spritesprite 1 of 256 university of maryland enae484 pdr march 14, 2005 preliminary design review...

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University of MarylandENAE484 PDR March 14, 2005

Preliminary Design Review

March 14, 2005

Small Pressurized Rover for Independent Transport and

Exploration

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What is SPRITE?

• SPRITE is a pressurized rover designed primarily for use on the moon. It can be used, with only minor changes, on the Martian surface.

• It would serve as the primary exploration vehicle for astronauts living at a lunar base.

• It accommodates two astronauts for a week-long scientific expedition

Charles Bacon

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Why SPRITE?

• With the new exploration initiative being undertaken by NASA for human presence on the Moon and Mars, there must be a way for humans to traverse long distances from the base.

• This is primarily because ideal sites for landing and base construction (flat, open terrain) are not the same as those most interesting for scientific exploration (geologically diverse regions).

Charles Bacon

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Why SPRITE?

• Though many pressurized rovers have been suggested, none have been fully developed mainly because of cost.

• To constrain this problem, SPRITE will be launched on a single Delta IV Heavy vehicle, including all systems needed for nominal and emergency use. The only thing not to be included on the launch will be consumables required. They will be provided by the lunar base.

Charles Bacon

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Launch to Landing

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CONOPS Overview(Delta-IV Heavy Separation to Landing)

Separate from Delta-IV Heavy

Perform lunar orbit insertion burn

Perform descent orbit insertion burn

Perform powered descent burn

Land on the Moon

Chris Hartsough

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Orbit Design Objectives

• Requirements– Accurately land anywhere on the Moon

• Powered descent for soft landing– Launch on a Delta IV Heavy

• Initially in a 185 km altitude LEO

• Optimization Parameters– Flight Time– Mission ∆V (proportional to landed mass)

Daren McCulley

Requirement M7: The SPRITE vehicle shall be capable of independent deployment to the lunar surface with a single Delta IV Heavy class launch

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Translunar Orbit Options• Low Thrust

– Advantage• Unmanned and mass constrained mission

– Disadvantages• Insufficient maximum thrust

– Flight times grossly exceeding reasonable limits• Requires two propulsion system reliability• Payload fairing constraints• High power requirements

– Latest advances require 7-20 kW for .5-1 N of thrust• Electromagnetic interference

• Delta IV Second Stage TLI– Advantages

• Flight time between 4.5 and 5.5 days• Presumably will be flight tested by 2016

– Disadvantage• Highly inefficient ratio between propellant and payload mass

– Over 50% of the mass in LEO is consumed during TLI

Daren McCulley

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Depiction of Translunar Orbits

http://sbir.gsfc.nasa.gov/SBIR/successes/ss/5-075text.html

Low Thrust Transit High Thrust TLI

Daren McCulley

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Apogee of Translunar OrbitRadius of Apogee TLI ∆V

356,000 km

Moon at Perigee

3.128 km/s

407,000 km

Moon at Apogee

3.140 km/s

* Additional V of only 12 m/s* Additional day of flight time

Daren McCulley

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Payload vs. ApogeeSample Delta IV Performance Curve

0

1000

2000

3000

4000

5000

6000

0 20 40 60 80 100

Altitude of Apogee [x 1e3 km]

Sep

arat

ion

Mas

s [k

g]

Delta-IV Payload Planners Guide

Daren McCulley

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Selenocentric Orbits

Options Total ∆V

Direct Descent 2.79 km/s

L1 Layover 3.10 km/s

Elliptical Lunar Orbit Insertion 2.83 km/s

Circular Lunar Orbit Insertion 2.85 km/s

Larson, Wiley J. and Pranke, Linda K ETD. Human Space Flight, Mission Analysis and Design

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Considered Approaches • Direct Descent

– Engine failure results in lunar impact (risk to base)– Lower landing accuracy– Limited landing site access

• L1 Layover– Nullified by ability to perform accurate trajectory analysis– Increased complexity

• Elliptical Lunar Orbit Insertion– Risk to spacecraft– Negligible ∆V savings

Daren McCulley

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Circular Orbit Insertion

• Safe Orbital Altitude (100 km)

• Constant Orbital Velocity– Congruent ∆V requirements for descent orbit insertion

• Control over argument of periselenium

• Standard Lunar Insertion/Descent Profile– Learning curve

Daren McCulley

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Descent Orbit AnalysisAltitude of

Periselenium

(km)

DOI ∆V

(km/s)

Tangent Velocity

(km/s)

Normal Velocity

(km/s)

Total ∆V

(km/s)

10 .0206 1.696 .1808 1.897

20 .0183 1.688 .2557 1.962

30 .0159 1.681 .3132 2.010

40 .0136 1.674 .3617 2.050

50 .0113 1.667 .4044 2.083

Daren McCulley

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Insertion and Landing Concept

• Lunar Orbit Insertion (LOI)– Retro burn at closest point of approach– 100 km altitude circular orbit

• Descent Orbit Insertion (DOI)– Retro burn at descent orbit aposelenium– 15 km periselenium above landing site

• Powered Descent Landing (PDL)– Retro burn near periselenium– Continue controlled burn to soft landing

Daren McCulley

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Depiction of Selenocentric Orbits

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3D Orbit Design In Reverse

Daren McCulley

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Gravity Assist

• Prevent the spacecraft from leaving Earth orbit in the event the retro engine fails to fire.

• Unmanned mission, makes this a low level requirement.

Chobotov, Vladimir. Orbital Mechanics

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Dynamic Simulations

• Translunar Injection Simulation– Controllable variables (Time of TLI, ) – Out of plane bending

• Perifocal Lunar Orbit Transfer Simulation– Nonimpulsive analysis of orbit transfers

• Powered Descent Simulation– Sets requirements on propulsion system– Ideal estimate of landed mass

Daren McCulley

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Analysis of Control Variables

Daren McCulley

all axes in km Daren McCulley

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Powered Descent Simulation

Daren McCulley

km

km

km

km

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PDL Simulation Results

Daren McCulley

Time after Aposelenium (s)

Vel

ocity

(km

/s)

Burn Altitude: 16.2 km

Burn Time: 283.5 s

Thrust: 42.9 kN

Residual Velocities:Negligible

Height: 4 m

Landed Mass: 5435 kg

Max Acc: 6.3 m/s2

∆V: 1.83 km/s

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Burn ProfileBurn / Maneuver Engine ∆V (km/s) ∆M (kg)

PL Fairing Evasion RCS Negligible Negligible

Delta IV SS TLI RL-10B-2 3.14 N/A

Midcourse Correction TBD 0.01 50

Lunar Orbit Insertion RETRO .816 1670

Circular Orbit Correction RCS Negligible Negligible

Descent Orbit Insertion RETRO 0.02 40

Descent Orbit Correction RCS Negligible Negligible

Powered Descent RETRO 1.89 2860

Daren McCulley

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Guidance Navigation & Control

• Derived Requirements– The GNC system shall provide:

• state vector estimations• attitude determination• attitude control systems• landing control systems• landing point localization

Aaron Shabazz

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• Critical GNC Hardware– Inertial Measurement Units (IMU)

• Senses pitch, yaw, roll & acceleration rates– Star Trackers

• Detects star patterns & magnitudes• Precisely aligns IMUs

– Guidance Computers (GC)• Uses IMU data to:

– Compute state vector estimation– Compute attitude estimation

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• IMU accuracy is vital to mission success– IMU drift bias is 0.0003 deg/hr *– Star trackers are re-aligned to compensate for IMU drift bias

• Star tracker to be re-aligned within 1.4 deg error• Star trackers require calibration after about 4667 hours

– IMU Reliability is > 0.996 *• Use 2 IMUs on spacecraft and rover• Probability that at least 1 IMU works > 0.9999

Aaron Shabazz

* Data from Honeywell IMU spec sheetSpec Sheet - http://content.honeywell.com/dses/assets/datasheets/fog.pdf

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Guidance Navigation & Control• Attitude Control System

– Pre-loaded trajectory/attitude data in guidance computer (GC)– IMUs provide actual estimate of attitude– GC uses residual of nominal and actual attitude data to:

• Run data through filter for best data• Convert error data to steering & thrust commands• Desired attitude is achieved

Aaron Shabazz

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Center of Gravity - Landing

• Center of gravity determined by worst-case dynamic conditions on landing

• The “tripping scenario” is the most difficult scenario to maintain stability upon landing

Mike Sloan

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• Using a rigid landing structure, the critical limit for CG height is 3.4 m

• The safety limit is 1.1 m• This height is achievable if the rover is placed

horizontally on the landing structure

Mike Sloan

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• For a landing-on-wheels scenario, the CG tolerances are much tighter

• Primary danger comes from descent engines hitting the surface

• Critical limit for CG height is 1.9 m• Safety limit is 0.1 m

Mike Sloan

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Center of Gravity - Driving

• Center of Gravity determines the vehicle’s propensity to roll over while driving

• Lunar required CG height - 1.1 m• Martian required CG height - 2.5m

Mike Sloan

Requirement I9: SPRITE shall be able to actively traverse terrain safely with 20o cross-slope and 30o direct slope

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Center of Gravity - Driving

• Mars CGrequired height > Moon CGrequired height

• Any vehicle geometry that can safely drive on the Moon can safely drive on Mars

Mike Sloan

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Transit Configuration 1

Mike SloanDaren McCulley

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Propulsion System Requirements

• Launch a specified payload to the moon• Expend practically all its fuel upon arrival• Landing engine must be able to restart 2 or 3

times• The total mass of the propulsion system must be

as low as possible• Maximum thrust of the landing engine must be

45 kN

Reuel Smith

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Assumptions Made

• Changes in Velocity– Retro Engine

• LOI: 816 m/s

– LOI - Lunar Orbit Insertion• DOI: 20 m/s

– DOI - Descent Orbit Insertion• PDL (tangent): 1792 m/s• PDL (hover): 60.96 m/s

– PDL - Powered Descent Landing– RCS Thrusters

• RCS (landing): 150 m/s

– RCS - Reaction Control System

Reuel Smith

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Assumptions Made

• Propellants: The module runs on one specific fuel/oxidizer

mixture

• Other Assumptions– Payload: 3790 kg

– Ae/At: 54 for all propulsion stages

– Inert Mass Fraction: 0.08 for all propulsion stages– Max RCS Thrust: 445 N per thruster– RCS Thruster Count: 16

Reuel Smith

Spacecraft Apollo- <http://www.braeunig.us/space/specs/apollo.htm>

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Fuel Analysis

MLanding Engine

+ MRCS Thrusters

+ MGimbals

+ MAvionics

+ MWiring

+ MThrust Structure

__________________

MPropellant System

Propellant Isp vac

(s)Mixture Ratio

Total Mass (kg)

LOX/Kerosene 1.24 353 2.56 705

LOX/LH2 1.26 451 4 665

LOX/Hydrazine 1.25 365 0.9 698

LOX/RP-1 1.225 323 2.3 722

NTO/MMH 1.132 336 2.1 716

NTO/UDMH 1.235 315 1.75 727

Reuel Smith

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Fuel Analysis

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Propellants

To

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kg

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Reuel Smith

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Possible RCS Configurations

• RCS thrusters may be placed along the center of mass

• It may be possible to do a 12 thruster RCS by removing four roll thrusters

Reuel Smith

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RCS Thruster Risk Analysis

• Assumptions: 95% Mission reliability, no fault tolerance (crew survival not dependant on RCS)

• Two configurations considered: 12 engines and 16 engines

• Must be able to maintain complete 3-axis control of the landing vehicle

Jason West

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RCS Thruster Risk Analysis

• Scenario A: 12 engines, none fail• Scenario B: 16 engines, up to 2 engines can fail

