Introduction to the Altair Project Lauri N. Hansen, Project Manager

NASA’s Exploration Roadmap 05 06 07 08 09 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 Science Robotic Missions Mars Expedition Design Mars Expedition Design Lunar Outpost Buildup Lunar Robotic Missions Human Lunar Return Surface Systems Development Altair Development Early Design Activity Ares V Development We need to select a chart or modify this one to be able to show the scope of this assessment i.e ISS IOC/FOC Earth Departure Stage Development Orion Production and Operations Orion Development Initial Orion Capability Ares I Development Space Shuttle Ops Commercial Crew/Cargo for ISS ISS Sustaining Operations 2 2 SSP Transition… 2 2 2

Components of Program Constellation Earth Departure Stage Crew Exploration Vehicle Heavy Lift Launch Vehicle Lunar Lander Crew Launch Vehicle

Typical Lunar Reference Mission MOON Vehicles are not to scale. Ascent Stage Expended 100 km Low Lunar Orbit Lander Performs LOI Earth Departure Stage Expended Service Module Expended Low Earth Orbit CEV EDS, Lander Direct Entry Land Landing EARTH

Lunar Lander and Ascent Stage 4 crew to and from the surface Seven days on the surface Lunar outpost crew rotation Global access capability Anytime return to Earth Capability to land 15 to 17 metric tons of dedicated cargo Airlock for surface activities Descent stage: Liquid oxygen / liquid hydrogen propulsion Ascent stage: Hypergolic Propellants or Liquid oxygen/methane

Configuration Variants Sortie Variant 45,000 kg Descent Module Ascent Module Airlock Outpost Variant 45,000 kg Descent Module Ascent Module Cargo Variant 53,600 kg Descent Module Cargo on Upper Deck

Initial Project Structure Using a Smart Buyer approach Develop a preliminary government design Coming out of initial design effort, have independent reviews and solicit industry input on initial design Continue to refine design & requirements based on industry input Using knowledge gained from in-house design effort, create draft vehicle design requirements In FY10 have a vehicle requirements review, and baseline requirements Between 2009 – 2011, build hardware/test beds to mature confidence in path for forward design (lower risk of unknown surprises) Continue to mature design in-house until PDR timeframe (tentative)

Detailed Approach for Design Team Initial task was developing a preliminary in-house design: 6-9 mth duration Agency wide team Expert designers from across the agency Minimalist approach – add people on a case-by-case basis, only as needed Subsystems, not elements Approximately 20 – 25 people on the core team Co-located initially (approx 2 months) Working from home centers following initial co-location period Another 20-25 FTE distributed across the Agency (not co-located) Focused on Design (‘D’ in DAC) Developed detailed Master Equipment List (over 2000 components) Developed detailed Powered Equipment List Produced sub-system schematics NASTRAN analysis using Finite Element Models Performed high-level consumables and resource utilization analysis Sub-system performance analysis by sub-system leads Keep process overhead to the minimum required Recognizing that a small, dynamic team doesn’t need all of the process overhead that a much larger one does But…. It still needs the basics

“Minimum Functionality” Approach “Minimum Functionality” is a design philosophy that begins with a vehicle that will perform the mission, and no more than that Does not consider contingencies Does not have added redundancy (“single string” approach) Altair has taken a Minimum Functionality design approach Provides early, critical insight into the overall viability of the end-to-end architecture Provides a starting point to make informed cost/risk trades and consciously buy down risk A “Minimum Functionality” vehicle is NOT a design that would ever be contemplated as a “flyable” design! The “Minimum Functional” design approach is informed by: NESC PR-06-108, “Design Development Test and Evaluation (DDT&E) Considerations for Safe and Reliable Human Rated Spacecraft Systems CEV “Smart Buyer” lessons learned Recent CEV “Buyback” exercises

p711-B Lunar Lander* Altair Project Vehicle Concept Characteristics Lander Performance Crew Size: 4 LEO Loiter Duration: 14 days Surface stay time: 7 days (sortie) 180 days (outpost visit) Launch Shroud Diameter: 8.4m Lander Design Diameter: 7.5 m Launch Loads: 5 g axial, 2 g lateral Crewed Lander Mass (Launch): 45,586 kg Crewed Lander Mass (@TLI): 45,586 kg Crew Lander Payload to Surface: 500 kg Project Manager’s Reserve: 3009 kg Crew Lander Deck Height: 6.97 m Cargo Lander Mass (Launch): 53,600 kg Cargo Lander Mass (@TLI): Not applicable. Cargo Lander Payload to Surface: 14,631 kg Project Manager’s Reserve: 2227 kg Cargo Lander Height: 6.97 m EDS Adapter Mass: 860 kg (Not included in numbers above, includes growth and Manager’s Reserve) Crew Lander LOI Delta V Capability: 891 m/s Cargo Lander LOI Delta V Capability: 889 m/s Crew/Cargo Plane change and Loiter (Post CEV sep, 1 degree): 28.4 m/s PDI Delta V Capability: 19.4 m/s Crew Descent Propulsion Delta V Capability: 2030 m/s Cargo Descent Propulsion Delta V Capability: 2030 m/s TCM Delta V Capability (performed by RCS): 2 m/s Descent Orbit Insertion Capability (performed by RCS): 19.4 m/s Settling Burn Requirement (performed by RCS): 2.7 m/s Descent and Landing Reaction Control Capability: 11 m/s Ascent Delta V Capability 1881 m/s Ascent RCS Delta V Capability: 30 m/s Vehicle Concept Characteristics Ascent Module Diameter: 2.35 meters Mass (at TLI): 6,128 kg Main Engine Propellants: N2O4/MMH Useable Propellant: 3007 kg # Main Engines/Type: 1/Derived OME/RS18 (Pressure Fed) Main Engine Isp (100%): 320 sec Main Engine Thrust (100%): 5,500 lbf RCS Propellants: N2O4/MMH Useable Propellant: Integrated w/main # RCS Engines/Type: 16/100 lbf each RCS Engine Isp (100%): 300 sec Airlock Pressurized Volume: 7.5 m^3 Diameter: 1.75 m Height: 3.58 m Crew Size: 2+ Altair Project Descent Module (crewed) Mass (at TLI): 38,002 kg Main Engine Propellants: LOX/ LH2 Useable Propellant: 25,035 kg # Main Engines/Type: 1/ RL-10 Derived (Pump Fed) Main Engine Isp (100%): 448 sec Main Engine Thrust (100%): 18,650 lbf RCS Propellants: N2O4/MMH # RCS Engines/Type: 16/100 lbf each RCS Engine Isp (100%): 300 sec Descent Module (cargo) Mass (at TLI): 38,970 kg Useable Propellant: 26,611 kg *ENVISION parametrically sized “polar” lander concept informed by the LDAC-1 Starworks activity with selected additional redundancy and delta-v's that are representative of realistic trajectories, but not optimized for Thrust to Weight.