Lunar Capability Concept Review (LCCR)

Transcription

1 Lunar Capability Concept Review (LCCR) Transportation Systems Only June 18 20, 2008 Report to the PSS

2 LCCR Agenda Date Time Topic Presenter June 18 8:00 8:15 am Welcome / Introduction Hanley / Muirhead 8:15 9:00 am 01: LCCR Overview Knotts 9:00 9:15 am 02: Lunar Requirements Summary Knotts 9:15 10:00 am 03: CxAT_Lunar Study Process Joosten 10:00 11:00 am 04: LSS Concepts Culbert 11:00 11:30 am Lunch 11:30 am 2:00 pm 05: Altair System Hansen / Connolly 2:00 4:30 pm 06: Ares V System Creech 4:30 5:30 pm 07: Ground Operations Quinn June 19 8:00 8:30 pm 08: Ares V and Altair Margins Strategies and Basis Muirhead 8:30 10:30 am 09: Integrated Performance and Mission Design Martinez 10:30 am 12:30 pm 10: Strategic Analysis Falker 12:30 1:00 pm Lunch 1:00 3:00 pm 11: HLR POD Architecture Drake 3:00 5:00 pm 12: LCCR Product Summary and Forward Plan Parkinson 5:00 5:30 pm 13: Architecture Summary and Next Steps Muirhead June 20 8:00 9:00 am 14: MCR Wrap-up Altair Hansen / Graham 9:00 10:00 am 15: MCR Wrap-up Ares V Creech 10:00 10:30 am 16: LCCR Success Criteria Review Knotts 10:30 11:00 am Summary / Conclusions Hanley 11:00 11:30 am Board Discussion Section 00: LCCR Agenda Page 2

3 Components of Program Constellation Earth Departure Stage Crew Exploration Vehicle Heavy Lift Launch Vehicle Crew Launch Vehicle Lunar Lander Section 05: Altair System Page 3

4 Typical Lunar Reference Mission MOON Vehicles are not to scale. 100 km Low Lunar Orbit Lander Performs LOI Ascent Stage Expended Low Earth Orbit Earth Departure Stage Expended Service Module Expended EARTH EDS, Lander CEV Direct Entry Land Landing Section 05: Altair System Page 4

5 Background VSE Global Exploration Strategy ESAS LAT1 LAT2 CxAT Lunar Section 03: CxAT_Lunar Study Process Page 5

6 LCCR Scope LCCR will define a Point of Departure (POD)* transportation architecture for the CxP Lunar Capability including capabilities to: Deliver and return crew to the surface of the moon for short durations, i.e. Human Lunar Return (HLR) Enable establishment of a lunar outpost This review focuses on the conceptual designs and key driving requirements for Ares V and Altair (crewed and cargo) This review assumes the capabilities of Ares I and Orion for the lunar missions This review will show how the POD transportation architecture, including EVA and Ground Ops, supports a range of mission campaigns and possible surface architecture solutions Specific Lunar Surface Systems definition is not part of this review *This is a POD transportation architecture and NOT the final baseline Section 01: LCCR Overview Page 6

7 CxP L2 Lunar Design Reference Missions DRM 1 : Lunar Sortie Crew DRM This mission lands anywhere on the Moon, uses only on-board consumables, and leaves within ~1 week. This mission enables exploration of high-interest science sites, scouting of Lunar Outpost locations, technology development objectives, and the capability to perform EVAs. DRM 2 : Uncrewed Cargo Lander DRM Used to support an Outpost, help build one, or merely preposition assets for a subsequent Sortie Lander, this uncrewed mission lands anywhere on the Moon, and has enough resources to sustain itself until a component of the Lunar Surface Systems takes over. DRM 3 : Visiting Lunar Outpost Expedition DRM Analogous to an assembly flight to ISS, this mission lands at the site of a complete Outpost or one under construction, and allows crewmembers to extend their stay by using assets of the Outpost rather than only what is carried onboard their Lander. DRM 4 : Resident Lunar Outpost Expedition DRM Realizing one of the goals of US Space Policy, this mission allows a sustained human presence on the surface of the Moon, since it follows a single crew of four to the surface, transitions them to a habitat at an Outpost, and gets them back to Earth after transitioning over to a replacement crew. DRM 5 : Outpost Remote DRM This mission is separated in function from the other DRMs by focusing only on those Lunar Surface Systems which need to operate without human intervention, either because humans are not present to operate them, or the task is more easily performed in an autonomous or automatic manner. Section 02: LCCR Requirements Summary Page 7

