Small Pressurized Rover for Independent Transport and Exploration

Small Pressurized Rover for Independent Transport and Exploration
Preliminary Design Review March 14, 2005

What is SPRITE? SPRITE is a pressurized rover designed primarily for use on the moon. It can be used, with only minor changes, on the Martian surface. It would serve as the primary exploration vehicle for astronauts living at a lunar base. It accommodates two astronauts for a week-long scientific expedition Charles Bacon

Why SPRITE? With the new exploration initiative being undertaken by NASA for human presence on the Moon and Mars, there must be a way for humans to traverse long distances from the base. This is primarily because ideal sites for landing and base construction (flat, open terrain) are not the same as those most interesting for scientific exploration (geologically diverse regions). Charles Bacon

Why SPRITE? Though many pressurized rovers have been suggested, none have been fully developed mainly because of cost. To constrain this problem, SPRITE will be launched on a single Delta IV Heavy vehicle, including all systems needed for nominal and emergency use. The only thing not to be included on the launch will be consumables required. They will be provided by the lunar base. Charles Bacon

Launch to Landing

CONOPS Overview (Delta-IV Heavy Separation to Landing)
Chris Hartsough

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) Daren McCulley

Translunar Orbit Options
Low Thrust Advantage Unmanned and mass constrained mission Disadvantages Insufficient maximum thrust Flight times grossly exceeding reasonable limits Requires two propulsion system reliability Payload fairing constraints High power requirements Latest advances require 7-20 kW for .5-1 N of thrust Electromagnetic interference Delta IV Second Stage TLI Advantages Flight time between 4.5 and 5.5 days Presumably will be flight tested by 2016 Disadvantage Highly inefficient ratio between propellant and payload mass Over 50% of the mass in LEO is consumed during TLI Daren McCulley

Depiction of Translunar Orbits
Low Thrust Transit High Thrust TLI Daren McCulley

Apogee of Translunar Orbit
Radius of Apogee TLI ∆V 356,000 km Moon at Perigee 3.128 km/s 407,000 km Moon at Apogee 3.140 km/s * Additional DV of only 12 m/s * Additional day of flight time Daren McCulley

Payload vs. Apogee Delta-IV Payload Planners Guide Daren McCulley

Selenocentric Orbits Options Total ∆V Direct Descent 2.79 km/s
L1 Layover 3.10 km/s Elliptical Lunar Orbit Insertion 2.83 km/s Circular Lunar Orbit Insertion 2.85 km/s Larson, Wiley J. and Pranke, Linda K ETD. Human Space Flight, Mission Analysis and Design Daren McCulley

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

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

Descent Orbit Analysis
Altitude of Periselenium (km) DOI ∆V (km/s) Tangent Velocity Normal Velocity Total ∆V 10 .0206 1.696 .1808 1.897 20 .0183 1.688 .2557 1.962 30 .0159 1.681 .3132 2.010 40 .0136 1.674 .3617 2.050 50 .0113 1.667 .4044 2.083 Daren McCulley

Insertion and Landing Concept
Lunar Orbit Insertion (LOI) Retro burn at closest point of approach 100 km altitude circular orbit Descent Orbit Insertion (DOI) Retro burn at descent orbit aposelenium 15 km periselenium above landing site Powered Descent Landing (PDL) Retro burn near periselenium Continue controlled burn to soft landing Daren McCulley

Depiction of Selenocentric Orbits
Daren McCulley

3D Orbit Design In Reverse
Daren McCulley

Gravity Assist Prevent the spacecraft from leaving Earth orbit in the event the retro engine fails to fire. Unmanned mission, makes this a low level requirement. Chobotov, Vladimir. Orbital Mechanics Daren McCulley

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

Analysis of Control Variables
all axes in km Daren McCulley Daren McCulley

Powered Descent Simulation
km km Daren McCulley km km Daren McCulley

PDL Simulation Results
Burn Altitude: 16.2 km Burn Time: s Thrust: 42.9 kN Residual Velocities: Negligible Height: 4 m Landed Mass: 5435 kg Max Acc: 6.3 m/s2 ∆V: 1.83 km/s Velocity (km/s) Time after Aposelenium (s) Daren McCulley

Burn Profile Burn / Maneuver Engine ∆V (km/s) ∆M (kg)
PL Fairing Evasion RCS Negligible Delta IV SS TLI RL-10B-2 3.14 N/A Midcourse Correction TBD 0.01 50 Lunar Orbit Insertion RETRO .816 1670 Circular Orbit Correction Descent Orbit Insertion 0.02 40 Descent Orbit Correction Powered Descent 1.89 2860 Daren McCulley

Guidance Navigation & Control
Derived Requirements The GNC system shall provide: state vector estimations attitude determination attitude control systems landing control systems landing point localization Aaron Shabazz

Critical GNC Hardware Inertial Measurement Units (IMU) Senses pitch, yaw, roll & acceleration rates Star Trackers Detects star patterns & magnitudes Precisely aligns IMUs Guidance Computers (GC) Uses IMU data to: Compute state vector estimation Compute attitude estimation Aaron Shabazz

IMU accuracy is vital to mission success IMU drift bias is 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 > * Use 2 IMUs on spacecraft and rover Probability that at least 1 IMU works > * Data from Honeywell IMU spec sheet Spec Sheet - Aaron Shabazz

Attitude Control System Pre-loaded trajectory/attitude data in guidance computer (GC) IMUs provide actual estimate of attitude GC uses residual of nominal and actual attitude data to: Run data through filter for best data Convert error data to steering & thrust commands Desired attitude is achieved Aaron Shabazz

Center of Gravity - Landing
Center of gravity determined by worst-case dynamic conditions on landing The “tripping scenario” is the most difficult scenario to maintain stability upon landing Mike Sloan Mike Sloan

Using a rigid landing structure, the critical limit for CG height is 3.4 m The safety limit is 1.1 m This height is achievable if the rover is placed horizontally on the landing structure Mike Sloan Mike Sloan

