Crew Exploration Lander for Ganymede, Callisto, and Earth s Moon - Vehicle System Design

Transcription

1 45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit 2-5 August 2009, Denver, Colorado AIAA Crew Exploration Lander for Ganymede, Callisto, and Earth s Moon - Vehicle System Design Mark G. Benton, Sr. * The Boeing Company, Huntington Beach, CA Abstract. The Spaceship Discovery design was presented at the Space 2006 Conference and is cited in this paper s references. This conceptual architecture for human solar system exploration includes a crew exploration lander for airless moons, the Lander Module 1 (LM1). LM1 is designed to land on the Jovian moons Ganymede and Callisto, and can be flight tested on Earth s Moon. The LM1 utilizes only vacuum propulsive braking. Its design was first presented at the AIAA 2009 Aerospace Sciences Conference. The objective of this paper is to provide additional design details. The LM1 crew lander provides two-way transportation for a nominal two-person crew between orbit and the surface, and life support for a stay of up to 30 days. The two-stage LM1 is designed for abort to orbit during the powered descent. It is sized to carry three crew members, with reduced payload, for rescue missions. To reduce development cost, the LM1 lander features a geometrical layout, structural concept, and landing gear that are common to those of the Spaceship Discovery LM2 and LM3 Mars landers. These landers were presented at the AIAA 2008 Joint Propulsion Conference and are cited in this paper s references. The LM1 incorporates technologies from recent NASA Altair lunar lander concept definition studies. It utilizes LH 2 /LO 2 propellants in the Descent Stage (DS) and storable monomethyl hydrazine/ nitrogen tetroxide (MMH/NTO) propellants in the Ascent Stage (AS). The design incorporates a layer of water in the AS pressure cabin to shield the crew from Jupiter s intense radiation environment. Three Spaceship Discovery design reference missions are presented to develop lander design requirements: Missions to the Jovian moons Ganymede and Callisto and a mission to Earth s Moon that would be used to flight test the lander prior to using it on a deep space mission. Ganymede mission requirements were most stressing for the design and sized the vehicle. Propellant can be offloaded to perform the missions to Callisto and Earth s moon. Design requirements, mission profiles, mass properties, performance data, and configuration layouts are presented for the LM1 crew lander vehicle. This lander design is a proposed solution to land humans on the Jovian moons Ganymede and Callisto. It is based on reliable and proven technology and can be flight tested on Earth s Moon. The LM1 design and operations concept stress safety and redundancy and feature abort and rescue capabilities. Its design incorporates many features in common with the Spaceship Discovery Mars landers to reduce development cost. I. Introduction The Spaceship Discovery design 1 was presented at the Space 2006 Conference and is shown in Fig. 1. This conceptual architecture for human exploration of the solar system includes piloted crew landers for exploration of airless moons, designated Lander Module 1 (LM1). This lander is designed to land on the Jovian moons Ganymede and Callisto, and can be flight tested on Earth s Moon. The objective of this paper is to provide design requirements, design details, mission profiles, and flight performance for this lander. The LM1 crew lander provides two-way transportation for a nominal two-person crew between orbit and the surface, and provides life support for a surface stay of up to 30 days. The two-stage LM1 is designed for abort to orbit during the powered descent. The Ascent Stage (AS) is sized to carry three crew members from the surface to orbit for rescue missions. The LM1 Lander features a geometrical layout, structural concept, and landing gear that are common to those of the Spaceship Discovery LM2 and LM3 mars landers. 2 The LM1 lander incorporates technologies from recent NASA Altair lunar * Director, Operationally Responsive Space Systems, Boeing Phantom Works, Bolsa Chica Road, Mail Code H017-D833, Huntington Beach, CA Senior Member of AIAA, Vice Chairman of the AIAA Space Colonization Technical Committee, and Member of the AIAA Space Transportation Technical Committee. 1