Scenario A Scenario B

Required Engine Reliability

0.9957 0.9469

• 5% less required engine reliability for 16-engine system

Jason West

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Next Step

• Examine using two different sets of propellants for the RCS and Landing Engine

• Modify mixture ratio for NTO/UDMH to lower the propellant system’s mass

• Examine using monopropellants for RCS• In-Space Propulsion analysis

Reuel Smith

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Landing Requirements

Rahkiya Medley

Requirement I8: The SPRITE system shall be capable of successful landing and subsequent operations with any or all of the following conditions occurring

simultaneously at the point of touchdown: 10o slope in any direction, 0.5 m boulder anywhere in landing footprint, 1m/s residual vertical velocity, 0.5 m/s residual horizontal velocity

Requirement S1: All Systems shall be designed to provide a non-negative MOS for worst-case loading conditions incorporating Primary Structure: 2.0

Requirement S2: All structural systems shall provide positive MOS for all loading conditions

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Landing Structure

• Disposable • Absorb kinetic energy ~2 kJ• Slow landing package to minimize force

transferred to SPRITE• Worst case platform height is 3 m above surface

to accommodate fuel tanks and nozzle• Deployable ramps

Rahkiya Medley

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Lander Options

• Crushable legs–Honeycomb insert

–Pivot feet

–80 kg/leg (Al wrought 2024-T4 and SPIRALGRIDTM)

• Joint legs–Torsion spring joint

–Pivot feet

–TBD kg/leg

Rahkiya Medley

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Crushable Legs

• Modeled as a mass damping system• Impulse force ~81 kN • Increasing leg length increases landing footprint • As the leg length increases, critical buckling load

decreases Pcr α 1/L2

Rahkiya Medley

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Lander - Future Work

• Model of joint leg• Optimum placement of landing legs for both

configurations• Optimum crush strength of SPIRALGRIDTM

• Fuel tank/nozzle support structure

Rahkiya Medley

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Lunar Mapping

• The surface of the moon will be mapped by the 2008 Lunar Reconnaissance Orbiter

• Both optical and topographical maps will be taken• These maps can be used to assist in landing and surface

navigation– Optical resolution is 0.5 m per pixel

– Vertical (altimeter) resolution is 10 cm over a 5 m sample

Dr David Smith, Goddard Space Flight Center

Mike Sloan

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Guidance Navigation & Control• Landing Control System

– 3 Microwave Scan Beam Landing Systems (MSBLS)• Transponders/receivers that find slant range, azimuth, and elevation relative

to moon base• Gives very accurate position info to GC to compute state vector

– GC selects middle values of 3 ranges, azimuths and elevations• Angle and range data are used to compute steering commands

– 2 Radar Altimeters• Measures absolute altitude• Both measurements are averaged• Can derive vertical velocity and match with IMU measurements

– GC checks nominal and actual approach velocities to ensure safe & soft landing

Aaron Shabazz

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• Landing Point Localization– Assume Moon Base has 4 m

high antenna• LOS is about 3.73 km• A 3.73 km radius about the

moon base defines our desired landing zone

Aaron Shabazz

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• Landing Point Localization– Rough Estimate Landing Accuracy

• Average all off-target data after Apollo 12

Estimate landing accuracy = 0.234 km

Apollo 12 Apollo 14 Apollo 15 Apollo 16 Apollo 17

Off target data 0.16 km 0.05 km 0.21 km 0.20 km 0.55 km

Aaron Shabazz

Off target data – Spring 2004 ENAE 484 CDR Slide # 239

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• Distance Between Landing Target and Moon Base Roughly Twice the Estimated Landing Accuracy for Safety

• Even in Worst Case Scenario, Rover will have LOS Communication w/ Moon Base after Touchdown

Aaron Shabazz

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Landing Hazard Avoidance

• Landing requirements– Must be able to survive a 0.5 m boulder and a 10o slope

• Larger boulders and slopes must be detected and avoided– Digital elevation map (DEM) generation options

• Stereo camera system– 6 - 7 m error

• Stereo from lander motion (more reliable option)

Joanneum Research: Vision-Based Navigation for Moon Landing

Scott Walthour

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Stereo From Lander Motion

Scott Walthour

Scott Walthour

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Hazard Detection Hardware

• One CCD Camera

• 16 Mb memory for onboard processing

• DSP board–TBD

• Laser Altimeter–LaserOptronix ALTM400

–(2 - 400 m range, 10 - 20 cm accuracy)

Digital Elevation Map (image source: http://qso.lanl.gov)

<http://www.laseroptronix.com>

Scott Walthour

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Hazard Detection Performance*CCD Array 512 x 512 pixels

Focal Length 10 mm

Footprint (200 m) 100 m

Ground and DEM Resolution 0.2 m

Required Pointing Accuracy 1.4 deg

Processing Time ~ 10 to 30 sec

Required Inertial Sensing Accuracy (90% overlap)

10 m

*From similar lunar mission (2 hr orbital period, 0.5m obstacle requirement)

Joanneum Research: Vision-Based Navigation for Moon Landing

Scott Walthour

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Lander Stereo Considerations

• Hovering could cause errors in inertial navigation–Requires position recalibration

• Calibration from previous DEM

– Not likely without a DEM from orbit

• Self-calibration

– Errors not significant compared to DEM errors (at least 10 – 20 cm)

Scott Walthour Joanneum Research: Vision-Based Navigation for Moon Landing

Scott Walthour

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Landed Mass Analysis• Delta IV Heavy delivers 9950 kg into Lunar Transfer Orbit (LTO)

• Used Available Mass Estimating Relationships, Fuel Properties, ∆V Values, and Rocket Equation to determine rover’s mass when landed

• Rover Mass = Mass of Landed Package – Mass of Main Propulsion System (varies) – Mass of RCS (~ 250 kg) – Mass of Landing Equipment (~ 250 kg)

Timothy Wasserman

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Single Stage• Single main engine used for all

phases of flight• Standard Landing Structure

Propellant Combination Rover Mass (kg)

LOX/LH2 3000

N2O4/MMH 2710

N2O4/UDMH 2610

LOX/CH4 2530

LOX/RP-1 2360

Timothy Wasserman

Timothy Wasserman

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Two Stages: Land on Wheels

• 1st stage performs LOI and most of powered descent

• 2nd stage performs remaining 300 m/s of ΔV

• Two parallel outboard engines (each thrust ~ 6 kN)

• For:– Stage 1: LOX/LH2

– Stage 2: 2 x N2O4/MMH

Rover Mass = 2750 kgTimothy Wasserman

Timothy Wasserman

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Two Stages: Reuse Cryogenic Tanks

• Assumes SPRITE uses fuel cells• Assumes fuel cell reactant tanks (capacity ~

700 kg) can be used for storing 2nd stage propellants

1st Stage Prop 2nd Stage Prop Surface Mass (kg)

LOX/LH2 LOX/LH2 3030

N2O4/MMH LOX/LH2 2840

Timothy Wasserman

Timothy Wasserman

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Comparison of Best Two Staging Options

Option Rover Mass (kg)

LOX/LH2 Single Stage 3000

LOX/LH2 First StageLOX/LH2 Second Stage

(reuse cryotanks)3030

• While reusing the cryotanks yields the highest rover mass, the savings are small

• May introduce additional plumbing mass• Single Stage LOX/LH2 system is simpler/cheaper

to design, and delivers a high mass to the surface of the Moon

Timothy Wasserman

Akin, David. ENAE 483 Lecture on Mass Estimating Relationships

Fuel Properties from: www.astronautix.com

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Launch Mass Budget

Design Group MassTransit Power, Propulsion & Thermal 5800

Surface Power, Propulsion & Thermal 700

Loads, Structures & Mechanisms 1250

Crew Systems 700

Mission Planning & Analysis 300

Avionics 300

Timothy Wasserman

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Surface Operations

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Mission Planning RequirementsRequirement M2: SPRITE shall be designed to carry 2 crew on a normal sortie of 7 days

covering 250 km Requirement M3: SPRITE shall be capable of replicating the science from an Apollo J-class lunar EVA on each of the 5 EVA days of the sortie

Requirement M5: The SPRITE system shall include provision for safe return of the crew following a worst-case SPRITE failure without outside intervention

Requirement M7: The SPRITE vehicle shall be capable of independent deployment to the lunar surface with a single Delta IV Heavy class launch

Requirement I1: The SPRITE system shall be designed to operate on the lunar surface. No feature of the design shall preclude its adaptation for use on the Martian

surface

Requirement A2: Systems onboard SPRITE shall be capable of operating in any of the following control modes: manual, teleoperation, supervisory control, autonomous control

Chris Hartsough

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CONOPS Overview(Deployment to Nominal Operations)

Deploy from landing system

Autonomous return Remote operated return

Dock with base

Pre-mission check of systems

Supply SPRITE with consumables and fuel

Chris Hartsough

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CONOPS Overview(Nominal Mission)

• Day One– 100 km drive in 10 hr

• Day Two through Six– 10 km morning traverse in 1 hr– 8 hr EVA conducting TBD experiments

• Day Seven– Return 100 km to base in 10 hr

Chris Hartsough

Requirement M2: SPRITE shall be designed to carry 2 crew on a normal sortie of 7 days covering 250 km

*Possible robotic arm operations everyday

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Route Options

• Drive out 100 km• Drive in 8 km radius circle, with

stops every 10 km• Loop A

– Never more than 116 km from base

• Loop B– Never more than 100 km from base

• Both situations easier for emergency operations

Loop A Loop B

Daniel Zelman

Daniel Zelman

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Route Options

• Drive out 100 km• Drive along arc for 50 km• Return along different

100 km path• Arc

–Never more than 125 km from base

• Inverted Arc–Never more than 100 km from base

• More scientific possibilities than previous routes

Inverted ArcArc

Daniel Zelman

Daniel Zelman

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Base Services• Supplies and services from the base are required for

rover operation– Water, food, atmospheric consumables– Power generation– Power system reactants– Astronauts and Suits– Communications devices– Waste management capability– Maintenance tools

• The base must have certain aspects– SPRITE-compatible mating hatch– Airlock– 14.7 psi atmosphere

Mike SloanDaniel Zelman

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Structures

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Pressure Hull

• Sized to contain the astronauts, crew systems and avionics

• Designed to handle launch loads, pressure loads, and kick loads

• Two options considered: Prolate Spheroid and Cylinder with Ellipsoidal Endcaps

• Mass is the primary driving factor

Evan Ulrich

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Prolate Spheroid•Configuration

•Rib/Stiffener

•Optimal number of ribs is 4

•Need for external mounts may increase number of ribs

•Stringer

•8 allows for ease of hatch/window placement

•provides sufficient structural support

•All stringers have hollow circular cross sections

•Shear panel

•Stringer

•Rib/Stiffener

Evan Ulrich

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Prolate Spheroid: Analysis• Applied Loads:

• Internal Pressure (2 atm)Punching force (3 kN)

Method of Analysis:• Skin idealized shell, 4 mm thickness• Point constraint

• Applied Loads: – 6g axial, 2.5g lateral

– Internal Pressure (2 atm)

• Method of Analysis:– Skin, Rib, Stringer Approximated by ~ 1.2

million finite elements

Component O.D (m) I.D (m) Length (m) Mass (Kg) Design Load (Mpa) S.M S.F Material Failure modeStringer 0.084 0.083 4.8 24 380 0.0 2 Ti-6Al-4V CompressionRib/stiffener inner0.115 0.114 6.6 4 380 0.0 2 Ti-6Al-4V BendingRib/stiffener outer0.075 0.072 4.2 6 380 0.0 2 Ti-6Al-4V BendingSkin (4mm) 4.808 4.800 639 550 2 Ti-6Al-4V local bucklingTota Mass (Kg) 673

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Cylinder with Ellipsoidal Endcaps (CEE)

• Optimal number of ribs is 4

• Need for external mounts may increase number of ribs

• 8 stringer configuration allows for ease of hatch/window placement

• Provides sufficient structural support

• All stringers have hollow circular cross sections

-Shear panel

-Stringer

-Rib/Stiffener

2m

2 m

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CEE: Analysis1.2e+81.0e+86.2e+73.2e+72.0e+7