For a landing-on-wheels scenario, the CG tolerances are much tighter Primary danger comes from descent engines hitting the surface Critical limit for CG height is 1.9 m Safety limit is 0.1 m Mike Sloan

Center of Gravity - Driving
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 m Martian required CG height - 2.5m Mike Sloan

Center of Gravity - Driving
Mars CGrequired height > Moon CGrequired height Any vehicle geometry that can safely drive on the Moon can safely drive on Mars Mike Sloan

Transit Configuration 1
Mike Sloan Daren McCulley

Transit Configuration 2
Mike Sloan Daren McCulley

Propulsion System Requirements
Launch a specified payload to the moon Expend practically all its fuel upon arrival Landing engine must be able to restart 2 or 3 times The total mass of the propulsion system must be as low as possible Maximum thrust of the landing engine must be 45 kN Reuel Smith

Assumptions Made Changes in Velocity LOI - Lunar Orbit Insertion
Retro Engine LOI: m/s LOI - Lunar Orbit Insertion DOI: 20 m/s DOI - Descent Orbit Insertion PDL (tangent): m/s PDL (hover): m/s PDL - Powered Descent Landing RCS Thrusters RCS (landing): 150 m/s RCS - Reaction Control System Reuel Smith

Assumptions Made Propellants: The module runs on one specific fuel/oxidizer mixture Other Assumptions Payload: kg Ae/At: 54 for all propulsion stages Inert Mass Fraction: for all propulsion stages Max RCS Thrust: 445 N per thruster RCS Thruster Count: 16 Spacecraft Apollo- < Reuel Smith

Fuel Analysis MLanding Engine + MRCS Thrusters + MGimbals + MAvionics
+ MWiring + MThrust Structure __________________ MPropellant System Propellant 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 Reuel Smith

Fuel Analysis Reuel Smith

Possible RCS Configurations
RCS thrusters may be placed along the center of mass It may be possible to do a 12 thruster RCS by removing four roll thrusters Reuel Smith

RCS Thruster Risk Analysis
Assumptions: 95% Mission reliability, no fault tolerance (crew survival not dependant on RCS) Two configurations considered: 12 engines and 16 engines Must be able to maintain complete 3-axis control of the landing vehicle Jason West

RCS Thruster Risk Analysis
Scenario A: 12 engines, none fail Scenario B: 16 engines, up to 2 engines can fail Scenario A Scenario B Required Engine Reliability 0.9957 0.9469 5% less required engine reliability for 16-engine system Jason West

Next Step Examine using two different sets of propellants for the RCS and Landing Engine Modify mixture ratio for NTO/UDMH to lower the propellant system’s mass Examine using monopropellants for RCS In-Space Propulsion analysis Reuel Smith

Landing Requirements 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 Rahkiya Medley

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

Lander Options Crushable legs Joint legs Honeycomb insert Pivot feet
80 kg/leg (Al wrought 2024-T4 and SPIRALGRIDTM) Joint legs Torsion spring joint TBD kg/leg Rahkiya Medley

Crushable Legs Modeled as a mass damping system Impulse force ~81 kN
Increasing leg length increases landing footprint As the leg length increases, critical buckling load decreases Pcr α 1/L2 Rahkiya Medley

Lander - Future Work Model of joint leg
Optimum placement of landing legs for both configurations Optimum crush strength of SPIRALGRIDTM Fuel tank/nozzle support structure Rahkiya Medley

Lunar Mapping The surface of the moon will be mapped by the 2008 Lunar Reconnaissance Orbiter Both optical and topographical maps will be taken These maps can be used to assist in landing and surface navigation Optical resolution is 0.5 m per pixel Vertical (altimeter) resolution is 10 cm over a 5 m sample Dr David Smith, Goddard Space Flight Center Mike Sloan

Landing Control System 3 Microwave Scan Beam Landing Systems (MSBLS) Transponders/receivers that find slant range, azimuth, and elevation relative to moon base Gives very accurate position info to GC to compute state vector GC selects middle values of 3 ranges, azimuths and elevations Angle and range data are used to compute steering commands 2 Radar Altimeters Measures absolute altitude Both measurements are averaged Can derive vertical velocity and match with IMU measurements GC checks nominal and actual approach velocities to ensure safe & soft landing Aaron Shabazz

Landing Point Localization Assume Moon Base has 4 m high antenna LOS is about 3.73 km A 3.73 km radius about the moon base defines our desired landing zone Aaron Shabazz

Landing Point Localization Rough Estimate Landing Accuracy Average all off-target data after Apollo 12 Estimate landing accuracy = km Apollo 12 Apollo 14 Apollo 15 Apollo 16 Apollo 17 Off target data 0.16 km 0.05 km 0.21 km 0.20 km 0.55 km Off target data – Spring 2004 ENAE 484 CDR Slide # 239 Aaron Shabazz

Distance Between Landing Target and Moon Base Roughly Twice the Estimated Landing Accuracy for Safety Even in Worst Case Scenario, Rover will have LOS Communication w/ Moon Base after Touchdown Aaron Shabazz

Landing Hazard Avoidance
Landing requirements Must be able to survive a 0.5 m boulder and a 10o slope Larger boulders and slopes must be detected and avoided Digital elevation map (DEM) generation options Stereo camera system 6 - 7 m error Stereo from lander motion (more reliable option) Joanneum Research: Vision-Based Navigation for Moon Landing Scott Walthour

Stereo From Lander Motion
Scott Walthour Scott Walthour

Hazard Detection Hardware
One CCD Camera 16 Mb memory for onboard processing DSP board TBD Laser Altimeter LaserOptronix ALTM400 ( m range, cm accuracy) Digital Elevation Map (image source: < Scott Walthour

