Lunar Settlement, Transportation, and Exploration Program (L-STEP)

Lunar Settlement, Transportation, and Exploration Program (L-STEP)
John Bowen Jon Castelli Bryan Giglio Jose Hernando Duarte Ho Kayvan Madani Najad Paul Nizenkov Chris Ragsdale Vinny Ramachandran Kate Stambaugh ENAE 788D: Advanced Space Systems Design Fall 2011 Term Project Background: Plum Crater, 1972 Photo Credit: NASA

Outline Introduction Orbital Mechanics Launch Vehicle Selection
Propulsion Module Landing Gear Crew Module Human Factors and Life Support Habitat and Power Systems

Introduction - Goals and Limitations
Goal: Use the lessons discussed in ENAE788D for the design of a low-cost, sustainable mission architecture for human missions to the Moon and beyond. Limitations Return humans to the Moon by Use only existing, proven launch vehicles $3 billion annual program budget (“Use or lose”) Guidelines Develop and analyze an innovative, extensible mission architecture and systems capable of supporting NEO missions, but aiming at human lunar return. Look for innovative high-payoff technologies and approaches at minimum cost. “Spend the money flying, not building.” Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Outline Introduction Mission Overview Orbital Mechanics
Launch Vehicle Selection Propulsion Module Landing Gear Crew Module Human Factors and Life Support Habitat and Power Systems

Human Factors & Life Support Habitat & Power Systems
Mission Statements MS-1 To develop an architecture that provides cost-effective human missions to the Moon and NEOs within the timeframe. MS-2 To develop a sustainable and adaptive multi-mission architecture for landing payload and maximizing science return on the Moon and NEOs MS-3 To develop an architecture that is affordable and sustainable within the expected budget of $3 billion per year MS-4 To develop an architecture that supports a safe long-term human presence on the Moon MS-5 To develop missions that leverage existing space exploration assets and capabilities Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Mission Requirements M-1 L-STEP shall be capable of sending humans to the surface of the Moon and returning them safely to Earth MS-1, MS-4 M-2 L-STEP shall be designed to launch missions using only existing launch vehicle systems MS-5 M-3 L-STEP shall not exceed $3 billion per year in development, launch, and mission (including resupply) costs MS-3 M-4 L-STEP shall be designed to accommodate modular science instrument and payload packages MS-2 M-5 L-STEP shall be designed to support lunar surface operation missions of increasing duration MS-2, MS-4 M-6* L-STEP shall be designed to transport humans in planned human-rated exploration vehicles (e.g. Dragon, MPCV, Soyuz) MS-1, MS-4, MS-5 M-7 L-STEP shall be capable of teleoperated and limited autonomous operations when no human presence is available * TBR: self-imposed and may be too stringent Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

System Requirements (1/2)
L-STEP shall be designed in accordance with all of NASA's safety guidelines and requirements for human-rated missions M-1 S-2 L-STEP shall operate with critical systems that are TBD fault tolerant M-1, M-7 S-3 L-STEP shall be capable of controlled descent and landing to the lunar surface M-1, M-5, M-7 S-4 L-STEP shall be capable of ascent from the lunar surface and return to Earth at any point during all human missions in the case of emergency M-1, M-5, M-6 S-5 L-STEP shall be designed to meet launch requirements for likely launch vehicles M-2 S-6 L-STEP shall be designed to dock all major flight systems in space with and without human presence M-6, M-7 S-7 L-STEP shall have the computational power and communications necessary to be teleoperated and to operate autonomously when no human presence is available M-7 Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

System Requirements (2/2)
L-STEP human missions shall maintain at all times an emergency plan for the crew to return to Earth within TBD days of a mission-threatening emergency M-1 S-9 L-STEP shall provide security for command data links M-1, M-4, M-6, M-7 S-10* L-STEP shall have sufficient consumables and power to support a minimum of three days on the lunar surface during the lunar day (human missions only) M-1, M-5 * TBR: self-imposed and may be too stringent Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Project Components & WBS
Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Homestead Mission Habitat delivery mission: soft landing, no ascent Launch Manifest: Launch 1: 3 Propulsion Modules (PMs), 2 delivered to LLO and 1 used for TLI Launch 2: Cargo Module (containing habitat) with 1 Lander-Propulsion Module (LPM) delivered to LLO and 1 PM used for TLI Autonomous (or remotely operated) docking in LLO 2 PMs and LPM expended in descent (expended LPM remains attached) Mission completes with Cargo Module atop LPM on lunar surface. Total, FY13$ 5 Propulsion Modules (incl. 1 LPM config) $193 M 2 Falcon Heavy Launches $320 M 1 Cargo Module $105 M Total $618 M Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Homestead CONOPS Diagram
Moon LLO TLI Earth Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

TaxiCab Mission Crew mission: soft landing with ascent stage Reduced PM/LPM cost estimate due to Homestead heritage Launch Manifest: Launch 1: 3 PMs, 1 used for TLI and 2 delivered to LLO Launch 2: 1 LPM and 2 PMs, 1 used for TLI Launch 3: 1 Crew Module and 2 PMs, 1 used for TLI PMs and LPM autonomously (or via remote operation) assemble in LLO Crew Module deposits 1 PM in LLO Crew Module docks with LPM / PM stack 2 PMs and LPM expended in descent to surface Remaining PM expended in ascent to LLO Crew Module docks with orbiting PM, expended in Earth return maneuver Total, FY13$ 7 Propulsion Modules (incl. 1 LPM config) $180 M 3 Falcon Heavy Launches $480 M 1 Crew Module $105 M Total $765 M Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

TaxiCab CONOPS Diagram (1)
Moon Autonomous / Remote-Operated Docking LLO TLI 1 PM expended TLI 1 PM expended Launch #1 3 PMs Launch #2 2 PMs + 1 LPM Earth Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

TaxiCab CONOPS Diagram (2)
11/8/201811/8/2018 Moon Docking with LPM/PM Stack Ascent 1 PM expended LLO Separation PM remains in LLO Descent 2 PM+LPM Expended Earth Return 1 PM expended TLI 1 PM expended Launch #3 Crew + 2 PM Splashdown Earth Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

