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Small Pressurized Rover for
Independent Transport and
Exploration
Preliminary Design Review
March 14, 2005
E
University of Maryland
ENAE484 PDR March 14, 2005
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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
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Charles Bacon
University of Maryland
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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).
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Charles Bacon
University of Maryland
ENAE484 PDR March 14, 2005
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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.
E
Charles Bacon
University of Maryland
ENAE484 PDR March 14, 2005
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Launch to Landing
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University of Maryland
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CONOPS Overview
P
(Delta-IV Heavy Separation to Landing)
R
Separate from Delta-IV Heavy
Perform lunar orbit insertion burn
I
Perform descent orbit insertion burn
T
Perform powered descent burn
E
Land on the Moon
Chris Hartsough
University of Maryland
ENAE484 PDR March 14, 2005
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Orbit Design Objectives
Requirement M7: The SPRITE vehicle shall be capable of independent deployment
to the lunar surface with a single Delta IV Heavy class launch
• 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)
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Daren McCulley
University of Maryland
ENAE484 PDR March 14, 2005
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Translunar Orbit Options
•
– Advantage
•
•
•
•
•
I
E
Unmanned and mass constrained mission
– Disadvantages
R
T
Low Thrust
•
•
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
University of Maryland
ENAE484 PDR March 14, 2005
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Depiction of Translunar Orbits
Low Thrust Transit
High Thrust TLI
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http://sbir.gsfc.nasa.gov/SBIR/successes/ss/5-075text.html
Daren McCulley
University of Maryland
ENAE484 PDR March 14, 2005
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Apogee of Translunar Orbit
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Radius of Apogee
TLI ∆V
R
356,000 km
Moon at Perigee
3.128 km/s
407,000 km
Moon at Apogee
3.140 km/s
I
T
* Additional DV of only 12 m/s
* Additional day of flight time
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Daren McCulley
University of Maryland
ENAE484 PDR March 14, 2005
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Payload vs. Apogee
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I
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6000
Separation Mass [kg]
R
Sample Delta IV Performance Curve
E
5000
4000
3000
2000
1000
0
0
20
40
60
80
100
Altitude of Apogee [x 1e3 km]
Delta-IV Payload Planners Guide
Daren McCulley
University of Maryland
ENAE484 PDR March 14, 2005
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Selenocentric Orbits
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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
Daren McCulley
University of Maryland
ENAE484 PDR March 14, 2005
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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
University of Maryland
ENAE484 PDR March 14, 2005
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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
University of Maryland
ENAE484 PDR March 14, 2005
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Descent Orbit Analysis
Altitude 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
I
20
.0183
1.688
.2557
1.962
30
.0159
1.681
.3132
2.010
T
40
.0136
1.674
.3617
2.050
50
.0113
1.667
.4044
2.083
R
E
Daren McCulley
University of Maryland
ENAE484 PDR March 14, 2005
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S Insertion and Landing Concept
P • Lunar Orbit Insertion (LOI)
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– 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
University of Maryland
ENAE484 PDR March 14, 2005
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Depiction of Selenocentric Orbits
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Daren McCulley
University of Maryland
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3D Orbit Design In Reverse
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Daren McCulley
University of Maryland
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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.
E
Chobotov, Vladimir. Orbital Mechanics
Daren McCulley
University of Maryland
ENAE484 PDR March 14, 2005
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Dynamic Simulations
• Translunar Injection Simulation
– Controllable variables (Time of TLI, w)
– 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
University of Maryland
ENAE484 PDR March 14, 2005
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Analysis of Control Variables
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Daren McCulley
all axes in km
Daren McCulley
University of Maryland
ENAE484 PDR March 14, 2005
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Powered Descent Simulation
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km
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km
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Daren McCulley
km
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km
Daren McCulley
University of Maryland
ENAE484 PDR March 14, 2005
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PDL Simulation Results
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Burn Altitude: 16.2 km
R
Burn Time: 283.5 s
I
Residual Velocities:
Negligible
Velocity (km/s)
T
Thrust: 42.9 kN
Height: 4 m
Landed Mass: 5435 kg
E
Max Acc: 6.3 m/s2
Time after Aposelenium (s)
Daren McCulley
∆V: 1.83 km/s
University of Maryland
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Burn Profile
Burn / Maneuver
Engine
∆V (km/s)
∆M (kg)
PL Fairing Evasion
RCS
Negligible
Negligible
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
T
Descent Orbit Insertion
RETRO
0.02
40
RCS
Negligible
Negligible
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Powered Descent
RETRO
1.89
2860
R
I
Delta IV SS TLI
Descent Orbit Correction
Daren McCulley
University of Maryland
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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
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Aaron Shabazz
University of Maryland
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P • Critical GNC Hardware
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– 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
E
Aaron Shabazz
University of Maryland
ENAE484 PDR March 14, 2005
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P • IMU accuracy is vital to mission success
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– 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
* Data from Honeywell IMU spec sheet
Spec Sheet - http://content.honeywell.com/dses/assets/datasheets/fog.pdf
Aaron Shabazz
University of Maryland
ENAE484 PDR March 14, 2005
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P • Attitude Control System
R
I
– 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
T
E
Aaron Shabazz
University of Maryland
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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
T
E
Mike Sloan
Mike Sloan
University of Maryland
ENAE484 PDR March 14, 2005
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Center of Gravity - Landing
• 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
T
E
Mike Sloan
Mike Sloan
University of Maryland
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Center of Gravity - Landing
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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
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Mike Sloan
University of Maryland
ENAE484 PDR March 14, 2005
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Center of Gravity - Driving
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Requirement I9: SPRITE shall be able to actively traverse terrain safely with
20o cross-slope and 30o direct slope
• 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
T
E
Mike Sloan
University of Maryland
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Center of Gravity - Driving
• Mars CGrequired height >
Moon CGrequired height
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• Any vehicle geometry that
can safely drive on the Moon
can safely drive on Mars
Mike Sloan
University of Maryland
ENAE484 PDR March 14, 2005
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Transit Configuration 1
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Mike Sloan
Daren McCulley
University of Maryland
ENAE484 PDR March 14, 2005
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Transit Configuration 1
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Mike Sloan
Daren McCulley
University of Maryland
ENAE484 PDR March 14, 2005
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Transit Configuration 1
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Mike Sloan
Daren McCulley
University of Maryland
ENAE484 PDR March 14, 2005
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Transit Configuration 2
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Mike Sloan
Daren McCulley
University of Maryland
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Transit Configuration 2
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Mike Sloan
Daren McCulley
University of Maryland
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S Propulsion System Requirements
P • Launch a specified payload to the moon
R
I
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E
• 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
University of Maryland
ENAE484 PDR March 14, 2005
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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):
• PDL (hover):
1792 m/s
60.96 m/s
– PDL - Powered Descent Landing
– RCS Thrusters
• RCS (landing): 150 m/s
– RCS - Reaction Control System
Reuel Smith
University of Maryland
ENAE484 PDR March 14, 2005
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Assumptions Made
P • Propellants: The module runs on one
specific
fuel/oxidizer
mixture
R
I
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E
• Other Assumptions
–
–
–
–
–
Payload:
Ae/At:
Inert Mass Fraction:
Max RCS Thrust:
RCS Thruster Count: 16
3790 kg
54 for all propulsion stages
0.08 for all propulsion stages
445 N per thruster
Spacecraft Apollo- <http://www.braeunig.us/space/specs/apollo.htm>
Reuel Smith
University of Maryland
ENAE484 PDR March 14, 2005
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Fuel Analysis