Hazard Detection Performance*
CCD Array 512 x 512 pixels Focal Length 10 mm Footprint (200 m) 100 m Ground and DEM Resolution 0.2 m Required Pointing Accuracy 1.4 deg Processing Time ~ 10 to 30 sec Required Inertial Sensing Accuracy (90% overlap) 10 m Joanneum Research: Vision-Based Navigation for Moon Landing *From similar lunar mission (2 hr orbital period, 0.5m obstacle requirement) Scott Walthour

Lander Stereo Considerations
Hovering could cause errors in inertial navigation Requires position recalibration Calibration from previous DEM Not likely without a DEM from orbit Self-calibration Errors not significant compared to DEM errors (at least 10 – 20 cm) Scott Walthour Joanneum Research: Vision-Based Navigation for Moon Landing Scott Walthour

Landed Mass Analysis Delta IV Heavy delivers 9950 kg into Lunar Transfer Orbit (LTO) Used Available Mass Estimating Relationships, Fuel Properties, ∆V Values, and Rocket Equation to determine rover’s mass when landed Rover Mass = Mass of Landed Package – Mass of Main Propulsion System (varies) – Mass of RCS (~ 250 kg) – Mass of Landing Equipment (~ 250 kg) Timothy Wasserman

Single Stage Single main engine used for all phases of flight
Standard Landing Structure Propellant Combination Rover Mass (kg) LOX/LH2 3000 N2O4/MMH 2710 N2O4/UDMH 2610 LOX/CH4 2530 LOX/RP-1 2360 Timothy Wasserman Timothy Wasserman

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

Two Stages: Reuse Cryogenic Tanks
Assumes SPRITE uses fuel cells Assumes fuel cell reactant tanks (capacity ~ 700 kg) can be used for storing 2nd stage propellants Timothy Wasserman 1st Stage Prop 2nd Stage Prop Surface Mass (kg) LOX/LH2 3030 N2O4/MMH 2840 Timothy Wasserman

Comparison of Best Two Staging Options
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: Timothy Wasserman

Launch Mass Budget Design Group Mass
Transit Power, Propulsion & Thermal 5800 Surface Power, Propulsion & Thermal 700 Loads, Structures & Mechanisms 1250 Crew Systems Mission Planning & Analysis 300 Avionics Timothy Wasserman

Surface Operations

Mission Planning Requirements
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 Chris Hartsough

CONOPS Overview (Deployment to Nominal Operations)
Chris Hartsough

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 *Possible robotic arm operations everyday Chris Hartsough

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 Daniel Zelman Loop A Loop B Daniel Zelman

Route Options Drive out 100 km Drive along arc for 50 km
Return along different 100 km path Arc Never more than 125 km from base Inverted Arc Never more than 100 km from base More scientific possibilities than previous routes Daniel Zelman Arc Inverted Arc Daniel Zelman

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

Structures

Pressure Hull Sized to contain the astronauts, crew systems and avionics Designed to handle launch loads, pressure loads, and kick loads Two options considered: Prolate Spheroid and Cylinder with Ellipsoidal Endcaps Mass is the primary driving factor Evan Ulrich

Prolate Spheroid Configuration Rib/Stiffener Stringer Shear panel
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 Evan Ulrich

Prolate Spheroid: Analysis
Applied Loads: Internal Pressure (2 atm) Punching force (3 kN) Method of Analysis: Skin idealized shell, 4 mm thickness Point constraint 6g axial, 2.5g lateral Skin, Rib, Stringer Approximated by ~ 1.2 million finite elements Evan Ulrich

Cylinder with Ellipsoidal Endcaps (CEE)
Optimal number of ribs is 4 Need for external mounts may increase number of ribs 8 stringer configuration allows for ease of hatch/window placement Provides sufficient structural support All stringers have hollow circular cross sections -Shear panel -Stringer -Rib/Stiffener Evan Ulrich

CEE: Analysis Applied Loads: Method of Analysis Method of Analysis:
Internal Pressure (2 atm) Punching force (3 kN) Method of Analysis: Skin idealized shell, 4 mm thickness Point constraint 6g axial, 2.5g lateral Method of Analysis Skin, Rib, Stringer Approximated by ~ 1.2 million finite elements Evan Ulrich

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 Michael Koszyk

Micrometeoroid Flux vs. Mass Calculate micrometeoroid flux Surface area ~ 36 m2 Mission duration ~ 10 days PNP ~ 0.996 Flux = (impacts/m2/yr) Critical mass ~ g [Vanzani, et al. Micrometeoroid Impacts on the Lunar Surface. Lunar and Planetary Science XXVIII, 1997.] Michael Koszyk

Design variables Hull properties MLI properties Hull/MLI spacing Michael Koszyk

Window Materials Material Density (kg/m3) Elastic Modulus (GPa)
Flexural Strength (MPa) Compressive Strength (MPa) CTE (10-6/°C) High-Strength 2010 37.2 18.6 50 0.6 Ultra High-Strength 38.3 56.2 207 0.5 Castable 220 2090 - 11.35 1.7 Ceradyne Thermo-Sil® Fused Silica Materials < Michael Koszyk

Window Requirements Curvature of material required
Filter out harmful radiation 0.1% Iron Oxide fused into glass Anti-reflective coating necessary Structural analysis underway Michael Koszyk

Other Required Structures/Mechanisms
Fairing structure Propulsion system structures All secondary structures Antennae Thermal regulation Mechanisms/Special Structures Hatches/suit interface Surface deployment On-orbit deployment Stage separation Emergency/Rescue Steering David Gruntz

Structures Summary SF Structure Loading Condition Applied Load (MPa)
MOS 2 Ribs/Stringer Launch 380 Pressure Hull 2 atm 550 Landing Structure 80 kN (I) TBD Wheels 3 kN (PL) 530 0.127 Chassis/Suspension 225 kN (I) 330 0.03 1.5 Avionics Support Thermal Regulation Support Structure 3 Pressure Vessels Primary Structure Secondary Structure Launch – 6g axially along Delta IV, 2.5g laterally (I) – impulse load (PL) – point load David Gruntz Rahkiya Medley