CarePkg Mission Resupply mission with up to 5,000 kg payload capacity 1000kg required per 90 days Food Solids (200 kg) Fresh Water (30 kg) Nitrogen Replenishment and Tanks (430 kg) Oxygen for Airlock (40 kg) Disposable Clothing/Dishes (300 kg) Soft landing, no ascent Identical in mission design to Homestead mission Reduced PM cost estimate due to Homestead/TaxiCab heritage Total, FY13$ 5 Propulsion Modules (incl. 1 LPM config) $108 M 2 Falcon Heavy Launches $320 M 1 Cargo Module $105 M Total $533 M Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Launch Vehicle Cost Analysis
The use of a commercial launch vehicle, crew capsule and cargo capsule significantly reduce overall mission costs. Prices have been inflated 30% above published / reported prices for Falcon Heavy and Dragon capsule to account for unexpected changes. Propulsion Module / Lander-Propulsion Module (PM/LPM) development and production costs based on CER function* for unmanned spacecraft and inflated an additional 30% for margin. 80% Learning Curve applied to PM / LPM production. Nonrecurring FY13$M Per-Unit FY13$M 1st PM / LPM 714 52 2nd 41 5th 31 10th 25 Dragon Capsule (Crew / Cargo) 105 Falcon Heavy Launch 160 * From “Cost Estimation and Engineering Economics” lecture slides, Akin 10/6/2011 Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Operational Yearly Budget Sample
11/8/201811/8/2018 Year 1 Cost, FY13$M Yr Running Total, FY13$M 1 Homestead Mission 618 1 TaxiCab Mission 765 1383 3 CarePkg Missions 2955 (Total PM/LPMs produced: 27) ($45M Margin Remaining) Year 2 Cost, FY13$M Yr Running Total, FY13$M 1 TaxiCab Mission 705 1 CarePkg Mission 506 1211 693 1904 499 2403 (Total PM/LPMs produced: 51) ($597M Remaining For Add’l Mission or R&D) Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Delta-V Requirements ΔV (km/s) LEO GTO GEO L1 L2 L4/L5 LLO Lunar Surface WSB 3.93 3.75 4.17 (3.45)* 3.96 3.95 6.01 3.20 1.48 1.29 1.71 (1.19)* 1.51 1.50 3.56 0.75 1.34 1.87 (1.23)* 1.62 1.80 3.86 1.21 0.66(0.14)* 0.35 0.98 2.79 0.32 0.95 2.70 0.68 2.74 2.09 5.65 3.19 3.50 2.43 2.34 2.38 1.73 0.62 2.67 Limiting assumptions Two body dynamics / Patched conics Hohmann transfer between orbits Moon orbit is assumed to be circular 200km LEO, 100km LLO Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Delta-V Requirements Observations Less dV required to get to the moon from higher altitude orbits Associated payload mass penalties with launching into higher orbits 0.35km/s additional dV required on descent Corrections dV to L2 can be reduced using gravity assist at the Moon Savings on the order of 8% of dV to L1 Cost from L1 to L2 is about 140m/s Descent extra dV based on Apollo 11 premission powered descent event ~60 nautical miles (111 km) Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Times of Flight TOF (days) LEO GEO L1 L2 L4/L5 LLO Lunar Surface WSB 0.219 3.826 6.167 4.933 5.013 5.011 45.000 4.465 6.912 5.626 5.710 5.708 0.878 12.016 0.336 0.335 15.300 0.313 4.843 4.841 0.004 Calculated assuming half period of Hohmann transfer ellipse Nominal case of LEO to LLO takes days Shorter transfers available at a cost of additional dV Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Other Transfer Options
Faster than 5-day Hohmann transfer to LLO: (Values are in addition to Hohmann transfer requirements) 4 days trajectory = 0.02 km/s dV = kg propellant 3 days trajectory = 0.12 km/s dV = kg propellant 2 days trajectory = 0.56 km/s dV = kg propellant 1 day trajectory = 2.92 km/s dV = kg propellant Weak Stability Boundary (WSB) Transfer: Increased time of flight – 90 days, viable only for cargo missions Savings of km/s of dV = kg of propellant saved Propellant savings is a fraction of a single propulsion module Loss of mission flexibility Requires 90 days of ground control Descent extra dV based on Apollo 11 premission powered descent event ~60 nautical miles (111 km) Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Commercial Launch Capabilities
Vandenberg Air Force Base (VAFB) – 5 active and 7 inactive Space Launch Complexes (SLC) Cape Canaveral Air Force Station (CCAFS) – 3 active and 3 inactive SLC’s John F. Kennedy Space Center (KSC) – 2 inactive SLC’s SpaceX Falcon 9 & Heavy CCAFS SLC 40, and VAFB SLC 4-East CCAFS SLC 40 (Falcon 9) completed in approx. 2 years, cost unknown SpaceX spending $30M to refit VAFB SLC-4E for Falcon 9 and Heavy Expected capability of 16 Launches per year by 2015 (9 & H) United Launch Alliance (ULA) Delta IV Since 2000, 17 Delta IV Launches (M+ & H) Delta IV-H, 1st launch in 2004, 5 total VAFB SLC 6 (1 Launch), CCAFS SLC 37B (4 Launches) Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Payload Capabilities Delta IV Heavy Falcon Heavy Mass Limits to LEO [kg] 22,500 53,000 Mass Limit to TLI [kg] 10,600 16,000 Fairing Length [m] Fairing Diameter [m] 5.080 5.200 Ask Vinny how to tweak backgrounds Ref: Delta IV Heavy Payload Users Guide Ref: Falcon 9 Payload Users Guide Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Delta IV Heavy Homestead/CarePkg Launch # Propulsion Modules % Filled Lander Modules Crew Modules Cargo Mass to LLO (kg) 1 2 70%, 92% 70%, 75% 1,000 3 70% 4 5,900 Total 6 - 6,900 TaxiCab Launch # Propulsion Modules % Filled Lander Modules Crew Modules Cargo Mass to LLO (kg) 1 3 70%, 29%, 53% 2 70%, 92% 87%, 75% 4 5 70% 6 -550 Total - Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Falcon Heavy Homestead/CarePkg (16,000* kg total launch payload) Launch # Propulsion Modules % Filled Lander Modules Crew Modules Cargo Mass to LLO (kg) 1 3 70%, 100%, 60% 900 2 70% 5,000 Total 4 - 5,900 TaxiCab (16,000* kg total launch payload) Launch # Propulsion Modules % Filled Lander Modules Crew Modules Cargo Mass to LLO (kg) 1 3 70%, 100%, 76% 320 2 70%, 76% 70%, 74% Total 6 - *If Falcon Heavy is assumed to be capable of delivering only 14,000 kg to TLI then we add one launch for unmanned and one launch for manned, for a total of 2 additional launches Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Reliability Analysis: Launch
One Falcon Heavy 1st stage = 27 Merlin Engines in parallel 1 Merlin Engine has an assumed 97% reliability* Probability that all engines work: 43% Partial failure (without launch failure): Single-engine thrust: 556 kN Thrust to achieve liftoff: 13,720 kN  At least 25 working engines needed out of 27 Revised probability of launch success: 95% * Merlin Engine reliability is derived from “Design Reliability Comparison for SpaceX Falcon Vehicles”, Futron Corp, 2004 ( Stated reliability is expected value for the Falcon 1 (which uses a single Merlin Engine) based on historical average subsystem failure rates due to all causes (2.845%). This is an entirely speculative estimate. Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Reliability Analysis: Launch
Reliability of the launch vehicle based on the reliability of the single engine, which we estimated to be of around 97% Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Lunar Ascent Module Mass
Propellant required using Dragon’s built-in propulsion Dry Ascent Mass 6,500 kg Ascent ∆v 1730 m/s Specific Impulse (NTO/MMH) 320s Propellant Mass 4,778 kg Total Ascent Mass (+30%) 14,662 kg Propellant mass is bigger than the maximum possible for Dragon Capsule (6,000 kg*) * Possible propellant mass was estimated based on the maximum up-capability of Dragon Capsule Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Propulsion Module Configuration
Overall Mass: kg Propellant Mass: kg Based on Lunar Ascent of 6500 kg Crew Module with Samples V Capability: km/s 1 Main Engine Storable Propellant Gimbaled Regeneratively Cooled 16 RCS Engines Storable Hypergolic Propellant 3 Main Fuel Tanks 3 Main Oxidizer Tanks 3 Main Pressurant Tanks 4 Solar Arrays: W On-board Navigation Autonomous Docking Capability Launch Vehicle Stackable Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Propulsion Module Components
# Component Mass [kg]: 1 3x Docking Adapter 18.6 2 4x RCS Propellant Tanks 15.3 3 3x Pressurant Tanks 8.12 4 Structure and Gimbals 115.23 5 Propellant Management Devices Part of tank mass 6 Main Engine 98 7 3x Main Fuel Tanks 90.15 8 4x RCS Thruster Pods 9.12 9 4x Solar Arrays 18 10 3x Main Oxidizer Tanks 11 Avionics 20.34 Total Mass: 483.01 Allotted Mass: 606.15 Margin: =20.32% Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Propulsion Systems AJ K Main Engine Thrust: 43.4 kN Isp: sec Storable Hypergolic Propellant Fuel: Aerozine 50 (50/50 Hydrazine and UDMH) Oxidizer: N2O4 Chamber Pressure: 884 kPa Max Burn Duration: 444 sec 2 Solenoid Triggered Propellant Valves (Multiple Restart Capable) Thermal Management Regeneratively cooled thrust chamber Radiatively cooled nozzle extension R-21 Reaction Control Thruster Thrust: 21 N Isp: 285 sec Storable Hypergolic Propellant Fuel: MMH Oxidizer: N2O4 Chamber Pressure: 740 kPa Max Burn Duration: ~1500 sec 2 Solenoid Triggered Propellant Valves Multiple Restart Capable Thermal Management Radiatively cooled nozzle Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Reliability Analysis: Propulsion Module
Engine AJ10-118K has flown over 200 times and never failed. 98% Reliability with a 98% Confidence Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Structural Analysis: Propulsion Module
Material: Aluminum alloy E= Mpa σE = 140 Mpa =23 µm/(m*K) ρ=2700 kg/m3 Dimensions Height: 3 m Radius: 1.75 m Thickness: 2.5 mm Shell mass = 231 kg Module mass = 6205 kg Loads for Falcon 9 (worst case: Crew vehicle + 2 prop. Modules) Steady load (FOS=1.4) = Mpa Thermal load (FOS=1.4) = Mpa (ΔT=55º) Random vibration load (FOS=3) = 46.9 Mpa Total load= Mpa Safety factor MS = 174 % Further analysis needed to include other effects (like buckling) and actual data from Falcon Heavy. Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