4.2e+83.6e+82.0e+81.1e+85.8e+7

Component Length (m) Mass (Kg) Design Load (Mpa) S.F Material Failure modeStringer 4.0 65 420 2 Titanium Ti-6Al-4V CompressionRib/stiffener inner 6.3 TBD 420 2 Titanium Ti-6Al-4V BendingRib/stiffener outer 6.3 TBD 420 2 Titanium Ti-6Al-4V BendingSkin (4mm) 851 120 2 Titanium Ti-6Al-4V local buckling

Total Mass 916

• Applied Loads:• Internal Pressure (2 atm)

Punching force (3 kN)

Method of Analysis:• Skin idealized shell, 4 mm thickness• Point constraint

• Applied Loads: – 6g axial, 2.5g lateral

– Internal Pressure (2 atm)

• Method of Analysis– Skin, Rib, Stringer Approximated by ~ 1.2 million finite elements

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Micrometeoroid Protection

• High velocity dust particles– Average velocity ~ 13 – 18 km/s– Average size ~ 10-8 – 10-2 g

• Inadequate protection can lead to catastrophic failure

• Probability analysis needed to design for sufficient protection

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• Calculate micrometeoroid flux–Surface area ~ 36 m2

–Mission duration ~ 10 days

–PNP ~ 0.996

• Flux = 0.00406 (impacts/m2/yr)

• Critical mass ~ 0.0002 g

Micrometeoroid Flux vs. Mass

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

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• Design variables–Hull properties

–MLI properties

–Hull/MLI spacing

Critical Micrometeoroid Mass vs Hull/MLI Spacing

0.0000

0.0001

0.0002

0.0003

0.0004

0.0005

0.0006

0 0.01 0.02 0.03 0.04 0.05

Spacing (m)

Ma

ss

(g

)6 mm hull thickness

5 mm hull thickness

4 mm hull thickness

3 mm hull thickness

2 mm hull thickness

Critical Design Mass

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Window Materials

Material

Density (kg/m3)

Elastic Modulus (GPa)

Flexural Strength (MPa)

Compressive Strength (MPa)

CTE (10-

6/°C)

High-Strength 2010 37.2 18.6 50 0.6

Ultra High-Strength

2010 38.3 56.2 207 0.5

Castable 220 2090 - 11.35 50 1.7

Michael Koszyk

Ceradyne Thermo-Sil® Fused Silica Materials <http://www.ceradyne.com>

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Window Requirements

• Curvature of material required• Filter out harmful radiation

– 0.1% Iron Oxide fused into glass

• Anti-reflective coating necessary• Structural analysis underway

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Other Required Structures/Mechanisms

• Fairing structure• Propulsion system structures• All secondary structures

– Antennae – Thermal regulation

• Mechanisms/Special Structures– Hatches/suit interface– Surface deployment– On-orbit deployment– Stage separation– Emergency/Rescue– Steering

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Structures Summary

SF StructureLoading

ConditionApplied Load

(MPa)MOS

2

Ribs/Stringer Launch 380 0

Pressure Hull 2 atm 550 0

Landing Structure 80 kN (I) TBD TBD

Wheels 3 kN (PL) 530 0.127

Chassis/Suspension 225 kN (I) 330 0.03

1.5

Avionics Support

Structure

TBD TBD TBD

Thermal Regulation Support Structure

TBD TBD TBD

3 Pressure Vessels TBD TBD TBD

David GruntzRahkiya Medley

Prim

ary

Str

uctu

re

Sec

onda

ry

Str

uctu

re

•Launch – 6g axially along Delta IV, 2.5g laterally•(I) – impulse load•(PL) – point load

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Mobility Systems

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Drive System• Overview

– Suspension– Tires– Engines, Drive-Train, Steering, and Brakes

• Surface propulsion’s Level 1 requirements– (M2) - Traverse 100 km in 10 hours, but overcompensated to 150

km → 15 km/hr (4.2 m/s)– (I9) – Capable to drive over terrain with 30° direct slope and 20°

cross slope– (I10) – Capable of turning in a 10 m radius– (M5/L7) – Safe return of crew following SPRITE failure (surface

propulsion needs to make this possible)– (I12) – Capable of towing a 2nd SPRITE 100 km to base

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Surface Propulsion Calculation• Calculate the frictional forces due to

tire roll based onFf = [0.87 / (b*k)1/2 ] * [W3/2 / D3/4] –b – Tire width–k – Average soil cohesion coefficient–W – Weight on each tire–D – Diameter of each tire–Multiply by number of tires

• Calculate force of gravity on incline of 30° (for peak power)

–Maximum load is the sum of friction on tires and normal force–Meets Level 1 requirement (I9)

Raja Krishnamoorthy

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Surface Propulsion– Power requirements:

• Continuous force ~ 8 kN– Assumes a constant velocity (4.2 m/s) on level ground

with each wheel ← Level 1 requirement (M2)

– Power Required ~ 36.5 kW (49 hp)• Maximum ascent force ~ 12 kN

– Assumes a constant velocity (4.2 m/s) up the slope of 30 degrees ← Level 1 requirement (M2) and (I9)

– Power Required ~ 55.5 kW (74 hp)– This represents peak locomotive power requirements,

but are conservative because of a safe estimate for velocity up an incline

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Drive System Requirements

92m 1.4ft 4.59m 0.30ft 0.98N 2041.3lb 459.28

Width of Wheel

Weight on Each Wheel

Diameter of Wheel

Engine Efficiency (%)

• Average engine efficiency is about 92% for an electric motor on the order of the power level required

• Weight, Torque and Power distribution on each wheel is about the same *These are rough estimates and will be refined throughout the course of the design process

Engine Requirements Units Peak Continuous

Nm 8169 5311.46lb-ft 6025 3917.20Nm 2042 1327.86lb-ft 1506 979.30kW 55.12 36.48hp 73.86 48.89kW 13.32 8.66hp 17.85 11.60

Torque per Wheel

Total Power Req

Power per Wheel

Total Torque Req

Assumptions Calculations

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AC vs. DC MotorPower (kW) RPM DC Motor Types

Mass Moment of Inertia

(kgm2) - DC

Ramp-up Time (s)

15 2000 DMP112-4L 0.05 0.619125 2000 DMP180-4LB 0.69 1.15329 2000 DMI225S 3 1.73560 1500 DMA+315M 10.68 1.57

Power (kW) RPM AC Motor TypesMass Moment

of Inertia

(kgm2) - AC

Ramp-up Time (s)

15 2000 180M4 0.161 0.946125 2000 315SMA4 2.3 1.73329 2000 355SMA4 8.2 2.42560 1500 450LG4 25 2.2

• Further analysis to be done with a wider range of motors

Raja Krishnamoorthy

Data from DC or AC Drives? A guide for users of variable-speed drives

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AC - DC Mass ComparisonMass vs Power

0

500

1000

1500

2000

2500

3000

3500

4000

4500

0 100 200 300 400 500 600

Power (kW)

Weig

ht

(kg

)

DCAC

• For a 15 kW motor the masses are as follows: AC – 175 kg DC – 110 kgEngines studied: DMP112-4L, DMP180-4LB, DMI225S, DMA315M, 180M4,

315SMA4, 355SMA, 450LG4

Raja Krishnamoorthy

Data from DC or AC Drives? A guide for users of variable-speed drives

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Other AC-DC Considerations

• Motor Controller types and setups (Pulse Width Modulation, Direct torque control, Vector Modulation, Phasing)

• Efficiency during variable speed operation and torque capabilities (TBD)

• Efficiency loss due to Temperature changes (TBD)• Other Drive-Train parts (Motor and Shaft sizing, Brake

Systems, Steering control and setup)

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Axle vs. Individual Wheel Drive

• A motor for each axle–Used 2-axle case

• Requires more power

• Provides more torque

• A motor for each wheel–Used 4-wheel case

• Requires less power

• Provides less torque

# of MotorskW hp kW hp

Continuous 13.00 17.42 35.00 46.90Peak 25.00 33.50 55.00 73.70

Nm lb-ft Nm lb-ftContinuous 1327.86 979.30 2655.73 1958.60Peak 2042.30 1506.20 4084.60 3012.40

4 motor 2 motor

Torque per Motor

Power per Motor

Raja Krishnamoorthy

4 MOTOR 2 MOTOR

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Risk Analysis for Drive Setups

• Can tolerate 2 failures*: 4 ways ↔ A-C, B-D, A-D, C-B

• R4 + 4R3(1-R) + 4R2(1-R)2

*Considered simple failure without wheel lock

• Can tolerate only 1 failure: A or B

• R2 + 2R(1-R)

2 MOTOR

4 MOTOR

R = e-t/MTBF = 0.999375 t = 25 hrs, MTBF = 40,000 hrs

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Future Analysis• Steering Systems

– Hydraulic or Electronically controlled or other– Meet Level 1 requirement (I10) – for 10 meter turning radius

• Braking Systems (derived requirement for braking distance at top speed)– Dynamic braking and regenerative braking incorporation

• Final drive-train setup – Dependent on number of wheels/axles– Disengaging clutch, gear setup, shaft sizing

• Motor Control (Level 1 requirement (M2) – speed min. of 10 km/hr)– Motor type determines controller type– Interface with avionics for speed control

• In-depth risk analysis - for number of motors and sizing – Dependent on final power numbers, number of wheels/axles, setup of motors– Need to find scenarios for different types of failures (i.e. wheel lock, locked

steering, brake lock)• Emergency systems

– Meet Level 1 requirement (I12) – Design to be able to tow a second SPRITE– Propulsion system design for emergency return of crew to base

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References• DC or AC Drives? A guide for users…

– https://www.abb-drives.com/StdDrives/RestrictedPages/Marketing/ Documentation/Documents/DCorAC.pdf

• Motor Formulas, 1997– http://www.elec-toolbox.com/formulas/motor/mtrform.htm

• Torque Capabilities of AC and DC Drives– http://www.powerqualityanddrives.com/torque_constant_ horsepower/

• Adjustable Speed Drives– http://www.hq.nasa.gov/alsj/lrvhand.html

• Lunar Rover Operations Handbook– http://www.hq.nasa.gov/alsj/lrvhand.html

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Wheels

Assumptions– Diameter > 1 m– Max point load = 3 kN

• Width vs. Power– Total power requirement for the locomotive changes with the width of

the wheel– Rolling friction is a function of width and length of the wheel.– Worst Case

• Vmax = 15km/hr• 30° incline

Requirement S1: All Systems shall be designed to provide a non-negative MOS for worst-case loading conditions incorporating Primary Structure: 2.0Requirement S2: All structural systems shall provide positive MOS for all loading conditionsRequirement I9: SPRITE shall be able to actively traverse terrain safely with a 30o slope Requirement I11: SPRITE shall be able to drive safely over 0.5 m obstacles in worst case

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Power vs. Width

• Width = 0.3 m

Power vs. Width (30 degree incline)

0

20

40

60

80

100

120

0 0.2 0.4 0.6 0.8 1 1.2

Width(m)

Po

we

r(k

W)

4 wheels

6 wheels

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Spokes

• Load is applied axially along the spoke (3 kN)

• Using aluminum

Length(m) 0.6

Width(m) 0.3

Thickness(m) 0.005

Mass (kg) 0.443

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Outer Rim• Force applied to the rim• Modeled as curved beam

under elastic bending• Assumptions

–Rectangular cross section–Constant radius of curvature–Bending moment due to point load remains perpendicular to the radius of curvature

Pyungkuk Choi

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Number of Spokes vs. Rim Thickness

• Titanium (10% Vanadium)–Density = 4650 kg/m3

–Tensile Strength = 1193 MPa

• Safety Factor = 2

Spokes Rim Thickness

(mm)

Inner σ (MPa)