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MLanding Engine
+ MRCS Thrusters
+ MGimbals
+ MAvionics
+ MWiring
+ MThrust Structure
__________________
MPropellant System
Reuel Smith
g
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
Propellant
University of Maryland
ENAE484 PDR March 14, 2005
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Fuel Analysis
P
720
710
700
690
680
670
660
650
640
Propellants
Reuel Smith
H
DM
M
H
University of Maryland
ENAE484 PDR March 14, 2005
/U
TO
N
N
TO
/M
P1
LO
X/
R
yd
ra
z
in
2
LO
X/
H
en
ro
s
LO
X/
Ke
E
e
630
e
T
730
LO
X/
LH
I
Total propulsion system mass (kg)
R
740
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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
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Reuel Smith
University of Maryland
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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
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Jason West
University of Maryland
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RCS Thruster Risk Analysis
• Scenario A: 12 engines, none fail
• Scenario B: 16 engines, up to 2 engines can fail
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Required Engine
Reliability
Scenario A
Scenario B
0.9957
0.9469
• 5% less required engine reliability for 16-engine
system
Jason West
University of Maryland
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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
E
Reuel Smith
University of Maryland
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Landing Requirements
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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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Rahkiya Medley
University of Maryland
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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
University of Maryland
ENAE484 PDR March 14, 2005
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Lander Options
• Crushable legs
–Honeycomb insert
–Pivot feet
–80 kg/leg (Al wrought 2024T4 and SPIRALGRIDTM)
• Joint legs
–Torsion spring joint
–Pivot feet
–TBD kg/leg
Rahkiya Medley
University of Maryland
ENAE484 PDR March 14, 2005
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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
E
Rahkiya Medley
University of Maryland
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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
E
Rahkiya Medley
University of Maryland
ENAE484 PDR March 14, 2005
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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
E
Dr David Smith, Goddard Space Flight Center
Mike Sloan
University of Maryland
ENAE484 PDR March 14, 2005
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S Guidance Navigation & Control
P • Landing Control System
R
I
T
– 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
E
Aaron Shabazz
University of Maryland
ENAE484 PDR March 14, 2005
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S Guidance Navigation & Control
P • Landing Point Localization
R
I
T
– 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
E
Aaron Shabazz
University of Maryland
ENAE484 PDR March 14, 2005
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S Guidance Navigation & Control
P • Landing Point Localization
R
– Rough Estimate Landing Accuracy
• Average all off-target data after Apollo 12
I
T
Off target data
Apollo 12
Apollo 14
Apollo 15
Apollo 16
Apollo 17
0.16 km
0.05 km
0.21 km
0.20 km
0.55 km
Off target data – Spring 2004 ENAE 484 CDR Slide # 239
E
Estimate landing accuracy = 0.234 km
Aaron Shabazz
University of Maryland
ENAE484 PDR March 14, 2005
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P • Distance Between Landing
R
I
T
E
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
University of Maryland
ENAE484 PDR March 14, 2005
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P
R
I
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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)
E
Joanneum Research: Vision-Based Navigation for Moon Landing
Scott Walthour
University of Maryland
ENAE484 PDR March 14, 2005
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Stereo From Lander Motion
P
R
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Scott Walthour
Scott Walthour
University of Maryland
ENAE484 PDR March 14, 2005
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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)
<http://www.laseroptronix.com>
Scott Walthour
University of Maryland
ENAE484 PDR March 14, 2005
Digital Elevation Map (image source: http://qso.lanl.gov)
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S Hazard Detection Performance*
CCD Array
512 x 512 pixels
P
Focal Length
10 mm
Footprint (200 m)
100 m
I
Ground and DEM Resolution
0.2 m
Required Pointing Accuracy
1.4 deg
T
Processing Time
~ 10 to 30 sec
Required Inertial Sensing Accuracy
(90% overlap)
10 m
R
E
Joanneum Research: Vision-Based Navigation for Moon Landing
*From similar lunar mission (2 hr orbital period, 0.5m obstacle requirement)
Scott Walthour
University of Maryland
ENAE484 PDR March 14, 2005
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Lander Stereo Considerations
P
• Hovering could cause
errors in inertial
navigation
R
–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)
I
T
E
Scott Walthour
Scott Walthour
Joanneum Research: Vision-Based Navigation for Moon Landing
University of Maryland
ENAE484 PDR March 14, 2005
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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)
E
Timothy Wasserman
University of Maryland
ENAE484 PDR March 14, 2005
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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
Timothy Wasserman
E
LOX/CH4
2530
LOX/RP-1
2360
Timothy Wasserman
University of Maryland
ENAE484 PDR March 14, 2005
65 of 256
S
P
R
I
T
E
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 kg
Timothy Wasserman
Timothy Wasserman
University of Maryland
ENAE484 PDR March 14, 2005
66 of 256
S
Two Stages: Reuse Cryogenic Tanks
P
R
I
T
E
• 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
University of Maryland
ENAE484 PDR March 14, 2005
Timothy Wasserman
67 of 256
S
Comparison of Best Two Staging Options
P
R
I
T
E
Option
Rover Mass (kg)
LOX/LH2 Single Stage
3000
LOX/LH2 First Stage
LOX/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
Akin, David. ENAE 483 Lecture on Mass Estimating Relationships
Fuel Properties from: www.astronautix.com
Timothy Wasserman
University of Maryland
ENAE484 PDR March 14, 2005
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S
Launch Mass Budget
P
R
I
T
E
Design Group
Transit Power, Propulsion & Thermal
Surface Power, Propulsion & Thermal
Loads, Structures & Mechanisms
Crew Systems
Mission Planning & Analysis
Avionics
Timothy Wasserman
University of Maryland
ENAE484 PDR March 14, 2005
Mass
5800
700
1250
700
300
300
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P
R
I
Surface Operations
T
E
University of Maryland
ENAE484 PDR March 14, 2005
70 of 256
S Mission Planning Requirements
P
R
I
T
Requirement 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
E
Chris Hartsough
University of Maryland
ENAE484 PDR March 14, 2005
71 of 256
S
P
CONOPS Overview
(Deployment to Nominal Operations)
R
Deploy from landing system
Autonomous return
I
Remote operated return
Dock with base
T
Pre-mission check of systems
E
Supply SPRITE with consumables and fuel
Chris Hartsough
University of Maryland
ENAE484 PDR March 14, 2005
72 of 256
S
P
R
I
T
CONOPS Overview
(Nominal Mission)
Requirement M2: SPRITE shall be designed to carry 2 crew on a normal sortie of 7 days
covering 250 km
• 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
E
*Possible robotic arm operations everyday
Chris Hartsough
University of Maryland
ENAE484 PDR March 14, 2005
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S
P
R
I
T
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
E
Daniel Zelman
Loop A
Daniel Zelman
University of Maryland
ENAE484 PDR March 14, 2005
Loop B
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S
P
R
I
T
E
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
Daniel Zelman
–Never more than 100 km
from base
• More scientific
possibilities than previous
routes
Daniel Zelman
Arc
University of Maryland
ENAE484 PDR March 14, 2005
Inverted Arc
75 of 256
S
P
R
I
T
E
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 Sloan
Daniel Zelman
University of Maryland
ENAE484 PDR March 14, 2005
76 of 256
S
P
R
I
Structures
T
E
University of Maryland
ENAE484 PDR March 14, 2005
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S
P
R
I
T
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
E
Evan Ulrich
University of Maryland
ENAE484 PDR March 14, 2005
S
P
R
I
T
E
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
University of Maryland
ENAE484 PDR March 14, 2005
78 of 256
S
Prolate Spheroid: Analysis
•
Applied Loads:
•
P
79 of 256
Internal Pressure (2 atm)
Punching force (3 kN)
Method of Analysis:
•
•
R
•
I
•
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
T
E
Component O.D (m) I.D (m) Length (m) Mass (Kg) Design Load (Mpa) S.M S.F Material Failure mode
Stringer
0.084
0.083
4.8
24
380 0.0 2 Ti-6Al-4V Compression
Rib/stiffener inner0.115
0.114
6.6
4
380 0.0 2 Ti-6Al-4V Bending
Rib/stiffener outer
0.075
0.072
4.2
6
380 0.0 2 Ti-6Al-4V Bending
Skin (4mm)
4.808
4.800
639
550
2 Ti-6Al-4V local buckling
Tota Mass (Kg)
673
Evan Ulrich
University of Maryland
ENAE484 PDR March 14, 2005
80 of 256
S
Cylinder with Ellipsoidal Endcaps (CEE)
P
•
Optimal number of ribs is 4
•
Need for external mounts may
increase number of ribs
R
•
8 stringer configuration allows for
ease of hatch/window placement
•
Provides sufficient structural support
I
•
All stringers have hollow circular cross
sections
T
E
2m
2m
-Shear panel
-Stringer
-Rib/Stiffener
Evan Ulrich
University of Maryland
ENAE484 PDR March 14, 2005
S
P
R
I
•
Applied Loads:
•
CEE: Analysis
Internal Pressure (2 atm)
Punching force (3 kN)
1.2e+8
1.0e+8
6.2e+7
3.2e+7
2.0e+7
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
4.2e+8
3.6e+8
2.0e+8
1.1e+8
5.8e+7
T
E
81 of 256
Component Length (m)
Stringer
4.0
Rib/stiffener inner
6.3
Rib/stiffener outer
6.3
Skin (4mm)
Total Mass
Evan Ulrich
Mass (Kg) Design Load (Mpa) S.F
Material Failure mode
65
420
2 Titanium Ti-6Al-4V Compression
TBD
420
2 Titanium Ti-6Al-4V Bending
TBD
420
2 Titanium Ti-6Al-4V Bending
851
120
2 Titanium Ti-6Al-4V local buckling
916
University of Maryland
ENAE484 PDR March 14, 2005
82 of 256
S
P
R
I
T
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
E
Michael Koszyk
University of Maryland
ENAE484 PDR March 14, 2005
83 of 256
S
P
R
I
T
Micrometeoroid Protection
Micrometeoroid Flux vs. Mass
• 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
E
[Vanzani, et al. Micrometeoroid Impacts on the Lunar Surface. Lunar and Planetary Science XXVIII, 1997.]