All Cryogenic Moon Mission Example
LM Design (Single Descent/Ascent Stage + Dragon) Mass Summary [kg] Gross 42088 Propellant 26387 Dry 9201 Payload (Ascent Stage) 6500 Total Margin 5021 Possible Features Isp = 445+ sec (P&W RL10 CECE Engine) One Descent /Ascent Stage (Re-useable if Refueled) Mass on Order of Altair LM (45000kg) 1 Falcon Heavy to LEO 4.3 km/s Δ V Opportunities to Optimize PM Design (22500kg and 53000kg Variants) V1 Mass Summary [kg] Gross 22500 Propellant 15912 Dry 6587 Total Margin 2689 V2 Mass Summary [kg] Gross 53000 Propellant 39896 Dry 13100 Total Margin 5348 Possible Features Flexibility to LEO per launch w/ Delta IV-H or Falcon Heavy Mixing and Matching 1LM Round Trip (2V1 PM’s and 6 V2 PM’s) – 8 Falcon Heavy Launches LEO Staging, Single DM, Dragon Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Cryogenics Development
Boil-off Rate Passive Thermal Management Boil-off Venting Thermal Venting System (TVS) Settling/Stratification of LH2/GH2 and LO2/GO2 Axial (TRL9, DIV Stage 2, Atlas Stage 2) Radial (TRL-8, Demo on Atlas Stage 2, 2008) Boil-off Applications Heat sinking/shielding RCS propellant Cooling/Replacement (TVS) Image from Mclean et al., NASA/TP— Image from Tomsik, et al., NASA/TM— Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Cryogenics Development
Image from Mclean et al., NASA/TP— Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Landing Gear The optimal design will: Absorb the most energy in a simple and cost effective manner. It will have the lowest mass and the smallest volume maximize the ratio of energy absorption to density. minimize launch cost maximize available room in the launch payload and lunar landing module itself. Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Previous Missions Apollo 4 legs Surveyor 3 legs Surveyor 1 Landing Gear Images from: Rogers, William F. Apollo Experience Report - Lunar Module Landing Gear Subsystem. Manned Spacecraft Center, National Aeronautics and Space Exploration. Houston : s.n., 1972. Apollo 11 Landing Gear Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Previous Missions 5 Leg Landing Gear Concept For Apollo Image from: Rogers, William F. Apollo Experience Report - Lunar Module Landing Gear Subsystem. Manned Spacecraft Center, National Aeronautics and Space Exploration. Houston : s.n., 1972. Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Energy Absorption System
Options Hydro-pneumatic Suspension Plastic Deformation Structure Shock absorber Plastic deformation structure was selected because Provides one of the best energy absorption density ratio The simplest design (testing, modeling and design time), Least mass. Reliable used in Apollo missions flawlessly Cheap because we use sacrificial disposable material “Apollo Primary Strut” crushable aluminum honeycomb cartridge Image from: Rogers, William F. Apollo Experience Report - Lunar Module Landing Gear Subsystem. Manned Spacecraft Center, National Aeronautics and Space Exploration. Houston : s.n., 1972. Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Structural Design of Landing Gear
Angled Dynamic Strut Design: Stability exceeds that of vertical strut designs. Uses the least volume in payload due to its stow-ability. Less materials and size are required due to strength. Reliability is reduced due to added complexity of design. Harder to model than vertical strut designs Assumptions: Height of Engine Cutoff = 1m Velocity at Impact = 1.79m/s Number of legs = 3 Friction coefficient of lunar regolith = 0.3 therefore minimum angle for θ = tan-1(0.3) = 16.7° Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Structural Free Body Diagram
Aluminum alloys are used in the construction Apollo: The struts are straight tubular Aluminum members approximately 53 inches in length, 3.5 inches in diameter, and inch in wall thickness Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Structural Design of Landing Gear
Theta=22 degrees Outside diameter =10 cm Inside diameter = 7.5 cm Our Selection (3 legs) Aluminum Outside Ø =10 cm Inside Ø =7.5cm θ =22 ° Compared to Apollo (4 legs) Outside Ø=8.75cm Inside Ø=7.5cm the minimum main strut outside diameter for given inside diameters vs. theta is shown. This plot will help to find the minimum wall thickness acceptable for the primary struts. Our Selection (3 legs): Aluminum/ outsideØ=10 cm / insideØ=7.5cm / theta=22 degrees Compared to Apollo (4 legs): Aluminum / outsideØ=8.75cm / insideØ=7.5cm / Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Crew Module: Block Diagram