Tensile

MOS Mass of one wheel

Total Mass

(kg) 4-wheels

Total Mass

(kg) 6-wheels

3 10.5 570.5 0.046 73.4 293.5 333.3

16 7 565.6 0.055 38.6 154.3 231.5

20 6.5 529.5 0.127 36.4 145.6 214.0

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Wheels - Future Work

• Tires• Cross slope loading• Different wheel configuration• Wheel protection

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Chassis / Suspension

• Struts connect to rib/stringer primary structure– External chassis if necessary

Chassis

Spring / Shock AbsorberWheel Mount

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Chassis / Suspension

• Factors considered– Load transferred by suspension– Vertical displacement of the vehicle

• Must absorb landing with residual velocity of 1 m/s (vertical) and 0.5 m/s (horizontal)

• Must absorb impulse resulting from a 0.5 m “fall” (~65 kN impulse)

• Must absorb impulse resulting from a collision (~225 kN impulse)

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Suspension Models

Linear Spring

•Modeled as 5,000 kg mass atop a linear spring

Lateral Torsion Bar

•Modeled as 5,000 kg mass attached to a 2 m moment arm

Axial Torsion Bar

•Modeled as 5,000 kg mass attached to a 0.25 m moment arm

David Gruntz

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Torsion Bar vs. Linear Spring

• Torsion bars transfer similar loads• Linear spring looks like ideal choice at this point

0

20

40

60

80

100

120

140

160

180

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7

Displacement (m)

Tra

nsm

itte

d F

orc

e (k

N)

Linear

Lateral

Axial

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Loads Transferred to Chassis

Type

Vertical Displacemen

t (m)

“Fall” Force (kN)

Landing Spring ConstantForce

(kN)Deflection

(m)

Linear 0.10.20.3

452620

362116

0.080.180.27

450 kN/m120 kN/m60 kN/m

Lateral Torsion

0.10.20.3

1005545

514032

0.140.160.23

1500 kN-m/rad1000 kN-m/rad550 kN-m/rad

Axial Torsion

0.10.20.3

856040

745035

0.10.150.27

50 kN-m/rad20 kN-m/rad8 kN-m/rad

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Initial Suspension Sizing• Titanium Ti-6Al-4V

– High specific strength (σyld/ρ) allows for a strong, lightweight chassis

• Initial chassis/suspension sizing with Titanium structure and steel springs– 20 kg – 140 kg

Load ConditionMax Stress

(MPa)MOS

Collision 330 0.03

“Fall” 280 0.2

Landing 200 0.7

Launch* TBD TBD

David Gruntz

* Will depend on how rover is integrated w/ fairing

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Crew Systems

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Consumables Summary

• Oxygen – 23.0 kg– Nominal usage ~ 0.85 kg/person-day– EVA usage ~ 0.63 kg/EVA– Leakage rate ~ 1% per day

• Nitrogen – 1 kg– Leakage rate ~ 1% per day

Larson, Wiley J., ed. Human Spaceflight: Mission Analysis and Design.

Requirement L4: SPRITE shall accommodate daily EVAs by a two-person team over a 5-day period, plus 2 contingency EVAs

Requirement L5: In case of the need to mount a rescue mission from base, SPRITE shall stock sufficient crew consumables to support the nominal crew at a subsistence level for

3 days following the normal sortie duration

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Consumables Summary

• Water – 250 kg– Drinking ~ 1.6 kg/person-day– Food hydration ~ 0.75 kg/person-day– Personal washing ~ 4.1 kg/person-day– Waste flushing ~ 0.5 kg/person-day– EVA cooling ~ 7.3 kg/person-EVA

• Food – 40 kg– ~ 2 kg/person-day required


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Atmospheric Composition

Lunar Base

•14.7 psi total pressure

•21% Oxygen

•79% Nitrogen

SPRITE Rover


•37% Oxygen

•63% Nitrogen

EVA Suit


•100% OxygenJohn Frassanito and Associates – <http://msnbc.msn.com/id/5990828>

<http://www.smallartworks.ca/PS/Space1999/AlphaMoonbase/AlphaMoonbase.html>

Alan Bean - <www.alanbeangallery.com/ab-artist.html> & www.andrew.cmu.edu/user/jplee/miscellaneous/new%20sprite%20bottles.jpg

R=1.4R=1.4

Larson, Wiley and Linda Pranke, ed. Human Spaceflight: Mission Analysis and Design

Requirement L8: SPRITE crew shall be capable of safely initiating extravehicular operations with no pre-breathe time beyond that required for suit donning and checkout

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Atmospheric Composition

Base – 14.7 psi SPRITE – 8.3 psi EVA – 3.5 psi

21% Oxygen 37% Oxygen 100% Oxygen

Man-Systems Integration Standards – NASA-STD-3000 <http://msis.jsc.nasa.gov/>


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Storage of Consumables

O2 tank N2 tank H20 tank

State Gas* Liquid Gas Liquid** Liquid

Mass 45 kg 320 kg 1 kg 319 kg 25 kg

Volume 0.09 m3 0.02 m3 0.004 m3 0.001 m3 0.025 m3

• All tanks assumed to be spherical

• Liquid tank specifications include required insulation

• Liquid storage would require power for cryogenic cooling

*Will be consolidated with Main Oxygen Tank to save mass

**Calculations assuming Liquid Nitrogen ~ LOX in properties

Requirement L2: All crew interfaces shall adhere to NASA-STD-3000, Man-Systems

Integration Standards

Michael Badeaux

Akin, David. ENAE483 Lectures Fall 2004 <http://spacecraft.ssl.umd.edu/academics/483F04Glatt, C.R. “WAATS – A Computer Program for Weights Analysis of Advanced Transportation Systems.” NASA CR-2420. Aerospace Research Corporation

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Temperature/Humidity Control

• Ideal Temperature ranges from 18-27 oC –SPRITE Cabin Temperature – 23 °C

• Ideal Humidity ranges from 4-16 oC

• Excess heat can be used to heat water

Wieland, Paul. Designing for Human Presence in Space NASA RP-1324 - <http://flightprojects.msfc.nasa.gov/book/rp1324.pdf>



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Temperature Control

Passive

Simple

Small Scale

Little Maintenance

Insulating Materials

Electric Heaters

Heat Pipes

Active

Complex

Large Scale

High Maintenance

Cold Plates

Heat Exchangers

Re-router

Heat RejectionFreudenrich, Craig “How Space Stations Work” - <http://science.howstuffworks.com/space-station4.htm>

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Carbon Dioxide RemovalRequirement L2: All crew interfaces shall adhere to NASA-STD-3000, Man-Systems


Removal Reduction

Regenerable Open Loop

2BMS EDC LiOH Sabatier

Weight 48.1 kg 44.4 kg 40 kg 76 kg

Volume 0.26 m3 0.071 m3 0.005 m3 0.14 m3

Heat N/A .336 kW N/A .268 kW

Power

Required

0.23 kW -0.148 kW AC -0.106 kW DC

0.012 kW .05 kW

Temperature 10 - 65 oC 18 - 24 oC 23 oC 427 oC

*EDC and LiOH have best overall qualifications for SPRITE

•Eckart, Peter. Spaceflight Life Support and Biospherics. Torrance, California: Kluwer Academic, 1994.

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Carbon Dioxide Removal• EDC

– Regenerable system

• Reacts H2 and O2 with CO2 inside and electrochemical cell

• CO2 + 0.5O2 + H2 CO2 + H20 + electrical energy + heat

– Products similar to H2-O2 fuel cell (H20 and DC power)

• CO2 concentration capacity may be regulated by current adjustment (capacity to handle large CO2 overload situation)

• Charges at base, generates usable 0.148 kW AC, 0.106 kW DC

• Mass = 44.4 kg; Volume = 0.071 m3

• Requires supply of H2 and O2

• Generates heat

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Carbon Dioxide Removal• LiOH

–Non-regenerable open loop–2LiOH + CO2 Li2CO3 + H20–The theoretical capacity of LiOH for CO2 is 0.92 kg CO2 per kg sorbent–Amount of LiOH required to remove one person’s daily average output of CO2 is about 2 kg

• Mass = 40 kg; Volume = 0.005 m3

–Power required = 0.012 kW

Lunar Module Environmental Control System. Historic Space Systems. <http://www.space1.com/Artifacts/Lunar_Module_Artifacts/LM_LiOH_Canister/lm_lioh_canister.html>

Shawn Butani

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Caution & Warning System

• Keeps crew aware that the current status of critical factors are within tolerable limits

• Important critical factors:– Fire/Smoke and particulate contamination– Pressure loss inside crew cabin– Pressure loss in tanks

– Atmospheric constituents (O2, N2, CO2)

– Power Generation and Electronic Cooling– Propulsion system operating conditions



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Caution & Warning System

• Interfaced with Environment Control, GNC, Power, Propulsion, Thermal, and Avionics

• Crew notified both audibly and visually– Audibly: Consists of a buzzer/siren

• Buzzer through headset• Siren at frequencies between 500 - 700 Hz

– Visual: Consists of a light array panel

Red – Emergency

Yellow – Cautious

Green – Nominal

<http://science.ksc.nasa.gov/shuttle/technology/sts--newsref/sts-caws.html>

<http://www.shuttlepresskit.com/scom/22.pdf>

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Acoustic Environment

• Noise generation should be controlled to reduce chance of personnel injury, communication interference, fatigue, or ineffectiveness of overall man-machine relationship- Equipment shall be designed to satisfy MIL-STD-1474B- Placement of all equipment should minimize noise at crew stations- C/W system should be integrated to monitor acoustic noise levels to verify that exposure limits are not being exceeded

• Safe Noise Limits- Maximum Noise Exposure - 115 dB is allowable, duration 2 min- Hearing Protection Devices - Provided for noise levels 85 dB

• Maximum Noise Level - Change in sound pressure level 10 dB 1 sec- Impulse noise shall not exceed 140 dB peak pressure level



Man-Systems Integration Standards – NASA-STD-3000 <http://msis.jsc.nasa.gov/>

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Contamination and Particulate Control

• Air filters– High Efficiency Particulate Arrestance (HEPA) filter – 99.7% efficiency on 0.3 microns

• NASA Standards 3000 - Section 13.2.3.1– Surfaces smooth, solid, nonporous– Grids easy to clean– No narrow openings– Areas must be covered when they are too narrow to

clean

Michelle Zsak

“Man-Systems Integration Standards.” NASA STD-3000. <http://msis.jsc.nasa.gov/>

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Contamination Control Wipes

• Biocide–Disinfecting food and waste systems

• Biofilm Control–Controls formation of Biofilm inside surface of fluid lines

• Cleaning Implements–Provides means for dislodging and collecting dirt/debris

• Detergent –Indoor cleaning

• Dry –Toilet tissue

• Utensil Cleaning–Sanitizers for post meal cleaning

• Vacuum

Michelle Zsak


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Waste Collection System (WCS)

• Internal system similar to shuttle• Presence of gravity eliminates vacuum• Urine stored in tanks under the system• Fecal matter is freeze dried and stored in tanks

under the system• Air filter used to eliminate odor and bacterial

contamination

Michelle Zsak


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Trash Management

• Ways to store trash–Free standing trash receptacle–Storage compartment built into structure–Trash compactor to minimum trash space

Michelle Zsak

2-Man Crew,

1-wk Mission

Mass (kg) Volume (m3)

Total 9.1 0.202

Food 4.5 0.16

WCS Supplies 4.6 0.042

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Radiation SourcesGalactic Cosmic Rays Solar Particle Event

Duration Near Constant 1-3 days

Composition85% Protons14% Alpha

1% Nuclides

90% Protons

10% Alpha

Flux Density

(photons/cm2-sec)

0 - 1

max ~2

0 - 104

max ~106

Energy Levels

(MeV)

102 - 104

max ~1011

10 - 103

max ~104

Michelle Zsak


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Radiation Limits

Lifetime Limits: Blood-Forming Organs (BFO) 5 cm depth

Gender Age

25 35 45 55

Male 150 rem 250 rem 325 rem 400 rem

Female 100 rem 175 rem 250 rem 300 rem

Exposure Interval

BFO

5 cm

Eye

0.3 cm

Skin

0.01 cm

10 days 8.33 rem 33 rem 50 rem

30 days 25 rem 100 rem 150 rem

Requirement L6: Radiation dosages shall, under all conditions, conform in all respects to

the current NASA standards for astronaut radiation limits

Michelle Zsak

Wilson, John, Francis Cucinotta, Lisa Simonsen, and Judy Shinn. “Galactic and Solar Cosmic Ray Shielding in Deep Space.” NASA Technical Paper. Dec 97.

