AS4102RPT00002 Return to the Moon, T 12 years and Counting

1
AS-4102-RPT-00002Return to the Moon,T- 12
years and Counting

• Andrews Space, Inc.
• January 12th, 2009

2
Altair Study Overview

• We were given NASAs Lunar Design Architecture
  Configuration 1 (LDAC-1) and supporting data,
  and were told to analyze, verify, and recommend
  improvements.
• Emphasis was to be on
• Improving crew safety (and mission reliability)
• Improving payload capability
• We added
• Improving operations capability (man-hours on
  surface/day)
• Not every issue could be explored with the
  limited funds available so different teams
  emphasized different areas. Andrews Space pursued
  propulsion and human factors as areas of emphasis.

3
NASA LDAC-1 Lander (Sortie Mission)
4
Altair LDAC-1 Configuration Sizing
10.582 m
5
Altair Lander (Right view)
Ascent Module
Airlock
Descent Module
6
Altair LDAC-1 Masses

• Mass at Separation from Earth Departure Stage
  (EDS) 45,000 kg
• Orion Mass during Midcourse and LOI
  Maneuvers 20,185 kg
• LDAC-1 Nominal Landed Mass 21,300 kg
• LDAC-1 Nominal Usable Payload 3,450 kg

7
Andrews Altair BAA Lunar Lander Study Plan
Hamilton Sundstrand
Draper Lab
Ball Aerospace
Pratt Whitney Rocketdyne
Phase IA Inputs
LDAC-1 Performance / Mass Assessment
Reliability Process Assessments LDAC-1
LDAC-1A
Trades
2nd Set Trades / Configurations
Minimum Functional Assessment of LDAC-1
Engines RL-10B2,Habitability, Avionics, Thermal
Control, etc.
Engines AM, RL-10B2, Propellants,Tank
Structures, Habitability, Power, Landing,
Thermal, etc.
STUDY INITIATION
INTERIM BRIEFING
FINAL BRIEFING
Final Mass Reliability Numbers
LDAC1-B Final Design Recommendations
Management Approaches
Phase IIA
Phase IIB
Andrews Iterated Specifics of the Study, Based on
Information Gained as the Study Progressed
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Study Overview

• Phase I Interim Study
• Minimum functionality assessment
• We used a systematic functional analysis method
  to analyze lunar lander for minimum
  functionality, and to iterate the LDAC-1 design.
• Minimum mass assessment / Minimum mass
  recommendations
• We used the functional allocation and minimum
  functionality definition to identify and
  recommend both a minimized mass option.
• Safety / Reliability Analysis
• We assessed all phases of the mission and
  identified design safety drivers, assessed the
  mission and made reliability improvements
• Phase II Final Report. Based on input from
  Interim Report
• Executed Several Different Trades to refine
  design LDAC-1B for performance and LDAC-1C for
  safety
• Study Configurations
• Tank Study
• Engine Trades
• Habitability Assessment
• LOC Buy Back / Improvements
• Produced Final Mass and Reliability Assessment
  with Recommendations

9
Andrews Functional Analysis Process
Requirements
Constraints
Environments
Function, Interface ID
System top-level segments and operations
Refined Architecture
1. Functional Hierarchy
2. Functional Flow
3. Functional Allocation
Ops flow for CONOPSFMEA / RBD
Allocated Requirements
Requirements Allocation is a systematic process
to derive segment and element requirements
10
Functional Flow Earth to LEO Operations
11
Lander Functional Allocation Matrix
12
MEL Minimum Function Non-Compliance

• Ascent Module
• MEL7.2.4-8 Human Factors Clothing, Stowage,
  Science Items, Mission Items, and Food
• Items should be considered consumables as not
  required to be lifted back to LLO
• Eliminate windows
• 2 is minimum
• No Crew
• Crew and Crew gear is never accounted for
• Descent Module
• S-Band antenna whip (MEL1.1.2)
• No unique function
• Avionics Instrumentation (MEL1.2)
• Purpose is never listed
• Boil off (MEL8.2.5)
• Mass is vented therefore not part of Inert Mass
• DMRCS
• Requires 8 motors
• Allocation for TCM
• Not minimal functional
• Air Lock

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Vehicle Configuration Minimal Functional