2 lander concept definition studies. 3 Three Spaceship Discovery design reference missions are presented to develop design requirements: Missions to the Jovian moons Ganymede and Callisto and a mission to Earth s Moon that would be used to flight test the lander prior to using it on a deep space mission. Ganymede mission requirements were most stressing for the design and sized the vehicle. Propellant can be offloaded to perform missions to Callisto and Earth s moon. Design requirements, mission profiles, mass properties, performance data, and configuration layouts are presented for the LM1 crew lander. This design is a proposed solution to land humans on the Jovian moons Ganymede and Callisto based on reliable and proven technology. It can be flight tested on Earth s Moon. The LM1 design and operations concept stress safety and redundancy and feature abort and rescue capabilities. Engineering Module (EM) Main LH 2 Propellant Core Tank (CT) (4) Main LH 2 Propellant Drop Tank (DT) (12) Service Module (SM) Crew Module (CM) Closed Brayton Cycle (CBC) Electrical Generation Systems (3) Very Low Boil-off Cryo. Retention Sys. a. Main Ship Side View Artificial Gravity (AG) Centrifuge 5-Port Docking Module (DM) Mars Lander Modules LM2, LM3 (DRMs 2, 3) Bimodal Nuclear Thermal Rocket (NTR) Engines (3) Deployable Solar / Thermal Shades (2) Abort Propulsion System (APS) Engine Galactic Cosmic Ray (GCR) Biological Shield (LH 2 & H 2 O) b. Side View Cutaway Key Design Features Ganymede, Callisto, Earth s Moon Lander Module LM1 (DRMs 1, 5, 6) FIGURE 1. Spaceship Discovery Conceptual Architecture for Human Solar System Exploration. Reentry Module (RM) II. Mission Requirements Figures 2a-2d depict the various Spaceship Discovery configurations corresponding to the design reference missions that were used to develop design requirements for the LM1 Crew Exploration Lander. A. Design Reference Mission 1 (DRM 1) Lunar Flight Test/Crew Training This DRM is used to flight test spaceship Discovery prior to flying deep space missions and train flight crews. LM1 landers can be flight tested in the similar environment of Earth s Moon prior to Jupiter missions. Spaceship Discovery departs the 556 km circular Earth parking orbit with the Trans-Lunar Injection (TLI) burn using the three nuclear thermal rocket (NTR) engines on the main ship as shown in Fig. 2c. The main ship is reusable for this mission. Six crew members and three redundant LM1 landers are carried. After a 3.4 day transit, the Lunar Orbit Insertion (LOI) burn places the vehicle into a 185 km circular lunar orbit. A stay time of 180 days in lunar orbit was chosen to approximate DRMs 5-6 mission durations at Jupiter and size Discovery s on-board consumables. With three landing missions in succession, two landing parties could explore the surface of the moon for a total of 90 days to train crews prior to Ganymede or Callisto surface exploration missions. At the conclusion of the landing missions, the TEI burn propels Spaceship Discovery into the 3.4 day return transfer orbit, and the Earth Orbit Insertion (EOI) burn places the vehicle into Earth parking orbit as shown in Fig. 2d. Crew and samples return to Earth in the RM. B. Design Reference Mission 5 (DRM 5) Callisto Exploration This mission has a duration of four years, with one year, nine month transits and a six month wait time in the Jovian system. DRM 5 has a crew complement of four. Spaceship Discovery escapes the 556 km circular Earth 2

