though the percentages used are based on total propellant mass including both LH2 and LOX. Propulsion Two candidate LH2/LOX engines were selected to satisfy the requirements of proven, near-term, lowrisk propulsion technology and to provide the range of thrust that will be necessary to accomplish the different missions being considered. These engines are the Next Generation Engine (NGE) and the MB- 60. Performance summaries of these two engines are shown in Table 9. The NGE is a notional cryogenic upper-stage engine Figure 7. Deployable Sun Shield (cutaway) 10

11 Parameter NGE MB-60 Propellants LOX/LH2 LOX/LH2 Vacuum Thrust 30,000 lbf 60,000 lbf Vacuum Isp 465 sec 465 sec Mixture Ratio (O:F) 5.8: :1.0 Thrust-to-Weight Area Ratio Length 9.2 ft 10.8 ft Table 9. NGE and MB-60 Performance design currently in requirements development and planning stages. It is targeted for initial operations by 2017, with a goal to utilize modern design and manufacturing to minimize life cycle cost. Likely suppliers for this engine are Aerojet, Pratt & Whitney Rocketdyne, or XCOR Aerospace. [45] The MB-60 is a joint development between Pratt & Whitney Rocketdyne and Mitsubishi Heavy Industries, Ltd. The MB-60 is a restart-able, high performance, in-space expander cycle LOX/LH2 engine. [46] VI. CPS DESIGN segment of the NEA missions. Block 2, with an average boil-off rate of 0.015%/day achieved by adding a sun shield, is used for the Mars arrival and departure maneuvers. An optional third vehicle, CPS Gamma, is an alternative for the Earth departure segment of the Mars mission. To reduce CPS mass and take full advantage of the exploration elements required, the propulsive element of the MPCV will be used to assist with final Earth return maneuver for all missions. In the case of the lunar mission, this element alone can provide the necessary return ΔV. StageSizer SpaceWorks s StageSizer tool is a conceptual level mass and sizing tool intended for the design and analysis of in-space stages including launch vehicle upper stages, cis-lunar transfer stages, lunar ascent and design stages, and deep space transfer stages. The tool is based on a combination of historical mass estimating relationships, physics-based equations, and empirical data for each major subsystem in the The insights gained from the parametric study and the design philosophies, approaches, and technologies selected were the foundation for developing the CPS designs used in the generation of a full architecture solution. The result was two required unique CPS designs, with one additional, optional CPS design. The CPS vehicle family designed for the full architecture solution is shown in Figure 8 and Table 10. CPS Alpha is required for Moon and NEA missions and has two versions. Block 1, with an average boiloff rate of 0.03%/day, is used for Moon mission and the Earth departure segment of the NEA mission. Block 2, with an average boil-off rate of 0.015%/day achieved by adding a sun shield, is used for the NEA arrival and departure maneuvers. CPS Beta is required Mars missions and optionally available for NEO missions. It also has two versions. Block 1, with an average boil-off rate of 0.03%/day, is used for the Earth departure segment of the Mars missions, and optionally for the Earth departure Figure 8. Proposed CPS Vehicle Family 11

12 CPS Alpha Block 1 CPS Alpha Block 2 CPS Beta Block 1 CPS Beta Block 2 CPS Gamma Max Diameter 5.0 m 5.0 m 6.5 m 6.5 m 8.0 m Total Length 19.2 m 21.2 m 28.5 m 31.0 m 32.1 m Propellant Mass 97.1 t 97.1 t t t t Wet Mass t t t t t Total PMF 92.9% 91.7% 92.8% 91.8% 94.4% Engines 4x NGE 4x NGE 4x MB-60 4x MB-60 5x MB-60 Avg. Boil-Off Rate 0.030%/day 0.015%/day 0.030%/day 0.015%/day 0.030%/day Table 10. CPS Vehicle Characteristics vehicle. The general inputs into the model are mission and level design parameters. StageSizer outputs a multilevel mass breakdown statement that details the contributions of subsystems, margin, propellants, payload, etc., to