• This is NASAs LDAC1 with modified subsystems
• Minimum Function design
• Focus on Engineering function
• Includes Subsystem changes
• No major changes were done
• Relocation of systems
• No changes in pressure shells
• No change in function
• Intended to be the refined LDAC1 with
  conservative and historical basis

14
First Design Iteration LDAC-1 to LDAC-1MF

• Keep same layout but move habitation equipment
  from Ascent Module to Airlock to minimize Ascent
  Inert Mass
• Resize Ascent Tankage for reduced propellant mass

15
Descent Module Systems Engineering

• Primary FOMs for the Descent Module are
• Payload Delivery Capability
• Minimum Dry Mass
• Minimum Propellant Mass
• High Isp and Reduced DV
• High Mission Completion Probability
• Descent and Landing Reliability
• High Probability of Crew Survivability
• Abort capability during DM landing
• Design Trades
• Landing Maneuver Landing Systems
• Number of Engines
• Propellant/Engine selection
• Outside Vehicle diameter and tank configuration
• Amount of volume allocated to side mounts

16
Lander Engine Trade Summary

• Number of Engines
• Lander Engine Shutdown Reliability 0.9995
• Duplicating Baseline Lander Engine Installation
  adds 350 kg
• Adding 350 kg for one in 2000 Loss of Mission
  Probability makes no sense
• Go with a single engine
• Propellant Selection
• LOX/LH2 Engines the right size are reliable and
  in production
• Thermal studies indicate LH2 boiloff will not be
  a problem
• LOX/LCH4 Engines might be advantageous for Mars
• Materials technology is advancing very fast
  (better insulators)
• Go with RL10 Derivative
• Lander Diameter
• Wider Landers are shorter and less prone to
  tipover
• Wider Landers offer more options for payload
  mounting
• Go with maximum diameter possible ( 8.6 m)

17
LDAC-1A Features Benefits
LDAC-1A

• Increases diameter to fit 10 m shroud
• Halves number of tanks to save tank mass
  residuals
• Packages RL10-B2 Engine for DV and Propellant
  savings
• Puts AM in position for best landing visibility
  and escape
• Lowers center of gravity to reduce chance of
  tipover
• Allows Airlock/Hab to be rotated to lie flat on
  surface
• Separates surface suits from living quarters
  (Suitlocks)

18
LDAC-1A to LDAC-1B Changes
LDAC-1B
LDAC-1A

• LIDS mounted on deck and Orion airlock stowed
  below deck
• AM propulsion switched from Biprop to
  LOX/Methane
• Suitlock/Airlock submerged into Hab Module

• LDAC-1A had engine buried up inside tanks while
  LDAC-1B has engine buried only as far as cooled
  portion of nozzle
• Tridyne used for AM tank pressurization
• LDAC-1B tankage optimized to shorten stack

19
LDAC-1B to LDAC-1F Changes
LDAC-1F
LDAC-1B

• LM Engine switched from LOX/LH2 to LF/LH2
  (RL10-A4 version)
• LDAC-1F tankage optimized to further shorten
  stack
• AM propulsion switched from LOX/Methane to
  LF/Ammonia

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Sortie Mission Performance Comparisons
21
Trajectory Shaping and DV Costs

• Approach Classes for LDAC2

Vehicle shown every 10 seconds
?60 HDsensor-view
No Approach Phase
?30 Apollo-steep
Trajectory Angle 16 for Apollo 11-14 25 for
Apollo 15
?15 Apollo-shallow
Main Engine PDI to Ground

• No-Approach comes in very fast and does not pitch
  up until the very end good for no landing
  hazards
• Apollo-shallow comes in shallow and slow to allow
  more time to view landing site and make decisions
• Apollo-steep comes in slightly steeper providing
  increased terrain clearance and better viewing
  conditions
• HDsensor-view comes in steep and slow allowing
  sensor visibility and processing time of images