3 parking orbit (Trans-Jupiter Injection, TJI) using nine NTR engines on main ship and boosters. The twin, strap-on boosters depicted in Fig. 2a enable high-energy missions to Mars and Jupiter. For design commonality, each booster utilizes the same modules as the main ship, with a single EM and two CTs. Main propellant DTs can be clustered around booster CTs, with the number of DTs tailored to mission requirements. During the TJI burn, the boosters will cross-feed propellant to the three NTR engines on the main vehicle as well as supplying the six NTR engines on the boosters. After booster propellant is expended, the boosters are jettisoned and the main ship briefly accelerates using its three NTR engines to ensure that the boosters miss the rendezvous with Jupiter. The boosters are disposed of into a solar orbit, and residual booster propellant is vented over time to raise their orbit periapsis above 1.0 AU. The main ship transfer orbit apoapsis is 6.7 AU. Callisto Orbit Insertion (COI) includes two burns: A propulsive capture burn into an elliptical Jupiter orbit with periapsis at the radius of Callisto (1.884 M km), and a propulsive capture into a 185 km circular parking orbit around Callisto with a degree plane change. Callisto s orbit is the farthest from Jupiter of the Galilean moons, but lander shielding will still be needed for radiation dose rates to be tolerable to the crew. Three redundant LM1 landers are carried. The first LM1 lands with two astronauts, leaving two astronauts in orbit. After a stay of 30 days, the crew departs the surface and docks with the main ship. The other two astronauts then take their 30 day turn on the surface in a second LM1. The third LM1 is always in standby and can be flown by one astronaut to rescue either landing party if it became stranded. The TEI burn is a direct escape from Callisto orbit and Jupiter, and propels the ship into the inbound transfer orbit. (Fig. 2d). Discovery has insufficient propellant left for propulsive capture at Earth. The Earth Trajectory Injection (ETI) burn is performed at 900,000 km (35.3 hrs) from Earth after the ship is within Earth s 924,133 km activity sphere radius. The crew enters the RM and separates from the main ship at 800,000 km (31.3 hrs) from Earth, aimed at a periapsis altitude of 122 km. The crew performs a hyperbolic reentry and landing in the RM. The ETI burn reduces RM entry velocity from 14.6 km/s to 13.0 km/s. After RM departure, the main ship performs a 434 m/s burn to increase periapsis altitude to 1,000 km, enabling a posigrade hyperbolic pass high above the atmosphere. The main ship is disposed of into solar orbit with apoapsis greater than 3.0 AU and periapsis greater than 1.0 AU to prevent any chance of it impacting Earth. C. Design Reference Mission 6 (DRM 6) Ganymede Exploration Ganymede is the largest moon of Jupiter, and at 5,262 k m in diameter, is the largest moon in the solar system. It is larger than Mercury or Pluto. DRM 6 has the same overall mission profile as DRM 5 except that Ganymede is the targeted moon. It is postulated that Ganymede could be briefly visited with sufficient radiation shielding. Ganymede Orbit Insertion (GOI) includes two maneuvers: A propulsive capture into an elliptical Jupiter orbit with periapsis at the radius of Ganymede (1.070 M km), and a propulsive capture into a 185 km circular parking orbit around Ganymede with a degree plane change. Three redundant LM1 landers are carried and the four crewmembers explore the surface of Ganymede. The TEI burn is a direct escape from Ganymede orbit and Jupiter. The inbound transit, ETI burn, RM hyperbolic entry, and disposal of the main ship into solar orbit are the same as for DRM 5. III. Design of LM1 Lunar Crew Exploration Lander Vehicle A. Design Overview and Assumptions The LM1 crew lander utilizes only vacuum propulsive braking and is designed to land on the airless Jovian moons Ganymede and Callisto. It is also designed to be flight tested on Earth s Moon. It provides two-way transportation for a nominal two-person crew between the parking orbit and the surface, and also functions as a habitat and rover vehicle. The design incorporates a layer of water in the crew cabin walls to shield the crew from Jupiter s intense radiation environment. The LM1 crew lander normally carries two astronauts, with an endurance of 30 days on the surface (including margin) and seven days for ascent and contingencies. It can support three astronauts for rescue missions, albeit with reduced endurance. It can achieve orbit with three crew, three space suits, three emergency life support system units, and a contingency sample (20 kg vs. the normal sample allocation of 125 kg). The LM1 lander extensively utilizes lightweight composite structures to maximize performance and minimize overall system mass. It is envisioned that advanced composite materials used for crew habitation areas will be dualmode, and provide radiation protection shielding as well as structural integrity. 4 It utilizes liquid hydrogen and liquid oxygen (LH 2 /LO 2 ) propellants in the Descent Stage (DS) and storable monomethyl hydrazine/nitrogen tetroxide (MMH/NTO) propellants in the AS. The LM1 lander has a 6-axis reaction control systems (RCS) for attitude control and rendezvous and docking translation maneuvers. The lander RCS engines utilize storable, hypergolic MMH/NTO propellants. The LM1 has functionality and flight profiles similar to the Apollo Lunar Module, but is larger, with higher performance and much higher endurance. Its design incorporates Altair, 3 Space Shuttle, 5, 6 MMH/NTO engines, 7 and Apollo Lunar Module 8-11 design data with upgrades for recent advances in materials and subsystems. The LM1 service life is three years (one-year assembly in low-earth orbit plus one-half of 3