Mobility Systems

Drive System Overview Surface propulsion’s Level 1 requirements
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

Surface Propulsion Calculation
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 Calculate force of gravity on incline of 30° (for peak power) Maximum load is the sum of friction on tires and normal force Meets Level 1 requirement (I9) Raja Krishnamoorthy Raja Krishnamoorthy

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

Drive System Requirements
Assumptions Calculations 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

AC vs. DC Motor 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

AC - DC Mass Comparison For a 15 kW motor the masses are as follows:
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, SMA4, 355SMA, 450LG4 Raja Krishnamoorthy

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) Raja Krishnamoorthy

Axle vs. Individual Wheel Drive
4 MOTOR 2 MOTOR A motor for each wheel Used 4-wheel case Requires less power Provides less torque A motor for each axle Used 2-axle case Requires more power Provides more torque Raja Krishnamoorthy

Risk Analysis for Drive Setups
4 MOTOR Can tolerate 2 failures*: ways ↔ A-C, B-D, A-D, C-B R4 + 4R3(1-R) + 4R2(1-R)2 2 MOTOR Can tolerate only 1 failure: A or B R2 + 2R(1-R) R = e-t/MTBF = t = 25 hrs, MTBF = 40,000 hrs *Considered simple failure without wheel lock Raja Krishnamoorthy

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

References DC or AC Drives? A guide for users… Motor Formulas, 1997
Documentation/Documents/DCorAC.pdf Motor Formulas, 1997 Torque Capabilities of AC and DC Drives horsepower/ Adjustable Speed Drives Lunar Rover Operations Handbook Raja Krishnamoorthy

Wheels Assumptions Width vs. Power Diameter > 1 m
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

Power vs. Width Width = 0.3 m Pyungkuk Choi

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

Outer Rim Force applied to the rim
Modeled as curved beam under elastic bending Assumptions Rectangular cross section Constant radius of curvature Bending moment due to point load remains perpendicular to the radius of curvature Pyungkuk Choi Pyungkuk Choi

Number of Spokes vs. Rim Thickness
Titanium (10% Vanadium) Density = 4650 kg/m3 Tensile Strength = MPa Safety Factor = 2 Spokes Rim Thickness (mm) Inner σ (MPa) Tensile MOS Mass of one wheel Total Mass (kg) wheels (kg) 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

Wheels - Future Work Tires Cross slope loading
Different wheel configuration Wheel protection Pyungkuk Choi

Chassis / Suspension Struts connect to rib/stringer primary structure
Spring / Shock Absorber Wheel Mount Struts connect to rib/stringer primary structure External chassis if necessary David Gruntz

Chassis / Suspension Factors considered
Load transferred by suspension Vertical displacement of the vehicle Must absorb landing with residual velocity of 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

Suspension Models Lateral Torsion Bar Axial Torsion Bar Linear Spring
Modeled as 5,000 kg mass attached to a m moment arm Axial Torsion Bar Modeled as 5,000 kg mass attached to a 0.25 m moment arm Linear Spring Modeled as 5,000 kg mass atop a linear spring David Gruntz David Gruntz

Torsion Bar vs. Linear Spring
Torsion bars transfer similar loads Linear spring looks like ideal choice at this point David Gruntz

Loads Transferred to Chassis
Type Vertical Displacement (m) “Fall” Force (kN) Landing Spring Constant Force (kN) Deflection (m) 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 Lateral Torsion 100 55 51 40 32 0.14 0.16 0.23 1500 kN-m/rad 1000 kN-m/rad 550 kN-m/rad Axial Torsion 85 60 74 50 35 0.15 50 kN-m/rad 20 kN-m/rad 8 kN-m/rad David Gruntz

Initial Suspension Sizing
Titanium Ti-6Al-4V High specific strength (σyld/ρ) allows for a strong, lightweight chassis Initial chassis/suspension sizing with Titanium structure and steel springs 20 kg 140 kg Load Condition Max Stress (MPa) MOS Collision 330 0.03 “Fall” 280 0.2 Landing 200 0.7 Launch* TBD * Will depend on how rover is integrated w/ fairing David Gruntz

Crew Systems

Consumables Summary Oxygen – 23.0 kg Nitrogen – 1 kg
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 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 Larson, Wiley J., ed. Human Spaceflight: Mission Analysis and Design. John Mularski

Consumables Summary Water – 250 kg Food – 40 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 Larson, Wiley J., ed. Human Spaceflight: Mission Analysis and Design. John Mularski

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 8.3 psi total pressure 37% Oxygen 63% Nitrogen R=1.4 R=1.4 < Lunar Base 14.7 psi total pressure 21% Oxygen 79% Nitrogen Alan Bean - < & EVA Suit 3.5 psi total pressure 100% Oxygen John Frassanito and Associates – < Larson, Wiley and Linda Pranke, ed. Human Spaceflight: Mission Analysis and Design Michael Badeaux

Atmospheric Composition
Man-Systems Integration Standards – NASA-STD-3000 < 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

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** 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 < Glatt, C.R. “WAATS – A Computer Program for Weights Analysis of Advanced Transportation Systems.” NASA CR Aerospace Research Corporation Michael Badeaux

Temperature/Humidity Control
Requirement L2: All crew interfaces shall adhere to NASA-STD-3000, Man-Systems Integration Standards Ideal Temperature ranges from oC SPRITE Cabin Temperature – 23 °C Ideal Humidity ranges from 4-16 oC Excess heat can be used to heat water Wieland, Paul. Designing for Human Presence in Space NASA RP < Michael Badeaux

Temperature Control Simple Small Scale Passive Insulating Materials
Little Maintenance Insulating Materials Electric Heaters Heat Pipes Active Complex Large Scale High Maintenance Cold Plates Heat Exchangers Re-router Heat Rejection Freudenrich, Craig “How Space Stations Work” - < Michael Badeaux