RL10 (SPAR) (RL10-B2)
LDAC2 Reference
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ALTAIR DESCENT LANDING Profile
Final approach Landing will require continued
analysis to balance sensor and pilot viewing
along with hazard avoidance divert
capabilities.
1 hr
Braking Phase
1.5 3 min
Powered Descent Initiation (PDI)
Efficiently reduce velocity from orbital speeds
Pitch-up Maneuver
Hazard Detection
Short pitch-up and throttle-down maneuver
Human Interaction
Hazard Avoidance
Approach Phase
View landing site while approaching at a low
throttle and relatively constant attitude
15 km
Terminal Descent Phase
30 m
Vertical descent to surface
Touchdown
300-600 km (8-10 min)
NOTE Not to scale
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Lander Propellant Consumption Comparison
Selected Engine Configuration uses tank head idle
for propellant settling, engine chill-down, and
RCS DV burns

• Enhanced Landed Mass (RL10-B2 vs LDAC Engine)
  710 kg
• Tank Head Idle Propellant Savings 470 kg
• Vent Gas Savings (Tank Head Idle plus improved
  MLI) 556 kg

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LDAC-1B Core Stage Three-View Drawings
HAB/Airlock (Stowed)
Tandem Tank AM (Propulsion Stage Only)
RL10-B2 Engine w/ Nozzle Permanently Extended
Centerline Docking Station for Orion
LOX Tank (2 _at_ 1.7 m Dia )
Liquid Hydrogen Tank (2 _at_3 m Dia)
Two Side-Mount Cargo Locations 3 m diameter
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LDAC-1B C Vehicle Configuration
Inflatable Airlock stays attached to Orion
(LDAC-1B only)
Airlock/HAB Module
Ascent Module (w/o crew accommodations)
LH2 Tank (1 of 2)
LOX Tank (1 of 2)
RL10-B2 Engine
Landing Direction
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DM tank structures

• Packing Analysis combined with mass analysis
  determines the optimum number of tanks
• At 8.6 m diameter with 24 t of LOX/LH2 the
  optimum number of tanks is two LOX and two
  hydrogen.
• Slosh and CG control require engineering, but
  mass penalties seem minimal.
• Level control will require active flow control
  system, but again impact seems minimal.

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Propellant Feed - LOX
Slosh Baffles
Propellant Control Valve (varies relative flow
from each tank)
Flow between tanks normalized with variable flow
valves similar to existing RL10 designs
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Propellant Feed LH2
Slosh Baffles
Propellant Control Valve (varies relative flow
from each tank)
Flow between tanks normalized with variable flow
valves similar to existing RL10 designs
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Composite Tank Comparison

• Tank Assumptions and Process
• Constant 0.010 thick Al liner (Welds covered in
  30 pad-up)
• No minimum gage overwrap (0.005 minimum
  thickness used)
• 0.010 2090 Al-Li minimum gage for metallic tank
• Elongation limited to .68 to preclude inelastic
  liner deformation
• Carbon fiber limited to 150 kpsi due to volume
  fraction
• 79 kpsi max stress for metallic tank
• 300 ksi Pan Type I fibers
• 50 Carbon/Matrix volume fraction
• 30 of Al tank weight added for pad-ups, mass
  sensors, vortex and slosh baffling (No benefit to
  slosh and vortex baffling foreseen)
• 1.0 (0.2 per applied maximum g) of loaded tank
  added for structural support
• Each system re-optimized for feed line diameter
  and flow velocity
• Composite over wrap systems used slightly smaller
  lines
• Break even pump power to weight increased
  significantly

30
Gr Overwrap Tanks Offer Significant Weight Savings
31
LDAC-1C Maximizes Useful Payload
Lander Design Features
LDAC-1C

• Switch from LDAC RL10 to RL10-B2
• More Thrust Less Landing DV
• Increase Isp from 452 to 465.5
• Add Tank Head Idle to RL10
• Burn vented boiloff gas for thrust
• Replace propulsive RCS burns
• Switch from 8 propellant tanks to 4
• Reduces tank mass
• Reduces boiloff (wetted area)
• High Performance Insulation
• Reduces boiloff
• Add active propellant management
• Controls mixture ratio
• Reduces unusable propellants
• Replace Aluminum with GrEpxy
• Reduces cryogenic tank masses

32
Ascent Module Systems Engineering

• Primary FOMs for the Ascent Module are
• Crew Safety
• Ascent Reliability
• Abort capability during DM landing
• Crew support under acceleration
• Landing visibility
• System performance
• Minimum Dry Mass
• Minimum Propellant Mass
• Design Trades
• Pressure shell versus open ride
• Number of Engines
• Propellant/engine selection
• Tandem versus in-line tanks configuration
• Centerline versus offset mounting