the total gross mass of the stage. StageSizer also outputs the required propellant volumes and dimensions for tanks, intertanks, interstages, and other structural elements. For the architecture analysis, individual StageSizer models were built for each CPS element. These models were used to size and compare the different CPSs required for each mission. The mass breakdown statements for the five CPS vehicles are shown in Table 11. VII. ARCHITECTURE ANALYSIS For the architecture analysis, the StageSizer models developed for each CPS were integrated into a new, high-fidelity MSAT model. The MSAT model determines the required propellant mass for each Item CPS Alpha Block 1 (kg) CPS Alpha Block 2 (kg) 12 CPS Beta Block 1 (kg) CPS Beta Block 2 (kg) CPS Gamma (kg) Structure 3,480 3,480 7,940 7,940 14,680 Propulsion 1,240 1,240 2,940 2,940 3,740 Thermal Control 450 1, ,860 1,270 Avionics Miscellaneous Margin (30%) 1,720 2,020 3,700 4,290 6,250 Dry Mass 7,450 8,750 16,020 18,610 27,080 Residuals, Reserves, Buildup 1,430 1,430 3,050 3,050 6,680 Consumables / RCS Propellants 0* 0* 0* 0* 0* Inert Mass 8,880 10,180 19,070 21,660 33,760 Usable Fuel 14,020 14,020 29,870 29,870 65,500 Usable Oxidizer 81,160 81, , , ,920 Startup Losses ,020 1,020 2,230 Wet Mass 104, , , , ,420 * IVF allows use of main propellants for pressurization and attitude control Table 11. CPS Vehicle Mass Breakdown Statements

13 mission based on required mission ΔVs and boil-off losses. An optimizer was used to minimize the total mass of each CPS design while achieving the required payload mass for each mission, based on the combination of CPS elements used for that mission. The optimizer can offload propellant from different stages for each mission to improve payload performance. Gravity losses were included in the MSAT trajectory model, estimated by generating a curve fit from representative trajectories flown in the Program to Optimize Simulated Trajectories (POST). These gravity losses were a function of burn time and stage thrust-to-weight. Using this MSAT model, the final CPS designs shown in Figure 8 and specified in Table 10 and Table 11 were developed. Lunar Mission Results The lunar mission results are summarized in Table 12. The fully assembled mated vehicle is shown in Figure 9. The lunar mission is the sizing case for the CPS Alpha propellant requirement, so all of the loaded propellant is required for this mission. With 4 days of boil-off, the tanked O/F ratio is slightly lower than the engine ideal O/F ratio. NEA Mission The NEA mission results are summarized in Table 13. Three mission options were considered. The first two depart from LEO but use a different configuration of CPS vehicles. The third option departs from E-M L1. All options carry a DSH-1, SEV, and MPCV including crew mass. The mated vehicle configurations for the three options are shown Figure 9. Lunar Mission Mated Vehicle Parameter Value Payload Mass 52,080 kg MPCV 21,250 kg Lunar Lander 30,000 kg Crew (4) + Suits 600 kg Payload Adapter 230 kg CPS Version Alpha B1 Fuel Offload 0 kg Oxidizer Offload 0 kg CPS Propellant Mass at Ignition 97,080 kg CPS Wet Mass at Ignition 104,530 kg Total Mass at Ignition 156,610 kg Initial Thrust-to-Weight at Ignition 0.35 Gravity Losses 1.5% Total Burn Time 13.3 min Videal 4,150 m/s Table 12. Lunar Mission Results in Figure 10. In LEO Departure Option 1, a small amount of propellant is offloaded from both the first and second stages. The final stage of the NEA mission suffers significant boil-off losses during its 150 day mission, even with the reduced boil-off rate offered by the Block 2 upgrades to the CPS. 15% of the available oxidizer mass can be offloaded to compensate for the fuel lost to boil-off. The MPCV is required to provide the final 500 m/s of the NEA departure / Earth return maneuver. It is assumed that the two Earth departure burns both occur at perigee. After the first CPS Alpha finishes its burn and is staged, the mated vehicle continues around the Earth for one full high Earth orbit period of roughly 4 hours. The second CPS Alpha then performs its burn. Performing the burns in this method reduces the gravity losses incurred, thereby reducing total mission Videal. In LEO Departure Option 2, CPS Beta shows similar performance in terms of propellant offload and total Initial Mass to Low Earth Orbit (IMLEO) to using two CPS Alphas. Using CPS Beta would provide advantages in reducing the total structural loads by shortening the mated vehicle at mission start and therefore reducing bending moments. 13