8 122 m (400 ft) Building on a Foundation of Proven Technologies - Launch Vehicle Comparisons - PPBE Submit ( ) Altair Crew Overall Vehicle Height, m (ft) 91 m (300 ft) 61 m (200 ft) 30 m (100 ft) Orion Upper Stage (1 J 2X) t (302.2k lb m ) LOX/LH 2 5-Segment Reusable Solid Rocket Booster (RSRB) Earth Departure Stage (EDS) (1 J 2X) t (517.0k lb m ) LOX/LH 2 Core Stage (5 RS 68B Engines) 1,435.5 t (3,164.8k lb m ) LOX/LH Segment RSRBs Lunar Lander S-IVB (1 J 2 engine) t (240.0k lb m ) LOX/LH 2 S-II (5 J 2 engines) t (1,000.0k lb m ) LOX/LH 2 S-IC (5 F 1) 1,769.0 t (3,900.0k lb m ) LOX/RP-1 DAC 2 TR 5 0 Space Shuttle Ares I Ares V Saturn V Height: 56 m (184 ft) Height: 99 m (325 ft) Height: 110 m (361 ft) Height: 111 m (364 ft) Gross Liftoff Mass: Gross Liftoff Mass: Gross Liftoff Mass: Gross Liftoff Mass: 2,041.2 t (4,500.0k lb 3,374.9 t (7,440.3k lb m ) m ) t (2,043.9k lb m ) 2,948.4 t (6,500.0k lb m ) Payload Capability: Payload Capability: Payload Capability: Payload Capability: 25.0 t (55.1k lb 63.6 t (140.2k lb m ) to TLI (with Ares I) m ) 25.6 t (56.5k lb m ) 44.9 t (99.0k lb m ) to TLI to Low Earth Orbit (LEO) to LEO 55.6 t (122.6k lb m ) to Direct TLI t (262.0k lb m ) to LEO t (316.1k lb m ) to LEO Section 06: Ares V System Page 8

9 Ares V Trade Space Core Booster 5 Segment PBAN Steel Case Reusable Standard Core W/ 5 RS t +5.0 t Opt. Core Length / 6 Core Engines 68.6 t Spacers: 1 Common Design Features Composite Dry Structures for Core Stage, EDS & Shroud Metallic Cryo Tanks for Core Stage & EDS 5 Segment HTPB Composite Case Expendable t 69.7 t +5.0 t t 74.7 t Spacers: 1 RS-68B Performance: I sp = sec Thrust = 797k lb vac J-2X Performance: I sp = sec Thrust = 294k lb vac 5.5 Segment PBAN Steel Case Reusable t 67.4 t +3.7 t t 71.1 t Spacers: 0 Shroud Dimensions: Barrel Dia. = 10 m Usable Dia. = 8.8 m Barrel Length = 9.7 m LCCR Initial Reference 1.5 Launch TLI Capability Section 03: CxAT_Lunar Study Process Page 9

10 Recommendations Altair Wet Mass Cargo Optimized Delta-V Crew Optimized Altair Wet Mass Post-LOI Loiter Time Approach Applied Margins/Reserves Methodology to Altair & Ares V (net loss of architecture performance ) Developed higher fidelity mission analysis techniques (net gain of architecture performance") Result Lunar Architecture still requires ~12% additional performance Higher performance Ares V options required Altair prop loading and loiter requirements determined Section 03: CxAT_Lunar Study Process Page 10

