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University of MarylandCDR 04/19/04

Modular RovingPlanetary Habitat,

Laboratory, and Base

Critical Design ReviewApril 19, 2004


“Returning to the moon is an important step for our space program… Themoon is home to abundant resources. Its soil contains raw materials that might beharvested and processed into rocket fuel or breathable air. We can use our time onthe moon to develop and test new approaches and technologies and systems thatwill allow us to function in other, more challenging environments. The moon is alogical step toward further progress and achievement.”

~ President George W. Bush

Mission Statement

• Restore human exploration to the moon,investigating a large portion of the lunar surface

• Maximize the search for natural lunar resourcesthrough a mobile base

• Develop the experience and the requiredtechnologies for extended human space travel

Kevin Eisenhower


Program Overview

• Launch small, autonomous, mobile modules with specializedpurposes to the moon

• Assemble the modules into a complete base• Send a crew of 4 to the assembled base to conduct experiments

for 3 lunar day-night cycles (months)• Disassemble base and send modules to next site of interest, up

to 1000km away, over an unmanned period of 3 months

Designed to meet externally applied constraints

Kevin Eisenhower

2005 2015

Design/Manufacture

2010

TestSystems

LaunchModules

Manned Missions(every 6 months)

2020


Water ice and ore depositsas potential resources

Deep lunar compositionVolcanic and impact history

Environmental variationsalong the surface

Muscle developmentand decay

Cardiovascular system

Skeletal system

Radiation

Nutrition and medical care

Human performanceand adaptations

Plant and animal biology

Crew dynamics

Science Objectives• Study the effects of the lunar environment on life,

especially long duration missions

• Study the moon and its resources

Kevin Eisenhower


Assembly Sites and Selection Criteria

http://geopubs.wr.usgs.gov/i-map/i2769/

Sites1) Marius2) Flamsteed3) Gassendi4) Tycho5) Meretus6) Drygalski7) Schrodinger8) Minnaert9) Mare Ingenii10) Gagarin

12

34

6 5

10

89

7 6

• Approximately 1000 km apart• Apollo site candidates• Surveyor landing sites

• Ore deposits and useful elements• Permanently shadowed regions• Age and appeal of lunar structures

Near Side Far Side

Kevin Eisenhower


Module Types• 6 Habitable Modules

– Size governed by launch vehicle payload size– Need 6 to achieve the necessary floor space and

volume to accommodate a crew of 4 for 3 lunar sols

• 6 Chassis with all-wheel drive– Provide locomotion for Habitable Modules– Equipped with transit connections to link vehicles

in the transit phase

• 4 Power Modules with all-wheel drive– Provides AC power to all modules– Equipped with navigation computers and leads

during transit

Kevin Eisenhower


• Transit configuration:1 Power Module,2 Habitable Modules, &2 Chassis Modules

Modularity – Transit Phase

Kevin Eisenhower

• After each mannedmission,MORPHLAB willdisassemble andtravel autonomouslyto a new base site

• Science will beperformed by themodules whiletraveling

• Enables exploration andanalysis of vast areas ofthe Lunar surface


Modularity – Manned Base Phase• Multiple base configurations possible• Subsystems split between modules

Kevin Eisenhower


Benefits of Modularity• Modules are less massive than a large-scale structure

– Can be launched on existing launch vehicles– Increased mobility

• Capable of traversing great distances– Multiple lunar base locations– Science may be performed while modules are in transit– Experiments may be subjected to long exposures to the

lunar environment• Feasible program expenditures

– Allows for cost spreading over the program duration• Utilizes a cost learning curve

Kevin Eisenhower


Module DetailsHabitable Module

Power ModuleChassis Module


Habitable Module


Habitable Module Overview• Habitable modules:

– provide the space where the crew will live and workwhile not out on EVA

– two types of modules:• Habitable-Living (3 of 6)• Habitable-Science (other 3 of 6)

• This breakup– allows minimal disruption if the astronauts have different

sleeping schedules– restricts lunar dust to the EVA areas

• Functions are distributed and redundant so that thefailure of any one module will not disrupt normalactivities (Level 1 requirements I3, I11)

Cat McGhan


Sizing Requirements• All crew interfaces are sized to accommodate 95th

percentile American males to 5th percentileJapanese females (Level 1 requirements L1, L2)

• The size of the modules is bounded by– maximum radius of Delta IV-Heavy payload area – 4.5 m– minimum amount of room necessary for four people in

the same module – 10.24 m2

• Habitable modules are 4.0 m in diameter, whichgives 12.56 m2 of floor space per module– accounts for wires and pipes behind walls and furnishings– leaves enough open space in the modules to move about– keeps the module mass reasonable

Cat McGhan &Jessica Seidel


Habitable-Living SubtypesCrew quarters (2) Galley (1)

• Both types include– Basement level radiation shelter sleeping areas– Two hatches, L-shape intersection

Cat McGhan

4m

3.7m

RadiatorAndrogynousDockingMechanism

ViewingWindow

O2 Tanks

N2 Tanks

H2O Tanks


Habitable-Science SubtypesEVA workspace (2) Experiment area (1)

• Both types include– Basement level storage compartments– Three exits, T-shape intersection

Cat McGhan

4m

3.7m

RadiatorAndrogynousDockingMechanism

EVA doorAndrogynousDockingMechanism

Ladder


Radiation Environment

Cat McGhan, John Albritton

Note: Drawing not to scale

96-99% p+, 1-3 days10-3 – 104, max ~10610-103, max ~104SPE

85% p+, 13% He+, 2% HZE, Near-constant10-5 – 1, max ~2102 – 104, max 1011GCR

Trapped p+, e-, Short-duration102 – 101010-1 – 103Van Allen Belts

Particles, TimeFlux (p/cm2-s)Energy Level (MeV)


Lifetime Limits: BFO dose (blood-forming organs), 5 cm depth

300 rem400 rem

55

250 rem325 rem

45

175 rem250 rem

3525

100 remFemale150 remMale

AgeSex

Necessity of Radiation Shielding

• SPE shielding is necessary• From the Level 1 requirements (M1, M6, I1, L2, L5, L6,

L7), shielding must prevent the radiation dose received fromexceeding these levels for the habitable mission length

Cat McGhan

1 year30 daysTime interval

25 rem50 rem

BFO dose limits

30 rem/yearGCR200 remSPEOn Lunar SurfaceDose Source


Radiation Shielding Comparison

• 5.7cm Polyethylene is used for SPE shielding• 4mm Aluminum pressure hull helps block GCR

Cat McGhan

SPE Dose, 4 August 1972 Flare

0

50

100

150

200

0 5 10 15 20 25 30 35 40 45 50

slab, x, g/cm^2

5 c

m B

FO

do

se, re

m

Regolith, H,LaRC

Polyethylene,H, LaRC

Aluminum, alt.reference

GCR Doses, Solar Minimum

0

5

10

15

20

25

30

0 5 10 15 20 25 30 35 40 45 50

slab, x, g/cm^2

5 c

m B

FO

do

se, re

m/y

ear Regolith,

H60, 1977

Polyethylene,H60, 1977

Aluminum,alt. reference


48.91 remTotal dose 84+180 days24.99 remTotal dose 30 days2.2200 kg/per4320180DIPS (5 modules)11.991 Al2805180GCR awake21.745.41 polyethylene72180SPE, August 1972 dosage0.011 Al~3.5180EVAs2.855.41 polyethylene1440180GCR asleep1.03200 kg/per201684DIPS (5 modules)4.311 Al100884GCR awake, no EVA1.335.41 polyethylene67284GCR asleep1.441 Al33684EVAs2------Van Allen Belts both ways

Totaldose (rem)

Shielding (g/cm2)Totalhours

Period

Radiation Dose Breakdown

Cat McGhan


Radiation Shelter Sleeping Areas

• Habitable-living basement level• Split in half by a noise-muffling

curtain for privacy• Dual-purpose• Emergency exit procedure:

– Move curtain aside– Exit via second door

• From the Level 1 requirements and trade studies:– Level 2 requirement L6: Radiation shielding capable of

suitable protection against a solar flare must be internal toat least three habitable modules.

Cat McGhan


Radiation Shelter Specifics• Shelter built into basement of Habitable-

Living modules– Maximizes daytime living floor area– Minimizes mass consumption by doubling shelter as

sleeping quarters– Basement also leaves room for storage and ECLSS

systems• High Density Polyethylene

– Superior radiation protection per amount of masscompared to other materials

– 0.057 m thickness based on radiation analysis– 568 kg total

Kevin Macko


Radiation Shelter Specifics• Can accommodate two

people each• Two doors built into

the top• Internal Dimensions

– Based on estimatednecessary volume for95th percentileAmerican male tosleep comfortably(derived from L1requirement)

– 1.9 m long, 1.95 mwide, 0.8 m tall

Kevin Macko

Dividing curtain Polyethylene shielding

Cutaway Shelter Layout

ECLSS andstorage volume

2.8 m

4.0 m

0.8 m


Procedure for Impending SPE• In the event of an impending SPE:

– SOHO (SOlar and Heliospheric Observatory) or itssuccessor will give 2-3 days notice of heightened activity

– The astronauts will be informed by Mission Control andEVAs will be cancelled for the duration

– Just before the event, SOHO and other Space Weathersatellites will give a few hours notice

– Astronauts will retreat to the shelter for the duration ofthe event

• The crew will be able to maintain contact withMission Control via their laptops, and they willhave access to whatever libraries and amusementsthat they can download from the NASA computers

Cat McGhan


• Micrometeoroids massive enough to cause damage arevery rare: probability analysis is necessary

• NASA Human rating requirement: The program shall bedesigned such that the cumulative probability of safecrew return over the life of the program exceeds 0.99

• We will assume a desired 0.996 Probability of NoPenetration (PNP) per mission - same as Apollorequirement

F = cumulative flux (hits/m^2-year) A = exposed surface area (m^2) = 41.6 m^2 Y = exposure time (year) = 0.75 years

• F = 1.28x10-4 hits/m2-year

Micrometeoroid Protection

tA

PNPF

*

)ln(!=

Kevin Macko


Micrometeoroid ProtectionCumulative Micrometeoroid Flux vs. Mass

[Vanzani, et al. Micrometeoroid Impacts on the LunarSurface. Lunar and Planetary Science XXVIII, 1997.]

• Micrometeoroid fluxof 1.28x10-4 hits/m2-year is off this chart

• Data curve wasgraphically extendedto this flux

• Criticalmicrometeoroid massestimated to be 3x10-4

grams

Kevin Macko


Micrometeoroid Protection

• Data based on heuristic from NASA Preferred Reliability Practices document onmicrometeoroid protection (Practice No. PD-EC-1107)

• Assumed .5 mm spacing between MLI and aluminum pressure hull• 4 mm Al skin chosen based on 13.3 km/s average lunar micrometeoroid velocity

Kevin Macko

Critical penetration mass vs. impact velocity

0

0.0001

0.0002

0.0003

0.0004

0.0005

0.0006

0.0007

0.0008

0.0009

0.001

0 5 10 15 20

micrometeoroid impact velocity (km/s)

cri

tical p

en

etr

ati

on

mass (

g)

1 mm Al

skin

2 mm Al

skin

3 mm Al

skin

4 mm Al

skin

design critical mass


Pressure Hull Design• Requirement S1 states that

pressurized tanks must havefactor of safety of 3.0

• Rounded shape chosen tominimize stresses due tointernal pressure

• Bottom is rounded tominimize pressure stress,but flattened to lay onchassis module correctly

Kevin Macko

Top View

Side View

Doors

Window


Pressure Hull Analysis• 4 mm thick Al-2014

pressure hull– Stress concentrations

around doors and windowsare assumed to be handledby their frames

– 2.8 margin of safety forhoop stress

• 2 mm thick bottomreinforcement– Where stresses are above

164 MPA with 4mmthickness

• Design based on FEAis conservative:stringers not includedin pressure model

240 Mpa

185 MPa

135 Mpa

80 MPa

25 MPa

Kevin Macko

Finite element analysis with57.2 kPa internal pressureapplied


Internal Support Structure• 3 critical loading scenarios considered

– Launch: 6g’s axially, 2.5g’slaterally (based on Delta-IV-Hpayload planner’s guide)

– Landing: 1 lunar g vertically andhorizontally, on one leg

– Hab/Science to Chassis attachment:4 lunar g’s

• All elements made of titanium alloywith hollow circle cross-section

• Primary structure designed for 2.0factor of safety based on S1requirement

Kevin Macko

8 Circular Stringers

8 Elliptical Stringers

3 Circular Frames

8 Vertical Stringers

157 kg

32 kg

24 kg

49 kg

Mass

524 MPa

542 MPa

524 MPa

394 MPa

Max Design Stress

.052

.018

.052

.40

Margin of Safety

Compression.026 m.031 mCircular Frames

Bending.014 m.018 mElliptical Stringers

Bending.032 m.033 mCircular Stringers

Compression.024 m.030 mVertical Stringers

Failure ModeI.D.O.D.Component


Internal Support Structure• FEM Assumptions

– Door and window frames that intersect with internal structure will act as thestringers in reality

– FEM analysis is conservative because skin is not included to take loads– Excessive load concentrations at joints can be ignored

• Launch Loading– Vertical stringers, circular

frames, and upper circularstringers sized based on thisscenario

– Constrained at 4 connectionpoints along bottom frame

– 3700 kg vehicle massdistributed to structuralconnection points

Highest stress concentrations

525 MPa395 MPa265 MPa135 MPa70 MPa

Kevin Macko


• Landing Loading– Worst case – landing on one leg– Constrained at leg attachment points– Max stress: 179 MPa - no internal

structural components sized forlanding

Internal Support Structure

• Hab to ChassisAttachment– .03 m fall onto chassis

when landing legs areremoved

– Max stress: 542 MPa inlower elliptical stringers

Displacement exaggerated by factor of 15;3 cm actual displacement

Attachment Loading FEM

Kevin Macko


Additional Structural Analysis

Kevin Macko

HoopStress

58.0033109Aluminum2014

NitrogenTank

HoopStress

65.0033109Aluminum2014

OxygenTank

Bending5683.61.53.2PolyethyleneRadiationShelter

Bending93.0492174Aluminum2014

DockingMechanismWalkway

Compression21.2.262437TitaniumAlloy

RadiatorSupport

Structure

Failure ModeMass(kg)