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Ascent Module Design Trade Summary

• Stripping the AM of all life support functions
  including the pressure shell increases the Landed
  Payload Mass by 1305 kg
• Selected AM uses PLSS for life support during
  descent and ascent, but included inflated
  pressure shell for safety (loss of suit
  pressure).
• Number of Engines (see following chart)
• Selected single gimbaled engine based on safety
  and performance.
• Propellant Selection
• Traded BiProp versus LOX/CH4 versus LF/NH3
• LOX/Methane Propulsion increased Landed Mass by
  280 kg
• Fluorine/Ammonia Propulsion increased Landed Mass
  by 865 kg
• Either BiProp or LOX Methane is viable. We show
  BiProp based on lower cost and risk, and
  propellant sharing with RCS. (What is 280 kg
  worth?)
• Tandem versus In-line tank mounting
• Little mass difference. We selected In-line
  tankage to maximum gimbal control authority and
  reduce RCS propellant usage.

34
Discussion of AM Crew Cab

• We broke with the LDAC-1 crew cab design for
  several reasons
• Move from LM centerline brought several
  advantages
• Direct visibility of landing site
• No more fire-in-the-hole issues
• Elongated AM configuration give control authority
  for TVC
• Opens up center for Orion structural dock (no
  LIDS loads limits)
• Separating the crew cab from the airlock matches
  Outpost operations
• No additional risks incurred beyond normal
  Outpost operations
• Eliminating or reducing the crew cab pressure
  shell adds payload
• Crew rides down and up in suits anyway
• Only additional risk is non-repairable suit
  malfunction
• Emergency backup inflatable sphere could be
  carried
• On-orbit EVA to gain access to Orion is practiced
  backup

When all was said and done we added a minimum
mass pressure shell
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Ascent Module GNC and Crew Interface

• Optimized design for safe abort
• Constantly calculating abort trajectories during
  landing
• After landing calculates minimum rendezvous
  launch windows for entire sortie mission period
• Flight is autonomous with high level human
  override capability
• Astronauts direct lander GNC to desired landing
  site using cursor in heads-up display (State of
  the Art in Military Aircraft)
• GNC optimizes ascent trajectory to minimize
  rendezvous time and propellants
• Final rendezvous and dock (grapple?) are
  automated with human intervention
• Flight controller trained for voice recognition
  and verbal commands
• Designed for redundant laptop Guidance,
  Navigation, Control Functions
• Bluetooth connection to stand-alone control
  system
• Use accelerometer chips in laptop as backup
  navigation aids
• Built-in AM PCS and TCS are only for control
  boxes plus Communications and Tracking
• PLSS unit provides backup computation, power, and
  thermal

36
Ascent Module Propulsion Design Space

• Traded Pressure Shell and three AM propellant and
  engine combinations
• 5500 Lbf Pressure-Fed Biprop
• 5500 Lbf Pressure-Fed LOX/Methane
• 4400 Lbf Pressure-Fed LF/Ammonia
• No difference in reliability if equal
  qualification program

• Removing the AM Pressure Shell increases Landed
  Mass by 1305 kg
• LOX/Methane Propulsion increased Landed Mass by
  280 kg
• Fluorine/Ammonia Propulsion increased Landed Mass
  by 865 kg

Reducing the mass of the pressure shell is much
less expensive than qualifying a new engine
37
AM propulsion- to gimbal or not to gimbal

• A trade was conducted to explore whether the
  ascent module engine should have a gimbal
  capability
• In general, eliminating a TVC system improves
  reliability
• Thrust vector/c.g placement is critical without
  TVC
• Decision was made to include gimbal to maximize
  abort opportunities, make the AM more robust, and
  minimize RCS.