Carbon Dioxide Removal
Requirement L2: All crew interfaces shall adhere to NASA-STD-3000, Man-Systems Integration Standards Removal Reduction Regenerable Open Loop 2BMS EDC LiOH Sabatier Weight 48.1 kg 44.4 kg 40 kg 76 kg Volume 0.26 m3 0.071 m3 0.005 m3 0.14 m3 Heat N/A .336 kW .268 kW Power Required 0.23 kW kW AC kW DC 0.012 kW .05 kW Temperature oC 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

EDC Regenerable system Reacts H2 and O2 with CO2 inside and electrochemical cell CO O2 + 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 kW AC, kW DC Mass = 44.4 kg; Volume = m3 Requires supply of H2 and O2 Generates heat Shawn Butani

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 = m3 Power required = kW Lunar Module Environmental Control System. Historic Space Systems. < Shawn Butani

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

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 Hz Visual: Consists of a light array panel Red – Emergency Yellow – Cautious Green – Nominal < < Michael Badeaux

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 dB is allowable, duration £ 2 min - Hearing Protection Devices - Provided for noise levels ³ 85 dB Maximum Noise Level - Change in sound pressure level ³ 10 dB £ 1 sec - Impulse noise shall not exceed 140 dB peak pressure level Man-Systems Integration Standards – NASA-STD-3000 < Michael Badeaux

Contamination and Particulate Control
Air filters High Efficiency Particulate Arrestance (HEPA) filter 99.7% efficiency on 0.3 microns NASA Standards Section 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 < Michelle Zsak

Contamination Control Wipes
Biocide Disinfecting food and waste systems Biofilm Control Controls formation of Biofilm inside surface of fluid lines Cleaning Implements Provides means for dislodging and collecting dirt/debris Detergent Indoor cleaning Dry Toilet tissue Utensil Cleaning Sanitizers for post meal cleaning Vacuum “Man-Systems Integration Standards.” NASA STD < Michelle Zsak

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

Trash Management 2-Man Crew, 1-wk Mission Mass (kg) Volume (m3) Total
9.1 0.202 Food 4.5 0.16 WCS Supplies 4.6 0.042 Ways to store trash Free standing trash receptacle Storage compartment built into structure Trash compactor to minimum trash space Michelle Zsak

Radiation Sources Galactic Cosmic Rays Solar Particle Event Duration
Near Constant 1-3 days Composition 85% Protons 14% Alpha 1% Nuclides 90% Protons 10% Alpha Flux Density (photons/cm2-sec) 0 - 1 max ~2 max ~106 Energy Levels (MeV) max ~1011 max ~104 “Man-Systems Integration Standards.” NASA STD < Michelle Zsak

Lifetime Limits: Blood-Forming Organs (BFO) 5 cm depth
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) cm depth Gender Age 25 35 45 55 Male 150 rem 250 rem 325 rem 400 rem Female 100 rem 175 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

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.” < Sonja, Baristic. “Making Mars Bricks for Long Term Red Planet Stays.” < Michelle Zsak

Shielding Options Possible Shielding Aluminum Currently used
Creates neutrons during nuclear interaction that increase exposure Polyethylene (CH2) without water Shields more than Aluminum since it is Hydrogen rich Polyethylene with water Shields 20% more than Aluminum since it is Hydrogen rich Must consider mass budget Michelle Zsak

Aluminum vs. Polyethylene
Solar Minimum 1977 Solar Maximum 1970 Thickness (g/cm2) Dose Equivalent (rem/yr) Al CH2 95 1 91 81 2 88 83 5 79 71 10 69 57 15 54 41 25 46 35 75 43 32 Thickness (g/cm2) Dose Equivalent (rem/yr) Al CH2 34.5 1 33.7 32.7 2 32.9 31.2 5 30.7 27.2 10 27.8 22.6 15 22.8 16.4 25 20.0 14.4 75 19.4 13.7 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

Possible Radiation Shielding Plan
Shield all sides exposed to radiation 0.4 cm aluminum hull provides shielding Polyethylene shielding specific mass ~10 kg/m2 - with surface area of 39 m2 ~390 kg 3 cm thickness of water from fuel cells provides additional shielding for Solar Particle Event (SPE) Michelle Zsak

Fire Suppression Type Liquid Density Volume Fraction Comments Halon 1301 1570 kg/m3 0.20 Highly effective CO2 758 kg/m3 0.62 Toxic in high concentration Can be cleaned by rover Oxygen masks required for crew during fire suppression Extra CO2 scrubber can be carried for post fire clean-up Halon 1301 decomposes into toxic products which must be filtered out post fire Friedman, Robert: “Fire Safety in Extraterrestrial Environments.” Lewis Research Center, May 1998. John Mularski

Internal vs. External Suits
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 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 < John Mularski

Layout John Mularski John Mularski

Layout Current cabin volume = 21 m3
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 John Mularski John Mularski

Layout Bunks fold to provide access to external suit and stowage
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 John Mularski

Visual Display (VD) VD must be at least 13 in, preferably > 20 in
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

Windows Inputs Output Problems…
Finding minimum window dimensions for navigational purposes Inputs Cabin height = 2.1 m vessel = 3 m diameter (tires add 0.5 m from ground) 95th male sitting height eye level = 135 cm Line of sight = 24.7o +/- 2.4o Eye movement laterally: 35o max, 15o optimum  25o (easily with head moment range) Output Navigator can see the ground m ahead of the rover Minimum window size (mass constraint) = 42 cm length, 40 cm width Problems… Stringers will divide window Curvature of rover Shawn Butani

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 Hull Michael Badeaux Shawn Butani