Margin ofSafety

SafetyFactor

Max DesignStress (MPa)

MaterialComponent

• Safety factors based on S1 requirement– 2.0 for primary structures– 1.5 for secondary structures– 3.0 for pressure vessels


Floor Panel Requirements• Composite sandwich panel used to sustain

compressive and bending loads of floor items• Level 1 requirements

– S1: Factor of Safety = 2– S2: Non-negative MOS (Margin Of Safety) for

launch loads (6g Earth)• Total load of interior items

– Must sustain launch loads– Assumed the load is distributed over entire panel– Future iterations will use composite laminate theory

for more precise numbers

Alan Felsecker


Floor Panel Design

Top-Down View

Cross-Sectional View

Mass of Interior Items Estimated 876.1 kg

Skin: High Modulus Carbon Fiber-Epoxy- 37 kg- 0.13 cm thick

Core: Aluminum 5056 Honeycomb- 37 kg/m3

- ¼ in. cell size- 24 kg- 7.24 cm thick

Total Panel:61 kg7.49 cm thick0.98 cm max. deflection

Module W

all

Module W

all

desk

Alan Felsecker


Floor Panel Deflection andFailure Modes

Buckling StressPotential FailurePhenomenon

1.24 GPa1.24 GPa

1.80 MPa

1.24GPa

5.20 MPa0.01 cm

Limit MOSEffect InducedLoading Effect

0.071.33 GPaSkin Wrinkling12.4420 GPaIntra-cell Buckling

9.690.168 MPaCore Shear Stress19.8359.5 MPaSkin Stress

451.7311500 PaLocal Compressive Stress0.020.98 cmMaximum Deflection

Alan Felsecker


• Types: Hatch and General Area• Level 1 Requirements

– S1: Factor of Safety = 3– S2: Nonnegative MOS in bending for launch loads (6g Earth)– L2: Adhere to NASA-STD-3000, delimiting the following:

• minimum window diameters, minimum blunt impact (550N) , heightof windows, optical and manufacturing requirements, ionizingradiation protection (GCR), non-ionizing radiation protection (UV),safety precautions

– L8: Must have emergency alternative access / bailout options• S2 met by using the window frame to channel launch loads

around the window and into a stringer• L2 is met using a sunshade 4 mm thick of Aluminum

2014. Sun shade also protects crew from earth’s magneticfield

• L8 is satisfied using hatches, where a porthole must bepresent to allow astronauts to see what is on the other sideof the hatch

Window Requirements

Alan Felsecker


Window Assembly• Material: Ceradyne Thermo-Sil® Ultra High Strength Fused Silica - radio wave, IR, microwave radiation transparent - low CTE (0.5*10-6/°C)

Hatch Window Assembly

Alan Felsecker

- high strength/weight ratio: flexural strength ~ 55 MPadensity ~ 2020 kg/m3


Window Assembly

• Similarities to ISS windows:– Same debris pane thickness– Anti-reflective coating on outer surface of outer window

pane– Blue-red reflective coating on inner surface of inner

window pane to filter out IR and UV– Air tight seal between pressure panes

Alan Felsecker

• Differences from ISS windows:–Different pressure pane thicknessfrom ISS

–MORPHLAB uses 4 mmAluminum 5056 sunshades over allwindows


Window Panes

6.180.835Mass (kg)0.9700.970Thickness (cm)

Fused Silica Debris Pane3.770.510Mass (kg)

0.4000.400Thickness (cm)

Aluminum Sunshade3.480.556Mass (kg)

0.01820.0182MOS0.6440.644Thickness (cm)

Fused Silica Kick Pane

50.820.3Diameter (cm)

0.01000.0100MOS8.760.559Mass (kg)

1.370.547Thickness (cm)

General AreaHatchPressure Pane

Alan Felsecker


Window Masses Per Module

1*0

Hatch Windows with PressurePanes, Kick Pane, AluminumSunshade, and Debris Pane

9.5735.3Total Window** Weight (kg)

01

General Area Windows withPressure Panes, Kick Pane,Aluminum Sunshade, andDebris Pane

3*2

Hatch Windows with PressurePanes, Kick Pane, andAluminum Sunshade

Habitable-ScienceHabitable-Living

* Walkways protect all hatch windows except outer EVA hatch window; only the outerEVA door’s window needs micrometeoroid and lunar dust protection during transit** Total Window Weight excludes TBD window frames

Alan Felsecker


Hatches & DockingInflatable tunnel

Extendable mating ringon each inflatable tunnel

Sliding interior doors

Russell Parker, Craig Nickel

• Provides travel between modules in shirtsleeve environment• Provides power, water, and air transfer between modules during base phase• More details forthcoming


Chassis Module


Chassis Overview• Single universal chassis for all MORPHLAB applications

4 power modules landed on Moon with chassis permanently attached

6 Chassis Module landed on Lunar Surface

• 4 wheel drive / 4 wheel steering

• Total Mass = 1483kg

• Drive platform for Power Module and Habitable Module

• Utilizes 24 hour battery power source until mated with Power Module

• Capable of supporting 3700kg load during transit and base phase

• Can be directly landed on lunar surface with a payload of 2000kg

• .7m minimum ground clearance and 1.4m diameter wheel to meet Level 1requirement : capability of traversing a .5m obstacle

Jason Seabrease


Chassis DimensionsRequirements:

Must be capable of traversingover a .5m boulder

Must fit within 4.5m dynamicenvelope of Delta IV-H payloadshroud

• Max wheelbase = 4.16m

• Max Wheel Stance (width between wheels) =3.2m

• Total Structural Mass = 305kg

• Materials used include Aluminum 2014 &Titanium Alloy (Ti -10V-2Fe-3AL)

Jason Seabrease

• Ground clearance = .7m

• Total Height = 1.7m

• Maximum width = 4.16m

• Closed Maximum length = 4.16m

• Open Maximum length = 7.6m


Chassis ExpansionRequirements:• Must be capable of traversing a 20 deg. cross slope.• Must be capable of traversing a 30 deg. Direct incline.• All components of MORPHLAB shall be launched on Atlas V or

Delta IV launching systems.- Launching on a Delta IV-H results in greatest payload in LEO,therefore, greatest landed mass on lunar surface.- To fit on Delta IV-H all components of MORPHLAB must fit inside4.5m diameter dynamic envelope of a 5m payload shroud.

To meet these requirements thechassis was designed to be in afolded position during transitatop a Delta IV-H and toexpand during lunar orbit to agreater wheelbase. Theextended wheelbase of 4.16m isnecessary to prevent tippingduring 30 deg. direct incline.

Jason Seabrease


Chassis Expansion• Expansion accomplished by two equal length ‘spreading bars’ rotating about a‘central pin’ to force two ‘sliding bars’ outward carrying with them the wheels andstruts that are attached.

• The spread bars are both driven by a single electric motor with a geared connectionto the bars.

• Once in final position the strut assembly is rigidly locked in place at the ends of theupper frame to prevent loads during landing and lunar surface transit to flow throughthe ‘spreading bars’.

• Expansion mechanism mass = 20kg.

Spread Bars About Central Pin

Sliding bar forcedoutward by 2cylindersconnected tospread bars withball joints

Jason Seabrease


Chassis Components

Transit Connections:• Connect modules and provide power and data connections during transitconfiguration• Initially stowed to fit in Delta IV-H payload

Jason Seabrease


Chassis Components

Decent Beams (2) :•Aluminum 2014

•Mass of each = 5.8kg

•Critical Load During Lunar Decent eachsustaining 3500N Maximum Stress =81MPa

•Margin of Safety = 1.02

•Failure mode : bending

Launch Beams (2):•Titanium Alloy Ti -10V-2Fe-3AL

•Mass of each = 53.7kg

•Critical Load During Delta IV-H main engine cutoff,max axial acceleration = 6g’s, max lateral acceleration= 2.5 g’s, supports maximum of 3500kg (includingchassis)

•Margin of Safety = .23

•Failure mode: bending

Jason Seabrease


Chassis Components

Wheel Axle Vertical StrutAngled Strut #1 Angled Strut #2

•Struts and axles affected by bending moments, compression, and tension.

•Compression, bending, and sheer forces most critical during lunar landing.

•Struts and axles also affected by dynamic loading during lunar transit and staticloading during base configuration.

•Components sized to support 3500 kg above chassis during all transit and static casesand to support 3500 kg total for landing loads.

•All components finalized through Finite Element Analysis with ProMechanica.

Jason Seabrease


Chassis ComponentsStruts and Axles:

27.80.0980.11.20.58350Ti AlloyWheel Axle

57.80.040.051.10.38400Ti AlloyAngled strut 2

25.00.0450.050.90.38400Ti AlloyAngled strut 1

22.20.0450.050.80.38400Ti AlloyVertical Strut

Totalmass

(kg)

InnerRadius

(mm)

OuterRadius

(mm)length

(m)Margin Of

Safety

MaxStress(MPa)Material

ComponentName

Vertical Strut (4) and Angled Strut 1 (4) :Critical loading during power modulelanding on lunar surface directly on wheels; worst case considered as landing directly onone wheel with all force acting in vertical direction. Failure due to combination ofcompression and bending.

Angled Strut 2 (4) and Wheel Axle (4) : Critical loading during power module landingon lunar surface directly on wheels; worst case considered as landing directly on onewheel with all force acting in chassis horizontal direction (first wheel touchdown withchassis perpendicular to lunar surface).

Jason Seabrease


Wheel Design• Requirement

– MORPHLAB mustovercome 0.5m obstacles.

– Wheel must support theweight of power moduleduring landing

– Safety Factor = 2

Kevin Mako/Abraham Daiub

vehiclelunarmg

Constraint

MS = 0.540

357MPa

Mass = 37.4kg dimensions: cm

• Assumptions– Axle constrain– Load =

glunar*mvehicle

– Wheel diameter= 1.4 m

– Ti alloy


Motor / GearWith a 20% margin:• The wheels need 655.4 N-m

torque on the Power module• The wheels need 810.2 N-m

torque on the Hab Life/Sciencemodules

• The chosen motor is theMF0210050 from EmoteqCorporation

Chosen for:• Decent torque output 305.7 N-m• Low Mechanical time constant

(time to get to 63.2% of maxspeed)

• Low weight 7.61 kg

• The torque is good, but a gearwill need to be used to get thetorque for the modules, thus agear with a gear ratio of 3 to 1 isused

www.emoteq.com

Picture of a MF0210 motor

Rob Winter


Braking System• Derived Requirements

– Stop while moving at maximum speed• Stopping distance < linkage bending tolerance• Stop full caravan assembly

– Stopped on maximum slope• 30º longitudinal slope• One end sitting on 0.5 meter obstacle

• Electrohydraulic Disc Brakes (shuttle)– Less torque variation during stop vs. drum brakes– Benefits of carbon brakes vs. steel brakes

• Lighter• Resistant to friction• Efficient whether hot or cold• Absorbs 2-3 times more heat

Emily Tai


Braking System

3920 WAverage Power Dissipated

358 JEnergy Dissipated

0.091 sStopping Time

165 N-mBrake Torque

236 NBraking Force

6370 N-mBrake Torque

9110 NBraking Force

Stopped on Slope

Moving at 1 km/hr

Emily Tai


Chassis / Habitable InterfaceRequirements:

•Provide rigid connectionbetween Habitable moduleand Chassis

•Transfer power betweenChassis and Habitablemodule

•Prevent Dust frominterfering with powerconnection

Blue cup fixed on 4 locations on bottom ofHabitable Module Skin.

Yellow sleeve spring loaded and designed toprevent dust from collecting on powerconnection.

Sleeve is driven down inside larger sleeveto expose supporting cylinder and powerconnection probe

Power connection probe

Supporting cylinder

As Habitable Module descends itautomatically aligns itself snuggly on topof supporting cylinder and allows powerprobe to reach into plug internal toHabitable Module

Cylinder Diameter = 5 cm

Cylinder length = 8 cm

Max Compression = 128 MPa

Material = Al 2014

Margin of Safety = .2

Jason Seabrease &John Albritton


Power Module


Power Module OverviewRequirements:• Supports a 5 year mission (M6)• Capable of producing power during the lunar night (M1)• Must not impose a health risk to the crew (M1, L7)• Multiple power sources to accommodate potential module loss (I3, I4)• Size must not impact MORPHLAB system mobility (M2)• Additional Factors: low system mass, high reliability

.9955No Single-Point Failure Reliability26%Thermal to Electric Efficiency

420 kgTotal DIPS Mass (with shielding)5.3 kWePower Output

4Number of Units

12.6 W/kgSpecific Power (with shielding)

Selected: Dynamic Isotope Power System (DIPS)

Joe Cavaluzzi


Power Module Components

Joe Cavaluzzi

Heat Source

Shadow Shield Radiator

Avionics suite

DIPS

Robotic Arm(stowed configuration)


DIPS Details

20.3 kWtThermal Output

19.5 kWtEOL ThermalOutput

95 kgMass

88 yearsHalf Life

64.3%Carnot Eff.

460 KTL

1300 KTH

7.4 kWWork Out

48.6%Actual Eff.