4 the duration of DRMs 5 or 6). The design of LM1 landers is in keeping with the design philosophy for high mission redundancy. Aborts are extensively considered in the design and operations concept. The LM1 crew lander is designed for abort-to-orbit (ATO) during all parts of the powered descent, where the DS is jettisoned and the AS returns to orbit. Multiple landers enable multiple exploration landing mission attempts and also enable rescue missions if a landing party were to become stranded on the surface. LM1 configuration layouts are presented in Figs. 3 and 4 and design data are presented in Tables 1 and 2. Lander ascent and descent flight performance was determined using a 2-D trajectory simulation with a spherical gravitational potential. Performance calculations included 1% flight performance reserve (FPR) on V. Flight performance data for Ganymede, Callisto, and Earth s Moon simulations are presented in Tables 3 and 4. The Ganymede mission requirements were most stressing for the design and sized the vehicle. Callisto and Earth s Moon missions had similar V requirements. The Callisto mission was the least stressing. Propellant can be offloaded to perform the Callisto and Earth s moon missions as shown in Table 4. Flight performance plots for the Ganymede sizing case are presented in Figs Performance plots for the Callisto and Earth s Moon simulations are very similar to the Ganymede case and are omitted for brevity. B. Design of Descent Stage (DS) The overall design of the DS is common to the LM2/LM3 Mars landers, with the same basic layout that utilizes a central thrust cylinder to efficiently carry loads. The DS airlock, cargo bay, and fuel cell consumables tanks are the same as those used on the LM2 Mars lander. The DS landing legs and traversing wheels utilize the same kinematics, geometry, and mechanisms as the LM3 Mars lander. Structural thicknesses have been reduced to account for reductions in flight loads between the Mars and moon missions. Landing gear struts have 0.6 m of shock absorbing stroke and another 0.6 m of adjustability to level the vehicle. The landing gear struts have wheels to enable the LM1 to traverse the surface at the relatively slow pace of 90 m per hour, using 1.0 kw for the drive motors. The LM1 can nominally traverse up to 1.08 km per 24 hour period, assuming 12 hours for stops and crew rest periods each day. Rover wheel motor drive power calculations are based on Mars Exploration Rover design data. 12 The DS houses propellant, pressurization, fuel cells and their consumables tanks (including breathing oxygen), and eight gimbaled, throttleable 12.2 kn RL-10 based descent engines that are spaced at 45-degree intervals around the base of the thrust cylinder. Engines are throttleable between 100% and 30% thrust and are sized for single engine-out considerations. In the event of an engine out, the corresponding opposite engine would be shut down to balance the thrust. Multiple descent engines are shut down during the powered descent, in groups of two, to maintain proper thrust and balance. The personnel airlock in the DS is accessible by tunnel from the AS crew cabin. There is a cargo bay on the opposite side of the DS as the airlock, with a capacity of 500 kg of payload for various types of exploration equipment and science experiments. Descent and surface electrical power is provided by fuel cells, whose design is based on Apollo fuel cells. 11 LM1 fuel cells produce up to 4.1 kw peak power. Lander wheel drive motors consume 1 kw, heaters 1 to 2 kw, and a maximum of 1 kw are consumed by the Equipment Cooling and Life Support System (ECLSS), Guidance, Navigation and Control (GN&C) System, communications, and lighting and habitat loads. Reactants are carried to support an average duty cycle of 3 kw. Eight storage tanks hold shielding water for the crew cabin. C. Design of Ascent Stage (AS) The design of the AS crew cabin, thrust cylinder, and airlock tunnel is common to the LM2 Mars lander, with the same basic layout that utilizes a central thrust cylinder to efficiently carry loads. The crew cabin has 29 m 3 of usable inner moldline volume. It has a pressurized tunnel connecting it to the airlock in the DS. After landing, 1,722 kg of water is pumped from the DS into a 4.0 cm thick shielding layer in the crew cabin inner wall. It provides 4 gm/cm 2 of radiation shielding, affording the same protection to the crew as that provided by the Crew Module on the main ship. This water must be drained into storage tanks in the DS to lighten the AS sufficiently for the ascent to orbit. The AS crew cabin functions as flight deck and habitat for the two person crew. The two crew members stand in front of the control console. Two large angled windows provide excellent visibility for approach and landing. The habitat portion of the cabin contains two bunks, a computer table, a kitchen with dining table and chairs, and hygiene facilities including a toilet, sink, and shower. The AS has four unpressurized external bays: A forward bay for the airlock tunnel and GN&C and communications equipment, an aft bay for batteries, ECLSS equipment and consumables tanks (breathing oxygen and water), and right and left side bays for main and RCS propellant tanks. The four unpressurized bays and the central crew cabin are covered with thermal blankets and heavily insulated against the cold environment. The central thrust cylinder, below the crew cabin, houses eight 3.33 kn fixed-thrust ascent engines and main and RCS helium pressurization tanks. Ascent engines are sized for single engine-out considerations. In the event of an engine out, the corresponding opposite engine will be shut down to balance thrust. Sixteen 0.77 kn RCS thrusters are arranged in four groups of four. Ascent electrical power is provided by solar arrays and batteries. Solar arrays and radar and communications antennas are attached to the exterior of the AS. 4

5 Booster Mated to Main Ship Boosters (2) Used for Trans-Jupiter Injection Burn a. DRMs 5 or 6 Assembly in Earth Parking Orbit in Preparation for Trans-Jupiter Injection b. DRMs 5 or 6 Outbound Transit to Jovian System and Insertion into Callisto or Ganymede Orbit c. DRMs 5 or 6 In Orbit around Callisto or Ganymede, Ready for Landing Missions DRM 1 Trans-Lunar Injection, Outbound Transit, Lunar Orbit Insertion, and Ready for Landing Missions d. DRMs 5 or 6 Trans-Earth Injection, Inbound Transit, and Earth Trajectory Injection DRM 1 Trans-Earth Injection, Inbound Transit, and Earth Orbit Insertion FIGURE 2. Design Reference Mission Configurations. 5