11 Recommended New Point of Departure - Vehicle m 10 m 10 m 21.7 m 23.2 m m 71.3 m Vehicle recommended 6 Engine Core, 5.5 Segment PBAN Steel Case Booster Provides Architecture Closure with Margin Recommend Maintaining Vehicle with Composite HTPB Booster as Ares V Option Final Decision on Ares V Booster at Constellation Lunar SRR (2010) Additional Performance Capability if needed for Margin or requirements Allows for competitive acquisition environment for booster Near Term Plan to Maintain Booster Options Fund key technology areas: composite cases, HTPB propellant characterization Competitive Phase 1 Industry Studies NOTE: These are MEAN numbers Section 06: Ares V System Page 11

12 Summary Ares V Initial 2008 Capability ( ) exceeds Saturn Capability by ~30% Ares V LCCR analysis focused on meeting lunar requirements and developing margin Ares V is sensitive to Loiter, Attitude, Power, and Altitude requirements in addition to payload performance Recommended new POD Ares V can meet current HLR requirements with ~6 t of Margin Additional budget required (~$1.7BRY) for the 5.5 Segment PBAN Booster and 6 Engine Core Plan to maintain new composite HTPB booster as an option Additional analysis required to determine Ares V PLOM and PLOC contributions for CARD recommendations Section 06: Ares V System Page 12

13 Altair Lunar Lander 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 14 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 Section 05: Altair System Page 13

14 Configuration Variants Sortie Variant 45,000 kg Descent Module Ascent Module Airlock Outpost Variant 45,000 kg Descent Module Ascent Module Mass Available for Payload Manager's Reserve Avionics Power Structures and Mechanisms Propulsion Thermal Control Life Support Other Non-Propellant Fluids Cargo Variant 53,600 kg Descent Module Cargo on Upper Deck Propellant Sortie Mission Lander Section 05: Altair System Page 14

15 Design Approach Project examined the multitude of concepts developed in the post-esas era, took lessons learned and began to develop a real design. Altair took a true risk informed design approach, starting with a minimum functionality design and adding from there to reduce risk. Lunar Design Analysis Cycle (LDAC) 1 developed a minimum functional vehicle. 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) 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! LDAC-2 determined the most significant contributors to loss of crew (LOC) and the optimum cost/risk trades to reduce those risks. LDAC-3 (current LDAC) is assessing biggest contributors to loss of mission (LOM) and optimum cost/risk trades to reduce those risks. Goal of the design process is to do enough real design work to understand and develop the requirements for SRR. Section 05: Altair System Page 15

16 LDAC-2 Overview P(LOC) The initial Lander Design and Analysis Cycles (May-November 2007) created a minimal functionality lander design that serves as a baseline upon which to add safety, reliability and functionality back into the design with known changes to performance, cost and risk. LDAC-2 completed in May Goal was to buy down Loss of Crew (LOC) risks. Spent approximately 1.3 t to buy down loss of crew (LOC) risks. Spent an additional 680kg on design maturity. 1.8E E E E E E E E E-02 1 in 6 Sum of System Contributions to LOC/ Mass Available for Payload 3652 kg 500 kg minimum payload 1 in kg Individual Subsystem Contribution to LOC: Events\Hazards Life Support Thermal Propulsion Structures and Mechanisms Power Avionics 0.0E+00 0 LDAC-1 LDAC-2 Section 05: Altair System Page mass (kg) Mass Available for Payload