38
Ascent Crew Module Accommodations

• Requirements
• Restrain crew for safe descent and ascent flight
  phases
• ECLSS for 2 hours descent, 5 hours ascent
• No reconfigurations for abort mode
• Assumptions
• Flight is autonomous with high level human
  override capability
• Implies some situational awareness and HMI
  provisions
• Crew is wearing pressure suits during transit
• Optimized crew accommodations design for minimum
  dry mass with reasonable safety
• Baseline is crew is ensconced in surface suits
  only
• Option one is inflatable cover only
• Option two is minimum hard pressure shell

Option 1
Option 2
39
LDAC-1C Ascent Module

• Ascent Module designed for safe abort during
  lander failure
• Gimbaled pressure-fed engine w/ hypergolic
  propellants assures fast separation
• Non-axisymmetric configuration allows good
  visibility and no fire-in-the hole issues
• Large moment arm and gimbal allows significant
  side-force generation for escape

40
Ascent Module use of Existing Assets

• Remove the seats, harnesses, and suit/seat
  interfaces from the Orion and install them onto
  the AM flight bridge.
• Provide each pair of crewmembers with a single
  PLSS during the descent and ascent stages.
• Utilize a third PLSS to provide circulating
  cooling water for AM Avionics.
• Utilize a fourth PLSS as a backup for either the
  crew or avionics cooling.
• Examine the possibility of using the PLSS
  batteries to provide peak and backup power for
  AM.

Inflatable or Convertible Version
Seats from Orion
PLSS Units
41
INFLATABLE ASCENT MODULE CREW CAB

• Mass includes
• inflatable pressure restraint shell
• triple-redundant bladders w bleeders
• proprietary hatch integration hardware
• Mass does not include hatches

42
LDAC-1C Orion Interface (w/ Pressure Shell)
Undocked then Re-docked for Crew Transfer)
Docked (for EDS or LM LLI Burns)
Inflated AM Cab
CEV Structural Dock (Attached to LM with
pyrotechnic bolts)
A
HAB/Airlock (Stowed in carry position)
HAB/Airlock Deployment Link
A
43
Surface Elements Systems Engineering

• Primary FOMs for the Sortie Surface Elements are
• Maximum EVA hours
• Crew Safety
• Crew Morale
• Design Trades
• Human Factors Analyses
• Airlock w/ or w/o Suitports
• High Pres GOX or LOX in PLSS
• HAB on LM or on lunar surface
• Polyethylene or in situ radiation shielding

44
Habitation Functional Analysis
Accounts for 2 bunks (bunk bed style) 0.71 x 2.03
is floor space for one/two bunks
45
LDAC-1B Airlock Design with Suitlocks
Habitat
Mark III Suits attached via suit-ports
Suit Entry
Bunks
Medical, Etc.
Privacy
1.75m
2.5m
Suit Entry
Bunks
Empty Mark III Suits attached to maintenance port
Mark III 20W x 23H .508m x .584m
4.25m
Top
Right
Expandable Version
2.5m
2.3m
2.2m
1.75m
1.0m
46
Surface Systems Trades

• High Pressure GOX or LOX in PLSS
• Abundant LOX available right after DM landing.
  (Pull out of feed lines into redundant evacuated
  dewars)
• Current EVA frequency limited by High Pressure
  GOX Pump
• LOX Dewars save 20 kg relative to High Pressure
  Pump system
• Selected LOX in PLSS to allow four EVAs/day
• HAB on LM or on lunar surface
• HAB on surface allows in situ radiation shielding
• HAB on surface reduces access time and falling
  risk
• Selected HAB on surface for safety and EVA
  efficiency
• Polyethylene or in situ radiation shielding
• Crew survival in 1010 proton/sec solar event
  requires 250 kg polyethylene shield (HAB mounted
  on lander)
• Crew survival in 1010 proton/sec solar event
  requires 6 kg of shovels and sandbags (HAB
  sitting in trench on surface)
• Selected surface siting and designed provisions
  into lander

47
Radiation Problem Statement

• Flares can be detected and operating procedure
  would entail retreating to protected
  areas/aspects for minutes to hours
• Warning time is days to minutes (50 predicted
  two days in advance)
• Protection must be designed around peak cycle
  years, NOT quiet years

48
Solar Particle Event Probability in 1-Week
Mission and BFO Exposure Level inside a Typical
Equipment Room in Free Space
From Space Radiation Risk Assessment for Future
Lunar Missions on NASA Web Site
49
Mission Timing is Everything for Radiation Hazard
Extrapolation
Lunar Sortie Missions do not occur over the
entire span of a Solar Cycle. What if they
Start Here?
2025 2030 2035
50
Solar Flare Solution