Michael Koszyk
University of Maryland
ENAE484 PDR March 14, 2005
84 of 256
P
R
I
Micrometeoroid Protection
• Design
variables
–Hull properties
–MLI properties
–Hull/MLI
spacing
T
Critical Micrometeoroid Mass vs Hull/MLI Spacing
0.0006
0.0005
6 mm hull thickness
0.0004
Mass (g)
S
5 mm hull thickness
4 mm hull thickness
0.0003
3 mm hull thickness
2 mm hull thickness
0.0002
Critical Design Mass
0.0001
E
0.0000
0
0.01
0.02
0.03
0.04
Spacing (m)
Michael Koszyk
University of Maryland
ENAE484 PDR March 14, 2005
0.05
85 of 256
S
Window Materials
P
R
I
T
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 HighStrength
2010
38.3
56.2
207
0.5
Castable 220
2090
-
11.35
50
1.7
Material
Ceradyne Thermo-Sil® Fused Silica Materials <http://www.ceradyne.com>
E
Michael Koszyk
University of Maryland
ENAE484 PDR March 14, 2005
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S
P
R
I
Window Requirements
• Curvature of material required
• Filter out harmful radiation
– 0.1% Iron Oxide fused into glass
• Anti-reflective coating necessary
• Structural analysis underway
T
E
Michael Koszyk
University of Maryland
ENAE484 PDR March 14, 2005
Other Required
Structures/Mechanisms
S
P
R
I
T
E
• 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
David Gruntz
University of Maryland
ENAE484 PDR March 14, 2005
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88 of 256
Structures Summary
S
P
Structure
Primary
R
SF
2
E
Structure
T
Secondary
I
1.5
3
Loading
Condition
Applied Load
(MPa)
MOS
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
Avionics Support
Structure
TBD
TBD
TBD
Thermal Regulation
Support Structure
TBD
TBD
TBD
Pressure Vessels
TBD
TBD
TBD
Structure
•Launch – 6g axially along Delta IV, 2.5g laterally
•(I) – impulse load
•(PL) – point load
David Gruntz
Rahkiya Medley
University of Maryland
ENAE484 PDR March 14, 2005
89 of 256
S
P
R
I
Mobility Systems
T
E
University of Maryland
ENAE484 PDR March 14, 2005
90 of 256
S
P
R
I
T
E
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
Raja Krishnamoorthy
University of Maryland
ENAE484 PDR March 14, 2005
91 of 256
S Surface Propulsion Calculation
P
• Calculate the frictional forces due to
tire roll based on
Ff = [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
R
I
• Calculate force of gravity on incline of
30° (for peak power)
T
–Maximum load is the sum of friction on tires
and normal force
–Meets Level 1 requirement (I9)
E
Raja Krishnamoorthy
Raja Krishnamoorthy
University of Maryland
ENAE484 PDR March 14, 2005
92 of 256
S
P
R
I
T
E
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
Raja Krishnamoorthy
University of Maryland
ENAE484 PDR March 14, 2005
93 of 256
S
P
R
I
T
E
Drive System Requirements
Assumptions
Calculations
Engine Efficiency (%)
Diameter of Wheel
Width of Wheel
Weight on Each Wheel
m
ft
m
ft
N
lb
92 Engine Requirements
1.4
Total Torque Req
4.59
Torque per Wheel
0.30
0.98
Total Power Req
2041.3
Power per Wheel
459.28
Units
Nm
lb-ft
Nm
lb-ft
kW
hp
kW
hp
Peak Continuous
8169
6025
2042
1506
55.12
73.86
13.32
17.85
5311.46
3917.20
1327.86
979.30
36.48
48.89
8.66
11.60
• 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
Raja Krishnamoorthy
University of Maryland
ENAE484 PDR March 14, 2005
94 of 256
S
P
R
I
T
E
AC vs. DC Motor
Power (kW)
RPM
DC Motor Types
15
125
329
560
2000
2000
2000
1500
DMP112-4L
DMP180-4LB
DMI225S
DMA+315M
Power (kW)
RPM
AC Motor Types
15
125
329
560
2000
2000
2000
1500
180M4
315SMA4
355SMA4
450LG4
Mass Moment
Ramp-up
of Inertia
Time (s)
(kgm 2) - DC
0.05
0.619
0.69
1.15
3
1.73
10.68
1.57
Mass Moment
Ramp-up
of Inertia
Time (s)
(kgm 2) - AC
0.161
0.946
2.3
1.73
8.2
2.42
25
2.2
Data from DC or AC Drives? A guide for users of variable-speed drives
• Further analysis to be done with a wider range of motors
Raja Krishnamoorthy
University of Maryland
ENAE484 PDR March 14, 2005
95 of 256
S
AC - DC Mass Comparison
P
Mass vs Power
4500
4000
I
3000
Weight (kg)
R
3500
2500
DC
AC
2000
1500
1000
T
E
500
0
0
100
200
300
400
500
600
Power (kW)
Data from DC or AC Drives? A guide for users of variable-speed drives
• For a 15 kW motor the masses are as follows:
AC – 175 kg
DC – 110 kg
Engines studied: DMP112-4L, DMP180-4LB, DMI225S, DMA315M, 180M4,
315SMA4, 355SMA, 450LG4
Raja Krishnamoorthy
University of Maryland
ENAE484 PDR March 14, 2005
96 of 256
S
P
R
I
T
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)
E
Raja Krishnamoorthy
University of Maryland
ENAE484 PDR March 14, 2005
97 of 256
S Axle vs. Individual Wheel Drive
4 MOTOR
2 MOTOR
P • A motor for each wheel
• A motor for each axle
R
I
T
E
–Used 4-wheel case
–Used 2-axle case
• Requires less power
• Provides less torque
• Requires more power
• Provides more torque
# of Motors
Power per Motor Continuous
Peak
Torque per Motor Continuous
Peak
Raja Krishnamoorthy
4 motor
kW
hp
13.00
17.42
25.00
33.50
Nm
lb-ft
1327.86
979.30
2042.30 1506.20
University of Maryland
ENAE484 PDR March 14, 2005
2 motor
kW
hp
35.00
46.90
55.00
73.70
Nm
lb-ft
2655.73 1958.60
4084.60 3012.40
98 of 256
S
Risk Analysis for Drive Setups
P
4 MOTOR
• Can tolerate 2 failures*:
4 ways ↔ A-C, B-D, A-D, C-B
R
• R4 + 4R3(1-R) + 4R2(1-R)2
2 MOTOR
I
• Can tolerate only 1 failure:
A or B
T
• R2 + 2R(1-R)
E
R = e-t/MTBF = 0.999375
t = 25 hrs, MTBF = 40,000 hrs
*Considered simple failure without wheel lock
Raja Krishnamoorthy
University of Maryland
ENAE484 PDR March 14, 2005
99 of 256
S
P
R
I
T
E
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
Raja Krishnamoorthy
University of Maryland
ENAE484 PDR March 14, 2005
100 of 256
References
S
P
• DC or AC Drives? A guide for users…
R
• Motor Formulas, 1997
I
T
– https://www.abb-drives.com/StdDrives/RestrictedPages/Marketing/
Documentation/Documents/DCorAC.pdf
– 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
E
Raja Krishnamoorthy
University of Maryland
ENAE484 PDR March 14, 2005
101 of 256
S
P
R
I
T
E
Wheels
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
Requirement 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
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
Pyungkuk Choi
University of Maryland
ENAE484 PDR March 14, 2005
102 of 256
S
Power vs. Width
P
I
T
E
120
100
Power(kW)
R
Power vs. Width (30 degree incline)
80
4 wheels
6 wheels
60
40
20
0
0
0.2
0.4
0.6
0.8
Width(m)
• Width = 0.3 m
Pyungkuk Choi
University of Maryland
ENAE484 PDR March 14, 2005
1
1.2
103 of 256
S
P
R
I
T
E
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
Pyungkuk Choi
University of Maryland
ENAE484 PDR March 14, 2005
104 of 256
S
P
R
I
T
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
E
Pyungkuk Choi
Pyungkuk Choi
University of Maryland
ENAE484 PDR March 14, 2005
105 of 256
S Number of Spokes vs. Rim Thickness
P
R
I
T
E
• 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
Pyungkuk Choi
University of Maryland
ENAE484 PDR March 14, 2005
106 of 256
S
P
R
I
Wheels - Future Work
•
•
•
•
Tires
Cross slope loading
Different wheel configuration
Wheel protection
T
E
Pyungkuk Choi
University of Maryland
ENAE484 PDR March 14, 2005
107 of 256
S
Chassis / Suspension
P
R
Chassis
I
Spring / Shock Absorber
Wheel
Mount
T
E
• Struts connect to rib/stringer primary structure
– External chassis if necessary
David Gruntz
University of Maryland
ENAE484 PDR March 14, 2005
108 of 256
S
P
R
I
T
E
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)
David Gruntz
University of Maryland
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Suspension Models
P
Lateral
Torsion Bar
Linear Spring
Axial
Torsion Bar
R
I
T
E
David Gruntz
•Modeled as
5,000 kg mass
atop a linear
spring
David Gruntz
•Modeled as
5,000 kg mass
attached to a
2 m moment arm
•Modeled as
5,000 kg mass
attached to a
0.25 m moment
arm
University of Maryland
ENAE484 PDR March 14, 2005
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Torsion Bar vs. Linear Spring
P
180
160
R
Transmitted Force (kN)
140
I
T
120
Linear
100
Lateral
80
Axial
60
40
20
0
E
0
0.1
0.2
0.3
0.4
0.5
0.6
Displacement (m)
•
•
Torsion bars transfer similar loads
Linear spring looks like ideal choice at this point
David Gruntz
University of Maryland
ENAE484 PDR March 14, 2005
0.7
111 of 256
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Loads Transferred to Chassis
P
R
Type
Vertical
“Fall”
Landing
Displacemen Force Force Deflection
t (m)
(kN)
(kN)
(m)
Spring
Constant
I
Linear
0.1
0.2
0.3
45
26
20
36
21
16
0.08
0.18
0.27
450 kN/m
120 kN/m
60 kN/m
T
Lateral
Torsion
0.1
0.2
0.3
100
55
45
51
40
32
0.14
0.16
0.23
1500 kN-m/rad
1000 kN-m/rad
550 kN-m/rad
E
Axial
Torsion
0.1
0.2
0.3
85
60
40
74
50
35
0.1
0.15
0.27
50 kN-m/rad
20 kN-m/rad
8 kN-m/rad
David Gruntz
University of Maryland
ENAE484 PDR March 14, 2005
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Initial Suspension Sizing
S
P
R
•
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
I
Load Condition
Max Stress
(MPa)
MOS
T
Collision
330
0.03
“Fall”
Landing
Launch*
280
200
TBD
0.2
0.7
TBD
E
* Will depend on how rover is integrated w/ fairing
David Gruntz
University of Maryland
ENAE484 PDR March 14, 2005
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Crew Systems
T
E
University of Maryland
ENAE484 PDR March 14, 2005
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Consumables Summary
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
• 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.