6 Tunnel to Airlock in DS Crew Cabin w/ 4 cm Water Shielding Layer LSS LO 2 Tank (2 pl) LSS H 2 O Tank (2 pl) Thermal Blanketing Over Stowed Landing Gear (4 pl) Main Prop. LH 2 Tank Cargo Bay ECLSS Equip. Bay Main Prop. LO 2 Tank Side View Cutaway, Ascent and Descent Stages Reflective Thermal Foil around Lander Base Descent Stage Cutaway through Landing Gear Bays (Gear Stowed) DS Main Propellant LO 2 Tank AS Main Propellant NTO Tank (2 pl) Crew Cabin Including Flight Deck & Habitat AS Main Propellant MMH Tank (2 pl) Fuel Cell/ LSS Supercrit. LH 2 & LO 2 Tanks (2 sets) DS Main Propellant LH 2 Tank Bio Shield Water Stg. Tanks (8 pl) DS Main Propellant & Bio Shield Water Pressurization Tanks (4 pl) DS Thrust Cylinder Primary Structure DS Main Propellant LO 2 Tank Front View Cutaway, Ascent and Descent Stages Gimballed Descent Engines (8 pl) Articulated Drive Wheel (8 pl) Descent Stage Cutaway through Landing Gear Bays (Gear Deployed) FIGURE 3. LM2 Crew Lander Configuration Drawings Elevation Views. 6

7 Crew Habitat Including Sleeping, Eating, and Hygene Facilities Crew Bunk (2 pl) LSS H 2 O Tks (2 pl) DS Main Propellant LH 2 Tank Fuel Cell/ LSS Supercrit. LH 2 Tanks (2 pl) DS Main Propellant & Bio Shield Water He Pressurization Tanks (4 pl) Cargo Bay Surface Access Door Pair (2 pl) Flight Deck and Main Windows Ascent Stage Upper Level showing Crew Cabin. Main Propellant and RCS Tank Bay (2 pl) Tunnel to Airlock in DS Air Lock Descent Stage Upper Level (Landing Gear Shown Stowed) DS-AS Explosive Separation Bolts (6 pl) Main Propellant Pressurization Tanks (4 pl) ECLSS Bay AS Main Engines (8 pl) DS Main Propellant LO 2 Tank DS Main Engines (8 pl) Cargo Bay Bio Shield Water Stg. Tanks (8 pl) Fuel Cell/ LSS Supercrit. LO 2 Tanks (2 pl) RCS Prop. Press. Tanks (2 pl) Fuel Cell (4 pl) Air Lock Main Propellant & RCS Tank Bay (2 pl) Ascent Stage Lower Level Propulsion and Systems Bays RCS Thruster Quad (4 pl) Descent Stage Lower Level (Landing Gear Shown Deployed) FIGURE 4. LM2 Crew Lander Configuration Drawings Cross Section Views. 7

8 TABLE 1. LM1 Design Assumptions for Life Support Consumables. Ascent Stage Descent Stage LM1 Overall Design Assumptions (kg/man-day) Breathing Oxygen Water Dried & Condensed Food Endurance Total Man-Days Total Days for 2 Crewmembers Consumables Mass (kg) Breathing Oxygen Water Dry Food Add l. Fuel Cell Reactants Available Total Consumables TABLE 2. LM1 Dimensions and Mass Properties (Propellants Shown for Ganymede Sizing Case). Descent Ascent Ascent Stage 1 Descent Stage 1 Total for Descent 1 Rescue Ascent 2 Nominal Ascent 3 Dimensions (m) Length Overall Diameter Masses (kg) Payload Crew Shielding Water 1,722 1,722 (2) PLSS in Airlock Crew, Suits, ELSS Structure & Insulation 605 1,079 1, Propellant & Press. Tanks Landing Gear Drive Motors and Wheels ECLSS, Power, & Avionics Main Engine RCS Dry Mass Margin (15%) LSS/Fuel Cell Consumables RCS Propellant (Usable) Operating Empty Mass 3,124 6,209 9,333 2,668 2,653 Main Propellant (Usable) 3,382 9,250 12,632 3,382 3,382 Total Mass 6,506 15,459 21,965 6,050 6,035 Non-Prop/Non-Cons. Mass Fract Nominal descent: Two crewmembers, space suits & ELSS units, and 500 kg payload. 2 Three crewmembers, space suits & ELSS units, and 20 kg contingency sample. 3 Two crewmembers, space suits & ELSS units, and 125 kg sample payload. 4 (Payload + Shielding + Struct./Insulation + Engines & Subsystems (Incl. Prop. Residuals)) / Total Mass 8