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Shielding Options

• Rejected Shielding– Lunar Shielding

• In research• Charged spheres that deflect protons and sift out electrons• Not enough information

– Mass– Power– Cost

– Mars Bricks• Under development• Produce radiation-resistant bricks with local materials on

surface• Not sure if possible on the moon surface

Michelle Zsak

Malik, Tarig. “Lunar Shields: Radiation Protect for Moon-Based Astronauts.” <http://www.space.com/businesstechnology/lunarshield_techwed_050112.html>Sonja, Baristic. “Making Mars Bricks for Long Term Red Planet Stays.” <http://www.space.com/sciencesastronomy/solarsystem/mars_bricks_wg_000816.html>

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Shielding Options

• Possible Shielding– Aluminum

• Currently used• Creates neutrons during nuclear interaction that increase

exposure

– Polyethylene (CH2) without water

• Shields more than Aluminum since it is Hydrogen rich

– Polyethylene with water• Shields 20% more than Aluminum since it is Hydrogen rich• Must consider mass budget

Michelle Zsak

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University of MarylandENAE484 PDR March 14, 2005Michelle Zsak

Aluminum vs. Polyethylene

Thickness

(g/cm2)

Dose Equivalent

(rem/yr)

Al CH2

0 95 95

1 91 81

2 88 83

5 79 71

10 69 57

15 54 41

25 46 35

75 43 32

Solar Minimum 1977

Thickness

(g/cm2)

Dose Equivalent

(rem/yr)

Al CH2

0 34.5 34.5

1 33.7 32.7

2 32.9 31.2

5 30.7 27.2

10 27.8 22.6

15 22.8 16.4

25 20.0 14.4

75 19.4 13.7

Solar Maximum 1970

Wilson, John, Francis Cucinotta, Lisa Simonsen, and Judy Shinn. “Galactic and Solar Cosmic Ray Shielding in Deep Space.” NASA Technical Paper. Dec 97.

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Possible Radiation Shielding Plan

• Shield all sides exposed to radiation• 0.4 cm aluminum hull provides shielding• Polyethylene shielding specific mass ~10 kg/m2

- with surface area of 39 m2 ~390 kg• 3 cm thickness of water from fuel cells provides

additional shielding for Solar Particle Event (SPE)

Michelle Zsak

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Fire Suppression

Type

Liquid Density

Volume Fraction

Comments

Halon 1301

1570 kg/m3

0.20 Highly effective

CO2 758 kg/m3 0.62 Toxic in high concentration

Can be cleaned by rover

• Oxygen masks required for crew during fire suppression

• Extra CO2 scrubber can be carried for post fire clean-up

• Halon 1301 decomposes into toxic products which must be filtered out post fire

Friedman, Robert: “Fire Safety in Extraterrestrial Environments.” Lewis Research Center, May 1998.

John Mularski

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Internal vs. External SuitsInternal External

Mass Airlock ~ 400 kg Suit Shields ~ 250 kg

Volume EVA Suits ~ 2 m3

Airlock ~ 4 m3

No internal space reduction

Power Pumping air out of airlock TBD

None

Habitability Airlock allows dust intrusion into cabin

None

Suit Condition Allows for crew maintenance of suits

Suits continuously exposed


Dumoulin, Jim: “Space Shuttle Coordinate System.” Kennedy Space Center, August 2000 <http://science.ksc.nasa.gov/shuttle/technology/sts-newsref/sts_coord.html>

John Mularski

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Layout

John Mularski

John Mularski

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Layout

• Current cabin volume = 21 m3

• Space surrounding cabin for pipes, wires and auxiliary equipment

Requirement L1: All crew interfaces shall accommodate 95th percentile American male to 5th percentile Japanese female

John Mularski

John Mularski

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Layout

• Bunks fold to provide access to external suit and stowage

• Food prep station used for stowage and hydration of food as well as personal hygiene

Requirement L1: All crew interfaces shall accommodate 95th percentile American male to 5th percentile Japanese female

Requirement L7: System shall provide for emergency alternative access and EVA “bailout” options

John Mularski

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Visual Display (VD)

• VD must be at least 13 in, preferably > 20 in• VD viewing distance: min = 16 in, max = 28 in• Navigation accomplished through use of

cameras and/or window, therefore require 6 or 7 monitors– 2 main multi-function displays (MFD) (2 - system

stats, for astronaut convenience)– 3 navigation displays (1 - primary view, 1 - data view,

1 - switch between auxiliary camera views)– 1 VD per robotic arm

Shawn Butani

Requirement L1: All crew interfaces shall accommodate 95% American male to 5% Japanese femaleRequirement L2: All crew interfaces shall adhere to NASA-STD-3000, Man-Systems Integration Standards

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Windows

• Inputs–Cabin height = 2.1 m

–vessel = 3 m diameter (tires add 0.5 m from ground)

–95th male sitting height eye level = 135 cm

–Line of sight = 24.7o +/- 2.4o

–Eye movement laterally: 35o max, 15o optimum 25o (easily with head moment range)

• Output–Navigator can see the ground 0.648 m ahead of the rover

–Minimum window size (mass constraint) = 42 cm length, 40 cm width

• Problems…–Stringers will divide window

–Curvature of rover

Finding minimum window dimensions for navigational purposes

Shawn Butani

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Window Solution

• Structures designed two windows evading the stringers

• Windows fit the curve of the rover

• Preliminary analysis and sector angle (33º per window) show ample room for navigation

• Length of window = 1.26 m

• Window separation = 0.24 m

• Future work includes performing thorough analysis of viewing range

Michael Badeaux

Shawn Butani

WindowSeats

Hull

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Front Display• Astronauts sit 16 in. from windows

and MFD

• MFD = .69 m (~27 in)

• NAV-PRI/AUX = .56 m (~22 in)

• NAV-DATA = .431 m (~17 in)

• Seat separation = .24 m

• Control panel includes :– Steering system : Throttle (SDOF), L &

R steer (SDOF), Lift Break

– Avionics : input from driver, indicators, sensors (wheels, pitch and roll, speed, etc.)

WindowSeats

Hull

Evan Ulrich

Shawn Butani

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Geographic SurveyRequirement M3: SPRITE shall be capable of replicating the science from an Apollo J-Class

lunar EVA, in terms of both instrument deployment and sample collection

• Cupola–During navigation, the second astronaut will be able to survey the area with 360° field of view–Mass estimates and structural design still in preliminary stages

Shawn Butani

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EVA Suit ShieldingRequirement L4: SPRITE shall accommodate daily EVA by a two-person team over a five day

period

• Shield serves to protect I-suit from micrometeoroid impact and dust storms

• Static Dissipative Polycarbonate – high impact strength and modulus of elasticity, absorbs little moisture, does not attract dust or other contaminants (surface resistivity (106 – 108 Ω/in2)

Strength (psi) Modulus (psi)

Tensile 9,500 320,000

Flexural 15,000 375,000

Compressive 12,000 240,000

Polycarbonate Specifications, www.boedeker.com

Shawn Butani

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Calculating Shield Dimensions

• Density = 0.043 lb/in3

= 1.2 g/cm2

• Designed one shield to fit two 95th percentile males with +/- 10 cm for each dimension

• Designed as a rectangular shaped enclosure to calculate maximum mass

• Mass = 260 kg• In the future will design to

better fit the suit and optimize mass

95th percentile male (cm)

A – Height 191.9

C - Width 66.0

D – Depth w/ PLSS 68.6

NASA-STD-3000, Volume 1 section 14. http://msis.jsc.nasa.gov/sections/section14.htm

Shawn Butani

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Crew Systems Future Work

• EVA checklist• Health monitoring• Interior stowage• Docking system• EVA support• Controls and displays

Shawn Butani

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Intermission

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Surface GNC

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• IMU accuracy is vital to mission success– IMU drift bias is 0.0003 deg/hr *– Star trackers are re-aligned to compensate for IMU drift bias

• Star tracker to be re-aligned within 1.4 deg error• Star trackers require calibration after about 4667 hours

– IMU Reliability is > 0.996 *• Use 2 IMUs on spacecraft and rover• Probability that at least 1 IMU works > 0.9999

Aaron Shabazz

* Data from Honeywell IMU spec sheetSpec Sheet - http://content.honeywell.com/dses/assets/datasheets/fog.pdf

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• Works Still in Progress– GNC Thermal Control– Determining which computers to use– Determining number of computers needed

Aaron Shabazz

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Navigation and Guidance on Moon Surface

• SPRITE shall be capable of navigating – Within 100 m of target

– Both day and night

• Absolute Navigation constraints on moon– Communication limited to

only base, earth and L2 satellite

– LOS, and natural landmark barriers

– No medium for sound to travel through

Navigation method w/ Moon Map

Trade study Accuracy

(m)

Method

Constraint

Celestial Sun and Earth Tracker

300 At least 600 obs.

Landmark VIPER 180 Needs assistance

at night

Low Frequency

Radio

Loran

Submarines

100 2 or more beacons

Ralph Myers

http://www.mit.edu/~ykuroda/research/iSAIRAS03Locali.pdf

http://www-2.cs.cmu.edu/~viper/Results/

Borenstein, Johann J., H.R. Everett, and Liqiang Fang. Navigating Mobile Robots. Wellesley, MA: AK Peters, 1996

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Navigation and Guidance on Moon Surface

• Use landmark for absolute reference and dead reckoning sensors for relative reference

• Errors in the dead reckoning sensor will determine the distance needed before a landmark is needed for correction update

Vehicle &Landmark

Latitude andLongitude

Compare values to

Lunar Map

Accel. X

Real time Calibration

Odometer

Accel. Y

Accel. Z

Gyro Roll

Gyro Pitch

Gyro Yaw

Slippage detection

AccelerometerCompensation

Torquesensor

Left Front Right Front

Left Rear Right Rear

Scan horizonor predetermined

landmark

Build DEM andCompare to lunar

Map surface

Myers, Ralph

Ralph Myers

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On Board Direct Human Control

• Drive by wire will control steering, acceleration, and braking through a feedback loop

• Have to reduce odometer errors caused by slippage

– Assuming driver has to control 4 independently motored wheels

• Assume Ackerman Steering to comply with 10 m turn radius requirement

• SPRITE shall incorporate sensors to allow positive diagnosis of credible failures in safety critical systems

Ralph Myers

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Minimize Odometry ErrorSpecifications for odometry accuracy

Encoders Resolvers

Controllable Speed Range 0.1 rpm to

10,300 rpm

30 rpm to

15,600 rpm

Counts Per Resolution 32,640 16,384

Signal Periods Per Revolution 2048 1

Accuracy Range (arc-minutes) 1 to 1.5 7 to 15

Tolerable Shock Level

(gs)

5 50

Operating Temperature Range

(ºC)

0 to 100 -55 to 175

Ralph Myers

http://www.heidenhain.com/Linear-2.htm

http://www.motec.co.uk/documents/ormec/encres.htm

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Robot Arm Control• SPRITE shall provide capability

for crew to interact with environment without EVA

• Teleoperator should be able to manipulate the arm

• Tactile sensors provide feedback to the operator

Sensor Parallel to human hand

Location

Tactile array sensor

Give feel of object’s shape

Outer surface of finger tip

Finger tip force-torque sensor

Determine how operator manipulates object

Near finger tip

Finger joint angle sensor

Position of robots manipulators

Finger joints or at motor

Actuator effort sensor

Motor torque as wrist movement

At motor or joint

Dynamic tactile sensor

Vibration, stress to tell if object is being fumbled

Outer surface of finger tip

Ralph Myers

http://www.biorobotics.harvard.edu/pubs/tac-manip.pdf

http://www.biorobotics.harvard.edu/pubs/tac-manip.pdf

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Surface Obstacle Avoidance

SPRITE must traverse 0.5 m obstacles, 20º cross-slope, 30º forward slope– Must detect hazardous terrain

• Derived detection requirements– Minimum look ahead distance - 4 m

• Based on minimum stopping distance– Maximum look ahead distance - 13 m

• Based on tightest turning radius

• Stereo camera strategy chosen

Scott Walthour

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Surface Obstacle Avoidance

• Camera parameter derivation assumptions– Maximum deceleration: 0.45g (comfortable automobile

deceleration)– Obstacle detection rate: 1 Hz (DEM updated every second)– Maximum velocity: 2.77 m/s (10 km/hr)– Resolve 0.5 m object at maximum look ahead distance– SPRITE width: 2 m

Scott Walthour

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Minimum Look Ahead Distance

Scott Walthour

Scott Walthour

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Maximum Look Ahead Distance/ Camera Horizontal Field of View (HFOV)

Scott Walthour

Scott Walthour

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Camera Vertical Field of View (VFOV)

VFOV dependent on:– Vertical location of sensor

• Negative obstacles* need sensor as high as possible– Assume = 3 m (located on top of SPRITE)

– Maximum obstacle size to be seen at 13 m• Assume = 1 m

*Negative obstacles – ditches, craters, etc.