14 unchanged from the LEO departure options. Figure 10. NEA Mission Options As with Option 1, a small amount of propellant is offloaded from the first stage. The requirements on the CPS Alpha final stage and MPCV are unchanged from Option 1. Departing from Earth-Moon L1 significantly reduces the Earth departure requirements for this mission. With the current architecture, the CPS Alpha used for Earth departure is mostly offloaded when departing from L1. It might be possible for this stage to be launched to LEO fully loaded, propel itself to L1, and then be used for the Earth departure leg without any refueling. More investigation is necessary. Alternatively, a much smaller propulsive element could be used for Earth departure. The requirements on the CPS Alpha final stage and MPCV are Mars Mission (orbit only) The Mars mission results are summarized in Table 14. Here, as with the NEA missions, three mission options were considered. The first two depart from LEO but use a different configuration of CPS vehicles. The third option departs from E-M L1. All options carry a DSH-2 and MPCV including crew mass. A second cargo leg could be used to bring Mars surface access and exploration elements but was not considered in this study. The mated vehicle configurations for the three options are shown in Figure 11. In LEO Departure Option 1, CPS Beta was sized specifically for the Earth departure segment of this mission. The first and second stages therefore have no propellant offload. The third stage of the mission suffers significant boil-off losses during its 750 day mission, even with the reduced boil-off rate offered by the Block 2 upgrades to the CPS. 38% of the available oxidizer mass can be offloaded to compensate for the fuel lost to boil-off. An additional 9% of the available fuel mass can be offloaded due to the stage being slightly oversized for this leg of the mission. The MPCV is required to provide the final 500 m/s LEO Departure Option 1 14 LEO Departure Option 2 L1 Departure Payload Mass (t) Total Initial Mass (t) Stage 1 Stage 2 Stage 3 Stage 1 Stage 2 Stage 1 Stage 2 Element Alpha 1 Alpha 1 Alpha 2 Beta 1 Alpha 2 Alpha 1 Alpha 2 Avg. Boil-Off Rate (%/day) 0.03% 0.03% 0.015% 0.03% 0.015% 0.03% 0.015% CPS Mass at Ignition (t) Ignition T/W Total Burn Time (min) Videal (m/s) 1,340 2,040 3,650 3,400 3, ,650 Gravity Losses (%) 1.0% 1.0% < 0.1% 1.5% < 0.1% < 0.1% < 0.1% Fuel Offload (%) 5.8% 5.8% 3.6% 6.1% 3.6% 78.1% 3.6% Oxidizer Offload (%) 5.6% 5.6% 15.5% 6.1% 15.5% 78.1% 15.5% Table 13. NEA Mission Results

15 the CPS Beta final stage and MPCV are unchanged from Option 1. Figure 11. Mars Mission Options of the Mars departure / Earth return maneuver. As with the similar NEA mission, it is assumed that the two Earth departure burns both occur at perigee. CPS Gamma was sized specifically for the Earth departure segment of the LEO Departure Option 2 mission, and therefore it requires no propellant offload. Using CPS Gamma would provide advantages in reducing the total structural loads by shortening the mated vehicle at mission start and therefore reducing bending moments. Using one large stage, rather than two smaller stages increases the total propellant mass required as a consequence of the rocket equation. This increases the total IMLEO required compared to Option 1. The requirements on Departing from Earth-Moon L1 significantly