9 TABLE 3. LM1 Mission Parameters. Target Planetary Body Ganymede Callisto Moon Planetary Parameters Radius at Surface (km) 2,631 2,410 1,738 Equatorial Rotation (m/s) Landing Gravitational Parameter (km 3 /s 2 Point ) 9,887 7,179 4,903 Gravity at Surface (m/s 2 ) Circ. Parking Orbit Parameters Parking Orbit Radius (km) 2,816 2,596 1,923 Parking Orbit Altitude (km) Descent Descent Orbital Velocity (m/s) 1,874 1,663 1,597 Transfer Transfer Burn Descent Parameters Orbit a. Descent to Surface V for Transfer Burn 1 (m/s) True Anomaly, Periapsis Altitude (PDI) (km) Launch to Inertial Velocity at PDI (m/s) 1,968 1,749 1,711 Circulariz. Burn Relative Velocity at PDI (m/s) 1,941 1,739 1,706 True Anomaly at Landing (deg) Ascent Parameters Relative Vel. at Burnout (m/s) 1,938 1,707 1,628 Rel. Vel. at 185 km Alt. (m/s) 1,876 1,663 1,600 Launch V for Circulariz. Burn 2 (m/s) Point Parking Ascent True Anomaly (deg) Orbit 1 Includes 25 m/s maneuvering, 5 deg. plane change, 1% flight perf. reserve (FPR). 2 Includes 100 m/s maneuv., 5 deg. plane change, +30 degree incl., and 1% FPR. b. Ascent to Orbit FIGURE 5. LM1 Mission Profiles. TABLE 4. LM1 Flight Performance Data. Target Planetary Body Ganymede (DRM 6) Callisto (DRM 5) Earth s Moon (DRM 1) Lander Mission Profile Major Propulsive Burns Nominal Descent 1 Deorbit & Descent Rescue Ascent 2 Ascent & Circularize Nominal Descent 1 Deorbit & Descent Powered Descent Initiation (PDI) Point Rescue Ascent 2 Ascent & Circularize 9 - True Anomaly, PDI to Landing Nominal Descent 1 Deorbit & Descent Orbit Circularization Burn Rescue Ascent 2 Ascent & Circularize Required Delta Velocity (km/s) Orbital Maneuvering Powered Descent or Ascent Gravity Losses Flight Performance Reserve (FPR) Total Performance Parameters Specific Impulse (I SP ) Mass Ratio (M i / M f ) Burn Propellant Fraction Initial Thrust / Weight Final Thrust / Weight Mass (kg) Operating Empty Mass 9,333 2,668 9,286 2,644 9,296 2,649 Main Propellant 12,632 3,382 10,334 2,819 10,824 2,906 Total Mass 21,965 6,050 19,620 5,463 20,120 5,555 1 Nominal descent (2 crewmembers) is sizing case. 2 Rescue ascent (3 crewmembers) is sizing case. 3 Includes a plane change of 5.0 degrees during both descent and ascent. 4 Referenced to the acceleration of gravity at surface: (1.428 m/s 2 Ganymede; m/s 2 Callisto; m/s 2 Earth s Moon).

10 20 Entry Interface Altitude Altitude (km) 4 0 1,800 1,600 1,400 1,200 1, Velocity (m/s) FIGURE 6. LM1 Ganymede Powered Descent Trajectory Parameters Axial Velocity vs. Altitude. 1, Altitude (m) m Hover Altitude Velocity (m/s) FIGURE 7. LM1 Ganymede Powered Descent Trajectory Parameters Axial Velocity vs. Altitude Altitude (km) m Hover Altitude Downrange Displacement from Entry Interface (km) FIGURE 8. LM1 Ganymede Powered Descent Trajectory Parameters Altitude vs. Downrange Displacement. 10

11 Thrust Vector or Flight Path Angle (Deg.) Flight Path Angle Thrust Vector Angle Timed Hover Phase Soft Landing Phase Time From Entry Interface (s) FIGURE 9. LM1 Ganymede Powered Descent Trajectory Parameters Flight Path and Thrust Vector Angles vs. Time Acceleration (Earth g's) Timed Hover Phase Soft Landing Phase Time From Entry Interface (s) FIGURE 10. LM1 Ganymede Powered Descent Trajectory Parameters Axial Acceleration vs. Time Timed Hover Phase Force (kn) Soft Landing Phase Time From Entry Interface (s) FIGURE 11. LM1 Ganymede Powered Descent Trajectory Parameters Thrust Force vs. Time. 11

12 200 Orbital Altitude Altitude (km) Vertical Velocity Burnout Horizontal Velocity - Includes 26.7 m/s Equatorial Rotation ,000 1,500 2,000 2,500 Velocity (m/s) FIGURE 12. LM1 Ganymede Ascent Trajectory Parameters Altitude vs. Inertial Velocity Components. 200 Orbital Altitude 150 Altitude (km) 100 Burnout Vertical Acceleration 50 Horizontal Acceleration Acceleration (m/s 2 ) FIGURE 13. LM1 Ganymede Ascent Trajectory Parameters Altitude vs. Inertial Acceleration Components. 90 Burnout 80 Angle from Horizontal (Degrees) Thrust Vector Angle Flight Path Angle Elapsed Time From Ignition (s) FIGURE 14. LM1 Ganymede Ascent Trajectory Parameters Flight Path and Thrust Vector Angles vs. Time. 12