17 Launch Shroud Packaging the lander within the Ares V launch shroud is akin to building a ship in a bottle Ares V descent stage structure and landing gear designed to package within a 10 meter launch shroud (8.8 meter diameter dynamic envelope) Key features Migration of Structural Configuration to 10m Shroud Assumed 8.8m Dynamic Envelope (but not a hard number) Scaled LDAC1 -Delta w/no Major Configuration Changes Incorporated Updates for 10m Shroud & For Increased ΔV Single CAD Update, Two Analysis Cases (10m & 10m + ΔV) AL and AM Global Geometry Unchanged Items that Did Change (or that were matured) Include Mass impact resized propellant tanks, engine mounts modified tank support scheme added realistic clearances for plumbing, radiators, insulation, struts Refined AM/DM separation and AM/AL separation/tunnel Details Some hatch details Descent stage is now Clocked 45 o with respect to AM Deck Height = 5.9m (upgrade to 10m shroud), 6.2m (shroud + ΔV) 10m Shroud Migration Adds +46 kg (DS Truss kg, EDSA kg) Increased ΔV Adds +217 kg ( kg DS Truss, EDSA kg) +47 kg Due to Combination of both 10m Shroud & Increased ΔV TOTAL MASS INCREASE: 310 kg 7-Apr-2008 NASA Internal Only Your Initials Here / 2 9.5m LDAC1-Delta & LDAC2 Configurations LDAC1-Δ AM Tanks with RCS Placeholder Structure Modified Cone Supports LDAC2 Airlock Modified Hatch 9.5m 7-Apr-2008 NASA Internal Only Your Initials Here / 3 Section 05: Altair System Page 17

18 Payload Shroud Current Design Concept Point of Departure (Biconic) Leading Candidate (Ogive) Mass: 9.1 t (20.0k lb m ) POD Geometry: Biconic Design: Quad sector Barrel Diameter: 10 m (33 ft) Barrel Length: 9.7 m (32 ft) Total Length: 22 m (72ft) Quad Sector Design Frangible Joint Horizontal Separation Composite sandwich construction (Carbon- Epoxy face sheets, Al honeycomb core) Painted cork TPS bonded to outer face sheet with RTV Payload access ports for maintenance, payload consumables and environmental control (while on ground) Thrust Rail Vertical Separation System Payload umbilical separation Section 06: Ares V System Page 18

19 Ares V Shroud Encapsulation Issues Shroud quad sector configuration will likely preclude partial encapsulation in SSPF Shroud Encapsulation risks are the same for all 51 Series Ares V variants GO and Ares V teams will continue to study shroud ground processing alternatives. Section 07: Ground Operations Page 19

20 Temporal Availability Contour Plots Temporal availability contour plots show the availability of lunar landing sites over the lunar nodal cycle. The following plots reflect both global sortie mission availability for the Altair alone as well as for the integrated Orion/Altair mission. The following proposed ESAS landing sites are indicated in the contours: Landing Site Latitude Longitude Notes A. South Pole 89.9 S 180 W (LAC 144) rim of Shackleton B. Far side SPA floor 54 S 162 W (LAC 133) near Bose C. Orientale basin floor 19 S 88 W (LAC 91) near Kopff D. Oceanus Procellarum 3 S 43 W (LAC 75) inside Flamsteed P E. Mare Smythii 2.5 N 86.5 E (LAC 63) near Peek F. W/NW Tranquilitatis 8 N 21 E (LAC 60) north of Arago G. Rima Bode 13 N 3.9 W (LAC 59) near Bode vent system H. Aristarchus plateau 26 N 49 W (LAC 39) north of Cobra Head I. Central far side highlands 26 N 178 E (LAC 50) near Dante J. North Pole 89.5 N 91 E (LAC 1) rim of Peary B Section 09: Integrated Performance and Mission Design Page 20

21 Global Access/LOI Δ-V/LLO loiter Access to all lunar landing sites ( global access ) requires a combination of additional LOI Δ-V, pre-descent LLO loiter, and post-ascent LLO loiter Minimum energy LLO maneuvers are sufficient for polar outpost missions Altair to size tanks for 1000 m/sec LOI maneuver, load consumables for 4 additional days of LLO loiter Section 05: Altair System Page 21