Crew Module: AOCS 3-axis stabilization with 2 fault tolerance Sun Sensors and Star Trackers for attitude determination Gyroscope for angular rates Number Mass per Item (kg) Power per Item (W) Total Mass (W) Total Power (W) Thrusters 18 2.7 2 48.6 36 Sun Sensor 1 Star Tracker 10 11 20 22 Processor 3 15 45 Gyroscope 40 22.5 120 67.5 Mass (kg) Power (W) Attitude and Orbit Control System 195.6 150 Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Crew Module: Thermal Control System
Assumptions Radiators Considering solar flux, Earth albedo radiation, Earth radiation, Moon albedo Visibility factor of Earth for 60° between local vertical and Sun’s rays Earth radiation and solar heat flux based on distances Efficiency and non-isotherm compensation factors applied Heaters Power required to heat to desired temperature (295K) MLI insulation 20 layers with εeff = 0.035, specific mass ρ = 1.2 kg/m2 Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Hot Case Cold Case Solar Heat Flux (W/m2) 1,371.12 Earth Albedo 0.39 0.1 Orbit (km) 300.00 384,405 Earth Radiation (W/m2) 216.18 0.06 Visibility Factor Earth 0.8 0.05 Moon Albedo 0.07 Effective Sink Temperature (K) 280.67 256.26 Internal Heat Crew (W) 3x180 Electronics (W) 2,500 Total Power (W) 3040 Heater Sizing Power Required (W) 166.39 Heater Power (W) 57 Heater Mass (W) 0.3 Number of Heaters 3 Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Liquid Loop Radiator Heat Pipe Radiator Condensing Radiator Typ. Efficiency 0.85 0.9 0.92 Absorptivitiy (BOL) 0.15 Emmisivity (BOL) Temperature of radiator surface (K) 295 Non-isotherm compensation factor 0.7 1 Specific Mass (incl. structure (kg/m2) 14.00 13.00 17.00 Sizing case (Hot) Needed radiator area (m2) 73.38 52.73 47.46 Radiator mass (kg) 685.53 806.82 Mass (kg) Power (W) Thermal Control System 737.54 167 e.g. 2 deployable heat-pipe radiators with 3m x 4.5m, 3 heaters and MLI insulation Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Crew Module: Thermal Protection System
Direct re-entry from Moon return Semi-ballistic re-entry with ablative heat shield Use of heuristic mass approximation Total Heat Load Qtot = 410 MJ/m2 mtps/mtot = 0.091*(Qtot) TPS Mass after iterations mtps = 1285 kg Laub, B. & Venkatapathy, E.: Thermal protection system technology and facility needs for demanding future planetary missions Proceedings of the International Workshop Planetary Probe Atmospheric Entry and Descent Trajectory Analysis and Science, 6-9 October 2003 Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Crew Module: Electrical Power System
Primary Solar Power - 2 Ultraflex solar arrays Combined 2500W output Physical Characteristics: Combined Mass: 18 kg Combined Area: m2 Contingency Power – Fuel Cells with consumable for 5 days at 2500 W Estimated Fuel Cell Mass: 100 kg Estimated Consumable Mass: Whr/kg Would require an equivalent 1700 kg of Lithium Ion batteries Supplemental Power – 50 kg rechargeable Lithium Ion batteries Provide supplemental power during power bursts above solar array capability 8750 Whr capacity (175 Whr/kg) Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Crew Module: Mass Budget
Mass [kg] % of Total Mass with Margin [kg] AOCS 195.6 4% 254.28 TCS 735.75 16% 956.47 Structure 1180 26% Avionics 255.39 6% 332.01 ECLSS 600 13% 780.00 EPS 270.00 351.00 Communication 22.75 1% 29.58 TPS 28% Subsystem Margin 30% Total Dry Mass (kg) Margin Total Dry Mass with Margin (kg) Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Crew Module: Power Budget
Power [W] % of Total Power with Margin [W] AOCS 150 10% 195.00 TCS 167.00 11% 217.10 Avionics 300 17% 390.00 ECLSS 650 44% 845.00 Communication 250 325.00 Subsystem Margin 30% Total Power (W) Margin Total Power with Margin (W) Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Dragon Capsule Crew Size up to 7 (ISS orbit) Diameter Height CM Height Trunk 3.6m 2.9m 2.3m Habitable Volume 10 m3 Pressure 13.9 – 14.9 psi Power W (average) 4000W (peak) GLOW CM Dry Mass CM 10,200 kg 4,200 kg Cost ca. 140 $M per flight (at least 4 flights per year, full capsule of 7) Heat shield capable of direct re-entry from Moon Landing gear and advanced thrusters for future land landing on Earth are planned Image credit: SpaceX Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Dragon Layout Current Layout Designed for 7 people Designed for hours Required Change: 3 people, for 5 days CTB Storage Water Storage Waste Storage Sink/Food Prep Source: Popular Mechanics Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Modified Layout Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Water Management Types of H2O Drinking Hygiene In food/food prep Trip Design 200 kg H2O for trip Sized as sphere with 75 cm diameter Required: 2 Tanks (Fresh, Used) Habitat Regenerable open-loop system VCD Disposables still recommended Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Nitrogen Replenishment
Considerations N2 is about 80% of Earth-like atmosphere N2 losses are about 1% per day Options Nitrogen-bearing compounds Cryogenic tanks (rejected due to temperature requirements) High pressure tanks (3000 psi) Density of N2 is 255 kg/m3 Tank Sizing Transport: 8.4 kg N2, 5.9 kg tank (0.033 m3) Habitat: Roughly 2 kg per day required between gas and tank Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Oxygen Recovery Requirement: 31 kg O2 for DRAGON Based on ISS figures (0.86 kg/person-day) Intended to maintain 79% N2, 21% O2 Includes 20% margin High Pressure Storage (3000 psi) Density of 292 kg/m3 Tank Sizing Volume required: 0.11 m3 Tank mass: 12.5kg Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