5.5 kWtQout

15.3 kWtQin

5.1 kWeEOL PowerOutput

5.3 kWePower Output

72%Efficiency

Radiator

PuO2

Power Out

5.3 kWe

Lithium HydrideShadow Shield200 kg (0.3 m3)

Stirling Cycle EngineHeat Source Alternator

Heat Source: Engine: Alternator:

(EOL: End of Life)

(Not to Scale)

Total System Mass: 420 kg

Joe Cavaluzzi

Heat transfer into engine estimated as 75% efficient


Common Sub-Systems

Communications

Thermal Control Systems

External Lighting


• Modules will be insulated with Multilayer Insulation (MLI)– Thin layers of Mylar coated with aluminum separated by a vacuum– The Mylar has low-conductivity of .20 W/m*K so it effectively isolates

the inner-workings of the module from the harsh lunar environment withfluctuations in temperature from 160K to 380K

– During the lunar day and night, the assumption is there will be enoughlayers of MLI (ie. 10) to act adiabatically and therefore keep the ambientinside reference temperature at 298K

– Moonshine and lunar albedo will not pose a heating problem to themodule, even though it will “see” 380K at some points, because of theMLI and its adiabatic assumption

– Aerogel will be used where heat leakage paths occur (ie. Panel seams,corners) because it insulates well (thermal cond.= .004 W/m*K) and hasa density of 10 kg/m3

– 10 layers @ .25 mm thickness implies volume of .14 m3 and with adensity of 1420 kg/m3 gives a mass of 193 kg

Thermal Control System

Matt Hughes


• A vertical radiator atop a module on the lunar surface will “see” solarflux (Is), moonshine (IR), and lunar albedo (alb) which means theenvironment temperature would be around 380K at the equator duringlunar noon.

• Since the top of the module is 4.5m off the surface, a horizontal radiatorwill never “see” moonshine or lunar albedo. The only radiation it will“see” is from solar flux, where the worst case would be noon at theequator when the entire radiator is illuminated.

• There will be no solar flux at night or at the poles because of the sun’sincident angle of zero degrees to the radiator

albedo

IR

Is IsVerticalRadiator

HorizontalRadiator

No albedo

NO IR

Radiator

Matt Hughes


• Ran thermal equilibrium analysis on horizontal radiator to find out what areawas needed to dissipate the allotted heat per module

• Since lunar albedo and moonshine are not considered, the thermalequilibrium equation then becomes:

α : Absorptivity (.08) ε : Emissivity (.92) As : Projected Area Facing Sun AR : Total Radiative Area Pint : Internal Heat Dissipated (Qin) σ : Stefan-Boltzman Constant (5.67e-8 W/m2K4) Trad : Temperature of Radiator Surface Tenv : Temperature of Environment (4K)

• Took case of sun directly overhead illuminating the whole area of the radiatorwith all energy needing to be dissipated through radiator. Therefore,

AR = As = A

Isdfsfdsf

)T(TåóAPáAI envradRss

44

int!=+ ( 1 )

Radiator Sizing

Matt Hughes


Radiator Sizing Continued• Taking the case of Lunar noon at the equator, the optical properties must have a low

absorptivity as to minimize solar heating effects and high enough emissivity toradiate internal heat effectively (ie. ε = .92 and α = .08)

• ^ Assuming Trad = 298K, and therefore in thermal equilibrium with inner cabin• * Assuming 5.3 kg/m2 for one-sided, horizontal radiators from Lunar Base

Handbook• ~ Assuming .04 m for thickness of typical radiator

4.65 x 1.5 x.04

3267PowerModule

2.77 x 2.77x .04

417.672.3HabitableScience

5.00 x 2.67x .04

7113.344.0HabitableLiving

Dimensions(m3) ~

Mass(kg)*

Rad Area(m2) ^

Qin(kW)


Radiator Sizing Cont.

• Using loop heat pipes (LHP) due to flexibility and fact that they dissipate moreheat then conventional heat pipes. A benchmark of 1 loop heat pipe being ableto dissipate 1.6kW was used.

• Habitable-Living Module has 4 LHP (two outer: 1350 kW/each, two inner:650 kW/each to handle lower night loads

• Habitable-Science Module has 4 LHP (two outer: 500 kW/each, two inner:650 kW/each to handle lower night loads as well

4.513.161.3.29HabitableScience

10.183.161.3.51HabitableLiving

Area(m2)

Area (m2)Low Level

Qin (kW)Low Level

Volume(m3)

Radiator Mass

Matt Hughes


Habitable-Living Radiator

Evaporators

Reservoirs

Qin via LHP

Thermistorsto monitorTrad

Trad = 298K

Low Conductance Mountings

Spring-loadeddeployable joints

Condensers

Mass = 71 kg

AluminumHoneycombbacking

α,ε = .08, .92

MLI undersideacts adiabatically

Deep Space = 4KHabitable Living Radiator

Matt Hughes


Communication Capabilities• Level 1 Requirements:

– 10Mbps of bidirectional digital data– Two uplink channels HDTV– One downlink channel HDTV

• Derived Requirements:

Mike Naylor

– EVA communications• Module – Rover

– Main comms. with EVA– Send one HDTV to base

• Module – Astronaut– Voice– Health monitoring

– Inter-Module communications• Nominally use only wireless• Backup physical connection• Receive commands during

transit• Processor/memory sharing

during peak usage times


Bandwidth Requirements• Lunar Surface

– Bandwidth determined by• Rover – Module• EVA/Suit - Module• Module – Module

– Inter-modulecommunications used forcomputer interaction andcommunications betweencrew members

• <1Mbps for voice comms• Remainder used for data

transfer and video link

• Total Uplink – 60Mbps• Total Downlink – 40Mbps

• HDTV Requirements– Uncompressed single

channel is 1.5Gbps– With MPEG-2 compression

single channel is 19.4 Mbps• Internationally accepted

industry standard• Depending on available

bandwidth high qualitycompression up to 80Mbps

– 4 – 8 standard videochannels on single HDTVband

– Two uplink – 40Mbps– One downlink – 20Mbps

• Total Uplink – 50Mbps• Total Downlink – 30Mbps

Mike Naylor


Halo Orbit• Far side communications will be accomplished

through satellites placed in halo orbits about theEarth-Moon L2 point– Number of satellites: 4

• Placed in the same orbit with 90° phase offset• Provides two fault tolerance regardless of base or satellite location

– Orbital parameters• Period: ~ 1 year• In-plane amplitudes:

– 32,250 km in line with the Earth-Moon vector– 90,000 km normal to the Earth-Moon vector

• Out-of-plane amplitude: 63,200 km

David Pullen


Halo Orbit Diagram

David Pullen


Moon – Earth Link Budgets

3.0 dB3.0 dB3.0 dB3.0 dBLink Margin

11.4 W4.0 W4.0 W1.2 WOutputPower

10 Mbps30 Mbps10 Mbps30 MbpsReceiveBit Rate

10 Mbps50 Mbps10 Mbps50 MbpsTransmitBit Rate

KaKaKaKaBand

10 cm50 cm10 cm50 cmDiameter

Moon FSPower-Halo

Moon FSHab-Halo

Moon NSPower-Earth

Moon NSHab-Earth

NS – Near Side FS – Far Side Hab – HabitableModule

Mike Naylor


Lunar Surface Link Budgets

3.0 dB3.0 dB3.0 dB3.0 dBLink Margin

37.5 W1 W18.2W(Rover only)

N/AOutputPower

40 Mbps<1 Mbps10 Mbps30 MbpsReceiveBit Rate

60 Mbps<1 Mbps30 Mbps30 MbpsTransmitBit Rate

UHFUHFUHFUHFBand

Omni-antennaOmni-antennaOmni-antennaOmni-antennaDiameter

Total ModuleBudget

EVA Suit-Module

10km range

Rover-Module20km range

Module-Module

20km range

Mike Naylor


Radiation-Hardened Computers• Level 1 Requirements

– Shall be capable of reaching a TRL of 6 by thetechnology cut-off date of Jan. 1, 2010

• TRL 6: System/subsystem model or prototype demonstration in arelevant environment (ground or space)

• Looked at 7 different radiation hardened (Rad-Hard) computers for trends

• Compared Rad-Hard vs. Commercial Off the Shelf(COTS) to identify technology gaps and trends:- Processor Speed -Bus Speed -Power Consumption

John Albritton

- RHPCM- RAD750- SCS750 - RHPPC

- RAD6000- MASS


COTS vs Rad-Hard: Clock Speed

• Trends suggest a 5 year technology gap in clockspeed between COTS and Rad-Hard• By 2010 Rad-Hard Processor speed of 3400 MHz

John Albritton


COTS vs Rad-Hard: Bus Speed

• Bus speed is positively related to clock speed for COTS• No correlation or trend could be established for Rad-Hard in 2010

John Albritton


COTS vs Rad-Hard: Max Power

• Used COTS to establish a heuristic• Rad-Hard plotted against heuristic as verification

John Albritton


Rad-Hard Computer Mass Rad-Hard Computers MHz Power Mass

MASS N/A 11 W < 1 kg

RAD6000 25 7.5 0.849

RHPPC 100 11 W 1kg

RHPCM(p) 120 11W 1kg

RAD750 166 10W .549 kg

RAD750*** 200 10W 1kg

Proton100K 400 7-9 W N/A

SCS750* 800 8-22 W < 1.5kg

• As technology increases, performance increases, but alsohardware is smaller and lighter, thus mass is relatively constant.• We adopted a mass estimation above the current heaviest andhighest performing Rad-Hard computer of 1.5 kg.

John Albritton


Rad-Hard Computer• Conclusion: By 2010:

• Other Features:– Single Event Latchup immune– Single Event Upset rate 1E-6 upsets/day– Direct Memory Access Bus 1GB/s for data storage– Selectable processor speed and power consumption

• Open Items: Look into ThinkPad 760X, ThinkPad755C, AP-101S

256 MB1.5 kg96 W3400-50 MHz

SDRAMMassPowerClock-Bus Speed

John Albritton


Dual Computer Interface

Non-volatilememory card

Solid staterecorder

Solid staterecorder

Non-volatilememory card

Storage

3

2

2

5

Transit

131Chassis

3-431Habitable

Living

1-532 (1)Habitable

Science

152Power

BaseLanding# of

ComputersModule

Mike Naylor

1 – Minimum processing power – sensor monitoring, basic communications3 – Medium processing power – GUI interfaces, full communications, calculations5 – High processing power – High rate data intake and analysis, autonomous tasks


Multi-Processor Issues• Each module will be equipped with up to two computer

systems that utilizes three CPUs each– Two of three processors used during healthy mode– Third processor monitors heartbeat of other two and takes control in

the event of CPU failure

• Measures must be taken to ensure correct handling of alltasks– Tasks must have locks to ensure only one processor attempts it at a

time– Local memory to perform current task– Shared memory to provide information to other processors about

state of completed tasks– Memory backups for recovery on separate CPU/computer

Mike Naylor


CPU Monitoring

– Hardware monitoring• Voltages throughout system• Memory parity checking• CPU temperature• Clock/timer input sync

– Software monitoring• Unexpected results trigger diagnostics• Checksum calculation verification on memory contents• Extra diagnostics run during idle periods• All functions return status codes• Timeouts on all loops• Event logging

Mike Naylor

Each set of computers will need to be monitored onboth a hardware and software level:


Computer System Fault Path

Mike Naylor


One CPU Failure…

Mike Naylor


Two CPU Failures…

Mike Naylor


External Lights• Made from super bright

LED’s (light emittingdiodes)

• LED’s will be bundled intoa “headlight” configurationconsisting of 19 individualdiodes

• The resulting bundle willbe 3-4 times brighter thana standard headlight, withat least as much range(~.15 km)

Close up of light

Rob Winter


External Lights• LED used LEDW47-66-

60-120 by Laser Optronix• Each LED produces 13

candles of light• Each LED weighs .037 kg

which makes each bundleweighs .71 kg

• Each LED needs 4.56 W ofpower

• Lights will be used duringtravel phases for far sideand night traversing

LED

• Lights will be used duringbase phase for perimeterlighting for safety

31.5 mm

Rob Winter


External Lights

Two types of lights made from the bundles

Non-movable

±125°±125°

Limited Movement lights

Limited Movement can also pitch up and down 35°

Rob Winter


External Light PlacementPower module:3 lights integratedwith sensor suite

Rover (not pictured):1 light mounted on a rotating mount

Robonaut onwheels: (notpictured)1 mounted light

Habitable module:2 limited movement lightsmounted on opposite sides

Power lights mounted with lidar suites

Lights mountedon Hab modules

Rob Winter


Module Mass Breakdown

Habitable Modules

Living Specialized:

Science Specialized:

Kevin Eisenhower, DanielSenai & Luke Twarek


Module Mass Breakdown

Power Module

Chassis Module

Kevin Eisenhower, DanielSenai & Luke Twarek

Avionics 100 kg

Structural 650 kg

Transit Connections 200 kg

Drive System 650 kg

Power Systems 90 kg

Total 1480 kg

Margin 60%

Avionics 200 kg

Structural 750 kg

Drive System 650 kg

Power Systems 240 kg

2 Robotic Arms 150 kg

Total 1980 kg

Margin 46%


Power Distribution• Combination of AC and DC power

– AC enables multiple base configurations whilereducing switchgear

– DC reduces development costs for electricalcomponents by utilizing already developedcomponents

– Exact voltages and frequencies are still beingdetermined to minimize power losses and reducearcing potential

David Pullen


Power Connections & Distribution

Chassis Modules:• AC/DC Rectifier• Rechargeable Battery

Habitable Modules:• AC/DC Rectifier• Rechargeable Battery

Power Modules:• DIPS• AC/DC Rectifier

AC Power Source

David Pullen


Power Module Power Layout

David Pullen


Habitable Module Power Layout

David Pullen


Chassis Module Power Layout

David Pullen


Unmanned Transit Phase

Vehicle Assembly


Transit Phase Overview• After each manned

mission, MORPHLABwill disassemble andtravel autonomously to anew base site

• Enables the explorationand analysis of vastareas of the Lunarsurface

• Science will beperformed by themodules while traveling

David Pullen


Transit Phase Topics• Transit assembly layouts• Mapping satellites• Position determination & obstacle avoidance• Transit power budgets• Transit connections• Steering• Module control scenarios & failure modes• Science performed in transit

David Pullen


Vehicle Assembly

• 3 Vehicle Assemblies:– Each vehicle assembly consists of a power module

and two habitable – chassis assemblies– The modules are joined through transit connections

on the chassis– Power and data are transferred through the transit

connections

Width:4mLength: 22.9m

Transit connection

David Pullen


Vehicle Assembly

• 1 power module isdedicated to carrytwo lunar modules

Width: 7.6m Length: 7.0m

• Vehicle assemblies can take different paths tothe next base site, expanding the area coveredfor scientific exploration