Scott Walthour

Scott Walthour

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Derived Obstacle Detection Requirements

Minimum Look Ahead Distance 4 m

Maximum Look Ahead Distance 13 m

Horizontal Field of View 103 deg

Vertical Field of View 29 deg

Angular Resolution* (mrad/pix) 1.88 (H) x 1.75 (V)

Minimum Image Resolution (pix) 954 (H) x 290 (V)

Update Rate 1 Hz

Stereo Camera Locations 3 m vertical

Camera Separation 2 m baseline

Night Operations Headlights

*Horiz:10 pixels on 5th %ile female width (24.5 cm) at 13 m Vert: 6 pixels on .5 m diameter ditch at 13 m

<http://msis.jsc.nasa.gov>

Scott Walthour

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Obstacle Detection Future Work

Choose COTS* cameras – Resolution– CCD, CID, Vision chips

Determine computational requirements

*COTS – commercial off the shelf

Scott Walthour

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Network Data Bus

• Network requirements– Data Rate – 50 Mbps

• HDTV requirement - 40 Mbps• Bidirectional transmission - 10 Mbps

• Serial vs. Parallel bus (serial reduces wiring)• Other busses (e.g. 1553a, 1773) have limitations:

– 1-20 Mbps data rate *too low– Node limitations– Half-Duplex

• Bus choice– Spacewire (std ECSS-E-50-12 A) – serial bus

Requirement A1: The SPRITE communications system shall be capable of supporting continual upload transmission of one channel of HDTV, download of two channels of HDTV and bi-directional transmission of 10 MB/sec direct to Earth when parked

<http://www.interfacebus.com>

Scott Walthour

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Network Data Bus

Advantages Disadvantages

High Data Rate (Mbps)

155-200 typical (400 max)

Lightweight

0.06 kg/m

Scalable

Radiation Tolerant

BER = 10-14*

Full Duplex

Not inherently redundant

• Requires routers to ensure

redundant paths

- Increases complexity of the

network

Spacewire

<http://www.estec.esa.nl>

Scott Walthour

*Bit Error Rate

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Example Network

Scott Walthour

Scott Walthour

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Network Data Bus

• Number of routers dependent on number and type (e.g. pressure sensor) of nodes– Desire redundancy

• Divide pressure sensors on multiple routers in case of router failure

• Future work:– Organize SPRITE’s data network

Scott Walthour

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Communications

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Communication Requirements

From Work Breakdown Structure• From SPRITE to Earth• From SPRITE to Base • From SPRITE to EVA • Contingency/Emergency

Requirement A1: The SPRITE communications system shall be capable of supporting continual upload transmission of one channel of HDTV, download of two channels of HDTV and bi-directional transmission of 10 MB/sec direct to Earth when parked

Jay Kim

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High Definition TeleVision

• HDTV specs–1920 pixels by 1080 lines–30 frames per second–3 primary colors (red, blue and white)–8 bits for each color–Uncompressed data rate at 1.5 Gbps

• Compression technique–MPEG 1: Standard for Video CD–MPEG 2: Standard for broadcast-quality television

• Compression rate up to 20 Mbps

Comparison of different displaysJay Kim

Jay Kim

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From SPRITE To Earth• Assumption: SPRITE is parked• Scenario 1

–SPRITE is on near side and Earth is in LOS–Communicate directly using antenna

• Transmission rate–20 Mbps at 1 channel–Bidirectional transmission of 10 Mbps of digital data

• Uplink = 30 Mbps (from Earth to SPRITE)• Downlink = 50 Mbps (from SPRITE to

Earth)

• Trade studies of link budgets–Frequency selection–Antenna selection

• Link budget constraints–Link margin 3 dB – 6 dB

High Gain Antenna Low Gain Antenna

Far sideNear side

Jay Kim

Jay Kim

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Link Budget• Initial assumption

–Ka band: widely used in spacecraft communication–Diameter of antenna: 1 m–High gain antenna: precision in targeting –Transmitter power: 20 W–Slant range: 400,000 km (Apoapsis of Moon)–Receiver antenna: Deep Space Network (34 m)

Effect of changing diameter David G. MacDonnell, “Communications Analysis of Potential Upgrades of NASA’s Deep Space Network”

Akin, Dave. ENAE483 Link Budget Spreadsheet

Jay Kim

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Link Budget

• Diameter of antenna size: 0.5 m• Transmitter power: 1 W• Mass: TBD

Operating frequency: 15 – 25 GHz

Link margin: 3dB – 6dB

Effect of changing transmitter powerJay Kim

Jay Kim

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SPRITE To Base• Transmitting antenna in SPRITE

– UHF Band: widely used in short distance communication

– Diameter of antenna: 0.5 meter– Transmitting power: 1 Watt– Slant range: 150 Km– Data rate: 50 Mb/s (HDTV)

• Receiving antenna in base– Same antenna as transmitting antenna– Takes advantage of learning curve

Operating frequency: 1 – 1.5 Ghz (UHF Band)

Link Margin: 3 – 6 dB

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Emergency

• In case of emergency– SPRITE communicates to

Base

– Use low gain antenna

• Reliable signals

• No pointing required

– Link budget (TBD)

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Flying Locator and Assistance Requesting Equipment (FLARE)

• To be used in the event of a regular communications pathway failure• Launch a small communications package (10 kg, 25 cm2) to provide temporary

link between the rover crew and base.–Small solid rocket motor for propulsion–Equipment based on amateur radio microsatellite technology

Motor MassWindow Duration

(150 km from base)Total Package Mass

1.5 kg 3.5 minutes 11.5 kg

4.5 kg 8 minutes 14.5 kg

• Small and lightweight communications solution• Still need to determine actual mass of electronics package, integration

with SPRITE, and communications window duration required for transmission of data/voice

ATK Retro/Separation Motors: <http://www.atk.com/starmotors/starmotors_retrooverview.asp>

AMSAT Echo Information: <http://www.skyrocket.de/space/doc_sdat/amsat-echo.htm>

Timothy Wasserman

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Future Work for CDR

• Rover to EVA communication– Need to work with Crew Systems– Determine requirements for EVA suit communication system

• Far side communication• Satellite communication

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Power Systems

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Power & Energy: Requirements and Budget

• Power and energy budget has been created to establish a buffer between requirements and available power and energy

• Current assumptions–Time for avionics, crew systems, thermal, and science missions power consumption have been estimated at full time usage

Jason West

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Power Requirements Overview Power [kW] Energy [kW-hr]

SPRITE Total 55.6 2276

Surface Propulsion

Nominal required 36.5 1277.5

Peak required 55.5 55.5

Continuous

Avionics

Communications (SPRITE to Earth) 0.02 3.8

Communications (SPRITE to Base) 0.02 3.8

Communications (SPRITE to EVA) 0.02 3.8

IMUs 0.032 6.1

Star Trackers 0.01 1.9

GNC Computers 0.015 2.9

Avionics total 0.117 22.5

Crew Systems

CO2 removal 0.012 2.3

Jason West

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Power: Requirements and BudgetRequired

Power

[kW]

Budgeted

Power

[kW]

Emergency

Power

[kW]

Surface Propulsion

(max)

36.5

(55.5)

40.0

(60.0)

0

(0)

Avionics .117 .250 TBD

Crew Systems 1 1 1.0

Science Mission TBD 1 0

Thermal TBD 1 TBD

Miscellaneous TBD 1 TBD

Total (max) 37.617(56.6) 44.25(64.3) 1.0(1.0)

Jason West

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Power: Requirements and Budget

Jason Mallare

55.5

0.117

4.25

60

0

10

20

30

40

50

60

Surface Propulsion Other

Po

wer

(kW

)Required Budgeted

Emergency Power, 1 kW

Nominal Power, 44.25 kW

Max Power, 64.3 kW

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Energy: Requirements and Budget

System

Power Req(kW)

Time(hr)

Energy Req(kW-hr)

PowerBudgeted

(kW)

Time(hr)

EnergyBudgeted(kW-hr)

SurfacePropulsion(cruising) 36.5 35 1277.5 40.0 35 1400.0

SurfacePropulsion(ascent) 55.5 1 55.5 60.0 1 60.0

Avionics .117 192 22.5 .250 192 48.0

CrewSystems 1 192 192.0 1 192 192

Thermal TBD TBD TBD 1 192 192

ScienceMission TBD TBD TBD 1 192 192

Misc. TBD TBD TBD 1 192 192

TotalEnergy 1357.8 2276.0

Jason Mallare

192 hours represent 8 day, 24 hour/day usage

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Energy and Power:Bottom Line

• Current bottom line energy/power budget for SPRITE– 2276 kW-hr of energy– 44.25 kW of nominal power with peak capabilities of 64.3 kW

• Current emergency power requirements– SPRITE

• 72 kW-hr of energy – meets L1 requirement for 3 day emergency

• 1 kW

Jason Mallare

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Power Management & Distribution

• Future Work– AC vs. DC– Centralized vs. Distributed power conversion

• Considerations: Ohmic losses in wires, hazard of 100+ V distribution throughout entire craft

– System Voltage• 28 V vs. 100 V system

Jason Mallare

Hyder, Wiley, Halpert, Flood, Sabripour. “Spacecraft Power Technologies”

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Energy Storage

• Technologies considered– Primary batteries– Secondary (rechargeable) batteries– Radio-isotope– Solar arrays– Fuel cells

Jason Mallare

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Primary Batteries

Chemistry

GravimetricSpecific Energy

(W-h/kg)

VolumetricSpecificEnergy(W-h/L)

SpecificPower(W/kg)

Minimum Temperature

(oC)

Maximum Temperature

(oC)

LiSOCl2 740.0 1241.4 0.04 -60 55

Li-Mn02 271.3 568.1 51.76 -30 72

Li-SO2 328.7 512.0 9.59 -60 70

Ni-MH 72.0 246.5 14.29 -10 40

• Advantages:• Primary cells offer higher specific energy then secondary batteries

• Disadvantages:• Non-rechargeable, low current, low specific power (W/kg)

Jason Mallare

<http://www.saftbatteries.com/010-Home/10-10_home.asp><http://www.varta.com/eng/><http://www.ulbi.com/>

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Secondary Batteries

Chemistry

GravimetricSpecific Energy

(W-hr/kg)

Volumetric Specific Energy(W-hr/L)

SpecificPower(W/kg)

Minimum Temperature

(oC)

Maximum Temperature

(oC)

Cycle Life

(cycles)