CO2 Reduction Trade System 5-Day Mass (kg) Volume (m3) Power (W) LiOH Cartridges 28 0.13 Sabatier Reaction 270 9 800 Bosch Reaction 2100 11.7 5000 SAWD 51 0.51 150 4BMS 90 0.33 680 2BMS 48 0.27 250 ACRS 180 0.3 400 METOX 106 0.17 1000 Consideration: Must remove 1 kg/person-day of CO2 LiOH System chosen for low mass/volume Rechargable METOX system for Habitat (106 kg with oven) Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

CO2 Removal – Long Term Trade

Consumables – Mass Budget
Amount per person, per-day (kg) For 3 people, 5 days (kg)* Food Solids 0.62 12 Drinking Water 1.6 29 Water in Food 1.15 21 Sanitary Water 7.3 132 Water (Food Prep) 0.75 14 Total H2O** 10.8 200 H2O Tanks (2) 50 Disposables 3 45 Oxygen 0.85 31 Oxygen Tanks psi) 12.5 Nitrogen (total) 0.1052 1.3 Nitrogen Tanks 0.90 CO2 Removal - LiOH Cartridge 1.75 32 *Includes a 20% Mass Margin **Doesn’t include 18 kg/day Disposable Dishes/Clothes *** Oxygen/Nitrogen masses are for 10 days Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Psychosocial Considerations
Notable Case Studies Hualta Cave Expedition, Ben Franklin Submersible Accommodations Windows for both DRAGON and habitat Crew Space Habitat: 34.4 m3 (3.4 m3 private) per person Dragon Capsule 3.3 m3 per person (50% more than Apollo) Source: NASA Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Radiation Physiology Concerns Must maintain levels <25 rem (0.25 Sv) per month Possible: Radiation carcinogenesis, mutagenesis, and cataracts Increased cancer risk +3% cancer risk above 50 rem/year Day-to-day levels: 1.2 mSv per day (as reported by Apollo astronauts) Would require 415 days in space Solar Particle Events Recommendation: 50 g/cm2 of regolith for lunar habitat Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