• Projected lunar missions will occur during
  2020-2024 peak in solar activity
• Current state-of-the-art in Solar Flare
  Prediction is 50 accuracy two days prior to
  flare (this is incorporated into our abort
  probability estimate).
• Therefore, plan is to protect HAB/Airlock using
  in situ materials shortly after landing, and
  abort back to CEV if major flare is predicted
  before protection is completed.
• Previous work indicates 250 kg of polyethylene
  blankets will reduce radiation inside the
  HAB/Airlock from a 1010 protons/sec-m2 SPE to
  acceptable levels (within 30 day limit for Blood
  Forming Organs).
• We are not recommending blankets be carried in
  LDAC-1C
• The polyethylene blankets may be preferred
  solution if rover carried

51
Micrometeoroid

• Statistical based hazard assessment
• Not shown to be a driver in design

52
Micrometeorite Risk Issue or Not?

• Micrometeorite risk based on LDEF impact craters
  corrected for lunar focusing (Ref V. Vanzani, et
  al., MICROMETEOROID IMPACTS ON THE LUNAR
  SURFACE ,1025.PDF, Lunar and Planetary Science
  XXVIII, 1994
• Flux for 0.0001 gm 0.006/ m2-yr
  6.8E-7m2-hr
• MTBF gt 114,000 hrs
• Micrometeorite risk based on Apollo Data (NASA
  CR-190014)
• Flux for 0.0003 gm 6.0E-11/ m2-sec
• MTBF gt 386,000 hrs
• Micrometeorite risk based on Astronaut photos of
  Lunar Surveyor
• Flux for 0.0003 gm 7.3 E-07 m2/hr
• MTBF gt 114,000 hrs

Ref V, Vanzani, et. al.
53
Discussion of AM Crew Cab

• We broke with the LDAC-1 crew cab design for
  several reasons
• Move from LM centerline brought several
  advantages
• Direct visibility of landing site
• No more fire-in-the-hole issues
• Elongated AM configuration give control authority
  for TVC
• Opens up center for Orion structural dock (no
  LIDS loads limits)
• Separating the crew cab from the airlock matches
  Outpost operations
• No additional risks incurred beyond normal
  Outpost operations
• Eliminating or reducing the crew cab pressure
  shell adds payload
• Crew rides down and up in suits anyway
• Only additional risk is non-repairable suit
  malfunction
• Emergency backup inflatable sphere could be
  carried
• On-orbit EVA to gain access to Orion is practiced
  backup

When all was said and done we added a minimum
mass pressure shell
54
Ascent Module LOC/LOM Comparisons
99.997
35,949
Higher reliability due to


pressurized cabin life support
which is back
-
up by suited
operation life support which is
back
-
up by
PLSS life support.


55
Safety Buyback Summary

• Safety Data Base Assessment Tool
• Weve used a simplified safety and hazards
  analysis to identify those parts of the design
  that have the greatest impact on safety.
• Only flight critical or mission critical failures
  were tracked. Noncritical failures were ignored.
  (Cliff Notes version of Buyback)
• An assessment of the LDAC-1 and LDAC-1MF designs
  were done to understand the safety level of a
  minimum functionality design.
• Safety Assessment of Improved LDAC-1B , LDAC-1C
  w/ Redundancy
• An assessment of the improved designs has been
  accomplished to understand how design for safety
  and abort impacts loss of vehicle.
• Abort is key to safe operation at low cost and
  low mass
• Design for safe abort proved to be very effective
  in keeping masses low
• Safety Buy-Back Process
• In order to bring the lander up to acceptable
  crew safety levels, a number of changes need to
  be made. Weve analyzed the most critical items
  to maximize their mass-effectiveness.

Note Cliff Notes version of failure modes will
over-exaggerate final crew and mission
reliabilities
56
Primary Failure Modes w/ Reliability Abort
Example
57
PLOC Buyback plus Improvements

• Basis of estimate is LDAC-1
• Abort design and analyses completed before
  buyback started
• Reliability estimates are from published data and
  subcontractor data bases
• Radiation protection traded against a 1010
  Proton/sec SPE
• Three published estimates of lunar micrometeorite
  flux indicates its not a design driver
• PLOM can benefit from abort capability only after
  surface exploration complete (Assume re-flight
  necessary if crew chooses to leave before
  exploration complete)
• Total mass increase from LDAC-1B to LDAC-1C for
  reliability improvements buyback are
• 168 kg for DM
• 116 for Hab/Airlock
• 150 kg for Ascent Module