David Pullen


Mapping Satellite• Current lunar

topographical information– Clementine LIDAR

mapping mission• Resolution

– Vertical: 40 m– Horizontal: 100 m

• Cross-track measurementspacing: 40 km

• In-track measurementspacing: 1 – 2 km

• Coverage limited to ± 60°latitude

• Photography– Lunar Orbiters

• Photographed 95% ofthe surface

• Resolution up to .3 m

• Derived requirements– Courses must be

plotted prior todeparture

– Polar regions need tobe analyzed fortopography and waterice deposits

David Pullen


Mapping Satellite• Specifications derived

from Mars GlobalSurveyor (MGS)– MGS Narrow Angle

Camera• Resolution: 1.4 m/pixel• Field of view (FOV) up to

2.9 km X 25.2 km– MGS LIDAR

• Resolution– Vertical: 2 m– Horizontal: 160 m

• 10 Hz sample rate

• Assumptions– Mission duration: 60 days– Circular Orbit– Tech capabilities scaled

by 2x compared to MGS– Photographs

• Taken on light side only• Resolution and FOV

scaled linearly vs. altitude– LIDAR ranging

• Taken on both sides• No across-track

measurements

David Pullen


Mapping Mission Analysis• Nominal mission

duration: 60 days• Altitude: 115 km• Camera specs

– Resolution: 20 cm– Field of View:

1.5km X 15 km• LIDAR specs

– 20 Hz sample rate– Measurements:

• In-track: 76m• Cross-track: 15 km

David Pullen


Position Determination• MORPHLAB must be capable of determining its position

and attitude during its entire mission duration– This will be accomplished with:

• Inertial mass units (IMU)• Star trackers• Communications with other modules and Earth

Tom Farrell


Position Determination• Currently, IMUs and star trackers are used in

many space applications for position andattitude determination– Current specifications will be sufficient to meet

requirements– Specific systems used in MORPHLAB will be

determined closer to the technology cutoff date• Design constraints

– Low Mass– High Reliability

Tom Farrell


Position Determination

Tom Farrell

Current IMU specifications– Source: Ball Aerospace– Mass: 1.4 kg– Power requirement: 6 W– Gyro

• Range: +/- 200 deg/sec• Bias: +/- 1 deg/hr

– Accelerometer• Range: +/- 12 g’s• Bias: +/- 0.3 mg’s• Clock: 10MHz• Activation time: 5 sec


Position Determination

Tom Farrell

Current star tracker specifications– Source: Ball Aerospace– Mass: 2.5 kg– Power requirements: 9W– Angular accuracy: 40 arc sec– Update rate: 10 Hz– Field of view: 20° x 20°– Catalog Size: 6000 stars


Obstacle Avoidance

Tom Farrell

• Level 1 requirements– MORPHLAB must be capable of traversing

obstacles smaller than 0.5 m– MORPHLAB must be capable of traversing

slopes up to 20° cross-slope or 30° direct slope• Derived requirements

– Sensors must be capable of resolving a 0.5 mobstacle at a distance of no less than 7 m away

– Transits must be plotted prior to departure toensure paths are free of large obstacles or slopesover the entire route


Obstacle Avoidance• Rough course will be determined from Earth by

analyzing mapping satellite data prior to missionstart

• Small obstacles will be located and avoided in real-time using on-board sensor suites– Power module sensor systems

• Laser imaging radar (LIDAR) – resolves obstacles regardless oflighting

• Video cameras – provide visual images for teleoperation /supervisory control

– All other modules will be controlled by the powermodules via communications links

Tom Farrell


Obstacle Avoidance• Power modules will have 3 complete sensor suites

– Mounted on gimbals in rotating assemblies• Azimuth range of vision: 360°• Elevation range of vision: 180°

– One sensor is located in the rear, dedicated torobotic arm operations

LIDAR/Camera Housing

Tom Farrell


Obstacle Avoidance

Tom Farrell

• Current LIDAR specifications– Mass: 3.6 kg x 3 = 10.8 kg– Power requirement: 16 W x 3 = 48 W– Pulse duration: 8 ns– Measurement rate: 12 kHz– Distance range: 0 – 250 m– Accuracy: +/- 6 mm– Acquisition time: 2 sec

• Current video camera specifications– Mass: 0.3 kg x 3 = 0.9 kg– Power requirements: 3 W x 3 = 9 W– Image: 1024 x 1024 pixels


Obstacle Avoidance

Tom Farrell

• The following conditions must be known or monitored:– Situational awareness

• State of own vehicle– Current speed / acceleration– Position on surface– Adherence to planned route– Reaction time of human or computer in control

• State / model of nearby objects / environment– Location and classification of object– Patterns of transit hindering objects and lighting distortions

– Processing and algorithms• What situation are we in?• How likely is a collision?• How dangerous is the situation?• Is an action needed?


Obstacle Avoidance

Tom Farrell

• System response– Human or on-board computer senses a transit

hindering obstacle via sensor systems– Collision warnings

• Audio and visual warning alert user of a likelycollision

• Vehicle comes to a stop

– Collision avoidance• Steering and speed of vehicle altered to avoid obstacle


Operational Requirements• Level 1 requirements:

– Systems onboard MORPHLAB shall be capable of operating in anyof the following control modes for any or all of the nominal missionsegments:- On-board direct human control- Teleoperation- Supervisory control- On-board autonomous control

• Derived requirements– A human operator shall be able to take over the control of the vehicle

from Earth and teleoperate if a critical situation arises (e.g. less than0.01m from an approaching boulder larger than 0.5 m in height)

– If an obstacle is observed, but the planned trajectory does not avoidthe obstacle, supervisory control shall be implemented to avoidobstacle.

Tom Farrell


Teleoperation and Supervisory Control

Tom Farrell

• Some situations may arise that will force MORPHLAB outof autonomous control and into either supervisory orteleoperational control

• Supervisory – operation via preprogrammed sequence of events fromEarth followed by the execution of these commands by MORPHLAB– Shall be implemented if MORPHLAB detects a potential problem that

cannot currently be avoided via autonomous control– Reliant on accuracy of on-board computers

• Teleoperation – operation of MORPHLAB from Earth– Shall be implemented if MORPHLAB senses an unavoidable object– Ground user shall take over control, navigate around object, and return

MORPHLAB to autonomous control– Reliant on accuracy of navigational equipment

• Teleoperation and supervisory control modeled from previous Earth, Mars andLunar exploration missions: Dante II, NOMAD, Sojourner, Lunakhod


Transit Peak Power System

Habitable Module: 101 kg Li-Ion Bank

Chassis: 44 kg Li-Ion BankShort TermPower

Li-Ion Battery Banks:

•Sized to provide 24 hours of standby power to an unmated module

•Under high system loads when modules are linked, battery resources can bepooled to meet peak power needs

Joe Cavaluzzi


Vehicle Assembly Power Budget

6.8 kWeTotal1.5 kWeBattery5.3 kWeDIPS

Generated Power

1 kWeAvionics5 kWeTotal

1 kWeThermal Control3 kWeLocomotion

Vehicle Assembly System Average

1.3 kWeAvionics6.7 kWeTotal

1.2 kWeThermal Control4.2 kWeLocomotion

Vehicle Assembly System Peak

Joe Cavaluzzi


Locomotion Resistive Forces•Soil Compaction Resistance Rc = func.(wheel width, soil properties, wheelsinkage)

•Bulldozing Resistance Rb = func.(wheel width, soil properties, terrain slope)

•Rolling Resistance Rr = func.(soil properties, wheel loading)

•Gravitational Resistance Rg = func.(wheel loading, terrain slope)

Assumptions:•Maximum chassis wheel load = 1962 N

•Maximum power module wheel load = 1430 N

•Terrain slope will not exceed 30 degrees

•Rolling friction is estimated as 0.2 (comparable to firm sand)

Joe Cavaluzzi

Forward Motion:

Tractive Force > Motion Resistance


Transit Power Usage

0º - 16 º slope : 16 º - 30 º slope :Transit Speed 1.2 - 1 km/h Transit Speed 1 - 0.7 km/h

1.2 kWe300 WPower Module4.2 kWe-Vehicle Assembly

1.5 kWe375 WChassisMax Motion Power (4 wheels)Max Power/WheelComponent

Joe Cavaluzzi


Transit Connection

Kevin Eisenhower

0.8 MPa

10,000N10,000N

Ball joint rotates± 45° left/right

Ball joint rotates± 30° up/down


Turning Capabilities• True Double-Ackerman Design:

– Back Wheel/Front Wheel turnangle ratio of 1

• Turning radius for single module:– Wheelbase Length: 4.2m.– Wheelbase Width: 3.3m.– Wheel turn Angle: 30 deg.– TURN RADIUS = 4m

• Turn radius for single vehicleassembly:– Total vehicle length: 20m.– Total vehicle width: 3.3m.– Wheel turn Angle: 30 deg.– TURN RADIUS = 8m

Daniel Senai

R = 8m

R = 4m

Note: Figures not to scale


Steering Mechanism• Modified rack-and-pinion system• Four-wheel steering with independent motors• Allows modules to rotate with no translation• Lunar dust effect on steering is a major concern

Abe Daiub, Daniel Senai

Top ViewNote: Figure not drawn to scale


Steering Mechanism

Abe Daiub, Daniel Senai

20Net Torque feedbackto Steering System

0Overturning Torque

0Rolling Resistance Torque30Aligning Torque50Tractive Force (Fx)0*Lateral Force (Fy)-48Vertical Force (Fz)

(N-m)Torque due to:

Fx

Fz

FyMx

My

Mz

* Assumed slow motion


Slope Traversal• Level 1 Requirements

– 30 degree direct slope– 20 degree cross slope– 0.5 meter obstacle

• Max direct angle: 36.8º• Max cross angle: 28.7º• Max C.G.: 2.80 m

Emily Tai


Hang-Up Failure

• Vehicle ground clearance: 1.35 m• Clearance Requirements

– 30º direct slope• 0.542 m (up-slope to flat)• 1.10 m (up-slope to down-slope)

– 20º cross slope• 0.2845 m (up-slope to flat)• 0.562 m (up-slope to down-slope)

Emily Tai


Nose-In Failure

Vehicle overhang from wheel base < wheel radiusNose-In Failure is not a concern

Emily Tai


Science in Transit

• Capable of collecting and storing soil samples• Equipped with soil analyzing instruments• Vehicle assemblies may travel separate to sample more of the surface• More details forthcoming

Kevin Eisenhower

2 robotic arms on each power module:

Arms stowed


Intermission


Manned Base Assembly Phase


Base Phase Topics• Science in base assembly• Docking Mechanism• Base power budget• Crew Systems & ECLSS• Logistics Module• Interior layout• Extra vehicular activity• Lunar roving vehicle

Kevin Eisenhower


Base Assembly Site• Crew support equipment is

supplied by a logistic depot,landed near the desiredassembly site

• The specific base locationand configuration will bedetermined whenMORPHLAB arrives

• Assembly will be within10km of logistic depot

• Astronauts land within 100mof the assembled base

• Tele-operated lunar roverspick up landed astronauts

Susumu Yamamoto


Base Assembly Overview• Transit connections

disconnect and modulesreconfigure into a base

• While reconfiguring,batteries supply power tothe habitable and chassismodules

• Once the modules arepositioned correctly, thedocking mechanismsextend and connect

Susumu Yamamoto


Base Assembly Overview• Power modules connect power

cords to the chassis moduleswith the robotic arms

• Power modules are driven 10maway from the base assembly toreduce radiation exposure to thecrew from the DIPS

• The modules arepressurized andcomputers are initializedto prepare for the arrivalof the astronauts

Susumu Yamamoto


Module Overview• 2 Habitable-Living: Crew Quarters

– 2 docking hatches– Primary used as sleeping quarters

• 1 Habitable-Living: Galley– 2 docking hatches– Primary used as a galley

• 2 Habitable-Science: EVA– 1 EVA Airlock– 2 docking hatches

• 1 Habitable-Science– 3 docking hatches

Bed

Liv:Galley

EVA

Sci:Galley

Susumu Yamamoto


Optimum Base Assembly• Base Assembly Requirement

– At least one habitable-science module is needed toprovide all habitable-living modules with complete lifesupport

– Each module should have two routes to reach EVAairlocks for an emergency evacuation

• Level 1 requirement L8 : System shall provide for emergencyalternative access and EVA "bailout" options

• Consideration for base assembly– Base assembly should be comfort for living and working

on missions for astronauts

Susumu Yamamoto


Base Assembly Configuration• Primary Base

Assembly

EVASci:Galley Bed

Bed

EVA

Liv:Galley

Susumu Yamamoto


Base Assembly Configuration• Optional Base Configuration

EVASci:Galley

Bed

Bed

EVA

Liv:Galley

Susumu Yamamoto


Assembly Site Selection Criteria

Kevin Eisenhower

• Apollo landing site candidates• Volcanic lava “tunnels” (rilles)• Permanently shadowed regions• Age of lunar structures• Highland material• Approximately 1000 km apart

– Helium– Nitrogen– Oxygen

– Carbon– Silicon– Thorium KREEP*

materials

Solar cellproduction

* Potassium, Rare Earth Elements, Phosphorus

Water ice and ore depositsas potential resources

Deep lunar composition

Volcanic and impact history

Environmental variationsalong the surface

ice

• Ore deposits and useful elements:

Level 1 Science Goals:


Assembly Site Selection

Kevin Eisenhower

12 3

46 5

10

89

7 6

Near Side Far Side

http://geopubs.wr.usgs.gov/i-map/i2769/* Site rankings done bytrade study of goals, sitecriteria, and feasibility


Assembly Site Distances

Kevin Eisenhower

Assembly Site

Straight line

distance

between (km)

Estimated

driving distance

(km) (+ 30%)