EVA Operations Rover Operations Provisions for 1 EVA/day Airlock Egress Losses: kg O2, kg N2 Must be maintained at: 8 psi 68/32 N2/O2 ratio High Pressure Storage of Gasses: 3-day Mission 15-day Mission 90-day Mission Mass of N2 (with tanks) [kg] 8.2 22 105 Volume of N2 Tanks [m3] 0.019 0.05 0.25 Mass of O2 3.5 8.6 41 Volume of O2 [m3] 0.01 0.021 0.1 Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Lunar Habitat Utilize existing HDU structural concept 5 m diameter vertical orientation cylinder with ellipsoidal end caps and four docking ports Maintain 3 active docking ports Two for rovers & one for airlock Provide an inflatable upper habitat 5 m diameter & ellipsoidal upper surface Support crew of four 95% American males Worst Case Support crew for 90-day mission between resupply Account for 2 crew EVA for 6 hours every day Account for 2 6-day rover sorties per day cycle Note: 1 Lunar day cycle is 14 Earth days Account for 1 6-day rover sortie per night cycle Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Living Space 11/8/201811/8/2018 Shower Studies Individual studies Desks can be stowed when not in use Pull down flap allows for privacy Shower and changing area Combined food preparation and common areas Allows for open floor plan Food prep area Common area CTB Shelves Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Living Space 11/8/201811/8/2018 Location Area (m2) Individual Studies (4) 1.07 (each) Food prep/common area 7.75 Shower/changing area 1.01 CTB Shelves 0.44 Total Floor Space 13.48 Section view of living area Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Sleeping Quarters 11/8/201811/8/2018 Type Pros Cons Volume Living Space Dome Loft Beds -Privacy -Controlled Darkness -Complex -Low Ceiling 10.9 m3 ~ 2.7 m3 per person 36.6 m3 Bunk Beds -Overhead Space -Simpler Radiation Contingency -Limited Privacy -Less wall space available 12 m3 ~ 1.7 m3 per person 24.6 m3 Hammocks -Simple -Open Space -No Privacy / Personal Space 2.1 m3 ~ 0.5 m3 per person 34.5 m3 Tent-like entrance flap Personal laptop 6 W LED dimmable lamp CTB Storage 20 per quarter when packed 9 nominal + 1 inhabitant 5 comfortably + 1 inhabitant Memory foam base Floor serves as sleeping area Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Long Term Lunar Power Requirements
Conceptual Habitat Requirements (W): Quiescent Average/Nominal Maximum Implicit Habitat 1315 3469 7038 Barely There Habitat 1135 3929 6189 Apollo Lunar Module* 74676Whr / 97hrs 770 Image from NASA Rover Power Requirements (W): Vehicle Driving Power MOONSTER 2200 Lunar Roving Vehicle 4x190+2x75 910 Image by Seth Napora Image by NASA/Dave Scott * Based on Apollo 15 Mission Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Energy Storage For Lunar Night (354 hrs)
Required Energy Storage (Whr) Quiescent Nominal/Average Maximum Implicit Habitat 465510 Barely There Habitat 401790 MOONSTER *88500 **259600 Lithium Ion Battery Mass (kg) Quiescent Nominal/Average Maximum Implicit Habitat 7017 14236 Barely There Habitat 7948 12519 MOONSTER *505 **1483 Fuel Consumables Mass (kg) Quiescent Nominal/Average Maximum Implicit Habitat 158 416 845 Barely There Habitat 136 471 743 MOONSTER 30 88 Image by James Humphreys * For 8 hours of operation per day at 750 W ** For 8 hours of operation per day at 2200 W Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

One 90 Day Mission per year
Option 1 Supplemental Ultraflex Solar Array 10000 W provided by 45 sq m, 71 kg array Fuel Cell Consumable Requirements Days Mass (kg) Quiescent 137.5 1370 Manned Operation 45 3382 4752 Option 2 Nuclear Reactor $ M - $1B SP – 100 SAFE-400 100 kW 5422 kg 512 kg SAFE 30 Reactor test setup. Image by David Poston Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Questions? Pinky, are you pondering what I’m pondering?
Um, I think so, Brain. We won’t have a final if we send Dr. Akin to the moon.

References Avionics AOCS Docking Landing Gear
EADS Astrium Satellite Equipment AOCS ADCS Design, Lecture 22, Mechanical, Materials and Aerospace Engineering, Armour College Docking International Docking System Standard (IDSS) Interface Definition Document (IDD) Revision A may 2011 Landing Gear Lunar and Planetary Institute. Lunar Mission Summaries. Lunar Science and Exploration. [Online] [Cited: October 10, 2009.] Rogers, William F. Apollo Experience Report - Lunar Module Landing Gear Subsystem. Manned Spacecraft Center, National Aeronautics and Space Exploration. Houston : s.n., 1972. Blanchard, Ulysse J. Evaluation of a Full-Scale Lunar-Gravity Simulator by Comparison of Landing-Impact Tests of a Full-Scale and a 1/6-Scale Model. Langley Reserach Center, National Aeronautics and Space Administration. Hampton : s.n., 1968. 5. Predicting the Impact Force on the Landing Gear of a Lander at a Lunar Landing. Tateyama, K., et al. 3, s.l. : National Office for Reserach and Technology, 2008, Vehicles and Mobile Machines, Vol. 1, pp Lunar Landing Gear, 2009, Queens University, team 27 final project.

References Electrical System
Apollo Lunar Module Electrical Power System Overview, Michael Interbartolo, Johnson Space Center, 2009 Battery Fundamentals and Operations, Doris L. Britton; Thomas B. Miller, NASA Glenn Research Center, 2000 Nuclear Marine Propulsion (ME 405), M. Ragheb, University of Illinois at Urbana-Champaign, 2011 The Safe Affordable Fission Engine (SAFE) Test Series, Mellisa Van Dyke et al., NASA/JPL/MSFC/UAH 12th Annual Advanced Space Propulsion Workshop, 2001 UltraFlex Solar Array Systems, ATK-Goleta Alliant Techsystems Inc., 2011 Three Newly Designed Tracking and Data Relay Satellites To Help Replenish Existing On-Orbit Fleet, FS GSFC, NASA Goddard Space Flight Center, 2001 D. MacDonnell, "Communications Analysis of Potential Upgrades of NASA's Deep Space Network." University of Maryland, April, 2000

References Thermal Control System
Larson, W.J.; Wertz, J.R. (Eds.): Space Mission Analysis And Design. 2nd Ed., Microcosm, Torrance, CL, 1992 Woodcock, G.R.: Space Stations And Platforms. Orbit Book Company, Malabar, FL, 1986 Chambliss, J. Et Al.: An Overview Of The Redesigned Space Station Thermal Control System. SAE Technical Paper , SAE International, Commonwealth Drive, Warrendale, USA, 1994. Mechanical Engineering Of Space Programmes - Thermal Control Standard. ECSS-E-30-00, ECSS Secretariat, ESA-ESTEC, Requirements & Standards Division, Noordwijk, The Netherlands, Thermal Control Handbook. RT523, DASA-RIT, Daimler-benz Aerospace, Bremen, Hanford, A.J.; Ewert, M.K.: An Assessment Of Advanced Thermal Control System Technologies For Future Human Space Flight. SAE , SAE International Conference On Environmental Systems, 26th, Monterey, California, USA, July 8-12, 1996. Spacecraft Thermal Control. Nasa-sp Space Vehicle Design Criteria. May 1973. Gilmore D.G. Et Al: Satellite Thermal Control Handbook, ISBN The Aerospace Corporation Press, El Segundo, CA, USA, 1994. Esa, Thermal Control & Life Support Divison, Spacecraft Thermal Control Design Data, Esa Pss , Nordwijk, The Netherlands, 1989.