Li-Ion 200 300 244 -40 60 500

Sodium Sulfur 240 304 200 300 350 2500

Li-Polymer 206 386 309 -20 60 500

• Advantages:• Secondary batteries generally allow a larger current, resulting in

greater specific power (W/kg) then primary batteries• Disadvantages:

• Secondary batteries have a lower specific energy (W-hr/kg) then primary batteries

Jason Mallare

<http://www.saftbatteries.com/010-Home/10-10_home.asp><http://www.varta.com/eng/><http://www.ulbi.com/>

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Batteries - Energy StorageBattery Specific Energy

386

1241

247

740

271240

329200 206

72

512

304300

568

0.0

200.0

400.0

600.0

800.0

1000.0

1200.0

1400.0

Li-Ion Li-Polymer

SodiumSulfur

Li-SOCl2 Li-Mn02 Li-SO2 Ni-MH

Gra

vim

etri

c S

pec

ific

En

erg

y(W

-hr/

kg)

Gravimetric Volumetric

Vo

lum

etric Sp

ecific En

ergy (W

-hr/L

)

Secondary Batteries

Primary Batteries

Jason Mallare

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Batteries - Power GenerationBattery Specific Power

244

200

0

52

10 14

309

0.00

50.00

100.00

150.00

200.00

250.00

300.00

350.00

Li-Ion Li-Polymer SodiumSulfur

Li-SOCl2 Li-Mn02 Li-SO2 Ni-MH

Sp

ecif

ic P

ow

er (

W/k

g)

Secondary Batteries

Primary Batteries

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Radio-isotope Power Systems

• Converts thermal energy generated from radioactive decay to electrical energy

• Rejected due to low power output per unit – At installation, power output is 110 W of electricity– After 14 years, power output is only 94-100 W of electricity

Phillip Adkins

<http://newfrontiers.larc.nasa.gov/newfrontiers/09_NF_PPC_Schmidt.pdf>

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Solar Cells

• Converts light to electrical energy– Estimated mass - 483 kg array

– Estimated area of 235 m2

• Reasonable efficiency with high specific power

• Not favorable: – Moon - restricts missions to the day side– Mars - restricts missions to the day side

• Additional area needed for same power output

Phillip Adkins

<http://spacecraft.ssl.umd.edu/academics/483F04/483L14.power_sys/483L14C.power.2004.pdf>

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Fuel Cells

Type

Specific Power (W/kg)

Efficiency Operating Temperature

(oC)

Alkaline 100-150 50-70% Below 80

Proton Exchange Membrane (PEM)

100-150 35-60% 75

Direct Methanol 100-150 35-40% 75

Phosphoric Acid TBD 35-50% 210

Molten Carbonate TBD 40-55% 650

Solid Oxide TBD 45-60% 800-1000

Phillip Adkins

<http://www.fuelcells.org><http://www.astronautix.com><http://www.utcfuelcells.com>Patel, Mukund R. Spacecraft Power Systems. Boca Raton: CRC Press, 2005<http://t2spflnasa.r3h.net/shuttle/reference/shutref/index.html>

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Fuel Cell Mass Calculations

• Max Power estimated at 64.3 kW– Assuming a specific power of 100 W/kg for the fuel cell reactor.

• Total Energy needed estimated at 2276 kW-hr– Using alkaline fuel cells and assuming 70% efficiency for the fuel cells.

Fuel Cell Reactor 640 kg

Reactants 860 kg

H2 and O2 tanks 420 kg

Total Mass 1920 kg*

* ~38% of total rolling mass

Phillip Adkins

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Power for Transit to Moon and to Base

• Only include enough reactants to power systems during the transit to the moon and for the drive to the base. – Mass of Reactants needed: 152 kg.– Total Mass estimate (with the fuel cell reactors and full size

tanks): 1222 kg.

Phillip Adkins

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Thermal Control

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Thermal Control

• Requirements– Maintain cabin temperature between 18.3 and

26.7ºC– Cool electronics and motors so that

equipment operates at peak efficiency

Evan Alexander

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Passive Thermal Control

• Multi-Layer Insulation System (MLI)– Several layers of thermal blankets used to insulate the cabin

• Advantages– Lightweight– Low thermal conductivity

• Disadvantages– Conductive properties diminished in areas where layers

meet

Evan Alexander

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MLI• Use layers of Mylar due to it its density as well as its

absorptivity and emmisitivity• Decron netting used to separate layers of Mylar

Material Features Thickness (µm) Emissivity Absorptivity

Mylar

Y9360-3M Aluminized TBD 0.03 0.19

Aluminized Backing 3.8 0.28 0.14

TeflonGold Backing 12.7 0.49 0.30

Kapton FilmAluminized Backing 2.0 0.24 0.23

Evan Alexander

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Aerogels

• Extremely lightweight form of insulation– Advantages

• Lighter than MLI system• Lower thermal conductivity

– Disadvantages• Structurally weak

• May be used in conjunction with MLI to improve insulation at joints

Evan Alexander

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MLI vs Aerogels

Category

Type Density (g/cm3)

Thermal Conductivity (W/m-K)

MLI

Kapton 1.42 0.12

Mylar 1.39 0.2

Teflon 2.15 0.195

Aerogels

Silica 0.01-0.3 0.004

Resorcinol 0.6 0.06

Carbon 0.9 0.04

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Active Thermal Control

• Use Heat Pipes to cool electronics• Radiators used to expel excess heat from cabin

<http://spacecraft.ssl.umd.edu>

Evan Alexander

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Heat Pipes• Use capillary motion in order to wick fluid

throughout the piping• Heat is transferred through the pipes to the fluid

around the sides which evaporate into the center of the pipes

• Heat flow through a pipe is a function of• k = Thermal conductivity• Te = Temperature of evaporator• Tc = Surface temperature of condenser• Tv = Temperature of vapor

Evan Alexander

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Heat Pipes (cont.)

• Properties of possible heat pipe fluids

Temperature Range (°C)Heat Pipe Working Fluid

Heat Pipe Vessel Material

-200 to -80 Liquid Nitrogen Stainless Steel

-70 to +60 Liquid AmmoniaNickel, Aluminum, Stainless Steel

+5 to +230 Water Copper, Nickel

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Heat Pipes (cont.)

• Properties potential metals used

MetalsDensity (g/cm^3)

Thermal Conductivity (W/m-K)

Aluminum 2.7 205

Nickel 8.91 90.7

Stainless Steel 8.03 50.2

Copper 8.92 394

Evan Alexander

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Radiators

• Condenses fluid from heat pipes• Expel excess heat from electronics at a rate

proportional to its area– A = Qrad / (σ * (T^4 – Ts^4))

• Qrad = Heat radiated • σ = Stefan-Boltzmann constant• Ts = Temp of heat sink• T = Temp of incoming fluid/vapor

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Science

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Suggested Landing/Mission Zones

• Crater Copernicus• Crater Tycho• Mare Orientale• South Pole-Aitken (SPA)

Basin

Lunar and Planetary Institute, 2005

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Crater Copernicus• Geographic Interest

–Diameter of ~90 km–Depth of ~4 km –Near side of Moon–Interesting central mountain range (~1 km above floor)–Ease of landing–Deeper inspection of the Moon’s crust


Lunar Orbiter image II-162H3

Chris Hartsough

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Crater Tycho

• Geographic Interest–Diameter of 85 km–Average depth of ~4 km–Central peak rising ~2.5 km–Ease of landing–Relatively young crater (one of the youngest

large craters on near side)–Deeper inspection of the Moon’s crust


Chris Hartsough

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Mare Orientale

• Geographic Interest–Diameter of ~950 km–Depth of ~3.2 km–Multi-leveled mare–Large iron concentration–Ease of landing–Half visible to earth



Chris Hartsough

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South Pole-Aitken (SPA) Basin• Geographic Interest

–Diameter of ~2500 km–Depth of ~12 km on average–Largest known impact crater on the

Moon–Deposits of iron and titanium–Possibility of water–Deeper inspection of the Moon’s crust


Lucey et al., 1998

Chris Hartsough

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Choosing Scientific Instruments for SPRITE

Completed steps1. Detail the mass and volume requirements for

scientific hardware used in previous J-Class missions.

Requirement M3: SPRITE shall be capable of replicating the science from an Apollo J- class lunar EVA on each of the 5 EVA days of the sortie

Ryan Livingston

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Instruments - Crew Experiments

Experiments *

Original Mass (kg)

Returned Mass (kg)

Stored Volume (m3)

Soil Mechanics Investigation** 15.7 15.7 TBD

Solar Wind CompositionExperiment 0.46 0.385 1.3e-3

Lunar Portable Magnetometer 0.46 0 1.18e-2

Far UltravioletCamera/Spectrograph 22 0 0.25

* Hand Tools to assist experiments = approx 50 kg

** includes ALSD (drill)

Ryan Livingston

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Experiments *

OriginalMass (kg)

ReturnedMass (kg)

Stored Volume (m3)

Cosmic Ray Detector 0.163 0.163 0.13e-3

Transverse GravimeterExperiment 14.6 0 0.0351

Lunar Neutron Probe 2.27 0.4 0.38e-3

Surface Electrical Properties 16 1 0.024

Instruments - Crew Experiments

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Instruments - Deployed

Experiments

Original Mass(kg)

Returned Mass

(kg)

Volume (m3)

Passive Seismic Experiment 11.5 0 0.012*

Heat Flow Experiment 9.9 0 0.023

Lunar Surface Magnetometer 8.6 0 0.044

Laser Ranging Retroreflector 36.2 0 0.135

Cold Cathode Gauge 5.7 0 0.012

Suprathermal Ion Detector Experiment 8.8 0 0.014

Solar Wind Spectrometer 5.3 0 0.007

* does not include foldable skirt

Ryan Livingston

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Experiments

Original Mass (kg)

Returned Mass (kg)

Volume (m3)

Lunar Dust Detector 0.27 0 TBD

Active Seismic Experiment 11.2 0 TBD

Lunar Seismic Profiling Experiment 25.1 0 TBD

Lunar Atmospheric Composition Experiment 9.1 0 0.018

Lunar Ejecta and Meteorites Experiment 7.4 0 0.02

Lunar Surface Gravimeter 12.7 0 0.027

Instruments - Deployed

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Choosing Scientific Instruments for SPRITE

Future Steps 1. Select scientific missions to be included.

2. Check for more advanced versions of chosen hardware.

3. Check for special requirements demanded by scientific hardware (i.e. storage temperature).

4. Locate storage location on SPRITE.

5. Select tools and storage suitable for EVA in I-Suits

Ryan Livingston

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Robotic Extendable Arm with Changeable Heads (REACH)

- Must reach entirety of SPRITE exterior

- Must have 100 kg payload capacity (suits)

- Perform specific science requirements TBD

- At least 6 DOF needed

Requirement M4: SPRITE shall provide the capability for the crew to interact with the local

environment and critical external vehicle systems without EVA

David Gruntz

David Gruntz

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REACH Configuration

• Several configurations considered– Single arm– Two arms (one on each end of rover)– Single arm on track

David Gruntz

David Gruntz

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REACH Material

• Carbon/Epoxy resin ideal choice – Lightweight and strong

Material

Density(kg/m3)

Yield Stress (MPa)

Elasticity(GPa)

Yield Stress/Density Ratio

Comments

Aluminum, wrought, 2024-T4 2800 325 73 0.12

Easy to machine

Titanium alloy, annealed 4460 1230 TBD 0.28

Expensive,Too strong

Carbon/Epoxy resin 1600 800 125 0.50Extremely lightweight

David Gruntz

Beer, Ferdinand. Mechanics of Materials

Werelety, Norman. ENAE423 Lectures - Composite Materials

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Initial Sizing & Mass• Static analysis performed to determine size and mass• 100 kg payload in Martian gravity (3.7 m/s2)

Configuration LengthMass

(per arm)

(kg)

Max Stress(per arm)

(MPa)

MOS

(per arm)

SF = 2

Material Al Resin Al Resin Al Resin

Single Arm three 2 m segments

17 8 150 280 0.10 0.43

Double Arm two 1.5 m segments

9 4 140 350 0.15 0.13

Tracked Arm two 2.5 m segments

28 16 125 330 0.33 0.20

David Gruntz

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Future Work…

• Finalize sizing & workspace• Dynamic analysis• Determine power requirements

• End-effector design– Gripper / Lifter– Shovel / Sample Collector / Drill– Other tools as needed for science/rover ops

David Gruntz

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Contingencies

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Emergency Return Vehicle

• To be used when the crew must return to base without the main rover

• Scenario 1: Rover becomes immobile– Drive system failure– Total electrical power failure

• Scenario 2: Immediate danger to crew– Critical pressure loss to hull– Medical emergency– Life support system failure

• Three options under consideration

Requirement M5: The SPRITE system shall include provision for safe return of the crew following a worst-case SPRITE failure without outside intervention.