reduces the Earth departure requirements for this mission. With the current architecture, the CPS Beta used for Earth departure has 44% of its propellant offloaded. It might be possible for this stage to be launched to LEO fully loaded, propel itself to L1, and then be used for the Earth departure leg without any refueling. More investigation is necessary. Alternatively, a much smaller propulsive element could be used for Earth departure. The requirements on the CPS Beta final stage and MPCV are unchanged from the LEO departure options. All of the Mars options shown only consider reached Martian orbit with a crew habitat. A separate cargo mission would be required to reach the surface. Though the propulsive elements proposed in this study would be sufficient to support a cargo mission, such a mission was not investigated in this study. VIII. SUMMARY This study has shown that through the use of innovative and purpose-driven design, an allchemical, common-element architecture solution is possible for human exploration missions to the Moon, NEAs, and Mars using proven, near-term, LEO Departure Option 1 15 LEO Departure Option 2 L1 Departure Payload Mass (t) Total Initial Mass (t) Stage 1 Stage 2 Stage 3 Stage 1 Stage 2 Stage 1 Stage 2 Element Beta 1 Beta 1 Beta 2 Gamma Beta 2 Beta 1 Beta 2 Avg. Boil-Off Rate (%/day) 0.03% 0.03% 0.015% 0.03% 0.015% 0.03% 0.015% CPS Mass at Ignition (t) Ignition T/W Total Burn Time (min) Videal (m/s) 1,650 2,810 4,250 4,620 4,250 1,770 4,250 Gravity Losses (%) 1.0% 1.5% < 0.1% 5.0% < 0.1% < 0.1% < 0.1% Fuel Offload (%) 0.0% 0.0% 9.0% 0.0% 9.0% 44.3% 9.0% Oxidizer Offload (%) 0.0% 0.0% 38.4% 0.0% 38.4% 44.3% 38.4% Table 14. Mars Mission (orbit only) Results

16 low-risk propulsion systems and passive boil-off management. For this type of architecture to be successful, the CPSs need to be designed to minimize inert mass while maintaining acceptable propellant boil-off rates. Engine Isp has a small impact over the range available to LOX/LH2 engines. In fact, engine thrustto-weight ratio, as a significant contributor to vehicle inert mass, is likely a more important engine parameter. To achieve these design goals, several innovative approaches need to be applied to CPS design. The primary structures mass can be reduced by launching the CPS empty as a non-loadbearing payload within a fairing. The subsystems masses can be reduced by adopting the IVF technology. Boil-off rates can be reduced passively through intelligent propellant tank design and the addition of MLI, minimizing the number of thermal paths associated with the support structure, and using a lightweight deployable sun shield when necessary. The all-chemical common element solution consists of two required stages, a 100t CPS and 220t CPS. Both have block upgrades available to reduce propellant boil-off losses with a deployable sun shield. An optional 480t CPS can help achieve Mars missions by reducing the length-to-diameter of the vehicle. This simplifies attitude control and reduces bending moments on structures. A brief summary of the required masses for each Figure 13. Comparison of LEO and E-M L1 as Mission Departure Point mission from architecture study is shown in Figure 12. Because these are high propellant mass fraction all-chemical stages, propellant mass dominates the total mission mass. Departing from E-M L1 instead of LEO reduces the total mission initial mass required, as shown in Figure 13. This can reduce the number of propulsive elements required and reduce the thrust requirements of those elements. Using E-M L1 as a deep space mission starting point rather than LEO may prove advantageous. All of the results generated in this study are highly dependent on the masses of the payloads selected. These payload masses are driven by several factors include number of crew and mission duration. The sensitivity of the propulsive elements to the payloads makes comparisons between this study and other work difficult. IX. CONCLUSIONS