John Mularski
University of Maryland
ENAE484 PDR March 14, 2005
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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
E
Larson, Wiley J., ed. Human Spaceflight: Mission Analysis and Design.
John Mularski
University of Maryland
ENAE484 PDR March 14, 2005
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P
Atmospheric Composition
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
SPRITE Rover
R
•8.3 psi total pressure
•37% Oxygen
I
T
E
•63% Nitrogen
<http://www.smallartworks.ca/PS/Space1999/AlphaMoonbase/AlphaMoonbase.html>
Lunar Base
Alan Bean - <www.alanbeangallery.com/ab-artist.html> &
www.andrew.cmu.edu/user/jplee/miscellaneous/new%20sprite%20bottles.jpg
•14.7 psi total pressure
EVA Suit
•21% Oxygen
•3.5 psi total pressure
•79% Nitrogen
John Frassanito and Associates – <http://msnbc.msn.com/id/5990828>
•100% Oxygen
Larson, Wiley and Linda Pranke, ed. Human Spaceflight: Mission Analysis and Design
Michael Badeaux
University of Maryland
ENAE484 PDR March 14, 2005
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Atmospheric Composition
P
R
I
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Man-Systems Integration Standards – NASA-STD-3000 <http://msis.jsc.nasa.gov/>
Base – 14.7 psi
SPRITE – 8.3 psi
EVA – 3.5 psi
21% Oxygen
37% Oxygen
100% Oxygen
Larson, Wiley and Linda Pranke, ed. Human Spaceflight: Mission Analysis and Design
Michael Badeaux
University of Maryland
ENAE484 PDR March 14, 2005
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Storage of Consumables
Requirement L2: All crew interfaces shall adhere to NASA-STD-3000, Man-Systems
Integration Standards
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
Akin, David. ENAE483 Lectures Fall 2004 <http://spacecraft.ssl.umd.edu/academics/483F04
Glatt, C.R. “WAATS – A Computer Program for Weights Analysis of Advanced Transportation Systems.” NASA CR-2420. Aerospace Research Corporation
Michael Badeaux
University of Maryland
ENAE484 PDR March 14, 2005
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Temperature/Humidity Control
Requirement L2: All crew interfaces shall adhere to NASA-STD-3000, Man-Systems
Integration Standards
• Ideal Temperature ranges from 18-27 oC
–SPRITE Cabin Temperature – 23 °C
• Ideal Humidity ranges from 4-16 oC
I
T
E
Wieland, Paul. Designing for Human Presence in Space NASA RP-1324 - <http://flightprojects.msfc.nasa.gov/book/rp1324.pdf>
• Excess heat can be used to heat water
Michael Badeaux
University of Maryland
ENAE484 PDR March 14, 2005
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Temperature Control
P
R
Insulating Materials
Passive
Simple
Small Scale
Little Maintenance
I
T
Electric Heaters
Heat Pipes
Cold Plates
Active
E
Complex
Large Scale
High Maintenance
Heat Exchangers
Re-router
Heat Rejection
Freudenrich, Craig “How Space Stations Work” - <http://science.howstuffworks.com/space-station4.htm>
Michael Badeaux
University of Maryland
ENAE484 PDR March 14, 2005
121 of 256
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P
Carbon Dioxide Removal
Requirement L2: All crew interfaces shall adhere to NASA-STD-3000, Man-Systems
Integration Standards
Removal
R
I
T
E
Regenerable
Reduction
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
•Eckart, Peter. Spaceflight Life Support and Biospherics. Torrance, California: Kluwer Academic, 1994.
*EDC and LiOH have best overall qualifications for SPRITE
Shawn Butani
University of Maryland
ENAE484 PDR March 14, 2005
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P•
R
I
T
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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
Shawn Butani
University of Maryland
ENAE484 PDR March 14, 2005
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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
E
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
University of Maryland
ENAE484 PDR March 14, 2005
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Caution & Warning System
Requirement L2: All crew interfaces shall adhere to NASA-STD-3000, Man-Systems
Integration Standards
• 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
Michael Badeaux
University of Maryland
ENAE484 PDR March 14, 2005
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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>
Michael Badeaux
University of Maryland
ENAE484 PDR March 14, 2005
126 of 256
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Acoustic Environment
Requirement L2: All crew interfaces shall adhere to NASA-STD-3000, Man-Systems
Integration Standards
• 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/>
Michael Badeaux
University of Maryland
ENAE484 PDR March 14, 2005
127 of 256
S Contamination and Particulate Control
P • Air filters
– High Efficiency Particulate Arrestance (HEPA) filter
R
– 99.7% efficiency on 0.3 microns
I
T
E
• 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
“Man-Systems Integration Standards.” NASA STD-3000. <http://msis.jsc.nasa.gov/>
Michelle Zsak
University of Maryland
ENAE484 PDR March 14, 2005
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Contamination Control Wipes
• Biocide
• Detergent
–Disinfecting food and waste
systems
• Biofilm Control
–Indoor cleaning
• Dry
–Toilet tissue
–Controls formation of Biofilm
inside surface of fluid lines
• Cleaning Implements
–Provides means for dislodging
and collecting dirt/debris
• Utensil Cleaning
–Sanitizers for post meal
cleaning
• Vacuum
“Man-Systems Integration Standards.” NASA STD-3000. <http://msis.jsc.nasa.gov/>
Michelle Zsak
University of Maryland
ENAE484 PDR March 14, 2005
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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
Larson, Wiley and Linda Pranke, ed. Human Spaceflight: Mission Analysis and Design
Michelle Zsak
University of Maryland
ENAE484 PDR March 14, 2005
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Trash Management
P
2-Man Crew,
1-wk Mission
R
Mass (kg) Volume (m3)
Total
9.1
0.202
I
Food
4.5
0.16
WCS Supplies
4.6
0.042
T
• Ways to store trash
–Free standing trash receptacle
–Storage compartment built into structure
–Trash compactor to minimum trash space
E
Michelle Zsak
University of Maryland
ENAE484 PDR March 14, 2005
131 of 256
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Radiation Sources
P
Galactic Cosmic Rays
Solar Particle Event
R
Duration
Near Constant
1-3 days
I
Composition
85% Protons
14% 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
T
E
“Man-Systems Integration Standards.” NASA STD-3000. <http://msis.jsc.nasa.gov/>
Michelle Zsak
University of Maryland
ENAE484 PDR March 14, 2005
132 of 256
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R
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Radiation Limits
Requirement L6: Radiation dosages shall, under all conditions, conform in all respects to
the current NASA standards for astronaut radiation limits
Lifetime Limits: Blood-Forming Organs (BFO)
5 cm depth
Age
Gender
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
Wilson, John, Francis Cucinotta, Lisa Simonsen, and Judy Shinn. “Galactic and Solar Cosmic Ray Shielding in Deep Space.” NASA Technical Paper. Dec 97.