Jason West

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• Astronauts leave caches of consumables while driving– In event of emergency, astronauts can walk back to base using caches

along the way for survival– Apollo astronauts completed a 10 km walk in 8 hrs– Separate caches every 10 km with oxygen, water, and food– Astronauts carry a 10 m3 inflatable habitat pressurized at 3.5 psi (same

as suits)– Six-hour rest period at each cache

• Deployment Mechanism– Use robot arm to remove packages from an external container on the

rover and drop them onto lunar surface

Portable Air, Nutrients, and Inflatables Cache (PANIC)

Samuel Schreiber

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PANIC - Habitat

• Habitat composed of space suit-like material for

insulation and pressurization ~ .4 kg/m2

• Habitat is inflated to 3.5 psi of 100% oxygen

• Provides an opportunity for astronauts to remove

space suits, eat, rest, and discard waste

• 10 m3 minimal habitable volume for two 95th percentile

American male astronauts with space suits.

• Reusable - Only one needed throughout return to base

Samuel Schreiber

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Consumable Mass Estimates

• Nominal usage of 0.95 kg/hr water, 0.1 kg/hr oxygen• Total Trip: 182 hrs at maximum distance – 125 km

Walking - 104 hrs; Resting – 78 hrs• 3.2 kg oxygen needed to pressurize habitat at each

stop (only 0.6 kg needed for respiration)• Each cache:

– 7.6 kg water for traverse 7.9 kg Oxygen Tank– 5.7 kg water for rest– 0.8 kg oxygen for traverse 1.3 kg Water Tank– 3.2 kg oxygen for rest

Samuel Schreiber

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PANIC - Mass Estimates• Estimated Total Masses*:

– 26.4 kg in each cache + food + habitat– 344 kg Total + food + habitat

• Habitat Mass: 7 - 12 kg depending upon geometry– Estimate using mass/area of space suit fabric– Only one needed, can be carried.

• Food/Nutrient mass TBD based upon length of return walk– Freeze dried food– Nutrient paste (emergency food supply)

*All consumable masses do not have to be launched with SPRITE - Can be picked up at lunar base

Samuel Schreiber

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PANIC - Concerns and Questions

• Overall reliability and probability of failure

• Astronaut exhaustion, malnutrition and overheating

• Probability of excessive radiation dosage due to solar flare

• Amount of time spent on return – upwards of 8 days

• Carbon dioxide build up in habitat

• Heating

• Oxygen leaks in habitat

• Different rover paths provide differences in difficulty of a walk return

Samuel Schreiber

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Transport Emergency Recovery by Rocket Operated Return (TERROR)

• Used for ballistic return• Rocket attached to panel with

restraints for astronauts• Would travel in a suborbital

trajectory back to base• Astronauts are in their suits• System lands near base and

astronauts walk to the nearest hatch

Timothy WassermanDaniel Zelman

Requirement M5: The SPRITE system shall include provision for safe return of the crew

following a worst-case SPRITE failure without outside intervention

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TERROR - Trajectory

D (km) v (rad) e V0 (km/s) ∆V (km/s) Apogee (km) TOF (min)25 0.0072 0.9928 0.201 0.401 45 5.850 0.0144 0.9857 0.283 0.566 91 8.275 0.0216 0.9787 0.345 0.690 140 10.0

100 0.0288 0.9716 0.397 0.794 191 11.5125 0.0360 0.9647 0.443 0.885 245 12.9

• D – Distance from base• v – Initial true anomaly of return trajectory• e – Eccentricity of return trajectory• V0 – Initial velocity• ∆V – Total delta-V• Apogee – Maximum altitude attained• TOF – Time of Flight

Timothy WassermanDan Zelman

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TERROR - Mass and Volume Estimates

Mass

Fuel 56 kg

Oxidizer 103 kg

Tank (Fuel) 4 kg

Tank (Oxidizer) 5 kg

Pressure Tank 5 kg

Wiring 10 kg

Engine 16 kg

Thrust Structure 1 kg

Avionics 10 kg

Seats 25 kg

Total Mass 224 kg

Volume

Fuel Tank 0.065 m3

Oxidizer Tank 0.065 m3

Engine 0.016 m3

Platform 0.016 m3

Total Volume 1.14 m3

Timothy WassermanDan Zelman

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Foldable Escape Assisting Rover (FEAR)

• Based on the original Apollo Rover• Lighter and Stronger

– New Material– Less Payload

• Higher Clearance– 0.5 m Requirement

• Faster and More Powerful – Newer engines

Laurie Knorr

Requirement M5: The SPRITE system shall include provision for safe return of the crew following a worst-case SPRITE failure without outside intervention

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FEAR - Mass and Material

Aluminum Alloy

2219Carbon Epoxy

Density 2.84 g/cm3 1.6 g/cm3

Tensile Strength 359 MPa 600 MPa

Yield Strength 248 MPa 600 MPa

Modulus of Elasticity 73.1 GPa 70 GPa

Shear Modulus 27 GPa 5 GPa

Shear Strength 230 MPa 90 MPa

Laurie Knorr

Aerospace Specification Metals Inc - <http://asm.matweb.com/search/SpecificMaterial.asp?bassnum=MA2219T37>

Goodfellow - <http://www.azom.com/details.asp?ArticleID=1995>

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FEAR - Height Change

• Increase the size of the wheels–Mass of new wheel would be 1.69 times the mass of old wheel if the diameter is increased by 20 cm

• Change the suspension–Mass increase minuscule

–Small loss in strength

Laurie Knorr

LRV Operations Handbook, 1973 Contract NASA-25145

LowerArm

Chassis Fitting

Upper Arm

Damper

Lunar Rover Operations Handbook - <http://www.hq.nasa.gov/alsj/irvhand.html>

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FEAR - Motors

• Four motors: One on each wheel• Old motors

–36 V Input–0.25 hp Power

–10,000 rpm

• New motors TBD–Lighter–More powerful

Laurie Knorr


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FEAR - Special Features

• Drive back to base in less than 10 hours

• Folds up into 0.9 m3 space

• Attaches to the outside of SPRITE

Laurie Knorr


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Emergency Return Operations

ERV Mass Mass Launched Volume

PANIC 344 kg* 13 kg TBD

TERROR 224 kg 76 kg 1.14 m3

FEAR 240 kg 210 kg 0.9 m3

Laurie Knorr

* Does not include food

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Safety Time Advantages Disadvantages

PANIC 182 hr

• Fairly simple

• Can be used in conjunction with other safety procedures

• Takes time to walk

• Very tiring on crew

• Increased probability of solar flare exposure

TERROR 13 min• Fast return to base

• Can work if one crew member is injured

• Very unsafe

• Complicated system

FEAR 10 hr• Crew exerts little energy

• Can work if one crew member is injured

• Complicated detachment procedures

Laurie Knorr

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Program Timeline and Costs

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Program Timeline

• Development/Production: 2005-2015• Launch: 2016

• Current Plan - 3-month program cycle – All costs will be calculated for a 3-month program– 6 SPRITE sorties will be completed during program

Charles Bacon

Requirement M8: The SPRITE shall be ready for initial lunar operations by 2016

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Sample 1-Month Timeline*• Day 1: Launch • Day 6: Lunar Landing• Day 8-14: 1st Sortie• Day 15-21: Prepare SPRITE for 2nd sortie

– Analyze Data Collected

• Day 22-28: 2nd Sortie• Day 29-35: Prepare SPRITE for 3rd Sortie

– Analyze Data Collected

Charles Bacon

**assumes 1 SPRITE Vehicle, 5 day trip to moon

*Timeline would repeat (except launch) approximately each month for a period of 3 months

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Program Timeline

• Deviations in this timeline could occur if an additional SPRITE vehicle is launched

• Plan TBD if 2 SPRITE’s are on the Moon– Both could be used to run normal missions

Charles Bacon

Requirement I12: The SPRITE design shall provide the necessary capabilities and interfaces for one SPRITE vehicle to tow a second inactive SPRITE 100 km to base for repairs

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Cost Analysis• No specified limitations for cost budget• Heuristics from NASA Cost Estimation site:

– C(FY04 $M)= ami[kg]b*

• Manned Spacecraft (SPRITE)– a = 20.738, b = .556

• Liquid Rocket Engine (TERROR, landing engine)– a = 32.391, b = .551

– Other system cost estimates derived uniquely for each system

Charles Bacon

*Derived from NASA Cost Models

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Other Sources of Cost

• Emergency Recovery Vehicles– FEAR – Very similar to original Apollo rover, cost of that was

converted to 2004 dollars using NASA Inflation Calculator

– PANIC – End product should be relatively low, development costs are still unknown

• Robotic Arms – Averaged from costs of different robotic arms already available

• Landing Structure• Delta IV Heavy Launch $254 Million (2004)

Charles Bacon

- Larson, Pranke Human Spaceflight: Analysis and Design, pg 755, table 23-10

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Cost TotalsSystem Cost (FY 04 $M)

SPRITE Vehicle (1) 1940

Landing Engine

(1-stage LOX/LH2)1460

Landing Structure TBD

FEAR 144

PANIC TBD

TERROR 320

Robotic Arm 185

Launch 254

Charles Bacon

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Cost Analysis

• Current total – $4.1 Billion– Estimated final cost to launch: 1 SPRITE + 1 ERV

• Worst Case Scenario – TERROR: most expensive– Cost will increase for another SPRITE, but not

significantly (production is only 2-6% of total cost)

• Other costs include consumables and fuels (relatively low cost)

Charles Bacon

Requirement I6: The SPRITE design shall be designed to minimize life cycle costs

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Mission Operation and Data Analysis Cost

• Mission Operational Costs - $154M/yr*– Includes

• maintaining and upgrading ground systems, mission control;• tracking; telemetry; command functions; mission planning;

data reduction and analysis; crew training and related activities

Charles Bacon

*assume investment price - $3.9B

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Cost Spreading

• Development and Production would occur from 2005-2015– Launch in 2016

• Beta Function – Non-Recurring costs over 11 years– Recurring Costs take over in 2016.

Charles Bacon

Requirement M8: The SPRITE shall be ready for initial lunar operations by 2016

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Cost Spreading

SPRITE's Estimated Annual Expenditures

0

100

200

300

400

500

600

2005 2006 2007 2008 2009 2010 2011 2012 2013 2014 2015

Year

FY

04

($M

)

Charles Bacon

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Cost Analysis

• NASA’s Advanced Missions Cost Model estimates the cost of SPRITE to be about $6 Billion….this is still more than we have already, but there is still more work to be done

Charles Bacon

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Systems Integration Future Work…

• Thorough itemized analysis for SPRITE to result in a reasonable projected cost

• Work breakdown timeline (2005-2015) to illustrate key systems, milestones, and deliverables with projected due dates

• Costs of major systems still unknown• Create System Block Diagrams

Charles Bacon

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spritesprite 1 of 256 university of maryland enae484 pdr march 14, 2005 preliminary design review...

Documents

exploration slide

sprite vehicle

single delta

heavy vehicle

lunar base

tli daren mcculley slide

moon chris hartsough

lunar orbit insertion