References Launch Vehicles Thermal Protection System
Falcon 9 User’s Guide 2009 “Design Reliability Comparison for SpaceX Falcon Vehicles”, Futron Corp, 2004 ( About the AJ10-118K engine: SpaceX Brochure V7 US Spacesuits. Chichester, UK: Praxis Publishing Ltd p. 32 Thermal Protection System Laub, B. & Venkatapathy, E.: Thermal protection system technology and facility needs for demanding future planetary missions Proceedings of the International Workshop Planetary Probe Atmospheric Entry and Descent Trajectory Analysis and Science, 6-9 October 2003, Lisbon, Portugal. Edited by A. Wilson. Lecture notes: Wiedereintrittsprobleme und Aerothermodynamik, University of Stuttgart, 2006. Jr. John D. Anderson. HYPERSONIC AND HIGH TEMPERATURE GAS DYNAMICS. McGraw-Hill Book Company, 1989. St. Leger Jumes E. Puulosky und Leslie G and Lyndon B. Apollo experience report - thermal protection subsystem. Technical report, Johnson Space Center, NASA, 1974.

References Orbital Mechanics
“Apollo Lunar Descent and Ascent Trajectories,” the AIAA 8th Aerospace Sciences Meeting, 1970 Wendell W. Mendell, “Strategic Considerations for Cislunar Space Infrastructure,” NASA Johnson Space Center E. A. Belbruno and J. P. Carrico, “Calculation of Weak Stability Boundary Ballistic Lunar Transfer Trajectories,” AIAA, 2000

Rover Image by Seth Napora Image by NASA/Dave Scott MOONSTER 200 kg
2200 W max power Lunar Roving Vehicle 210 kg 910 w Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

Orion Multi-Purpose Crew Vehicle
Crew Size up to 4 (Lunar orbit) Dimensions (CM) Dimensions (SM) 5.03 m diameter 3.30 m height Habitable Volume 10.22 m3 Pressure Power up to 9.15 kW Crew Module ∆v Service Module ∆v 50 m/s 1,855 m/s Gross Lift-Off Weight CM Gross Lift-Off Weight SM Spacecraft Adapter 8,500 kg 12,337 kg 581.2 kg Cost Lander and ascent capability (Lander required) Lunar sample return: 99.8 kg ATV considered as Service Module Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Image credit: NASA Budgets

Altair (Lunar Lander) Crew Size up to 4 Height Width (at tanks)
Width (footpads, diag.) 9.9m 8.8m 14.9m Habitable Volume 17.5 m3 Pressure Power Surface Duration 7 Days (Sortie missions) Up to 210 Days (Outpost mode) Ascent Stage Mass Descent Stage Mass Manned Variant Payload Cargo Variant Payload 6,141 kg 37,045 kg 500 kg (additional to crew) 14,500 kg Cost Image credit: NASA Capable of global landing sites on the moon Outpost mode: Altair brings crew down to the surface and stays unmanned till crew returns Cargo variant capable of automated landing on pre-selected site Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

ATV Development cost: €1.35B Unit cost: $300M
4 main engines, each 490 N thrust. Solar arrays: 22.3 span. (4800W) Manned Concept: Automated re-entry vehicle. Aproximated values Dry mass 9800 kg PL 6600 kg Propellant 2500 kg Launch mass 20400kg Lenght 10.3 m Diameter 4.5 m Payload Up to: Dry cargo 3200 kg Drinking water 840 kg Fuel for ISS 860 kg Support fuel 4000 kg Air 100 kg Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

HTV Development cost: $680M Unit cost: $220M
4 main engines, each 500 N thrust. Solar panels on main body Not automatic, needs robotic arm to dock. Vehicle mass 10500 PL 6000 kg Launch mass 16500kg Lenght 10 m Diameter 4.4 m Hatch 1.2x1.2 Payload Up to: Pressurized 5200 kg Unpressurized 1500 kg Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

Cygnus Docks with robot arm Height: 3.66 m /4.86 m Diameter:
Weight (dry): 1,500 kg/ 1,800 kg Volume (pressurized): 18.9 m3/ 26.2 m3 Delivered Payload: 2,000 kg/ 2,700 kg Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

Progress Automated Height: 7.23 m (23.72 ft) Diameter:
Progress M Progress M1 Total payload limit 2,350 kg 2,230-3,200 kg Maximum pressurized (dry) cargo 1,800 kg Launch mass 7150 kg Maximum water 420 kg up to 300 kg in cargo module Maximum air or oxygen 50 kg 40 kg Propellant for refueling 850 kg 1,700 -1,950 kg Propellant for ISS 250 kg kg Height: 7.23 m (23.72 ft) Diameter: 2.72 m (8.92 ft) Volume: 7.6 m3 (268 ft3) Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

General Requirements Minimum of a 685 mm transfer passageway at SCS
seal-on-seal mating at HCS Maximum Temperature difference between the two mating interfaces 55°C Materials and Surface Finishes should have stiffness and hardness comparable to commonly used metal alloys. Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

Communications – S Band
DSN TDRSS dB Frequency Hz f 2.15E+09 93.32 Wavelength m l 0.1395 -8.55 Diameter of Tx Antenna d(t) 3.00 4.77 6.00 7.78 Tx Gain G(t) 2.85E+03 34.55 1.14E+04 40.57 Tx Power W P 90.00 19.54 150.00 21.76 Tx Beamwidth deg 3.26E+00 1.63E+00 2.27E+07 1.14E+07 Slant Range D 4.00E+08 86.02 Diameter of Rx Antenna d(r ) 34.00 15.31 4.57 6.60 Data Rate bits/sec R(b) 3.00E+08 84.77 3.50E+07 75.44 Link Margin 2.97 3.11 Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