Michelle Zsak
University of Maryland
ENAE484 PDR March 14, 2005
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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
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>
Michelle Zsak
University of Maryland
ENAE484 PDR March 14, 2005
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Shielding Options
• Possible Shielding
– Aluminum
• Currently used
• Creates neutrons during nuclear interaction that increase
exposure
T
– Polyethylene (CH2) without water
• Shields more than Aluminum since it is Hydrogen rich
E
– Polyethylene with water
• Shields 20% more than Aluminum since it is Hydrogen rich
• Must consider mass budget
Michelle Zsak
University of Maryland
ENAE484 PDR March 14, 2005
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Aluminum vs. Polyethylene
Solar Minimum 1977
Thickness
(g/cm2)
Dose
Equivalent
(rem/yr)
Al
CH2
0
95
95
1
91
2
Solar Maximum 1970
Thickness
(g/cm2)
Dose
Equivalent
(rem/yr)
Al
CH2
0
34.5
34.5
81
1
33.7
32.7
88
83
2
32.9
31.2
5
79
71
5
30.7
27.2
10
69
57
10
27.8
22.6
15
54
41
15
22.8
16.4
25
46
35
25
20.0
14.4
75
43
32
75
19.4
13.7
Wilson, John, Francis Cucinotta, Lisa Simonsen, and Judy Shinn. “Galactic and Solar Cosmic Ray Shielding in Deep Space.” NASA Technical Paper. Dec 97.
Michelle Zsak
University of Maryland
ENAE484 PDR March 14, 2005
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S Possible Radiation Shielding Plan
P • Shield all sides exposed to radiation
R
I
• 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)
T
E
Michelle Zsak
University of Maryland
ENAE484 PDR March 14, 2005
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Fire Suppression
P
R
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Liquid
Density
Volume
Fraction
Halon
1301
1570
kg/m3
0.20
Highly effective
CO2
758
kg/m3
0.62
Toxic in high concentration
Can be cleaned by rover
Type
Comments
• 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
University of Maryland
ENAE484 PDR March 14, 2005
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Internal vs. External Suits
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R
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Internal
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
Larson, Wiley J., ed. Human Spaceflight: Mission Analysis and Design.
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
University of Maryland
ENAE484 PDR March 14, 2005
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Layout
P
R
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John Mularski
John Mularski
University of Maryland
ENAE484 PDR March 14, 2005
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Layout
Requirement L1: All crew interfaces shall accommodate 95th percentile American male to
5th percentile Japanese female
• Current cabin volume =
21 m3
• Space surrounding cabin
for pipes, wires and
auxiliary equipment
E
John Mularski
John Mularski
University of Maryland
ENAE484 PDR March 14, 2005
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Layout
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
• 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
E
John Mularski
University of Maryland
ENAE484 PDR March 14, 2005
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Visual Display (VD)
Requirement L1: All crew interfaces shall accommodate 95% American male to 5% Japanese female
Requirement L2: All crew interfaces shall adhere to NASA-STD-3000, Man-Systems Integration Standards
• 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
University of Maryland
ENAE484 PDR March 14, 2005
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Windows
Finding minimum window dimensions for navigational purposes
• Inputs
• Output
–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)
Shawn Butani
–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
University of Maryland
ENAE484 PDR March 14, 2005
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Window Solution
Window
Seats
•
•
•
•
•
•
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
Shawn Butani
Hull
Michael Badeaux
University of Maryland
ENAE484 PDR March 14, 2005
145 of 256
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Front Display
•
•
•
•
•
•
Astronauts sit 16 in. from windowsSeats
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 :
–
–
Window
Steering system : Throttle (SDOF), L &
R steer (SDOF), Lift Break
Avionics : input from driver, indicators,
sensors (wheels, pitch and roll, speed,
etc.)
E
Hull
Evan Ulrich
Shawn Butani
University of Maryland
ENAE484 PDR March 14, 2005
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Geographic Survey
P
Requirement 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
R
•
I
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
T
E
Shawn Butani
University of Maryland
ENAE484 PDR March 14, 2005
147 of 256
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EVA Suit Shielding
Requirement 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)
T
E
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
University of Maryland
ENAE484 PDR March 14, 2005
148 of 256
S Calculating Shield Dimensions
P • Density = 0.043 lb/in
3
R
•
I
•
T
E
•
•
= 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
University of Maryland
ENAE484 PDR March 14, 2005
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S
Crew Systems Future Work
P
R
•
•
•
•
•
•
I
T
E
Shawn Butani
EVA checklist
Health monitoring
Interior stowage
Docking system
EVA support
Controls and displays
University of Maryland
ENAE484 PDR March 14, 2005
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R
I
Intermission
T
E
University of Maryland
ENAE484 PDR March 14, 2005
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Surface GNC
T
E
University of Maryland
ENAE484 PDR March 14, 2005
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S Guidance Navigation & Control
P • IMU accuracy is vital to mission success
R
I
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– 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
* Data from Honeywell IMU spec sheet
Spec Sheet - http://content.honeywell.com/dses/assets/datasheets/fog.pdf
Aaron Shabazz
University of Maryland
ENAE484 PDR March 14, 2005
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S Guidance Navigation & Control
P
R
I
• Works Still in Progress
– GNC Thermal Control
– Determining which computers to use
– Determining number of computers needed
T
E
Aaron Shabazz
University of Maryland
ENAE484 PDR March 14, 2005
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S Navigation and Guidance on Moon
Surface
P • SPRITE shall be
R
I
T
E
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
Ralph Myers
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
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
University of Maryland
ENAE484 PDR March 14, 2005
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S Navigation and Guidance on Moon
Surface
landmark for absolute reference and dead reckoning sensors
P • Use
for relative reference
R
Scan horizon
or predetermined
landmark
Build DEM and
Compare to lunar
Map surface
Vehicle &
Landmark
Latitude and
Longitude
I
Accel. X
Gyro Pitch
Accel. Y
Gyro Yaw
Accel. Z
Real time
Calibration
Compare
values to
Lunar Map
T
E•
Gyro Roll
Myers, Ralph
Odometer
Slippage
detection
Left Front
Errors in the dead reckoning sensor will
Left Rear
determine the distance needed
before a landmark is needed for correction update
Ralph Myers
Accelerometer
Compensation
University of Maryland
ENAE484 PDR March 14, 2005
Torque
sensor
Right Front
Right Rear
156 of 256
S On Board Direct Human Control
P • Drive by wire will control steering, acceleration, and braking
R
I
T
E
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
University of Maryland
ENAE484 PDR March 14, 2005
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S
Minimize Odometry Error
P
R
I
T
E
Specifications for odometry
accuracy
Encoders
Resolvers
Controllable Speed Range
0.1 rpm to 30 rpm to
10,300 rpm 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
http://www.heidenhain.com/Linear-2.htm
http://www.motec.co.uk/documents/ormec/encres.htm
Ralph Myers
University of Maryland
ENAE484 PDR March 14, 2005
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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
T
E
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
http://www.biorobotics.harvard.edu/pubs/tac-manip.pdf
http://www.biorobotics.harvard.edu/pubs/tac-manip.pdf
Ralph Myers
University of Maryland
ENAE484 PDR March 14, 2005
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Surface Obstacle Avoidance
SPRITE must traverse 0.5 m obstacles, 20º crossslope, 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
University of Maryland
ENAE484 PDR March 14, 2005
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T
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
E
Scott Walthour
University of Maryland
ENAE484 PDR March 14, 2005
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S Minimum Look Ahead Distance
P
R
I
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E
Scott Walthour
University of Maryland
ENAE484 PDR March 14, 2005
Scott Walthour
162 of 256
S Maximum Look Ahead Distance/ Camera
P
Horizontal Field of View (HFOV)
R
I
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E
Scott Walthour
University of Maryland
ENAE484 PDR March 14, 2005
Scott Walthour
163 of 256
S Camera Vertical Field of View (VFOV)
P
R
I
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Scott Walthour
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
University of Maryland
ENAE484 PDR March 14, 2005
S
P
Derived Obstacle Detection
Requirements
Minimum Look Ahead Distance
4m
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
R
I
T
E