22 Global Sortie Mission Sequence IC: Post- Ascent LEO Epoch Specified TLI Earth-Moon Transfer LOI-1 LOI-3 DE-ORBIT PDI Landing Site Specified ASCENT TO DOCK Pre-TEI Extended Loiter Moon-Earth Transfer EI 3-Burn LOI 1-day Loiter 3-7 Day Surface Stay 1-day Loiter 3-Burn TEI RENDEZVOUS IN LEO GIVEN: Epoch Landing Site LAT/LONG TLI Window Duration Trans-Lunar Time of Flight Post-LOI Extended Loiter Pre-TEI Extended Loiter Trans-Earth Time of Flight Entry Interface Conditions DETERMINE: TCM s Required Propellant Mass for EDS, Altair, and Orion Vehicles. Post- LOI Extended Loiter ASCENT PLANE CHANGE LLO ORBIT MAINTENANCE TEI-1 TEI-3 TCM s ACTIVE VEHICLES: 1. ORION ACTIVE 2. ALTAIR ACTIVE 3. EDS ACTIVE Section 09: Integrated Performance and Mission Design Page 22

23 Greater Than Zero Temporal Coverage In order to ensure global lunar surface access, the following minimum mission architecture conditions were determined to be sufficient for providing access to all lunar landing sites at some epoch during the lunar nodal cycle*: 48 HR LOI and TEI flight times 5 days of extended LOI loiter 3 days of extended TEI loiter Altair Only 1000 m/s LOI ΔV Capability For these conditions the Altair can provide access to the worst case landing sites ~8% of the time. For the integrated capability, this provides for access to the worst case integrated landing sites ~5% of the time. Integrated Altair and Orion * To the resolution of the MAPP data and given the assumptions in the MAPP analysis (e.g., no Earth perturbations assumed in LOI and TEI 3-burn maneuvers). Section 09: Integrated Performance and Mission Design Page 23

24 LDAC-2 Configuration Section 05: Altair System Page 24

25 Vehicle Architecture Ascent Module Descent Module Airlock Three Primary Elements Descent Module Provides propulsion for TCMs, LOI, and powered descent Provides power during lunar transit, descent, and surface operations Serves as platform for lunar landing and liftoff of ascent module Ascent Module Provides propulsion for ascent from lunar surface after surface mission Provides habitable volume for four during descent, surface, and ascent operations Contains cockpit and majority of avionics Airlock Accommodates two crew per ingress / egress cycle Connected to ascent module via short tunnel Remains with descent module on lunar surface after ascent module liftoff Section 05: Altair System Page 25

26 Altair LDAC-2 Sortie Vehicle Configuration AM RCS Thruster Pod (x4) LIDS Star Tracker and Comm. Antenna (x2) Docking Window (x2) Forward Facing Window (x2) Avionics Platforms (x2) AM Fuel Tank (x2) AM Oxidizer Tank (x2) Pressurant Tank (x2) AM Main Engine AM Connecting Structure (Remains on DM) AM-Airlock Connecting Structure Airlock Airlock Egress Hatch Life Support Oxygen Tank Avionics boxes (x2) Thermal Insulation DM LH2 Fuel Tank (x4) Landing Leg (x4) Pressurant Tank (x2) DM RCS Thruster Pod (x4) LOX Tank Support Cone (x4) DM Main Engine RCS Tanks Radiator (x2) Problem: central location of capsule = bad visibility. Section 05: Altair System Page 26

27 Altair Key Messages Current design demonstrates a lander design that closes within the Constellation transportation architecture Altair has investigated a wide breadth of lander concepts, using lessons learned to influence the current design concept Altair has undertaken a process that begins with minimum functionality and buys back safety, reliability and additional capabilities with known performance, cost and risk impacts The Altair team is using this design process to help develop good requirements Design process (particularly risk based design approach) is resulting in a smart government design team Altair has used its bottoms-up design work to inform sizing of landers for transportation architecture trades Altair has developed a detailed bottoms-up cost estimate Section 05: Altair System Page 27