Communications – Ka Band
DSN TDRSS dB Frequency Hz f 3.20E+10 105.05 Wavelength m l 0.0094 -20.28 Diameter of Tx Antenna d(t) 0.39 -4.09 0.91 -0.41 Tx Gain G(t) 1.07E+04 40.28 5.81E+04 47.64 Tx Power W P 25.00 13.98 Tx Beamwidth deg 1.68E+00 7.21E-01 1.17E+07 5.03E+06 Slant Range D 4.00E+08 86.02 Diameter of Rx Antenna d(r ) 34.00 15.31 4.57 6.60 Data Rate bits/sec R(b) 3.00E+08 84.77 3.00E+07 74.77 Link Margin 3.14 3.07 Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

Storage Cargo Transfer Bags (CTB’s) Storables
Dimensions: 0.5 m x 0.5 m x 0.25 m Storables Food Disposable Clothes Disposable Dishes LiOH Canisters Personal Belongings Storage Requirements: 6 CTB’s

Considerations for NEO Exploration
“Taxi” to accommodate longer duration missions Include fitness accommodations More volume/astronaut (~15 m^3/person) Additional Radiation Shielding EVA for asteroids Different Subsystems METOX for CO2 Removal Regenerable Water Reclamation

Faster Transfers Overview Mission Objectives Program Mgmt
Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

Weak Stability Boundary (WSB) Transfer
Alternative to “Direct” Lunar transfer Trajectory to Earth-Sun L1 Naturally falls back into Earth-Moon system Increased time of flight – 90 days Apoapsis – 1.5 million km Can save 25% in required dV for lunar capture Transfer in ECI Frame Ref: E. A. Belbruno and J. P. Carrico, Calculation of Weak Stability Boundary Lunar Transfer Trajectories, AIAA, 2000 Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

Merits of WSB Time of flight of 90 days (compared to 5 days of direct trajectory) means it is valid for unmanned cargo missions only. Total dV from LEO to Moon (Direct): 6.01 km/s Total dV from LEO to Moon (WSB): km/s Savings of km/s in dV km/s of dV = kg of propellant saved While the payload mass savings is attractive, it was decided that this trajectory would not be used due to associated cost with mission control for 3 months. Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

Orbital Mechanics Asteroids
Available C3 as a function of quoted useful mass for our launch vehicles For our crew capsule mass, found viable launch windows to 2004 MN4 (Apophis) as an example Departure dates from Jan 1, 2013 to Jan 1, 2020 Demonstrating the ability of ELVs to deliver humans to NEO Assuming no return fuel Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

Orbital Mechanics NEA Launch Window: Delta IV Heavy
Scenario Launch Date (MM-DD-YYYY) Time of Flight (Days) 1 56 2 88 3 138 4 146 5 130 Cases 1-5 are the shortest times of flight for different launch windows Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

Orbital Mechanics NEA Launch Window: Falcon Heavy
Scenario Launch Date (MM-DD-YYYY) Time of Flight (Days) 1 56 2 88 3 138 4 146 5 130 The shortest times of flight from Falcon Heavy ranges from 12 to 24 days shorter compared to using Delta IV Heavy Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

Modified Layout <CAD>

Dragon Capsule Avionics
VxWorks Operating System Two-fault tolerant avionics system Dragon Capsule Fully autonomous rendezvous and docking with manual override capability in crewed configuration Reaction control system with 18 MMH/NTO thrusters designed and built in-house; these thrusters are used for both attitude control and orbital maneuvering Designed for water landing under parachute for ocean recovery Payload RS-422 serial I/O, 1553, and Ethernet interfaces (all locations) IP addressable payload standard service IMU Specifications: Attitude Determination: < 0.004° w.r.t. inertial frame Attitude Control: 0.012°/axis during station-keep Attitude Rate: <0.02°/sec/axis during station-keep Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

Dragon Capsule Communications
Links via TDRSS and ground stations Fault tolerant S-band telemetry & video transmitters Onboard compression & command encryption/ decryption Command uplink: 300 kbps Telemetry/data downlink: 300 Mbps (higher rates available) Inertial Measurement Units, GPS & Star Trackers Overview Mission Objectives Program Mgmt Launch Vehicle Power & Thermal Avionics Crew Thermal Budgets

Docking

International Docking System Standard (IDSS)
It is intended for: International Space Station Lunar exploration Crew rescue International cooperative missions

Vehicles that Qualify Light vehicles in the range of 5-8K kg
Medium vehicles in the range of 8-25K kg Variations: - dock to each other, - to large space complexes in the range of K kg -to large earth departure stages in the range of K kg

2-stage approach to docking
The first stage establishes the initial capture of the docking vehicles, and is performed by the Soft Capture System (SCS). The second stage of docking is performed by the Hard Capture System (HCS). The HCS performs structural latching and sealing The docking operation needs to be completed within a maximum time to ensure a safe docking operation.

SCS & HCS The SCS interface consists of a capture ring, guide petals, magnets, and or mechanical latches, magnetic striker plates, mechanical latch strikers, sensors and sensor strikers. The HCS uses active hooks to engage opposing passive hooks to provide the structural connection and pressure seal compression. The HCS interface consists of a tunnel, 12 active/passive hook pairs on each side, dual concentric pressure seals, fine alignment guide pins and holes, sensors, sensor strikers, separation system, and resource umbilicals.

Docking Interface - Axial View

Docking Interface - Cross Section

General Requirements Minimum of a 685 mm transfer passageway at SCS
seal-on-seal mating at HCS Maximum Temperature difference between the two mating interfaces 55°C Materials and Surface Finishes should have stiffness and hardness comparable to commonly used metal alloys.

HCS Maximum Mated Loads

Vehicles that Qualify Light vehicles in the range of 5-8K kg Medium vehicles in the range of 8-25K kg Variations: dock to each other, to large space complexes in the range of K kg to large earth departure stages in the range of K kg What??? Introduction Mission Overview Orbital Mechanics Launch Vehicle Propulsion Module Landing Gear Crew Module Human Factors & Life Support Habitat & Power Systems

Lunar Settlement, Transportation, and Exploration Program (L-STEP)

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