1) Marius Hills

515 670

2) Flamsteed Crater P

479 623

3) Gassendi Crater

966 1256

4) Tycho Crater

911 1184

5) Moretus Crater

553 719

6) Drygalski Crater

212 276

go through South Pole

546 710

7) Schrodinger Crater

562 731

8) Minnaert Crater

822 1069

9) Mare Ingenii

539 701

10) Gagarin Crater

Largest

distance

University of MarylandCDR 04/19/04 Kevin Eisenhower

1

1) Marius Hills(51W, 11N)• Steep volcanic hills• Large amounts of silica• Rilles nearby could be

collapsed lava tunnels2) Flamsteed Crater P

(43W, 4S)• 1m thick regolith infers

that it is young• Surveyor 1 landing site

Marius

Flamsteed


3) Gassendi Crater(38W, 19S)• Contains ancient

highland rocks• Lunar transient

phenomenonobservations

• Most likelycandidate forApollo 18

Flamsteed

GassendiCrater

Kevin Eisenhower


4) Tycho Crater(12W, 42S)• Rare materials• Lunar transient

phenomenon observations• Crater still has rays infers

that it is young• Youngest known crater by

current knowledge• Surveyor 7 landing site

GasendiCrater

TychoCrater

Kevin Eisenhower


5) Moretus Crater(9W, 72S)• Possibility of ice in

permanently shadowedcraters

TychoCrater

MoretusCrater

Kevin Eisenhower


6) Drygalski Crater(83S, 90W)• Possibility of ice in

permanently shadowedcraters

• Areas of constant sunlightnearby (MalapertMountains)

DrygalskiCrater

Moretus Crater


7) Schrodinger Crater(72S, 134W)• Contains recent

volcanic ejecta• Second youngest

impact basin

Schrodinger Crater

DrygalskiCrater


8) Minnaert Crater(67S, 171W)• Strongest possibility

of investigating lunarmantle

• Located withinAitken Basin

Minnaert Crater

Schrodinger Crater


9) Mare Ingenii(40S, 167W)• High concentration of

thorium• Possibility of investigating

lunar mantle• Located within Aitken basin

MinnaertCrater

MareIngenii


10) Gagarin Crater(29S, 150W)• Contains ancient highland rock

Mare Ingenii

GagarinCrater


Geological Science Instruments

Potential water ice

Ore deposits aspotential resources

Deep lunarcomposition

Lunar volcanic andimpact history

Lunar dust andenvironmental

changes

Level 1

Soil collection

Seismicinstruments

Mineralanalysis

Dust detectinginstruments

Surveyor instrumentretrieval

Level 2 Level 3• Mini Thermal Emission spectrometer• X-ray / Gamma-ray spectrometer• Mossbauer spectrometer• Integrated Dust/Soil Experiment mass spectrometer

• Passive Seismic Experiment Package• Active Seismic Profiling Experiment

Robotic arms

• Scoop• Hammer• Rake• Tongs• Electric drill

Lunar Dust Detector

Kevin Eisenhower


Soil Sample CollectionApollo Soil Collection Details

total rock largest rock

Apollo mass (kg) # of rocks returned (kg) mission hours EVA hours

11 21.6 68 5.6 195.5 2.25

12 34.3 69 2.4 31.5 7.5

14 42.3 150 9.0 33.5 9.5

15 77.3 370 9.7 67 18.5

16 95.7 731 11.7 71 20.25

17 110.5 741 3.6 75 22

Apollo

avg EVA hours /

mission hours

average rock

mass / EVA hour

average

rock # / EVA hour

11 0.012 9.6 30.2

12 0.238 4.6 9.2

14 0.284 4.5 15.8

15 0.276 4.2 20.0

16 0.285 4.7 36.1

17 0.293 5.0 33.7

average 0.275 4.6 23.0

MORPHLAB manned mission: 2160 hours (3 months)Total collected rock mass: 600 kg

Assumption:Collect at 1/4th the

rate of Apollo(larger stay times,

and more timespent traveling in

the rover)

Kevin Eisenhower

Apollo 11 discarded


Soil Sample CollectionApollo Rock Samples

0

5

10

15

20

25

30

0 - .2 .2 - .4 .4 - .6 .6 - .8 .8 - 1 1.0 - 2.0 2.0 - 4.0 4.0 - 6.0 6.0 - 8.0 8.0 - 10 10+

rock mass (kg)

nu

mb

er o

f ro

cks

11

12

14

15

16

17

Keep rocks for 7 days 46 EVA hours (28% mission hours)50 kg of rocks collected

+ 300 kg of rocks returned each mission (50% total collected)= 350 kg total rock storage mass

Want robotic armcapable of picking

up 6 kg rock

Kevin Eisenhower


Robotic Arm Capabilities• Pick up and store soil samples• Soil inspection• External module inspection / maintenance• Removal of descent engines• Plugging in power cords

lift at min. temp. max temp.

mass (kg) reach (m) extension (N) (deg C) (deg C)end effectors

Ranger Arm 77 1.35 133 -40 150interchangeable

Requirements: 2 59 -173 127

6 kg rock moon

http://ranger.ssl.umd.edu

• Different end effectors for each operation• Modifications necessary (1m extensions, temperature)

Kevin Eisenhower

Arms stowed

Arms extended


Life Sciences: Level 1 and 2

Muscle development/decay

Cardiovascularsystem

Skeletal system

Long term radiationexposure

Nutrition andmedical care

Human performance,adaptations

Plant biology in thelunar environment

Ultrasound

Questionnaires

EVARM system

Gas Analyzer Systemfor Metabolic AnalysisPhysiology (GASMAP)

Portable ClinicalBlood Analyzer

(PCBA)

AdvancedAstrocultureTM Chamber

Crew dynamics

Foot/Ground Reaction System

Ambulatory DataAcquisition System

Paula C. Herda


Life Sciences: Level 2• Most of these systems have been used on the

ISS. The reason for repetitions are:– Different gravitational environment– Find trends from mission to mission– Find a relationship between the Earth’s gravity,

the Moon’s gravity, and microgravity; theneducated predictions for other planetary gravitiescan be made

– Enhance technology for medical care on Earthand future planetary bases

Paula C. Herda


Racks• Each science module will

have one science rack

• These racks contain theinstruments needed toperform variousexperiments

• The rack method makes iteasy to changeexperiments & instrumentsfrom mission to mission

• Rack 1– PCBA (Portable Clinical Blood

Analyzer)– The PCBA Cartridge Kit– PCBA Control Kit– EVARM system

• Rack 2– The Advanced Astroculture– Foot Ground Interface - Flight Cal

Unit– Total Force - Foot Ground Interface– Lower Extremity Monitoring Suit– Ambulatory Data Acquisition System

• Rack 3– Ultrasound– GASMAP /calibration module

Paula C. Herda


Questionnaires

• “The objective of this study is to measure theimpact of cultural and language backgroundon space missions and to characterize changesover time in a number of importantinterpersonal factors, such as tension,cohesion, leadership role, and the relationshipbetween space crews and monitoringpersonnel on Earth” (http://lsda.jsc.nasa.gov )

• Questionnaires are submitted via computerback to Earth

Life Sciences: Level 2

Muscle decayCardiovascularsystem

Skeletal system

Radiation

Nutrition &medical care

Humanperformance,adaptations

Plant biology

Crew dynamics

Paula C. Herda


Life Sciences: Level 2• Bones have a steadydecay, even withexercise.

• Muscle strengthdecreases significantly

• In lunar gravity, thesenumber are notexpected to be asdrastic, but bone andmuscle decay still needto be studied

Muscle Decay in Space

020406080

100120

0 50 100 150 200 250 300 350Day

Perc

ent

max explosive powermuscle strength

Bone Decay in Space

95.000

96.000

97.000

98.000

99.000

100.000

0 20 40 60 80Day

Perc

ent

w/o exercise

w exercise

Paula C. Herda


Foot/Ground Reaction System– Lower Extremity Monitoring Suit– Total Force – Foot Ground Interface– Foot Ground Interface – Flight Calibration Unit

• Measures the forcesyour legs and feetexperience through

a normal work day

• Mass: 46 kg http://spaceflight.nasa.gov/station/science/experiments/fact_foot.html


Paula C. Herda


Skeletal system

Radiation



Plant biology

Crew dynamics


Ultrasound• 3D capability &

real time 2D• Respiratory

Trace Display

• Total Mass: 86 kg

http://hrf.jsc.nasa.gov/ultrasound.htm

Life Sciences: Level 2• Applications:

– Echocardiography

– Abdominal (deep organ)

– Vascular, muscle, tendon,trans-cranial ultrasound

Paula C. Herda


Skeletal system

Radiation



Plant biology

Crew dynamics


• Measures volume,temp, frequency,pressure of inhaled &exhaled air

• To be used duringexercise

• Total Mass: 244 kg

http://hrf.jsc.nasa.gov/gasmappics.htm

Life Sciences: Level 2Gas Analyzer System for MetabolicAnalysis Physiology (GASMAP)• Study effects of reducedgravity on the musculoskeletalaspects of breathing duringrest, heavy exercise, and deepbreathing

Paula C. Herda


Skeletal system

Radiation



Plant biology

Crew dynamics


Portable Clinical BloodAnalyzer (PCBA)

• Study red blood cell massand production.

• Mass PCBA: 0.54 kg• Mass Cartridge Kit: 0.27 kg

http://www.amedical.com/istat.html


Paula C. Herda


Skeletal system

Radiation



Plant biology

Crew dynamics


• Provides convenient dataacquisition and storage of datacollected by physiological sensors(such as FOOT). It will measurecore body temperature, bloodpressure, respiration and otherphysiological responses.



Skeletal system

Radiation



Plant biology

Crew dynamics

Paula C. Herda

Ambulatory Data Acquisition System


• Mounted inside theEVA suit

• Worn on the head,torso and leg


http://resources.yesican.yorku.ca/trek/radiation/final/index_EVARMS.html

EVARM System

Paula C. Herda

• Records radiationover different partsof the body duringan EVA


Skeletal system

Radiation



Plant biology

Crew dynamics


• Autonomous &continuous care forthe plant for anaverage 180 days.

http://wcsar.engr.wisc.edu/advasc.html



Skeletal system

Radiation



Plant biology

Crew dynamicsAdvanced AstrocultureTM (ADVASC)

•Mass: 4.4 kg

• Growing Space– Shoot area: 486 cm2

– Shoot height: 34.5 cm– Root area: 486 cm– Root height: 5 cm

• The top box contains thecomputer, electronics &other support systems

• The bottom box containsthe controlled plant growthchamber and ancillarysubsystems

Paula C. Herda


AA batteries2.31Ambulatory DataAcquisition System

Peak 300 W46.7AstrocultureTM Chamber

1574W426 kgTotals

--Questionnaires--EVARM system

Two 9 volt batteries.81 kg(PCBA)Peak 270 W244 kg(GASMAP)Peak 1004 W86 kgUltrasound

Batteries included 1.5 V dc46 kgFOOTPowerMass: kgInstrument

Summary of Instruments

Paula C. Herda


Habitat to Habitat Connections• To fulfill requirement M5, that all nominal movement

between modules be accomplished in shirtsleeve conditions,some form of pressurized mating tunnel is required.

• Space limitations in the launcher payload shroud dictatedthat this system be retractable.

• Mass limitations on the habitat modules dictated that thissystem be as light as possible.

• Concerns over tipping while in transit dictated that thissystem needed to be able to deploy and retract before andafter each crewed mission segment.

Luke Twarek


Paramount Tunnel Tradeoffs• Metallic vs. Composite• Male/Female mating vs.

Androgynous mating• Layered pressure bladder vs.

Abrasion cover• Active terminal mating vs.

Passive terminal mating• Passive latch connection vs.

Active bolt connection• Doors: normally open vs.

normally closed

Luke Twarek


Additional Tunnel Concerns• To modularize the air and water

purification systems– Airflow must be allowed between

habitable sections– Piping for water transport would need

to exist between habitable modules

• Power delivery between modulescould not be accomplished usingthe mechanical chassis linkagesduring Base Assembly Phase– Cables for power transfer would need

to exist between habitable modules

Luke Twarek

Water

Air

Power

Module

Module

Tunn

el


Extendable Tunnel• Foldable composite (Kevlar-

49) was chosen over atelescoping metal tube.– Mass savings of 8x– Lighter extension mechanism

than a large telescopingstructure

– Less likely to puncturepressure bladder

• Layers will be held together atthe ends by clampedaluminum hoops.

• Two lengths: 1 m and 0.5 m• Mass of 1 m tunnel: 10 kg• Mass of 0.5 m tunnel: 6 kg

Luke Twarek

MLI Layers

Kevlar-49 Cross-Weave Stiffness Layer

PVC Pressure Bladder

Nomex Cross-Weave Abrasion Layer

***Inside (P = 25.7 kPa)***

***Outside (P = 0)***


Mating Ring• Androgynous mating system

chosen over male/female systemto maximize possible baseconfigurations.

• Tunnel connected to mating ringvia clamping mechanism withself-sealing o-rings to ensurepressure seal.

• Power and water pipingconnections located in ring– Hard dock seals piping– Piping runs inside or parallel to

extension mechanism• Total mass of mating ring: 80 kg

Luke Twarek

Aluminum Ring

Dust Cover

Pressurized Seal Bladder

Pressurization InletA-A

A-A

A

A

Dimensions in meters


Hard Dock• Antisymmetric nut/bolt

configuration providesandrogynous interface.

• Inner and outer hard connectionpoints.

• 16 connections needed to ensureboth two-fault tolerance andantisymmetry.

• Active bolt hard dock chosen due to2x mass savings over passivelatches.

• Spring-loaded dust covers ensureno particles get into threads whennot in base assembly.

Luke Twarek

AA

A-A Soft Dock

A-A Hard Dock

B=BoltN=Nut

N

N

N

N B

BB

B


Pressure Seal• After hard dock, PVC pressure

seal bladder inflates in order toform a self-sealing pressure seal.

• In the event of a rupture in thebladder, the other pressure sealbladder can inflate to seal theleak.

• Spring-loaded dust coverprevents dust and UV radiationfrom degrading the PVC whennot in base assembly.

• Pressurization lines carried withextension mechanism.