13 IV. Mission Profiles The key events of the LM1 mission description are presented pictorially in Figs. 15a-15h. A. Launch from Earth to LEO Assembly Orbit The Spaceship Discovery design is modular and assembled in a circular, 556 km Earth orbit. Subassemblies are up to 33.0 m length x 8.4 m in diameter, with a maximum mass of 50 metric tons (MT), including fairing, adaptors, and airborne support equipment. Two 21.8 MT LM1 landers can be launched simultaneously as shown in Fig. 15a. Launch vehicles are assumed to be enhanced versions of the current Evolved Expendable Launch Vehicles (EELVs). 13, 14 Once in the parking orbit, landers dock to the Spaceship Discovery docking module as shown in Fig. 15b. Vehicle assembly and outfitting will be completed and the main propellant tanks will be topped off. The Jupiter mission crew will board, make preparations for the Trans-Jupiter Injection burn, and depart Earth orbit for Jupiter. B. Outbound Transit Figure 15c shows three LM1 landers and a Reentry Module docked to the main ship docking module. During the 21 month DRM 5/6 outbound transit from Earth to Jupiter the LM1 landers will be kept in a hibernation mode to conserve power. Periodic automated checkouts of lander systems will be performed. Spaceship Discovery utilizes a Very Low Boil-Off System, with refrigeration plants and space radiator panels, to remove heat from on-board cryogenic propellant and consumables tanks to minimize boil-off. Chilled helium gas is circulated through cryocoolers to remove heat from these tanks. This system will be connected by umbilicals to the three LM1 landers to cool liquid and supercritical H 2 and O 2 tanks in the descent stages. After insertion into the 185 km altitude circular parking orbit, the landers are powered up and thorough systems checks are performed in preparation for landing. C. Descent from Parking Orbit to Surface The crew enters the LM1 and the lander undocks from the main ship. Landing gear are deployed and locked. LM1 landers are designed to operate from parking orbits inclined up to 30 degrees to the equator, and can execute a plane change of up to 5 degrees during descent. This will permit significant flexibility in surface targeting. A deorbit burn is accomplished using the descent engines as shown in Fig. 15d. This maneuver inserts the lander into an elliptical transfer orbit whose periapsis is 15.2 km. The powered descent (PD) is initiated at this point and continues to 100 m above the surface at constant thrust/weight (T/W) of 2.5. PD continues with hovering (T/W = 1.0) at 100 m altitude, with sufficient propellant carried to hover for a maximum of 60 seconds to locate and avoid obstacles. PD concludes with a soft landing (Tables 3 and 4 and Figs. 5a, 6-11, and 15e). In the event of a landing abort, the LM1 descent stage will be jettisoned and the ascent engines will ignite and propel the ascent stage on an ascent trajectory that returns it to the 185 km altitude Spaceship Discovery parking orbit. An abort-to-orbit is available during the entire PD duration. The LM1 abort-to-orbit performance and trajectories will be detailed in future work. D. Surface Operations The LM1 lander provides a roving habitat for the crew. It provides biological shielding to the crew from the hazardous radiation environment near Jupiter using a combination of advanced dual-mode composite materials and a layer of shielding water in the crew cabin. It is designed to traverse up to 30 km during its 30 day surface exploration mission using its powered wheels. It is envisioned that the crew will spend most of their 30 day mission inside the LM1 shielded habitat and only make brief extra vehicular activity (EVA) sorties due to the high radiation environment. Figure 15f shows the LM2 crew lander in its landed configuration. A ladder has been extended from the airlock to permit the crew to access the surface for EVA. LM1 surface operations will be detailed in future work. E. Ascent to Parking Orbit, Rendezvous, and Docking Before liftoff from the surface, the crew drains the 1,722 kg of shielding water from the AS crew cabin shield tank into the DS storage tanks, and also leaves their portable life support system backpacks in the airlock in order to lighten the AS. The LM1 ascent stage launches from the descent stage as shown if Fig. 15g. The AS is designed to ascend to the parking orbit from latitudes of up to + 30 degrees from the equator. The eight ascent engines burn at constant thrust until sufficient velocity is achieved and they are shut down. The AS then coasts up to the parking orbit altitude of 185 km. At this point, the ascent engines fire again to circularize the orbit and, if required, execute a plane change of up to 5 degrees (Tables 3 and 4 and Figs. 5b and 12-14). The AS then maneuvers in orbit using the ascent engines and RCS to rendezvous and dock with the Spaceship Discovery main ship as shown in Fig. 15g. The main ship could also rendezvous and dock with the AS in the event of a failure of the AS after it has reached parking orbit. The crew transfers to the main ship and the ascent stage is deorbited using its residual propellants. 13