*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
University of Maryland
ENAE484 PDR March 14, 2005
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165 of 256
S Obstacle Detection Future Work
P Choose COTS* cameras
R
I
– Resolution
– CCD, CID, Vision chips
Determine computational requirements
T
E
*COTS – commercial off the shelf
Scott Walthour
University of Maryland
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Network Data 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
• 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
<http://www.interfacebus.com>
Scott Walthour
University of Maryland
ENAE484 PDR March 14, 2005
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Network Data Bus
P
R
I
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Spacewire
Advantages
Disadvantages
High Data Rate (Mbps)
Not inherently redundant
155-200 typical (400 max)
Lightweight
0.06 kg/m
Scalable
Radiation Tolerant
• Requires routers to ensure
redundant paths
- Increases complexity of the
network
BER = 10-14*
Full Duplex
*Bit Error Rate
<http://www.estec.esa.nl>
Scott Walthour
University of Maryland
ENAE484 PDR March 14, 2005
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S
Example Network
P
R
I
T
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Scott Walthour
Scott Walthour
University of Maryland
ENAE484 PDR March 14, 2005
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I
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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
E
Scott Walthour
University of Maryland
ENAE484 PDR March 14, 2005
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Communications
T
E
University of Maryland
ENAE484 PDR March 14, 2005
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Communication Requirements
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
From Work Breakdown Structure
• From SPRITE to Earth
• From SPRITE to Base
• From SPRITE to EVA
• Contingency/Emergency
I
T
E
Jay Kim
University of Maryland
ENAE484 PDR March 14, 2005
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S
High Definition TeleVision
P
•
–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
R
I
T
HDTV specs
•
Comparison of different displays
E
Jay Kim
Jay Kim
Compression technique
–MPEG 1: Standard for Video CD
–MPEG 2: Standard for broadcastquality television
• Compression rate up to 20
Mbps
University of Maryland
ENAE484 PDR March 14, 2005
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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
•
•
T
Uplink = 30 Mbps (from Earth to SPRITE)
Downlink = 50 Mbps (from SPRITE to
Earth)
•
Trade studies of link budgets
E
•
Near side
High Gain Antenna
–Frequency selection
–Antenna selection
Link budget constraints
–Link margin 3 dB – 6 dB
Jay Kim
Jay Kim
University of Maryland
ENAE484 PDR March 14, 2005
Far side
Low Gain Antenna
S
P
R
I
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)
T
E
David G. MacDonnell, “Communications Analysis of Potential Upgrades of NASA’s Deep Space Network”
Akin, Dave. ENAE483 Link Budget Spreadsheet
Jay Kim
Effect of changing diameter
University of Maryland
ENAE484 PDR March 14, 2005
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175 of 256
Link Budget
S
P
R
I
Jay Kim
Effect of changing transmitter power
T
•
•
•
Diameter of antenna size: 0.5 m
Transmitter power: 1 W
Mass: TBD
Operating frequency: 15 – 25 GHz
Link margin: 3dB – 6dB
E
Jay Kim
University of Maryland
ENAE484 PDR March 14, 2005
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S
SPRITE To Base
P
•
Transmitting antenna in SPRITE
–
R
–
–
–
–
I
•
E
Receiving antenna in base
–
–
T
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)
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)
T
E
University of Maryland
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S
P
Flying Locator and Assistance
Requesting Equipment (FLARE)
•
•
R
–Small solid rocket motor for propulsion
–Equipment based on amateur radio microsatellite technology
I
T
E
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.
•
•
Motor Mass
Window 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
University of Maryland
ENAE484 PDR March 14, 2005
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P
R
I
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
T
E
University of Maryland
ENAE484 PDR March 14, 2005
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P
R
I
Power Systems
T
E
University of Maryland
ENAE484 PDR March 14, 2005
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S Power & Energy: Requirements and Budget
P
R
I
• 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
T
E
Jason West
University of Maryland
ENAE484 PDR March 14, 2005
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S Power Requirements Overview
P
Power [kW]
Energy [kW-hr]
55.6
2276
Nominal required
36.5
1277.5
Peak required
55.5
55.5
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
0.012
2.3
SPRITE Total
Surface Propulsion
R
Continuous
I
Avionics
T
E
Crew Systems
CO2 removal
Jason West
University of Maryland
ENAE484 PDR March 14, 2005
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S
Power: Requirements and Budget
Required
Power
[kW]
P
R
I
T
E
Surface Propulsion
(max)
Budgeted
Power
[kW]
Emergency
Power
[kW]
36.5
(55.5)
40.0
(60.0)
0
(0)
.117
.250
TBD
1
1
1.0
Science Mission
TBD
1
0
Thermal
TBD
1
TBD
Miscellaneous
TBD
1
TBD
37.617(56.6)
44.25(64.3)
1.0(1.0)
Avionics
Crew Systems
Total (max)
Jason West
University of Maryland
ENAE484 PDR March 14, 2005
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S
Power: Requirements and Budget
P
Required
R
Max Power, 64.3 kW
60
60
Budgeted
55.5
50
T
E
Power (kW)
I
Nominal Power, 44.25 kW
40
30
20
Emergency Power, 1 kW
10
4.25
0.117
0
Surface Propulsion
Jason Mallare
Other
University of Maryland
ENAE484 PDR March 14, 2005
185 of 256
S Energy: Requirements and Budget
P
Power Req
(kW)
Time
(hr)
Energy Req
(kW-hr)
Power
Budgeted
(kW)
Time
(hr)
Energy
Budgeted
(kW-hr)
Surface
Propulsion
(cruising)
36.5
35
1277.5
40.0
35
1400.0
Surface
Propulsion
(ascent)
55.5
1
55.5
60.0
1
60.0
Avionics
.117
192
22.5
.250
192
48.0
Crew
Systems
1
192
192.0
1
192
192
Thermal
TBD
TBD
TBD
1
192
192
Science
Mission
TBD
TBD
TBD
1
192
192
Misc.
TBD
TBD
TBD
1
192
192
System
R
I
T
E
Total
Energy
1357.8
192 hours represent 8 day, 24 hour/day usage
Jason Mallare
University of Maryland
ENAE484 PDR March 14, 2005
2276.0
186 of 256
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I
T
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
E
Jason Mallare
University of Maryland
ENAE484 PDR March 14, 2005
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S Power Management & Distribution
P • Future Work
R
I
T
E
– 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
Hyder, Wiley, Halpert, Flood, Sabripour. “Spacecraft Power Technologies”
Jason Mallare
University of Maryland
ENAE484 PDR March 14, 2005
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S
Energy Storage
P • Technologies considered
– Primary batteries
R
– Secondary (rechargeable) batteries
I
T
– Radio-isotope
– Solar arrays
– Fuel cells
E
Jason Mallare
University of Maryland
ENAE484 PDR March 14, 2005
189 of 256
S
P
R
Primary Batteries
• Advantages:
• Primary cells offer higher specific energy then secondary batteries
• Disadvantages:
• Non-rechargeable, low current, low specific power (W/kg)
I
T
E
Gravimetric
Specific
Energy
(W-h/kg)
Volumetric
Specific
Energy
(W-h/L)
Specific
Power
(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
Chemistry
<http://www.saftbatteries.com/010-Home/10-10_home.asp>
<http://www.varta.com/eng/>
<http://www.ulbi.com/>
Jason Mallare
University of Maryland
ENAE484 PDR March 14, 2005
190 of 256
S
P
R
I
T
E
Secondary Batteries
• 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
Gravimetric
Specific
Energy
(W-hr/kg)
Volumetric
Specific
Energy
(W-hr/L)
Specific
Power
(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
Chemistry
<http://www.saftbatteries.com/010-Home/10-10_home.asp>
<http://www.varta.com/eng/>
<http://www.ulbi.com/>
Jason Mallare
University of Maryland
ENAE484 PDR March 14, 2005
191 of 256
S
Batteries - Energy Storage
P
Battery Specific Energy
Gravimetric
T
E
Gravimetric Specific Energy(W-hr/kg)
I
1241
1200.0
1000.0
800.0
740
568
600.0
400.0
512
Secondary Batteries
386
304
300
240
200
206
329
271
247
200.0
72
0.0
Li-Ion
LiPolymer
Sodium
Sulfur
Li-SOCl2
Li-Mn02
Li-SO2
Primary Batteries
Jason Mallare
University of Maryland
ENAE484 PDR March 14, 2005
Ni-MH
Volumetric Specific Energy (W-hr/L)
R
Volumetric
1400.0
192 of 256
S
Batteries - Power Generation
P
Battery Specific Power
350.00
R
T
309
300.00
Specific Power (W/kg)
I
Secondary Batteries
250.00
244
200
200.00
150.00
Primary Batteries
100.00
52
E
50.00
0
10
14
Li-SO2
Ni-MH
0.00
Li-Ion
Jason Mallare
Li-Polymer
Sodium
Sulfur
Li-SOCl2
Li-Mn02
University of Maryland
ENAE484 PDR March 14, 2005
193 of 256
S
P
R
I
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
T
E
<http://newfrontiers.larc.nasa.gov/newfrontiers/09_NF_PPC_Schmidt.pdf>
Phillip Adkins
University of Maryland
ENAE484 PDR March 14, 2005
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S
P
R
I
T
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
E
<http://spacecraft.ssl.umd.edu/academics/483F04/483L14.power_sys/483L14C.power.2004.pdf>
Phillip Adkins
University of Maryland
ENAE484 PDR March 14, 2005
195 of 256
S
P
R
I
T
E
Fuel Cells
Efficiency
Operating Temperature
(oC)
Type
Specific
Power
(W/kg)
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
<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>
Phillip Adkins
University of Maryland
ENAE484 PDR March 14, 2005
196 of 256
S
P
R
I
T
E
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 kWhr
– Using alkaline fuel cells and
assuming 70% efficiency for
the fuel cells.