As indicated in the parametric study, propulsive stages showed the highest growth sensitivity to propellant mass fraction. Therefore, investing in technologies that focus on reducing stage dry mass should be prioritized as these technologies are likely to benefit future human exploration missions more significantly than other technologies (i.e. boil-off, propulsion system performance, etc.). Figure 12. Summary of Mission Masses Vehicle configurations were proposed in the architecture analysis using stages predicated on the use of existing or near-term technologies. This will mitigate the cost and risk associated with technology 16

17 development programs. By using passive thermal control, reliability issues associated with complex, active thermal systems over a long duration mission can be avoided. On-orbit fueling can add a significant amount of mission flexibility. When launched empty, the CPSs can be delivered to LEO and supported by a variety of launch vehicles. X. REFERENCES Wagner, S., Wie, B., A Crewed 180-Dat Mission to Asteroid Apophis in , Iowa State University,IAC-09.D Constellation: Orion Crew Exploration Vehicle, FS GRC, Available from: 0_expl_vehicle.pdf. 4. Davis, T., et al, Environmental Assessment for NASA Launch Abort System (LAS) Test Activities at the U.S. Army White Sands Missile Range, NM, August NASA Facts : The Orion Crew Exploration Vehicle, FS JSC, Available from: heet.pdf 6. Hatfield, S., Project Orion Overview and Prime Contractor Announcement, August NASA s Exploration Systems Architecture Study: Final Report, NASA-TM , November Return to the Moon, T-12 Years and Counting, Andrews Space, Inc., AS-4102-RPT-00002, January Dorris, C., Altair: Constellation Returns Humans to the Moon, Available from: n_symposium_ _2d_dorris.pdf 10. NASA Facts: The Altair Lunar Lander, FS JSC, Available from: _lander.pdf 11. Hansen, L., Introduction to the Altair Project, Available from: t 12. Toups, L., et al, Deep Space Habitat Team: HEFT Phase 2 Efforts, Available from: _ pdf 13. Drake, B.G., Human Exploration of Mars Design Reference Architecture 5.0, NASA/SP , July Drake, B.G., Decadal Planning Team Mars Mission Analysis Summary, NASA/TM , July HSF TransHab, Updated 27 June 2003, Visited 18 July 2011, Drake, B.G., Reference Mission Version 3.0 Addendum to the Human Exploration of Mars: The Reference Mission of the NASA Mars Explotaiton Study Team, NASA/SP-6107-ADD, June Zubrin, Robert, The Case for Mars, Thronson, H., et al. Human Exploration Beyond LEO by the End of the Decade: Designs for Long-Duration Gateway Habitats, FISO Telecon Colloquium, December Available from: Kutter_ /Thronson% pdf 19. Thronson, H., et al. Post-ISS Human Operations in Free Space: Scenarios for Future Exploration Beyond LEO, FISO Working Group Telecon, September 2009, Available from: 09.pdf 20. Geffre, J., et al, Lunar L1 Gateway Conceptual Design Report, EX , October Chow, N., and Gralla, Erica, Low Earth Orbit Constellation Design Using the Earth-Moon L1 Point, Princeton University, May 2004, Available from: hesis_chow_gralla.pdf 22. Ashmore, M., et al, Clarke Station: An Artificial Gravity Space Station at the Earth-Moon L1 Point, University of Maryland, Available from: /maryland01b.pdf 23. Mendell, W. W., Hoffman, S., Cislunar Infrastructure, Viewed 7/18/2011, XLibrary/DOCS/EIC042.HTML 24. Lunar Orbit Insertion Targeting and Associated Outbound Mission Design for Lunar Sortie Missions, FltDyn-CEV-06-72, March Available from: _ pdf 25. Farquahar, R.W., A Halo-Orbit Lunar Station, June 1972, Available from: rstation.pdf 26. Foster, C., Daniels, M., Mission Operations for Human Exploration of Nearby Planetary Bodies, AIAA , AIAA Space 2010 Conference and Exhibition, Anaheim, CA, August Landis, R.R., et al, A Piloted Orion Flight to a Near- Earth Object: A Feasibility Study, AIAA , SpaceOps 2008 Conference, Heidelberg, Germany, May LeCompte, M.A., et al, Near-Earth Asteroid Rendezvous Missions with the Orion Crew 17