Transcript Slide 1

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
Science Robotic Missions
21
22
23
24
25
Mars Expedition Design
Mars Expedition Design
Lunar Robotic Missions
Lunar Outpost Buildup
Human Lunar
Return
Surface Systems Development
Altair Development
Early Design Activity
Ares V Development
Earth Departure Stage Development
Orion Production and Operations
Orion Development
Ares I Development
Initial Orion Capability
Commercial Crew/Cargo for ISS
ISS Sustaining Operations
Space Shuttle Ops
SSP Transition…
2
2
2
Components of Program Constellation
Earth Departure
Stage
Crew Exploration
Vehicle
Heavy Lift
Launch
Vehicle
Crew Launch
Vehicle
Lunar
Lander
3
Typical Lunar Reference Mission
MOON
Vehicles are not to scale.
100 km
Low Lunar
Orbit
Ascent Stage
Expended
Lander Performs LOI
Earth Departure
Stage Expended
Service
Module
Expended
CEV
EDS, Lander
Low
Earth
Orbit
Direct Entry
Land Landing
EARTH
4
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
5
Configuration Variants
Sortie Variant
Outpost Variant
Cargo Variant
45,000 kg
45,000 kg
53,600 kg
Descent Module
Ascent Module
Airlock
Descent Module
Ascent Module
Descent Module
Cargo on Upper Deck
6
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)
7
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
8
“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
9
p711-B Lunar Lander*
Lander Performance
Altair Project
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
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+
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
Ascent Delta V Capability 1881 m/s
Ascent RCS Delta V Capability: 30 m/s
*ENVISION parametrically sized “polar” lander concept informed by the LDAC-1 Starworks activity with selected additional redundancy and deltav's that are representative of realistic trajectories, but not optimized for Thrust to Weight.
10