Front Display Window Astronauts sit 16 in. from windows and MFD Seats
MFD = .69 m (~27 in) NAV-PRI/AUX = .56 m (~22 in) NAV-DATA = .431 m (~17 in) Seat separation = .24 m Control panel includes : Steering system : Throttle (SDOF), L & R steer (SDOF), Lift Break Avionics : input from driver, indicators, sensors (wheels, pitch and roll, speed, etc.) Seats Hull Evan Ulrich Shawn Butani

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

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) Strength (psi) Modulus (psi) Tensile 9,500 320,000 Flexural 15,000 375,000 Compressive 12,000 240,000 Polycarbonate Specifications, Shawn Butani

Calculating Shield Dimensions
Density = lb/in3 = 1.2 g/cm2 Designed one shield to fit two 95th percentile males with +/- 10 cm for each dimension Designed as a rectangular shaped enclosure to calculate maximum mass Mass = 260 kg In the future will design to better fit the suit and optimize mass 95th percentile male (cm) A – Height 191.9 C - Width 66.0 D – Depth w/ PLSS 68.6 NASA-STD-3000, Volume 1 section 14. Shawn Butani

Crew Systems Future Work
EVA checklist Health monitoring Interior stowage Docking system EVA support Controls and displays Shawn Butani

Intermission

Surface GNC

IMU accuracy is vital to mission success IMU drift bias is 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 > * Use 2 IMUs on spacecraft and rover Probability that at least 1 IMU works > * Data from Honeywell IMU spec sheet Spec Sheet - Aaron Shabazz

Works Still in Progress GNC Thermal Control Determining which computers to use Determining number of computers needed Aaron Shabazz

Navigation and Guidance on Moon Surface
SPRITE shall be capable of navigating Within 100 m of target Both day and night Absolute Navigation constraints on moon Communication limited to only base, earth and L2 satellite LOS, and natural landmark barriers No medium for sound to travel through Navigation method w/ Moon Map Trade study Accuracy (m) Method Constraint Celestial Sun and Earth Tracker 300 At least 600 obs. Landmark VIPER 180 Needs assistance at night Low Frequency Radio Loran Submarines 100 2 or more beacons Borenstein, Johann J., H.R. Everett, and Liqiang Fang. Navigating Mobile Robots. Wellesley, MA: AK Peters, 1996 Ralph Myers

Navigation and Guidance on Moon Surface
Use landmark for absolute reference and dead reckoning sensors for relative reference Errors in the dead reckoning sensor will determine the distance needed before a landmark is needed for correction update Vehicle & Landmark Latitude and Longitude Gyro Roll Accel. X Scan horizon or predetermined landmark Build DEM and Compare to lunar Map surface Gyro Pitch Accel. Y Gyro Yaw Accel. Z Real time Calibration Accelerometer Compensation Compare values to Lunar Map Odometer Slippage detection Torque sensor Myers, Ralph Left Front Right Front Left Rear Right Rear Ralph Myers

On Board Direct Human Control
Drive by wire will control steering, acceleration, and braking through a feedback loop Have to reduce odometer errors caused by slippage Assuming driver has to control 4 independently motored wheels Assume Ackerman Steering to comply with 10 m turn radius requirement SPRITE shall incorporate sensors to allow positive diagnosis of credible failures in safety critical systems Ralph Myers

Minimize Odometry Error
Specifications for odometry accuracy Encoders Resolvers Controllable Speed Range 0.1 rpm to 10,300 rpm 30 rpm to 15,600 rpm Counts Per Resolution 32,640 16,384 Signal Periods Per Revolution 2048 1 Accuracy Range (arc-minutes) 1 to 1.5 7 to 15 Tolerable Shock Level (gs) 5 50 Operating Temperature Range (ºC) 0 to 100 -55 to 175 Ralph Myers

Robot Arm Control Sensor Parallel to human hand Location
SPRITE shall provide capability for crew to interact with environment without EVA Teleoperator should be able to manipulate the arm Tactile sensors provide feedback to the operator Sensor Parallel to human hand Location Tactile array sensor Give feel of object’s shape Outer surface of finger tip Finger tip force-torque sensor Determine how operator manipulates object Near finger tip Finger joint angle sensor Position of robots manipulators Finger joints or at motor Actuator effort sensor Motor torque as wrist movement At motor or joint Dynamic tactile sensor Vibration, stress to tell if object is being fumbled Ralph Myers

Surface Obstacle Avoidance
SPRITE must traverse 0.5 m obstacles, 20º cross-slope, 30º forward slope Must detect hazardous terrain Derived detection requirements Minimum look ahead distance - 4 m Based on minimum stopping distance Maximum look ahead distance - 13 m Based on tightest turning radius Stereo camera strategy chosen Scott Walthour

Surface Obstacle Avoidance
Camera parameter derivation assumptions Maximum deceleration: 0.45g (comfortable automobile deceleration) Obstacle detection rate: 1 Hz (DEM updated every second) Maximum velocity: 2.77 m/s (10 km/hr) Resolve 0.5 m object at maximum look ahead distance SPRITE width: 2 m Scott Walthour

Minimum Look Ahead Distance

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

Camera Vertical Field of View (VFOV)
VFOV dependent on: Vertical location of sensor Negative obstacles* need sensor as high as possible Assume = 3 m (located on top of SPRITE) Maximum obstacle size to be seen at 13 m Assume = 1 m Scott Walthour *Negative obstacles – ditches, craters, etc. Scott Walthour

Derived Obstacle Detection Requirements
Minimum Look Ahead Distance 4 m Maximum Look Ahead Distance 13 m Horizontal Field of View 103 deg Vertical Field of View 29 deg Angular Resolution* (mrad/pix) 1.88 (H) x 1.75 (V) Minimum Image Resolution (pix) 954 (H) x 290 (V) Update Rate 1 Hz Stereo Camera Locations 3 m vertical Camera Separation 2 m baseline Night Operations Headlights *Horiz:10 pixels on 5th %ile female width (24.5 cm) at 13 m Vert: 6 pixels on .5 m diameter ditch at 13 m < Scott Walthour