Phillip Adkins
Fuel Cell Reactor
640 kg
Reactants
860 kg
H2 and O2 tanks
420 kg
Total Mass
1920 kg*
* ~38% of total rolling mass
University of Maryland
ENAE484 PDR March 14, 2005
197 of 256
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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.
T
E
Phillip Adkins
University of Maryland
ENAE484 PDR March 14, 2005
198 of 256
S
P
R
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Thermal Control
T
E
University of Maryland
ENAE484 PDR March 14, 2005
199 of 256
S
P
R
I
T
Thermal Control
• Requirements
– Maintain cabin temperature between 18.3 and
26.7ºC
– Cool electronics and motors so that
equipment operates at peak efficiency
E
Evan Alexander
University of Maryland
ENAE484 PDR March 14, 2005
200 of 256
S
P
R
I
T
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
E
Evan Alexander
University of Maryland
ENAE484 PDR March 14, 2005
201 of 256
S
P
R
I
T
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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
Mylar
Teflon
Features
Thickness (µm)
Emissivity
Y9360-3M
Aluminized
TBD
0.03
0.19
Aluminized
Backing
3.8
0.28
0.14
Gold Backing
12.7
0.49
0.30
2.0
0.24
0.23
Aluminized
Kapton Film Backing
Evan Alexander
University of Maryland
ENAE484 PDR March 14, 2005
Absorptivity
202 of 256
S
P
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T
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
E
Evan Alexander
University of Maryland
ENAE484 PDR March 14, 2005
203 of 256
S
MLI vs Aerogels
P
R
I
Type
Thermal Conductivity
(W/m-K)
Kapton
1.42
0.12
Mylar
1.39
0.2
Teflon
2.15
0.195
Silica
0.01-0.3
0.004
Resorcinol
0.6
0.06
Carbon
0.9
0.04
Category
MLI
T
E
Density
(g/cm3)
Aerogels
Evan Alexander
University of Maryland
ENAE484 PDR March 14, 2005
204 of 256
S
P
R
Active Thermal Control
• Use Heat Pipes to cool electronics
• Radiators used to expel excess heat from cabin
I
T
E
<http://spacecraft.ssl.umd.edu>
Evan Alexander
University of Maryland
ENAE484 PDR March 14, 2005
205 of 256
S
P
R
I
T
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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
University of Maryland
ENAE484 PDR March 14, 2005
206 of 256
S
Heat Pipes (cont.)
P
R
I
T
• 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 Ammonia
Nickel, Aluminum,
Stainless Steel
+5 to +230
Water
Copper, Nickel
E
Evan Alexander
University of Maryland
ENAE484 PDR March 14, 2005
207 of 256
S
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R
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Heat Pipes (cont.)
• Properties potential metals used
Metals
Aluminum
Density
(g/cm^3)
Thermal Conductivity
(W/m-K)
2.7
205
Nickel
8.91
90.7
Stainless Steel
8.03
50.2
Copper
8.92
394
Evan Alexander
University of Maryland
ENAE484 PDR March 14, 2005
208 of 256
S
P
R
I
T
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
E
Evan Alexander
University of Maryland
ENAE484 PDR March 14, 2005
209 of 256
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Science
T
E
University of Maryland
ENAE484 PDR March 14, 2005
210 of 256
S Suggested Landing/Mission Zones
P
R
I
T
•
•
•
•
Crater Copernicus
Crater Tycho
Mare Orientale
South Pole-Aitken (SPA)
Basin
E
Lunar and Planetary Institute, 2005
Chris Hartsough
University of Maryland
ENAE484 PDR March 14, 2005
211 of 256
S
P
R
I
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 and Planetary Institute, 2005
T
E
Lunar Orbiter image II-162H3
Chris Hartsough
University of Maryland
ENAE484 PDR March 14, 2005
212 of 256
S
Crater Tycho
P
R
I
T
• 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
E
Lunar and Planetary Institute, 2005
Chris Hartsough
University of Maryland
ENAE484 PDR March 14, 2005
213 of 256
S
Mare Orientale
P
R
I
Lunar and Planetary Institute, 2005
•
T
Geographic Interest
–Diameter of ~950 km
–Depth of ~3.2 km
–Multi-leveled mare
–Large iron concentration
–Ease of landing
–Half visible to earth
E
Lunar and Planetary Institute, 2005
Chris Hartsough
University of Maryland
ENAE484 PDR March 14, 2005
214 of 256
S South Pole-Aitken (SPA) Basin
• Geographic Interest
P
R
I
–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
T
Lunar and Planetary Institute, 2005
E
Lucey et al., 1998
Chris Hartsough
University of Maryland
ENAE484 PDR March 14, 2005
215 of 256
S Choosing Scientific Instruments for
SPRITE
P
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
R
I
T
Completed steps
1. Detail the mass and volume requirements for
scientific hardware used in previous J-Class
missions.
E
Ryan Livingston
University of Maryland
ENAE484 PDR March 14, 2005
216 of 256
S Instruments - Crew Experiments
P
R
I
T
E
Original
Mass (kg)
Returned
Mass (kg)
Stored
Volume (m3)
Soil Mechanics Investigation**
15.7
15.7
TBD
Solar Wind Composition
Experiment
0.46
0.385
1.3e-3
Lunar Portable Magnetometer
0.46
0
1.18e-2
22
0
0.25
Experiments *
Far Ultraviolet
Camera/Spectrograph
* Hand Tools to assist experiments = approx 50 kg
** includes ALSD (drill)
Ryan Livingston
University of Maryland
ENAE484 PDR March 14, 2005
217 of 256
S Instruments - Crew Experiments
P
R
I
T
Original
Mass (kg)
Returned
Mass (kg)
Stored
Volume (m3)
Cosmic Ray Detector
0.163
0.163
0.13e-3
Transverse Gravimeter
Experiment
14.6
0
0.0351
Lunar Neutron Probe
2.27
0.4
0.38e-3
16
1
0.024
Experiments *
Surface Electrical Properties
E
Ryan Livingston
University of Maryland
ENAE484 PDR March 14, 2005
218 of 256
S
Instruments - Deployed
P
R
I
T
E
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
Experiments
* does not include foldable skirt
Ryan Livingston
University of Maryland
ENAE484 PDR March 14, 2005
219 of 256
S
Instruments - Deployed
P
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
T
Lunar Atmospheric Composition
Experiment
9.1
0
0.018
E
Lunar Ejecta and Meteorites Experiment
7.4
0
0.02
Lunar Surface Gravimeter
12.7
0
0.027
R
I
Experiments
Ryan Livingston
University of Maryland
ENAE484 PDR March 14, 2005
220 of 256
S Choosing Scientific Instruments for
SPRITE
P Future Steps
R
I
T
1.
2.
3.
4.
5.
Select scientific missions to be included.
Check for more advanced versions of chosen hardware.
Check for special requirements demanded by scientific hardware
(i.e. storage temperature).
Locate storage location on SPRITE.
Select tools and storage suitable for EVA in I-Suits
E
Ryan Livingston
University of Maryland
ENAE484 PDR March 14, 2005
221 of 256
S
P
Robotic Extendable Arm with
Changeable Heads (REACH)
Requirement M4: SPRITE shall provide the capability for the crew to interact with the local
environment and critical external vehicle systems without EVA
R
I
T
E
- Must reach entirety of SPRITE
exterior
- Must have 100 kg payload
capacity (suits)
- Perform specific science
requirements TBD
- At least 6 DOF needed
David Gruntz
David Gruntz
University of Maryland
ENAE484 PDR March 14, 2005
222 of 256
S
P
R
I
REACH Configuration
• Several configurations considered
– Single arm
– Two arms (one on each end of rover)
– Single arm on track
T
E
David Gruntz
David Gruntz
University of Maryland
ENAE484 PDR March 14, 2005
223 of 256
S
REACH Material
P
R
I
T
Density
(kg/m3)
Yield
Stress
(MPa)
Elasticity
(GPa)
Material
Aluminum,
wrought, 2024-T4
Titanium alloy,
annealed
Carbon/Epoxy resin
2800
325
4460
1230
1600
800
73
TBD
125
Yield
Stress/
Density
Ratio
0.12
Easy to
machine
0.28
Expensive,
Too strong
0.50
Extremely
lightweight
Beer, Ferdinand. Mechanics of Materials
E
Werelety, Norman. ENAE423 Lectures - Composite Materials
• Carbon/Epoxy resin ideal choice
– Lightweight and strong
David Gruntz
Comments
University of Maryland
ENAE484 PDR March 14, 2005
224 of 256
S
Initial Sizing & Mass
P
• Static analysis performed to determine size and mass
• 100 kg payload in Martian gravity (3.7 m/s2)
R
I
T
E
Configuration
Length
Material
Mass
Max Stress
MOS
(per arm)
(per arm)
(per arm)
(kg)
(MPa)
SF = 2
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
University of Maryland
ENAE484 PDR March 14, 2005
225 of 256
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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
University of Maryland
ENAE484 PDR March 14, 2005
226 of 256
S
P
R
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Contingencies
T
E
University of Maryland
ENAE484 PDR March 14, 2005
227 of 256
S
P
R
Emergency Return Vehicle
Requirement M5: The SPRITE system shall include provision for safe return of the
crew following a worst-case SPRITE failure without outside intervention.