Luke Twarek


Extension Mechanism• Exact mechanism has not

yet been chosen, thecandidates are:– Telescoping tube

• Can also be used to transferpower and water

– Rotators and beams• Can be used for precision

alignment of mating rings– Extending beams

• Easily adapted from chassis

Luke Twarek


Mating Sequence• Initial alignment with

chassis wheels• Initial tunnel extension• Terminal alignment via

sensors• Terminal tunnel extension• Soft dock• Extend bolts• Hard dock• Pressure seal inflation• Pressure bladder inflation• Exterior walkway

deployment• Doors open to circulate air

Luke Twarek


Base Peak Power System

MORPHLAB Net Battery Totals

Mass: 870 kg

Storage Capacity: 108 kWhr

24 Hour Discharge: 4.5 kWe

Maximum Power Output: 20.4 kWe (for 24 hours) With 3 DIPS

Joe Cavaluzzi


Base Power Budget

20 kWeTotal4 kWeScience

1 kWeBattery Recharge

3 kWeThermal Control

1 kWeHouse Keeping3 kWeAvionics

8 kWeCrew Systems

Base System Peak

20.4 kWeTotal4.5 kWeBattery

15.9 kWeDIPS*

Generated Power

*Output of 3 Power Modules

15 kWeTotal3.5 kWeScience

0.5 kWeBattery Recharge

2 kWeThermal Control

1 kWeHouse Keeping2 kWeAvionics

6 kWeCrew Systems

Base System Average

Joe Cavaluzzi


Crew Systems

• All crew interfaces shall– Accommodate the 95th

percentile American male to the5th percentile Japanese female(L1)

– Adhere to NASA STD-3000 (L2)

• For optimal performance onlong duration missions– Minimum surface area = 61 m2

– Min. ceiling clearance = 2.2 m– Min. interior volume = 135 m3

Design Constraints:• Nominal mission consists of

– 4 member crew (L4)– Capable of daily 2 person EVAs(L5)– Zero pre-breathe time for EVAs(L9)

• Emergency requirements– 180 day subsistence food stock(L6)– At least 2 IVA exits from eachmodule–EVA bailout options(L8)

* IVA = Inter-Vehicular Activities

Cat McGhan &Jessica Seidel


MORPHLAB vs. LMLSTP• Lunar-Mars Life Support Test Project

(LMLSTP)– NASA/ Johnson Space Center project– 4 crew members spending 90 days in a closed-

loop life support system habitat– Habitable Surface Area = 58.44 m2

– Habitable Volume = 163.63 m3

• Total for MORPHLAB base configuration– Habitable Surface Area = 75.4 m2

– Habitable Volume =198 m3

Jessica Seidel


Module SubtypesHabitable-Living:• 2 Crew quarters

– Sunken SPE-shielded beds– Sanitary systems, main level

• 1 Galley– Food preparation area– Exercise area, lounge– Backup quarters (I3 & I11)

• ECLSS systems– Vapor Phase Catalytic

Ammonia Removal– Multi-filtration– Supercritical Waste

Oxidation

Habitable-Science:• 2 EVA areas

– Main airlocks, dust removal– Repair area, tool storage– Durable experimentation area

• 1 Experiment area– Finer specialized equipment– Backup galley (I3 & I11)

• ECLSS systems– Water Vapor Electrolysis– Electrochemical Depolarized

Cells– Sabatier– Trace Contaminant Control

Cat McGhan &Syed Hasan


Atmosphere Selection• Cabin pressure set at 57.1 kPa

– Design Considerations:• Pressure low enough to allow R value under 1.6

(NASA standard)• Pressure high enough to allow for minimum fire risk

• 40% O2, 59% N2, 1% CO2– Calculations of partial pressures of each of the

gas’ minimum and maximum levels for longduration exposure were used to determine thismake-up

– Levels for hyperoxia, hypoxia, CO2 toxicity, andfire prevention from NASA STD-3000 document

Syed Hasan


ECLSS – Overview

Syed Hasan

1.440.00040.100.150.050.15/0.111.60PowerReq. (kW)

2.120.00120.240.300.040.070.75Volume(m3)

694.003.9068.00100.0017.9044.40144.00Weight(kg)

SCWOMFVAPCARTCCSabatierEDCWVESubsystem

Supercritical Water OxidationSolid Waste Removal: MultiFiltrationWater Filtration: Vapor Phase Catalytic Ammonia RemovalWater Distillation: Trace Contaminant Control SystemTrace Contaminant Control: SabatierCarbon Dioxide Reduction: Electrochemical Depolarization ConcentratorCarbon Dioxide Removal: Water Vapor ElectrolysisOxygen Generation:

* All values for crew per day


ECLSS – Needs and Effluents

101.6Hygiene Water

2.48Solids9.48Supply Water

3.36OxygenSupply (kg)Needs

Product (kg)Effluents

0.44Solids

6.00Supply Water

101.6Hygiene Water

4.00Carbon Dioxide

Oxygen - Shortage of 0.40 kg per crew/day

Water vapor Electrolysis produces 2.96 kgper crew/day while 3.36 kg is required

Hygiene Water – no shortage.

Multifiltration produces 100% water resupplyup to flow of 114 kg day ( > 101.6 kg )

Supply Water – Shortage of 3.48 kg percrew/day

Vapor Phase Catalytic Ammonia RemovalSystem produces 100% water resupply up toflow of 13 kg/day (> 6 kg). 6.00 kg or supplywater is recycled while 9.48 kg is required.

Solids – 2.48 kg required per day

Supercritical Waste Oxidation has 0%resupply flow rate. All solids needed must beprovided using food.

* All values for crew per day

Syed Hasan


ECLSS – Interior LayoutsHabitat Module Basement –

Living/BedHabitat Module Basement –

Science/EVA

Water VaporElectrolysis

Sabatier

Electrochemical DepolarizationConcentrator

Trace ContaminantControl

Vapor PhaseCatalytic AmmoniaRemoval

Supercritical WasteOxidation

Multifiltration SleepingQuarters

Trap door fromupstairs

Syed Hasan


ECLSS Flow Diagram

Syed Hasan


Air Flow Diagram

Syed Hasan


Water and Waste Flow DiagramsWATER FLOW WASTE FLOW

Syed Hasan


ECLSS - Summary

347.0Supercritical Water Oxidation

2.0MultiFiltration

34.0Vapor Phase Catalytic AmmoniaRemoval

100.0Trace Contaminant Control System

17.9Sabatier

44.4Electrochemical DepolarizationConcentrator

144.0Water Vapor Electrolysis

Weight (kg)Subsystem

383.0Living/Bed

306.3Science/EVA

Weight (kg)Module

208.3Solids

292.3Supply Water

0.0Hygiene Water

33.6Oxygen

Amount(kg)

Product

Supply needed permission (84 days)

Syed Hasan


Logistics Depot• Will be launched before

every mission containingre-supply of provisions

• Exterior of habitablemodule, with provisions inplace of interior equipment

• First depot:– Supplies provisions for

two months– Contains mobile

Robonauts to performmaintenance throughoutprogram

– Contains two lunar rovers

LogisticsDepot 2-10

First mission– 56 days

Remaining missions- 84 days

3121Water tanks (6)

208139Food

TOTAL

Lunar Rovers (2)

Hull

Science instruments

Landing legs

DC battery

Robonaut on wheels (2)

Water

Nitrogen tanks (6)

Nitrogen gas

Oxygen tanks (6)

Oxygen gas

ITEM

25422833

-416

410410

563563

212212

4444

-312

292195

308205

5336

345230

7550

MASS (kg)MASS (kg)

Syed Hasan

LogisticsDepot 1


Habitable-Living: Crew Quarters

*2 modules

stowage

desk desk

chairchair

stowage

closet

closetshower

sink & cabinet

toilet

mirror

dockingmechanism

radiationshelterdoors

dockingmechanism

Top View

Jessica Seidel


Habitable-Living: Galley

stowage

storage& exercise equipment

table &4 chairs

sink

refrigerationunit

shower

toilet

cabinets,microwave& counter

storagecabinets& counter

radiationshelter doors

dockingmechanism

dockingmechanism

Top View

Jessica Seidel


Habitable-Science: EVA

loungechair

sciencerack 2

airlock pumpstorage

EMUStore

airlock

desk

EMUboxes &storage

loungechair

storagestowage (3)tv

chair

storage

dockingmechanism

dockingmechanism

*2 modules

Top View

Jessica Seidel


Habitable-Science

cabinets, microwave& counter

extendablework table& 4 chairs

refrigerationunit

sink

sciencerack 3

exerciseequipment storage

storage

cabinet & counter

dockingmechanism

dockingmechanism

dockingmechanismTop View

Jessica Seidel


Interior Acoustics

Human lethalitySudden onset167

Reduced visual acuity; chest wallvibrations; gag sensations;respiratory rhythm changes

2 minutes150

Hearing loss occursContinuous120Pain in the earsSudden onset114

Reflex response of tensing,grimacing, covering the ears, andurge to avoid or escape

Sudden onset100

Protective earwear must be wornLong85

Voices must be raised significantly;telephone use difficultContinuous60-70

No conversation interferenceContinuous30-40

Performance EffectsDurationSound PressureLevel in dB

Jessica Seidel


Computer User Interface• Will be LCD flat panel, touch screen mounted on the walls

in the habitable modules– Currently have 18” LCD touch screen panels that weigh less than 2kg– No external input devices reduces clutter inside module

• System of menus to browse different command options– Sensor monitoring

• ECLSS• Power grid• Computer status• Communication status

– Communications with Earth– Communications with other modules– EVA support

• Communications with Astronauts• Tele-operation of rovers

– Entertainment

Mike Naylor


Alerts for Sensor System• Integrated sensors for

– air and water composition– pressure levels– system integrity

• Monitoring Computer System (MCS)– sensors are all tied into it; tracks changes in the levels– issues commands to the appropriate system to make

adjustments if optimal levels are not met– communicates problems to the crew via yellow and red

alert alarms when problems occur

Cat McGhan


System Levels – ECLSS

• Note: Yellow levels are 10% in from Red levels

Cat McGhan, Syed Hasan

1023K,--

923K,25.3MPa

647K,22.1MPa

Supercritical Water Oxidation

328K,210kPa

----

289K,70kPa

Multifiltration

723K,200kPa

350K,80kPa

--50kPa

Vapor Phase CatalyticAmmonia Removal

----

723K,100kPa

----

Trace Contaminant Control811K700K366.5KSabatier

297K--291KElectrochemical DepolarizedConcentrator

Red,high

NominalRed, lowSystem


System Levels – Atmosphere• For internal working considerations only:

• For EVA considerations only:

• Psuit = 25.7kPa, Pmodule = 57.2kPa nominal• O2% = 100% - N2% - 1% CO2

Cat McGhan

----

18.62kPa,32.5%

Red, low

3.54kPa,3.5%

1.01kPa,1.7%

0.4kPa,0.69%

----

CO2

41.38kPa,72.28%

34.48kPa,60.2%

22.89kPa,40%

21.38kPa,37.3%

O2

Red, highYellow, highNominalYellow, lowGas

62.76kPa, 65%60.69kPa, 68%1.655.17kPa, 65%

52.41kPa, 63.5%Red, low Pmodule, high N2%

57.93kPa, 62%1.455.17kPa, 60.5%1.3

Yellow, low Pmodule, high N2%R-factor


Other Monitoring• Fire protection

– minimal because of the atmospheric composition– fire extinguishers are present to put out small fires

• Health monitoring– occurs on a daily basis– is part of the exercise routine for data collection

on the effects of a low-gravity environment onthe human physiology

Cat McGhan


Crew Activity• Crew Schedule Guidelines:

– The crew may adjust theschedule to that day’s science

– The following guidelinesshould be followed whenadjusting daily schedule (L2)

• Crew members must eat atleast once every 7 hours

• 2 hours delegated for exerciseand recreation each day

• 8 hours of sleep a day

• EVA rotation:– Two, two person teams will

alternate daily EVAs– Teams will alternate crew

members every weekSleep2200 - 2400

Downtime2100 - 2200Hygiene2000 - 2100Dinner1900 - 2000

Exercise & Recreation1700 - 1900

EVA/Tasks1400 - 1700Lunch1300 - 1400

EVA/Tasks0800 - 1300Daily Coordination0700 - 0800

Hygiene/Breakfast0600 - 0700Sleep0000 - 0600

Crewmember Task Daily schedule

Sample Crew Schedule:

Jessica Seidel


Food & Water• Level 1 requirements state that MORPHLAB must hold 180

days of food at a subsistence level in addition to the 90 daysnominal mission supply (L1)

• Subsistence level will contain mostly EVA food bars &nutritional supplements, estimated weight=100kg

Data based off ISS dailyconsumables

2230.62Food

2700.75Food preparation water

4141.15In food water

5761.6Drinking water

Single manned missionrequirement (kg)

kg/person-day

Jessica Seidel


Nutrition• NASA STD-3000 states crew member diet must meet FDA

nutritional requirement standards (L2)• Daily caloric intake based on weight in kg

– Men• 18-30 years: kcal/day = 1.7 (15.3*Weight + 679)• 30+ years: kcal/day = 1.7 (11.6*Weight + 879)• Range: 2791.9 - 3716.3 kcal/day

– Women• 18-30 years: kcal/day = 1.6 (14.7*Weight + 496)• 30+ years: kcal/day = 1.6 (8.7*Weight + 829)• Range: 1897.12 – 2244.8 kcal/day

• An additional 500 kcal/day should be added whenan EVA occurs

Jessica Seidel


Nutrition

100-200 µgChromium20 mgNiacin150 µgIodine400 µgFolate70 µgSelenium2 mgRiboflavin15 mgZinc1.5 mgThiamin4 mgFluoride2 µgVitamin B6

2.0-5.0 mgManganese2 µgVitamin B12

1.5-3.0 mgCopper100 mgVitamin C10 mgIron80 µgVitamin K3500 mgPotassium20 mgVitamin E1500-3500 mgSodium10 µgVitamin D350 mgMagnesium1000 µgVitamin A1000-1200 mgPhosphorus238-357 ml per MJ consumedFluid1000-1200 mgCalcium30-35% total energy consumedFat5 mgPantothenic Acid50-55% total energy consumedCarbohydrate100 µgBiotin12-15% total energy consumedProteinRequirementNutrientRequirementNutrient