28 Sample Return Mass Considerations Nominal return mass: 100 kg. Note that Apollo 17 returned kg of sample. The PSS views the sample mass allocation in the current exploration architecture for geological sample return as too low to support the top science objectives. We are asking that CAPTEM undertake a study of this issue with specific recommendations for sample return specifications. Tempe Workshop, CAPTEM: Minimum of 230 kg total return mass, but noted that on the basis of simple extrapolation of the Apollo 17 mission, sample return mass could be as much as 800 kg. The recent Lunar Surface Scenarios workshop estimated that a 7-day sortie mission with 4 crew and 8 EVAs could collect 306 kg of samples. 100 kg IS NOT ENOUGH Section 05: Altair System Page 28

29 Sample Return Mass Considerations LCCR Action Items: Stochastic modeling of sample return mass on the Altair ascent stage; Study of Orion volume capacity for increasing return sample mass. At the moment there is 1.6 metric tons of reserve mass on Altair, not including PMR. Section 05: Altair System Page 29

30 LCCR & The Surface Architecture (not part of this LCCR) Can the Constellation transportation system (Ares, Orion & Altair) support the deployment and operations of a lunar outpost? Is the POD cargo capacity to the lunar surface enough? Can Do the surface systems fit in the Ares-5 shroud and on the Altair deck? Have solutions for unloading cargo to the surface been identified? Element Mass (t) LCT (Lunar Communications Terminal) Crew Mobility Chassis Pressurized Crew Cab 6.13 Mobile Power Unit 0.85 OTSE - Davit 0.16 OPS - ISRU system - H2 Reduction (0.5t) - TS Excavation - H2 Reduction (0.5t) - TS Science Logistics Total Capability Difference Yes Yes Yes Section 04: LSS Concepts Page 30

31 ISRU Functional Description: Perform lunar regolith excavation and handling, oxygen extraction from regolith, and oxygen storage and delivery, and support lander propellant scavenging and water production. For flexibility, two 1/2- scale plants will be delivered and 2 sets of excavation tools. Total O 2 Produced = 1000 kg/yr Mass per O 2 plant = 219 kg Power per plant = 3.93 kw Total Regolith = 415 kg/day Excavation Tools = 42.7 kg (each) Excavation Time = <1 hr/day Large hoppers hold 1 day s regolith Hoppers raised to allow dumping of spent regolith into CMC Oxygen Extraction from Regolith Radiator panels fold down for launch TS 2/3 O 2 Production System with Storage and Thermal Control 1 st Gen O 2 Production System (660 kg/yr ) for Field Demo, Nov. 08 Regolith Excavation and Movement Scoop lifts 11 kg regolith/scoop (38 scoops to fill hoppers for day) Cratos Excavator O 2 Storage H 2 O Processing Excavation and O 2 Plant mounted on mobile chassis 1 t of oxygen per year requires a regolith excavation rate of <1/2 cup per minute! (1% efficiency - 70% light) RESOLVE Subscale O 2 Extraction and Volatile release reactor Regolith H 2 Processing Area Clear Blade on CMC Bucketwheel Excavator Section 04: LSS Concepts Page 31

32 Surface Architectures Assessed Three surface architectures were developed in support of LCCR: Rapid Outpost Buildup (TS-1) Deliver as much outpost capability as soon as transportation system permits Full-up outpost based on the recommendations from LAT-2. Substantial robustness through element duplication Initial Mobility Emphasis (TS-2) Temper outpost build-up based on affordability with initial emphasis on mobility capabilities Full-up outpost has less volume and limited eclipse operating capability than TS1 Robustness achieved through functional reallocation Assumed water scavenging Initial Habitation Emphasis (TS-3) Temper outpost build-up based on affordability with initial emphasis on core habitation & exploration capabilities Full-up outpost has less volume and limited eclipse operating capability than TS1 Robustness achieved through functional reallocation Assumed water scavenging Surface Systems Review in 2010 Section 04: LSS Concepts Page 32