• Suitlocks were already included in the LDAC_1B
  mass buildup Eliminated the suitlock function
  within the existing airlock will save
  approximately 100 kg
• Extra LOX scavenging also included for potential
  doubling number of EVA hours

• LDAC-1MF has no abort capability (by definition)
• LDAC-1B is designed for abort, but has no
  redundancies
• LDAC-1C is LDAC-1B with buyback improvements

58
Selected Safety Buyback Curve for LDAC-1C
59
Prioritized List of Safety Buyback Candidates

• Last six items have marginal worth to PLOC, but
  benefit PLOM
• Improvements for landing needs to be quantified
  (beyond the scope of this study)
• Abort is the far and away best buy for PLOC, but
  mass cost is hard to quantify

60
Probability of Loss of Mission Worksheet
61
Mass Buildup Comparison
62
Performance Conclusions

• Performance Trades showed Sortie Payloads up to
  10.5 tons possible.
• No Pressure shell on AM
• Fluorine/Ammonia Propellants on AM
• Fluorine/Hydrogen Propellants on DM
• Graphite-wrapped Cryogenic Tankage
• We recommend a less risky approach for Payloads
  up to 7.5 tons.
• Inflated pressure shell on AM
• Gimbaled NTO/MMH Pressure-fed engine on AM
• Gimbaled RL10-B2 Engine on the DM
• Tank-Head Idle added to the RL10-B2
• Graphite-wrapped Cryogenic Tankage

Our recommendation is a compromise between more
payload versus less risk
63
Altair Lunar Lander Final Briefing 11/19/2008

• 0830 Introduction PI / Dana Andrews
• 0845 Major Trade Results
• Lander Trade Overview PI / Dana Andrews
• Refined Landing Simulations Draper Lab
• GNC Sensor Update Ball Aerospace
• Thermal Control Refinements Ball Aerospace
• RL10B-2 Reliability / Tank Head Idle PWR
• Ascent Module Trade Overview PI / Dana Andrews
• Surface Systems Trade Overview PI / Dana
  Andrews
• Reliability / LOC Buyback Overview PI / Dana
  Andrews
• Reliability Near Surface, LSS Power Hamilton
  Sundstrand
• 1115 Final Design Recommendations
  PI / Dana Andrews
• 1130 Study Summary / Where do we go from here? PI
  / Dana Andrews

64
Summary of Study Results

• The NASA LDAC-1 Altair baseline works as
  advertised
• Our masses, performance and reliabilities match
  LDAC-1 data
• Reference LDAC-1 technologies are a good
  compromise between performance and risk
• There is room for improvement
• Take advantage of the 10 m shroud
• Move the ascent module outboard for visibility
  and enhanced abort
• Mount the airlock/HAB (or pressurized Rover)
  outboard for easy ground deployment
• Use the larger RL10-B2 engine (better T/W and 13
  sec Isp)
• Put a reinforced docking station for Orion on
  Altair centerline
• Use SOA technology to replace some structures
• Replace aluminum tanks with composite-wrapped
• Replace composite AM pressure shell with
  engineered-fabric inflatable (or with nothing at
  all)
• Design so that some components are dual use
• PLSS units serve as backups for critical AM
  subsystems
• Orion seats transferred to AM for descent and
  ascent
• Emergency equipment moved from Orion to Altair
  for surface stay

65
Summary/Recommendations

• Trade needs to be done comparing cost of Altair
  technology developments versus cost of upsizing
  ARES V.
• Many of the higher performance options we did not
  choose are probably justifiable when compared to
  the cost of increasing the ARES V payload.
• Sharing resources between Orion and Altair is a
  win, win area.
• PLSS units as backup for computing, ECLSS, Power,
  and Thermal
• Seats, batteries, fire suppression, emergency
  supplies, etc. transferred from Orion to Altair
  and back
• No-approach landing matches cargo operations,
  minimizes throttling requirements, and minimizes
  landing propellants, while still giving
  astronauts over-ride capability.
• Designing Altair to easily deploy habitats or
  rovers during a sortie mission costs little mass
  and opens up mission growth options.

66
Questions?

• Andrews Space, Inc.
• November 19, 2008