ISS Daily Nutritional Recommendations

Jessica Seidel


Food Storage• Three levels of storage

– Frozen: includes most entrees, vegetables, baked foods, grains, anddesserts

– Refrigerated: fresh-treated fruits and vegetables, extended-shelf-liferefrigerated foods, dairy products

– Ambient: thermostabalized, irradiated, aseptic-fill, shelf-stablenatural-form foods, and rehydratable beverages

• 180 subsistence level kept on board at all times, split upevenly between modules– Diet will consists of EVA food bars and nutritional supplements

• Nominal manned mission food supply will be sent up viaLogistics Depot before each manned mission– Food will be stored in 2 week supply containers– Containers will have straps so it can be carried onboard during the

first EVA

Jessica Seidel


EMU Selection

Late 1970s/early 80s1970s1980s19881998Year developed

Bags & HungBagsHungHungBagsStoring

WaistWaistBackBackWaistEntryILC Dover

ILCDoverNASA AmesILC Dover

ILCDoverManufacturer

Soft w. hardupper torsoSoftHardHybridSoftSuit type

1.141.30.591.14 (at 4.3 psi)1.3R value

4.3 psi3.7psi8.3 psi

8.3 psi, testedat 4.3 psi3.75 psi

OperatingPressure

65 kg48 kg83.25 kg82.5 kg78.9 kgTotal Weight15 kg26 kgnever fitted15 kg15 kgWeight- PLSS50 kg22 kg83.25 kg67.5 kg63.9 kgWeight- Suit

Shuttle EMUA7LBAX5MK IIII-Suit

Jessica Seidel


EVA Support• I- Suit

– Manufactured by ILC- Dover– Soft bodies suit increases mobility and

decreases weight– Waist entry– 8 suits for redundancy (I4)

• 4 come up with crew, fitted specifically for them• 4 kept on board, general sizes with

interchangeable section– 80 kg weight with Portable Life Support System

(PLSS)– 8 hours of nominal breathing, 45 minutes of

emergency life support– 25.7 kPa, 100% O2 atmosphere– Zero pre-breathe (L9)

• R value of 1.3 http://strangeblue.iwarp.comspacesuits/ILC.html

Jessica Seidel


Emergency EVA Access• Level 1, L8: System shall provide emergency

access and EVA bailout alternatives• Airlock connections between modules will serve as

pressure doors in case of decompression and also asemergency bailout hatches

• Procedure for emergency egress:– Put on EVA suit– Make sure all doors are closed and sealed– Disengage umbilical walkway– Open hatch door to outside– Deploy flexible ‘rope’ ladder from floor and egress

Cat McGhan


Airlock• 101 kg total mass

– 50 kg structure– 22 kg inner door

(pressure hatch)– 10 kg ladder– 14 kg pump– 5 kg tubing

• Accommodates one person in I-suit– 0.8 m x 0.8 m x 2.0 m high

• Depressurized and Re-pressurized– Most of airlock air pumped back into module before

outer door is opened– Air shower from above when air returns to airlock

Kate Castell

Airlock

Habitable- Science: EVA (Top View)


Airlock• Pump similar to Dekker Vacuum

Technologies DuraVane RVD007L– Chosen based on: mass, ultimate vacuum,

speed to depressurize– 14 kg mass– 12 kPa ultimate vacuum– 0.2 m3/min pumping speed, depressurizes

airlock in 7 minutes

Kate Castell


Lunar Roving Vehicle• Level 1 requirements lead

us to selection of a rover– MORPHLAB shall have

daily EVAs by twoastronauts (L5), and that theymust be able to travel withina radius of 10 km from thebase (M4)

• A 10 km-radius circle is toomuch area for astronauts toexplore on foot, so they need arover

– A rover will allow them to bring back soil and rocksamples, as in the Level 1 requirements (M3)– The rover can be used to transport the astronautsfrom the landing vehicle to MORPHLAB (M5)

Kate Castell


Lunar Roving Vehicle• Based on Apollo Lunar Roving Vehicle

– 208 kg mass– 3.1 m long, 1.8 m wide, 1.2 m high– Carries up to 490 kg payload including astronauts,

samples, and equipment– Folds up to 0.85 m3 for storage– Top speed 14 km/h– Can go over obstacles 30.5 cm high, crevasses 70 cm

wide– Travels on 25° slopes, parks on 35º slopes– Pitch and roll stability of ±45º

• Will have two modules to meet redundancyrequirement (I3 & I11)

Kate Castell


• Original Apollo LRV batteries not acceptable– Two 36 V silver-zinc batteries– 78 hour lifetime

• MORPHLAB LRV will have rechargeablebatteries– Two 36 V lithium ion– 112 W constant power– 11 kg mass– Rechargeable

• Plug into habitat module (currently drawing least power)• Budgeted 200 Watts per hour to recharge in 8 hours

• Batteries aren’t sufficient for 3 month traverse, sothe rovers are transported on the power modules

Lunar Roving Vehicle

Kate Castell

Lunar Rover 1

Lunar Rover 2

Power Module


Lunar Roving Vehicle• Payload Breakdown

– 2 astronauts, 95 percentile American males:197 kg

– 2 Mark III Suits & PLSS: 148 kg– Lunar samples and soil instruments: 95 kg– Communication/control/camera equipment: 10

kg Total: 450 kg• Speed is useful

– 14 km/h for an 8-hour EVA gives 112 km oftravel total

– Can return from 10 km in 43 minutes if needed

Kate Castell


Getting to the Moon


Getting to the Moon Topics• ΔV Calculations• Launch Options• Multi-Stage System• RCS• Landing Autonomy• Landing Zone• Landing Legs• Chassis Transit


ΔV Requirements

Russell Parker

.01 km/sLLO to LD3

.84 km/sLTO to LLO2

3.14 km/sLEO to LTO1ΔVBurn DescriptionBurn # LEO = Low Earth Orbit

(alt =185km, 28.5º)LTO = Lunar Transfer Orbit (6560km X 386200km)LLO = Low Lunar Orbit (alt = 50km)LD = Lunar Descent (50km x 10km)

Burn 1

Burn 2Burn 3

*NOTE: Not to Scale


ΔV Requirements

.26 km/sLanding5

1.88km/sBraking andApproach4

ΔVBurnDescriptionBurn #

Total ΔV = 6.13 km/s

Burn 4 Burn 5

*NOTE: Not to Scale

Russell Parker


Launch Vehicle Options• Level 1 Requirement:

All components for MORPHLABsystem installation andoperation shall be designed forlaunch on the Atlas V, Delta IVHeavy

• LEO Mass: Delta IV vs. Atlas V– Delta IV-H ~ 25,800 kg– Atlas V-551 ~ 20,000 kg

• Delta IV-H 2nd Stage to LEO vs.LTO.– Mass after LTO insertion burn:

• Delta IV to LEO 8350kg• Delta IV to LTO 12980kg

• One Stage vs. TwoStages– Assuming:

– Payload (LandedUsable Mass)

• One Stage 4200 kg• Two Stage 3800 kg

Russell Parker

.26.073502nd

5.87.014651st

ΔVInert MassFractionIspStage


Two Stage System• Despite lower mass still going

with a two stage system.• Benefits:

– Allows us to takeadvantage of a high Ispengine for majority of burns

– Easily dispose of largepropellant tanks and mainengine

– Decrease ground clearance(smaller landing legs)

– Landing stage can bemoved out of the way byPower Module dexterousarm. (largest mass = 45kg)

Lunar Landing Stage

(Second Stage)

with Habitable module

ΔV = .26 km/s

Lunar Transfer andApproach Stage

(First Stage)

ΔV = 5.87 km/s

Russell Parker


Lunar Transfer and Approach Stage

Modified Pratt andWhitney RL-10B-2

• Isp: 465s• Mass : 280 kg• Thrust can be throttled

Titanium Thrust Structure

• Support loads of:• 6g’s vertical• 2.5g’s lateral

• Low Thermal Conductivityto decrease boil off

• Mass = 920kg

LOX Tank• Propellant Mass: 15960 kg• Tank Mass: 560 kg• Insulation Mass: 50 kg• Variable Density Multi-Layer Insulation to reduceboil off from radiation

LH2 Tank• Propellant Mass: 2720 kg• Tank Mass: 550 kg• Insulation Mass: 200 kg• VD-MLI to reduce boiloff from radiation

Russell Parker


Structural Analysis• Launch loads:

– 6 g’s vertical– 2.5 g’s lateral

• Material comparison– Al ~ 1340kg– Titanium ~ 920 kg

Euler Buckling

Euler Buckling

Bending

Compression

Local Buckling

Failure Mode

3.40

3.38

.30

2.42

4.31

.198

.198

.250

.100

.346

LengthI.D.M.S.

Dimension (m)Truss Member

O.D.

.10.105LOX

2.7.300LOX Connect

.25.200Module

.009.200LH2

.013.350Engine

450MPa

50MPa

Module

LH2

LOXConnect

LOX

Engine

Ti FEM Analysis

Russell Parker


Thermal Analysis

Q12 Q23 Q34 270 KModule

4 KLH2 Tank

80 KLOX Tank

270 KEngine

AssumedTemperatureComponent

2.10W27.5 m2LH2SunQradiation

3.24W.007 m2LH2ModuleQ34

4.12W14.0 m2LOXSunQradiation

.13 W.007 m2LH2LOXQ23

16.2 W.031 m2LOXEngineQ12

QAreaSinkSourceFlow

Qradiation

•Given:•KTi = 6.7 W/m-K•LO2 = 213 kJ/kg•LH2 = 446 kJ/kg

•Boil Off•LOX = 29.7kg (.19%)•LH2 = 5.7kg (.21%)

Russell Parker


Descent Engine• Descent Engine’s Responsibility

– Hovering• Modules hovers for 60 seconds• During hovering the lunar surface will be surveyed

– Landing• The engine used to carry the module within the designated landing

zone• The engines are throttled to low thrust for soft landing on landing

legs onto the lunar surface– Descent Engines Considered

Thrust(N) Isp(sec) length(m) diameter(m) gimbaled(deg)throttled mass(kg)

S3K (DASA/EADS) 3500 352 1.03 0.53 no 14.5

OMV Variable Thrust Engine (TRW) 578 308 0.7005 0.2713 10-100 6.8

RS-28 (Rocketdyne) 2670 220 ? ? ?no 12.7

RS-21 (Rocketdyne) 1330 294 ? ? ?no 8.4

Tia Burley


Landing Stage

Modified

EADS S3K (6)• Isp: 350s• Mass : 14.5 kg• Full Throttle : 3500 N• Thrust can be throttled

MMH Tank(2)• Propellant Mass : 95 kg/tank

• Tank Mass : 45 kg/tank

MON3 Tank(2)• Propellant Mass : 72 kg/tank

• Tank Mass : 22.5 kg/tank

He Pressure Tank• Not shown on model.

• He and Tank Mass < 5 kg

Tia Burley & Russell Parker

Support Structure• Still being determined.

• Conservative mass estimateof 150 kg.


Landing Engine Schematic

MMH

Tank

MON3

Tank

helium isolation

valve

pressure relief

valve

fill/vent

S3K Engine

check valve

supply

line

pressurized

helium tank Pressure Fed System

Tia Burley


Reaction Control• Reaction Control System’s (RCS)

Responsibility– Attitude control during orbital maneuvering– Small translation maneuvers– Assist descent engine with landing maneuvers

• Reaction Control Thrusters ConsideredRocket Thrusters Thrust(N) Isp(sec) length(m) diameter(m) throttled % mass(kg)

CHT 20 (EADS) 20 230 0.195 0.033 40-100 0.395

S22 (EADS) 22 290 0.212 0.055 ? 0.65

R-1E (Marquardt) 110 280 0.279 0.15 60-140 3.7

MR-107 (Rocket Research) 178 236 0.218 0.066 28-140 0.885

Tia Burley


Reaction Control System• Assume RCS only necessary to

prevent rotational drift duringtransit.

• RCS uses Kaiser MarquardtR-1E (68N) located on thruststructure at LOX connection andon Module

• Total burn time + 50% Margin =192s

After 1st Burn, approximatevalues

•Ixx = 1.62 x 105 kgm2

•Iyy = 1.62 x 105 kgm2

•Izz = 4.2 x 104 kgm2

•c.g. = 8.3 m (z-direction)

ZY

X

+/- 5°

+/- 5°

+/- 5°

AllowableDrift

.01 °/s

.01 °/s

.01 °/s

RotationalVelocity

63s268N-mZ

33s532N-mY

33s532 N-mX

DutyCycleMax TAxis



Reaction Control System• Kaiser Marquardt R-1E

– Variable Thrust (68N – 158N)– MMH/N2O4

– Isp = 280s– Mass Flow Rate

• MMH = .0354 kg/s• N2O4 = .0256 kg/s

– Mass = 3.7 kg• Mass Breakdown


Cluster of 4

Cluster of 3

Location ofModule cluster of 4

151 kgTOTAL20 kg*Structures2 kgTanks

118 kgEngines

11 kgPropellant

MassComponent

*Structural Mass still being determined. 20kg is an approximation


Launch System Mass Summary

20 kg150 kg920 kgStructure

2 kg130 kg1350 kgTanks andInsulation

N2O4/MMH11 kg

MON3/MMH330 kg

LOX/LH218675 kg

Propellant

R-1E (32)3.7 kg each

S3K (6)14.5 kg each

RL-10B-2280 kg

Engine

ReactionControlSystem

LandingStage

LaunchTransfer and

Approach Stage

Resultant available landed mass: 3700 kg



Landing Requirements• Level 1 Requirements:

– Each MORPHLAB module shall be capable ofsuccessful landing and operations with any or all ofthe following conditions occurring simultaneously:

- 10° slope, any direction- 0.5 m diameter boulder anywhere in landing footprint- residual vertical velocity 1 m/sec- residual horizontal velocity 0.5 m/sec

– Landing systems onboard MORPHLAB modulesshall be capable of operating in on-boardautonomous control

• Derived Requirements:– Upon landing chassis, habitable module must

connect to power module within 24 hours

John Albritton


Landing Determination• Prior to descent

– Mapping satellite will provide preliminary terrainanalysis of 1.4m resolution to determine ideal landingzone of .10km

– Ideal landing zone will be relatively flat, and have fewterrestrial hazards, cliffs or craters which can disturbcommunications

– Trajectory corrections will keep us in the target footprintof 100m with terrestrial hazards greater than 1.4m known

• Autonomous Descent– Power Modules:

• LIDAR will distinguish objects down to .5m , determine slopesof >10 deg, determine any zone considered unsafe

– Habitable and chassis modules:• Power modules guide to safe landing

John Albritton


Landing Order and SensorsPower Module Chassis Modules

Landing Order

Habitable Module

John Albritton

Power Module Landing Sensors Chassis & Habitable Sensors

1,2,3,4 5,6,7,8,9,10 11, 12,13,14,15,16


Landing Footprint• Upon touchdown, LIDARrotates forward to be used asobstacle avoidance

• Power modules reposition100m apart in a diamondformation

• The diamond area, .01km2,will be scanned using LIDARto construct a safe landing zone

• If unsafe landing zone, powermodules will reposition untilsafe area found

“Safe Zone”.01 km2

John Albritton


Chassis Landing• Power module determines

chassis’s relative positionand descent velocity throughtriangulation

• Power modules continuouslyupdate and communicate thedata to the chassis moduleproviding real timenavigation and guidance

• Chassis module is guidedand landed within the “SafeZone”

John Albritton


Chassis Recovery• After successful landing of

chassis, power module retrievesthe chassis

• In the retrieval process the powermodule removes the landingengines first and then connects tochassis for power up.