Obstacle Detection Future Work
Choose COTS* cameras Resolution CCD, CID, Vision chips Determine computational requirements *COTS – commercial off the shelf Scott Walthour

Network Data Bus Network 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 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 A) – serial bus < Scott Walthour

Network Data Bus Spacewire Advantages Disadvantages
High Data Rate (Mbps) typical (400 max) Lightweight 0.06 kg/m Scalable Radiation Tolerant BER = 10-14* Full Duplex Not inherently redundant Requires routers to ensure redundant paths - Increases complexity of the network *Bit Error Rate < Scott Walthour

Example Network Scott Walthour Scott Walthour

Network Data Bus Number of routers dependent on number and type (e.g. pressure sensor) of nodes Desire redundancy Divide pressure sensors on multiple routers in case of router failure Future work: Organize SPRITE’s data network Scott Walthour

Communications

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 Jay Kim

High Definition TeleVision
HDTV specs 1920 pixels by 1080 lines 30 frames per second 3 primary colors (red, blue and white) 8 bits for each color Uncompressed data rate at 1.5 Gbps Compression technique MPEG 1: Standard for Video CD MPEG 2: Standard for broadcast-quality television Compression rate up to 20 Mbps Comparison of different displays Jay Kim Jay Kim

From SPRITE To Earth Near side Far side Assumption: SPRITE is parked
Scenario 1 SPRITE is on near side and Earth is in LOS Communicate directly using antenna Transmission rate 20 Mbps at 1 channel Bidirectional transmission of 10 Mbps of digital data Uplink = 30 Mbps (from Earth to SPRITE) Downlink = 50 Mbps (from SPRITE to Earth) Trade studies of link budgets Frequency selection Antenna selection Link budget constraints Link margin 3 dB – 6 dB Near side Far side High Gain Antenna Low Gain Antenna Jay Kim Jay Kim

Link Budget Initial assumption Effect of changing diameter
Ka band: widely used in spacecraft communication Diameter of antenna: 1 m High gain antenna: precision in targeting Transmitter power: 20 W Slant range: 400,000 km (Apoapsis of Moon) Receiver antenna: Deep Space Network (34 m) Effect of changing diameter David G. MacDonnell, “Communications Analysis of Potential Upgrades of NASA’s Deep Space Network” Akin, Dave. ENAE483 Link Budget Spreadsheet Jay Kim

Link Budget Effect of changing transmitter power
Jay Kim Effect of changing transmitter power Diameter of antenna size: 0.5 m Transmitter power: 1 W Mass: TBD Operating frequency: 15 – 25 GHz Link margin: 3dB – 6dB Jay Kim

SPRITE To Base Transmitting antenna in SPRITE
UHF Band: widely used in short distance communication Diameter of antenna: 0.5 meter Transmitting power: 1 Watt Slant range: 150 Km Data rate: 50 Mb/s (HDTV) Receiving antenna in base Same antenna as transmitting antenna Takes advantage of learning curve Operating frequency: 1 – 1.5 Ghz (UHF Band) Link Margin: 3 – 6 dB

Emergency In case of emergency SPRITE communicates to Base
Use low gain antenna Reliable signals No pointing required Link budget (TBD)

Flying Locator and Assistance Requesting Equipment (FLARE)
To be used in the event of a regular communications pathway failure Launch a small communications package (10 kg, 25 cm2) to provide temporary link between the rover crew and base. Small solid rocket motor for propulsion Equipment based on amateur radio microsatellite technology Motor 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: < AMSAT Echo Information: < Timothy Wasserman

Future Work for CDR Rover to EVA communication Far side communication
Need to work with Crew Systems Determine requirements for EVA suit communication system Far side communication Satellite communication

Power Systems

Power & Energy: Requirements and Budget
Power and energy budget has been created to establish a buffer between requirements and available power and energy Current assumptions Time for avionics, crew systems, thermal, and science missions power consumption have been estimated at full time usage Jason West

Power Requirements Overview
Power [kW] Energy [kW-hr] SPRITE Total 55.6 2276 Surface Propulsion Nominal required 36.5 1277.5 Peak required 55.5 Continuous Avionics Communications (SPRITE to Earth) 0.02 3.8 Communications (SPRITE to Base) Communications (SPRITE to EVA) IMUs 0.032 6.1 Star Trackers 0.01 1.9 GNC Computers 0.015 2.9 Avionics total 0.117 22.5 Crew Systems CO2 removal 0.012 2.3 Jason West

Power: Requirements and Budget
Required Power [kW] Budgeted Emergency Surface Propulsion (max) 36.5 (55.5) 40.0 (60.0) (0) Avionics .117 .250 TBD Crew Systems 1 1.0 Science Mission Thermal Miscellaneous Total (max) 37.617(56.6) 44.25(64.3) 1.0(1.0) Jason West

Power: Requirements and Budget
Max Power, 64.3 kW Nominal Power, kW Emergency Power, 1 kW Jason Mallare

Energy: Requirements and Budget
System Power Req (kW) Time (hr) Energy Req (kW-hr) Power Budgeted Energy Surface Propulsion (cruising) 36.5 35 1277.5 40.0 1400.0 (ascent) 55.5 1 60.0 Avionics .117 192 22.5 .250 48.0 Crew Systems 192.0 Thermal TBD Science Mission Misc. Total 1357.8 2276.0 192 hours represent 8 day, 24 hour/day usage Jason Mallare

Energy and Power: Bottom Line
Current bottom line energy/power budget for SPRITE 2276 kW-hr of energy 44.25 kW of nominal power with peak capabilities of 64.3 kW Current emergency power requirements SPRITE 72 kW-hr of energy – meets L1 requirement for 3 day emergency 1 kW Jason Mallare

Power Management & Distribution
Future Work AC vs. DC Centralized vs. Distributed power conversion Considerations: Ohmic losses in wires, hazard of 100+ V distribution throughout entire craft System Voltage 28 V vs. 100 V system Hyder, Wiley, Halpert, Flood, Sabripour. “Spacecraft Power Technologies” Jason Mallare