14 a. Launch Configuration b. Orbital Assembly c. Outbound Transit d. Deorbit Burn FIGURE 15. Mission Configurations. 14

15 e. Powered Descent f. Surface Operations g. Ascent to Parking Orbit h. Rendezvous and Docking with Main Ship FIGURE 15. Mission Configurations, Continued. 15

16 V. Enabling Technologies Continued development of key enabling technologies will be necessary for the implementation of the LM1 crew lander vehicle. These key enabling technologies are listed below: (1) Reliable launching of 50 MT subassemblies, 33.0 m long x 8.4 meters in diameter, into LEO parking orbits. (2) Lightweight dual-mode advanced composite materials for structures that incorporate radiation protection. (3) Highly reliable retention systems for long-term, low-loss storage of cryogenic liquids (LH 2 and LO 2 ). (4) Support equipment for long-duration human habitation and surface exploration of airless moons, including space suits, power generators, life support systems, communications gear, and scientific equipment. This will enable the crew to conduct extensive and detailed scientific explorations of Ganymede and Callisto. VI. Conclusion The Spaceship Discovery LM1 lander design is a proposed solution to land humans on the large Jovian moons Ganymede and Callisto. The LM1 can be flight tested on Earth s Moon. All maneuvers necessary for the Ganymede or Callisto landing can be practiced in the similar environment of the nearby Moon. The LM1 functions as lander vehicle, habitat, and rover. It provides two-way transportation for a nominal two-person crew between orbit and the surface, life support for a surface stay of up to 30 days, and surface mobility of up to one kilometer per day. It also provides biological radiation shielding for the crew. The LM1 design is based on reliable and proven technologies from the Space Shuttle and the Apollo Lunar Module. It incorporates the results of recent NASA Altair lunar lander design studies. To reduce development cost, the LM1 lander features a geometrical layout, structural concept, crew cabin, airlock, and airlock access tunnel, and landing gear that are common to those of the Spaceship Discovery LM2/LM3 Mars landers. The LM1 lander design and operations concept stresses safety and redundancy and is in keeping with the Spaceship Discovery design philosophy for high mission redundancy: Aborts modes and rescue capabilities are extensively considered in the design. The LM1 is designed for abort-to-orbit (ATO) during all parts of the powered descent. Multiple landers enable multiple exploration landing mission attempts during high value, deep space missions and also enable rescue missions if a landing party were to become stranded on the surface. References 1 Benton, Sr., M. G., Spaceship Discovery Vehicle Architecture for Human Exploration of Moon, Mars, and Beyond, AIAA , AIAA Space 2006 Conference, San Jose, CA, Benton, Sr., M. G., Crew and Cargo Landers for Human Exploration of Mars Vehicle System Design, AIAA , AIAA 2006 Joint Propulsion Conference, Hartford, CT, Benton, Sr., M.G., Caplin, G., Reiley, K., Donahue, B., Messinger, R., and Smith, D.B., Boeing Design Trades in Support of the NASA Altair Lunar Lander Concept Definition, AIAA , AIAA Space 2008 Conference, San Diego, CA, Sen, S., Schofield, E., Carranza, S., O Dell, S., Development of Multifunctional Radiation Shielding Materials for Long Duration Human Exploration Beyond the Low Earth Orbit, 58th International Astronautical Congress, Hyderabad, India, (IAC- 07-C2.4.02), Morita, W.H., (ed.), Space Shuttle System Summary, Rockwell International Corp., Downey, CA, May, Goree, J.F., Shuttle Systems Weight & Performance Monthly Status Report, (NASA-TM-84748), NASA, May 18, Wade, M., References for Aerojet HiPAT Thrusters and Space Shuttle OME, Encyclopedia Astronautica Online Reference, URL: [cited 20 April 2008]. 8 Apollo Program Summary Report, App.C Apollo Spacecraft Weights, (NASA-TM-X-68725), NASA, 1975, pp. C1-C4. 9 Apollo 11 Lunar Landing Mission Press Kit, (NASA Release No K), NASA, 1969, pp Apollo 7 Mission Press Kit, (NASA Release No K), NASA, 1968, pp Heitchue, R.D., (ed.), Space Systems Technology, Reinhold Book Corp., New York, 1968, pp (No author listed) Mars Exploration Rover: Rover Design, Drive System and Power and Electronic Systems, Wikipedia Online Reference, URL: [cited 22 June 2008]. 13 Covault, C., Launch Vehicles: Trial by Fire, Aviation Week and Space Technology, Volume 162, No. 8, February 21, 2005, pp Scott, W.B., Morphing Rockets: Lockheed Martin's Atlas V Could Evolve to Saturn V-Class Performance, Aviation Week and Space Technology, Volume 162, No. 25, June 20, 2005, pp