33 Lunar Transportation Architecture Recommendations Ares-V Maximize commonality between Lunar and Initial Capabilities: Ares-V engine core, 5.5 segment PBAN steel case booster Provides architecture closure with additional margin High commonality with Ares I Continue to study the benefits/risk of improved performance: Ares-V Final decision on Ares V booster at Program SRR (6/2010) Altair Additional performance capability if needed for margin or requirements Allows for competitive acquisition environment for booster Requires further study and technology investment funding Provide a robust capability to support Lunar Outpost Missions: Optimize for crew missions (500 kg + airlock with crew) Lander cargo delivery: ~ 14,500 kg in cargo only mode Size the system for global access while allowing future mission and system flexibility Size Altair tanks for 1,000 m/s LOI delta-v Size for an additional 4 days of Low-Lunar Orbit loiter (site specific) Retain adequate margins: ~1,000 kg Program reserve at TLI Minimum of 40% total Altair margin/reserve Orion Continue to mature Orion vehicle concept Maintain strong emphasis on mass control Continue to hold Orion control mass to 20,185 kg at TLI Increase emphasis on evolution of Orion Block 2 to support lunar Outpost missions Section 13: Architecture Summary and Next Steps Page 33

34 DRMs/Mission Key Driving Requirements Mapping Lunar Sortie Design Reference Mission A TBD or TBR is associated with this requirement Multi-Mission Phase Requirements Anytime Abort LOC 1 in 100 Altair Crew of kg (1,102 lb m ) cargo Global Access Landing accuracy 100 m (328 ft) with 95% accuracy ( ) 373 ( ) hrs crew support Airlock functionality LOC 1 in 250 ( ) LOM 1 in 75 ( ) LOM 1 in 20 Descent ΔV 2,030 m/s (6,660 ft/s) LH2/LO2 descent engine restartable/throttleable 7 d MOON Ascent 1,881 m/s (6,171 ft/s) 100 kg (220 lb m ) pressurized return payload TBD hrs post lunar ascent LLO 100 km (54nm) Altair Performs LOI 1,000 m/s (3,281 ft/s) (Propellant load for 950 m/s) Altair ΔV for LOI 1,000 m/s (3,281 ft/s) 3-burn LOI 1-4 days Altair LLO loiter TEI 1,492 m/s (4,895 ft/s) (Tanks sized for 1, 560 m/s (5,118 m/s) Altair TLI Injected Control Mass 45 t (99,200 lb m ) EVA Mass Allocation kg (378.0 lbm) FCE Mass Allocation kg (295.0 lbm) EDS TLI Injection Capability 66.1 t (145,726 lb m ) + 5 t reserve EDS Performs TLI 3,175 m/s (10,417 ft/s) Orion Orion TLI Control Mass 20,185 kg (44,500 lbm) FCE & EVA Mass Allocation 675 kg (1,488 lbs) Orion 382 kg (842) unpressurized cargo 21.1 days crew support LOC 1 in 200 LOM 1 in 50 ERO 241km (130nm) -20x185 km (-11x100 nm), 29º Ares-I Delivered Mass 23.6 t (52,070 lb m ) 4 days LEO loiter Direct or Skip Entry 905 t (2M lbm) 3,698 t (8.2 Mlb m ) EARTH Ares V 4 launches per year (6 launches per year) Weather exclusive launch availability TBD sebment SRBs; 6 RS-68B LOC 1 in 37,000 LOM (vehicle) 1 in 125 Water Landing 90 min. 1-5 d ~4d 1-5d 7 d 1d Section 13: Architecture Summary and Next Steps <5.8d Page 34