• To be used when the crew must return to base without
the main rover
• Scenario 1: Rover becomes immobile
I•
T
E
– 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
Jason West
University of Maryland
ENAE484 PDR March 14, 2005
228 of 256
S
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R
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T
Portable Air, Nutrients, and Inflatables
Cache (PANIC)
• 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
E
Samuel Schreiber
University of Maryland
ENAE484 PDR March 14, 2005
229 of 256
S
P
R
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
I
• Provides an opportunity for astronauts to remove
T
• 10 m3 minimal habitable volume for two 95th percentile
E
• Reusable - Only one needed throughout return to base
space suits, eat, rest, and discard waste
American male astronauts with space suits.
Samuel Schreiber
University of Maryland
ENAE484 PDR March 14, 2005
230 of 256
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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
5.7 kg water for rest
0.8 kg oxygen for traverse
3.2 kg oxygen for rest
Samuel Schreiber
7.9 kg Oxygen Tank
1.3 kg Water Tank
University of Maryland
ENAE484 PDR March 14, 2005
231 of 256
S
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R
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T
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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
University of Maryland
ENAE484 PDR March 14, 2005
232 of 256
S
P
R
I
T
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
E
Samuel Schreiber
University of Maryland
ENAE484 PDR March 14, 2005
233 of 256
S
P
Transport Emergency Recovery by
Rocket Operated Return (TERROR)
Requirement M5: The SPRITE system shall include provision for safe return of the crew
following a worst-case SPRITE failure without outside intervention
R
I
T
E
• 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 Wasserman
Daniel Zelman
University of Maryland
ENAE484 PDR March 14, 2005
234 of 256
S
TERROR - Trajectory
P
D (km)
25
50
75
100
125
R
I
T
E
•
•
•
•
•
•
•
v (rad)
0.0072
0.0144
0.0216
0.0288
0.0360
e
V0 (km/s) ∆V (km/s) Apogee (km) TOF (min)
0.9928
0.201
0.401
45
5.8
0.9857
0.283
0.566
91
8.2
0.9787
0.345
0.690
140
10.0
0.9716
0.397
0.794
191
11.5
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 Wasserman
Dan Zelman
University of Maryland
ENAE484 PDR March 14, 2005
235 of 256
S
TERROR - Mass and Volume Estimates
P
R
I
Volume
Mass
Fuel
Oxidizer
56 kg
103 kg
Tank (Fuel)
4 kg
Tank (Oxidizer)
5 kg
Pressure Tank
5 kg
Wiring
10 kg
T
Engine
16 kg
Avionics
10 kg
E
Seats
25 kg
Thrust Structure
Total Mass
Timothy Wasserman
Dan Zelman
Fuel Tank
0.065 m3
Oxidizer Tank
0.065 m3
Engine
0.016 m3
Platform
0.016 m3
Total Volume
1 kg
224 kg
University of Maryland
ENAE484 PDR March 14, 2005
1.14 m3
236 of 256
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Foldable Escape Assisting Rover
(FEAR)
Requirement M5: The SPRITE system shall include provision for safe return of the
crew following a worst-case SPRITE failure without outside intervention
• Based on the original Apollo Rover
• Lighter and Stronger
– New Material
– Less Payload
T
• Higher Clearance
E
• Faster and More Powerful
– 0.5 m Requirement
– Newer engines
Laurie Knorr
University of Maryland
ENAE484 PDR March 14, 2005
237 of 256
S
FEAR - Mass and Material
P
Aluminum Alloy
2219
Carbon Epoxy
2.84 g/cm3
1.6 g/cm3
Tensile Strength
359 MPa
600 MPa
Yield Strength
248 MPa
600 MPa
T
Modulus of Elasticity
73.1 GPa
70 GPa
Shear Modulus
27 GPa
5 GPa
E
Shear Strength
230 MPa
90 MPa
R
I
Density
Aerospace Specification Metals Inc - <http://asm.matweb.com/search/SpecificMaterial.asp?bassnum=MA2219T37>
Goodfellow - <http://www.azom.com/details.asp?ArticleID=1995>
Laurie Knorr
University of Maryland
ENAE484 PDR March 14, 2005
238 of 256
S
P
R
I
T
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
Chassis Fitting
Lower
Arm
Upper
Arm
Damper
• Change the suspension
–Mass increase minuscule
–Small loss in strength
E
LRV Operations Handbook, 1973 Contract NASA-25145
Lunar Rover Operations Handbook - <http://www.hq.nasa.gov/alsj/irvhand.html>
Laurie Knorr
University of Maryland
ENAE484 PDR March 14, 2005
239 of 256
S
P
R
I
T
E
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
Lunar Rover Operations Handbook - <http://www.hq.nasa.gov/alsj/irvhand.html>
Laurie Knorr
University of Maryland
ENAE484 PDR March 14, 2005
240 of 256
S
P
R
FEAR - Special Features
• Drive back to base in less than 10 hours
I
• Folds up into 0.9 m3 space
T
• Attaches to the outside of SPRITE
E
Lunar Rover Operations Handbook - <http://www.hq.nasa.gov/alsj/irvhand.html>
Laurie Knorr
University of Maryland
ENAE484 PDR March 14, 2005
241 of 256
S Emergency Return Operations
P
R
I
T
E
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
* Does not include food
Laurie Knorr
University of Maryland
ENAE484 PDR March 14, 2005
242 of 256
S
Safety
P
R
I
T
PANIC
TERROR
FEAR
E
Laurie Knorr
Time
Advantages
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
13 min
• Fast return to base
• Can work if one crew
member is injured
• Very unsafe
• Complicated system
• Crew exerts little energy
• Can work if one crew
member is injured
• Complicated detachment
procedures
10 hr
Disadvantages
University of Maryland
ENAE484 PDR March 14, 2005
243 of 256
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R
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Program Timeline and Costs
T
E
University of Maryland
ENAE484 PDR March 14, 2005
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S
P
Program Timeline
Requirement M8: The SPRITE shall be ready for initial lunar operations by 2016
R • Development/Production: 2005-2015
I
T
E
• 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
University of Maryland
ENAE484 PDR March 14, 2005
245 of 256
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R
I
T
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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
*Timeline would repeat (except launch) approximately each month
for a period of 3 months
**assumes 1 SPRITE Vehicle, 5 day trip to moon
Charles Bacon
University of Maryland
ENAE484 PDR March 14, 2005
246 of 256
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P
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Program Timeline
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
• 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
E
Charles Bacon
University of Maryland
ENAE484 PDR March 14, 2005
247 of 256
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P
R
I
T
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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
*Derived from NASA Cost Models
Charles Bacon
University of Maryland
ENAE484 PDR March 14, 2005
248 of 256
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I
T
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)
E
- Larson, Pranke Human Spaceflight: Analysis and Design, pg 755, table 23-10
Charles Bacon
University of Maryland
ENAE484 PDR March 14, 2005
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S
Cost Totals
P
System
SPRITE Vehicle (1)
R
Landing Engine
(1-stage LOX/LH2)
Landing Structure
FEAR
I
T
E
Cost (FY 04 $M)
1940
1460
TBD
144
PANIC
TERROR
Robotic Arm
TBD
320
185
Launch
254
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Cost Analysis
P • Current total – $4.1 Billion
– Estimated final cost to launch: 1 SPRITE + 1 ERV
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• Worst Case Scenario – TERROR: most expensive
Requirement I6: The SPRITE design shall be designed to minimize life cycle costs
I
T
– 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)
E
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S Mission Operation and Data Analysis Cost
P
R
I
• 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
T
E
*assume investment price - $3.9B
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Cost Spreading
Requirement M8: The SPRITE shall be ready for initial lunar operations by 2016
• Development and Production would occur from
2005-2015
– Launch in 2016
• Beta Function
– Non-Recurring costs over 11 years
– Recurring Costs take over in 2016.
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Cost Spreading
P
SPRITE's Estimated Annual Expenditures
R
600
I
T
FY 04 ($M)
500
400
300
200
100
0
2005
E
2006
2007
2008
2009
2010
2011
2012
Year
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2013
2014
2015
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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
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S Systems Integration Future Work…
P • Thorough itemized analysis for SPRITE to result in a
R
I
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
T
E
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The End
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University of Maryland
ENAE484 PDR March 14, 2005