• The power module alongwith connected chassis willreturn to the diamondformation waiting for nextchassis’s descent

• This retrievalprocess will occurfor all 6 chassis

John Albritton

• Chassis andpower connectionmust occur within24 hours


Habitable Module Landing

• The habitable and sciencemodules will land using the sameexact process as the chassis.

John Albritton

• A power module-chassis moduleassembly will be sent to retrieve thehabitable module

• After All 6 chassis are successfullylanded and connected, the habitableand science modules follow


Chassis Recovery

John Albritton

• The powermodule will firstremove thedescent engines

• The chassis will thenbe commanded todrive under thehabitable module andestablish a permanentconnection

• This connection alsois required to takeplace within 24 hours

• This processwill repeat untilall habitablemodules haveestablishpermanentconnections withthe 6 chassis


Landing Summary• Power modules land using full sensor suite• Chassis and habitable modules landing

without LIDAR utilize navigation andguidance provided by the power modules

• Chassis and habitable modules must connectwith power modules within 24 hours

• All retrieval processes are guided andcommanded by the power modules

John Albritton


Landing Zone• Landing zone is based upon

Apollo actual off target data• The data was then filtered• A standard deviation and

average value was then found• Then a technology

advancement reduction taken(30% off Apollo numbers)

• A new standard deviation andaverage were then calculated

Apollo 11 6.86 km(filtered) Apollo 15

.55 km(filtered)

Apollo 16.21 km

Apollo 17.20 km

Apollo 12.16 km

Apollo 14.05 km

.155 km average

Apollo Distances from Target Landing Site

Standard deviation.155 kmAverage.068 km

New Standard deviation.0406 km

New Average.0204 km

Rob Winter


Landing Zone• The new Standard deviation

and average were normalizedand then applied to a normaldistribution curve to find a.9990 probability of success

• This value then applied to theaverage gives us theprobability of .9990 of beingwithin a circle of that radiusof the target site

• So, this leaves us with alanding zone with a circularradius of .109 km for the firstmodule (power) to land in

• Other modules will land closeto the first one, due to landingbeacons being placed afterpower module landing,allowing them to land withinmeters of the power module

This .9990 correspondsto .109 km

Normal distribution curve

Rob Winter


Landing Stability Analysis

• Assumptions– 2-D motion & rigid body dynamics– Collisions are perfectly inelastic– Habitable module modeled as a cylinder and

power module modeled as a box2mv2

1

• Level 1 Requirement– MORPHLAB modules must be able to land on a 10° slope with

residual 1.0 m/s vertical and 0.5 m/s lateral velocities

• Analysis– Energy req. for tipping =

potential energy of centerof gravity – kineticenergy of vehicle

Abraham Daiub


Landing Stability AnalysisCase A Case B

Case C

Abraham Daiub

No-failure of power or habitable module

7.30m = footprint of habitable module2.28m = footprint of power module

No-failure

7.10 = footprint of habitable module2.22 = footprint of power module

Habitable module footprint = 7.60m → stablePower module wheel base = 3.3m wide 4.2m long → stable


Design of Landing Legs forHabitable Module

Constraint

Accelerationvector

1/16*landed mass

Abraham Daiub

• Level 1 Requirement– MORPHLAB modules must

be able to land on 10° slopewith residual 1.0 m/s verticaland 0.5 m/s lateral velocities

• Assumptions– 63 m/s2 axial , 1.63m/s2 lateral– Loads through of CG carried by

stringers– Footpad constrained 1.

FEM Setup


Design of Landing Legs forHabitable Module

Bending

LocalBuckling

LocalBuckling

EulerBuckling

FailureMode

874.2.250.30.50Foot-pad

2900.1960.722.100.099SecondaryStrut

5260.0522.67.120.118PrimaryLowerStrut

3880.3791.45.130.128PrimaryUpperStrut

Max.Stress(MPa)

S.M.Length(m)

OD (m)ID (m)MemberPrimaryUpper Strut

Primary Lower Strut

Secondary Strut

Foot-pad

Abraham Daiub


Design of Landing Legs forHabitable Module• Requirement

– Absorb all kinetic energy withone load limiter

– Residual velocity 1.0 m/s axialand 0.5 m/s lateral

0.342Mass (kg)0.0129Cross section (m2)

1.40Length (m)

http://www.mcgillcorp.com/alcore/datasheets/spiral-0600.pdf

Parameters

Abraham Daiub


Chassis During Transit to Moon• Wheels retracted inward toward center of

thrust structure to allow chassis to fit withindynamic envelop of Delta IV-H 5m payloadfairing.

• Launch loads transferred two four points atends of 2 chassis ‘launch beams.’

• Maximum load acted on chassis during transitand launch occur during Delta IV – H mainengine cutoff. Axial acc. = 6g’s, later acc =2.5g’s. Chassis capable of supporting 3500 kg(including itself) during transit and launch.

Jason Seabrease


Chassis LandingMaximum landed mass = 3500kg

Maximum impact deceleration designed for= 6 lunar g’sMax deceleration is a very conservative estimatederived from our requirement that the maximumresidual velocity during landing cannot exceed1m/s and assuming that the deceleration duringimpact will not occur in less that 1/10th of a second.

Worst cases for struts and axles correspond to a power module initiallytouching down with only one wheel.

Landing loads can act in all three directions and create greater stresses in eachmember for different loading directions.

All cases closely examined and modeled using ProEngineer.

All structural member pass all worst case landing load conditions withpositive margins of safety.

Jason Seabrease


Mission Planning


Fault Tolerance

• Fail Safe Mode losses:– 1 Habitable-Science module– 1 Habitable-Living module– 2 Power modules

• Survival Mode losses:– 2 Habitable-Living modules– 2 Habitable-Science modules– 3 Power modules

Daniel Senai

• Level 1 Requirements:– Loss of 1 module should not end mission– Loss of 2 modules should support crew until launch

window occurs


Fault Tolerance/Reliability• NASA JSC - 28354 requires reliability of 0.99 for crew return• All sub-systems will have to meet these reliability requirements by

Jan. 1st 2010 with a NASA technology readiness level (TRL) of 6

Daniel Senai

= Fail Safe Mode

R-Power

R-Power

R-PowerR-Living

R-Living

R-Living

R-ScienceR-Power

R-Power =0.989 R-Living = 0.930 R-Science = 0.930

R-Power Sys = 0.99988 R-Living Sys = 0.995 R-Science Sys = 0.995

R-Science

R-Science


Required Sub-System Reliabilities

Daniel Senai

0.9978Thermal0.9978Avionics0.9955DIPS0.9978Chassis Module

Power Module = 0.989

0.9819Docking Mechanism0.9819ECLSS0.9819Thermal0.9819Avionics0.9819Power Sub-systems

Habitable-Living andHabitat-Science Modules =0.930

0.9966Avionics0.9966Power Sub-systems0.9966Drive System0.9966Steering System0.9966Suspension System0.9966Brake System

Chassis Module = 0.9978


Timeline

• Historical Precedent:– Apollo:

• 5 years from program start to first test• 2 years from testing to first human launch• 7 years total time from program start to human launch

– Shuttle:• 5 years from program start to first test• 4 years from testing to first human launch• 9 years total time from program start to human launch

2005 2015

Design/Manufacture

2010

TestSystems

LaunchModules

Manned Missions(every 6 months)

2020

Craig Nickel


Timeline• Launching Phase (2013 – 2014):

– All landings are scheduled forlocal lunar dawn + 3 days

– Launch #18 - #24 are launchwindows built into the timelinein case a launch must bedelayed and so that the mannedmissions will begin on schedule

– LVs can be ready in 22 days,each launch is scheduled 28-30days apart

– Mapping satellites to belaunched January 2011 to get24 months of data and analysisabout the lunar terrain

– Communications satellites willbe launched February 2017

Craig Nickel


Timeline• 17 required launches before first human

mission– Launches #1 - #4 will be the power modules– Launches #5 - #10 will be the chassis modules– Launches #11 - #16 will be the habitable modules– Launch #17 will be the first logistics depot

• Each logistics depot will be launched one monthbefore the next human arrival

Craig Nickel


Timeline• Mission Phase (2015 – 2019+):

– Each human mission will be 6 months apart startingJune 5, 2015

• Launch will be June 1, 2015• Level 1 Requirement: I10

– System shall provide autonomous checkout and unambiguousevidence of functionality prior to committing the next human

• The first 3 month mission will be unmanned for the firstmonth and land the first crew at local dawn on the second

– The human mission will be 2 months in order to stay onschedule

– Approximately 14 days will be lost of program timewhen the missions change over to far side

Craig Nickel


Timeline

Craig Nickel


Timeline• In the event that a mission must be delayed

the column labeled “Land Day before/afterFull Moon” will be used to get the newlanding date

• This factor will land humans at local lunardawn + 2 days

• All full moons from 2015 – 2020 are listed inthe detailed timeline document

Craig Nickel


Cost Analysis• Level I Requirements: $51.5 Billion budget

– Spread over 15 years– Derived from projected NASA spending on a lunar

exploration program

• Mass-based, vehicle-level cost estimatingrelationship– Non-recurring costs– Production costs with 80% learning curve– Constants differ for launch vehicle stage, liquid rocket

engine, scientific instrument, unmanned planetary,manned spacecraft

Emily Tai


Production CostsVehicle Assembly Launch Structure

Satellites

9.67Total1.999

Logistics Modules(2-10)

1.191Logistics Module 11.266Chassis Module1.913Habitable-Science1.993Habitable-Living1.334Power Module

Cost($B04)QtyComponent

1.91Total0.15952Small Engine0.78026Engine0.97326Thrust Structure

Cost($B04)QtyComponent

182Total1154Communications671Mapping

Cost($M04)QtyComponent

Emily Tai


Launch Costs

4.54Total0.122Atlas V1Mapping Satellite0.122Atlas V1Communication Satellites1.49Delta IV-Heavy9Logistics Modules 2-100.165Delta IV-Heavy1Logistics Module 10.990Delta IV-Heavy6Chassis Module

0.495Delta IV-Heavy3Habitable-Science Module0.495Delta IV-Heavy3Habitable-Living Module0.660Delta IV-Heavy4Power Module

Cost($B04)LV

Number LaunchesComponent

Emily Tai


Wrap Factors

Emily Tai

7.54Total1.180.10Fees1.180.10Program Management and Integration0.7770.066Launch and Landing1.770.15Operations Capability Development1.770.15Program Support0.4120.035Phase B Definition Studies

0.03530.003Phase A Conceptual Studies0.4120.035Advanced Development

Cost($B04)PercentageFactor


Cost Summary

Allowable reserve factor: 0.537Total budget with 0.4 reserve factor: $39.7 Billion

Emily Tai

23.8Total System Cost7.52Wrap factors4.53Launches0.182Satellite production1.91Launch structure production9.67Vehicle production

Cost ($B04)


Cost Summary

Emily Tai

Cost-Budget Comparison

0

500

1000

1500

2000

2500

3000

3500

4000

4500

5000

5500

Year/Phase

Do

llars

($M

2004)

Budget

Cost

2005

2006

2007

2008

2009

2010

2011

2012

2013

2014

2015

2016

2017

2018

2019

Development - Testing/Mapping Launch - Module Launches - Mission/Comm Launch


Summary• Designed to fulfill the goals of the

President’s new space initiative• Stepping stone for an Earth to Mars mission• Design utilizes existing launch vehicle

technologies• Design has great advantages over a large-

scale stationary lunar base• Inexpensive program


Open ItemsThere are a few subsystems that either have yet

to be determined or need the completion ofdetailed analysis, these systems are:– Radiator for the power module– Extension for docking mechanism– Specific computing requirements– Interior lighting layout– Steering and suspension mechanism– Lunar rover storage


Acknowledgments• Our advisors, Dr. Dave Akin and Dr. Mary

Bowden, for their help and support

• All of the people who helped by giving usfeedback at our Trade Study, and PreliminaryDesign Reviews

• Dr. Fourney and the University of Maryland’sAerospace Engineering Department

Documents

Modular Roving Planetary Habitat, Laboratory, and Basespacecraft.ssl.umd.edu/academics/484.archives/2004.MORPHLAB.CD… · Modular Roving Planetary Habitat, Laboratory, and Base Critical