Energy Storage Technologies considered Primary batteries
Secondary (rechargeable) batteries Radio-isotope Solar arrays Fuel cells Jason Mallare

Minimum Temperature (oC)
Primary Batteries Advantages: Primary cells offer higher specific energy then secondary batteries Disadvantages: Non-rechargeable, low current, low specific power (W/kg) Chemistry Gravimetric Specific Energy (W-h/kg) Volumetric (W-h/L) 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 70 Ni-MH 72.0 246.5 14.29 -10 40 < < < Jason Mallare

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 Chemistry Gravimetric Specific Energy (W-hr/kg) Volumetric (W-hr/L) Power (W/kg) Minimum Temperature (oC) Maximum Cycle Life (cycles) Li-Ion 200 300 244 -40 60 500 Sodium Sulfur 240 304 350 2500 Li-Polymer 206 386 309 -20 < < < Jason Mallare

Batteries - Energy Storage
Jason Mallare

Batteries - Power Generation
Secondary Batteries Primary Batteries Jason Mallare

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 W of electricity < Phillip Adkins

Solar Cells Converts light to electrical energy
Estimated mass kg array Estimated area of 235 m2 Reasonable efficiency with high specific power Not favorable: Moon - restricts missions to the day side Mars - restricts missions to the day side Additional area needed for same power output < Phillip Adkins

Operating Temperature
Fuel Cells Type Specific Power (W/kg) Efficiency Operating Temperature (oC) Alkaline 50-70% Below 80 Proton Exchange Membrane (PEM) 35-60% 75 Direct Methanol 35-40% Phosphoric Acid TBD 35-50% 210 Molten Carbonate 40-55% 650 Solid Oxide 45-60% < < < Patel, Mukund R. Spacecraft Power Systems. Boca Raton: CRC Press, 2005 < Phillip Adkins

Fuel Cell Mass Calculations
Max Power estimated at 64.3 kW Assuming a specific power of 100 W/kg for the fuel cell reactor. Total Energy needed estimated at 2276 kW-hr Using alkaline fuel cells and assuming 70% efficiency for the fuel cells. Fuel Cell Reactor 640 kg Reactants 860 kg H2 and O2 tanks 420 kg Total Mass 1920 kg* * ~38% of total rolling mass Phillip Adkins

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): kg. Phillip Adkins

Thermal Control

Thermal Control Requirements
Maintain cabin temperature between 18.3 and 26.7ºC Cool electronics and motors so that equipment operates at peak efficiency Evan Alexander

Passive Thermal Control
Multi-Layer Insulation System (MLI) Several layers of thermal blankets used to insulate the cabin Advantages Lightweight Low thermal conductivity Disadvantages Conductive properties diminished in areas where layers meet Evan Alexander

MLI Use layers of Mylar due to it its density as well as its absorptivity and emmisitivity Decron netting used to separate layers of Mylar Material Features Thickness (µm) Emissivity Absorptivity Mylar Y9360-3M Aluminized TBD 0.03 0.19 Aluminized Backing 3.8 0.28 0.14 Teflon Gold Backing 12.7 0.49 0.30 Kapton Film 2.0 0.24 0.23 Evan Alexander

Aerogels Extremely lightweight form of insulation
Advantages Lighter than MLI system Lower thermal conductivity Disadvantages Structurally weak May be used in conjunction with MLI to improve insulation at joints Evan Alexander

Thermal Conductivity (W/m-K)
MLI vs Aerogels Category Type Density (g/cm3) Thermal Conductivity (W/m-K) MLI Kapton 1.42 0.12 Mylar 1.39 0.2 Teflon 2.15 0.195 Aerogels Silica 0.004 Resorcinol 0.6 0.06 Carbon 0.9 0.04 Evan Alexander

Active Thermal Control
Use Heat Pipes to cool electronics Radiators used to expel excess heat from cabin < Evan Alexander

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

Heat Pipes (cont.) Properties of possible heat pipe fluids
Temperature Range (°C) Heat Pipe Working Fluid Heat Pipe Vessel Material -200 to -80 Liquid Nitrogen Stainless Steel -70 to +60 Liquid Ammonia Nickel, Aluminum, Stainless Steel +5 to +230 Water Copper, Nickel Evan Alexander

Heat Pipes (cont.) Properties potential metals used Metals
Density (g/cm^3) Thermal Conductivity (W/m-K) Aluminum 2.7 205 Nickel 8.91 90.7 Stainless Steel 8.03 50.2 Copper 8.92 394 Evan Alexander

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 Evan Alexander

Science

Suggested Landing/Mission Zones
Crater Copernicus Crater Tycho Mare Orientale South Pole-Aitken (SPA) Basin Lunar and Planetary Institute, 2005 Chris Hartsough

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 Lunar Orbiter image II-162H3 Chris Hartsough

Crater Tycho Geographic Interest Diameter of 85 km
Average depth of ~4 km Central peak rising ~2.5 km Ease of landing Relatively young crater (one of the youngest large craters on near side) Deeper inspection of the Moon’s crust Lunar and Planetary Institute, 2005 Chris Hartsough

Mare Orientale Geographic Interest Diameter of ~950 km
Lunar and Planetary Institute, 2005 Geographic Interest Diameter of ~950 km Depth of ~3.2 km Multi-leveled mare Large iron concentration Ease of landing Half visible to earth Lunar and Planetary Institute, 2005 Chris Hartsough

South Pole-Aitken (SPA) Basin
Geographic Interest Diameter of ~2500 km Depth of ~12 km on average Largest known impact crater on the Moon Deposits of iron and titanium Possibility of water Deeper inspection of the Moon’s crust Lunar and Planetary Institute, 2005 Lucey et al., 1998 Chris Hartsough

Choosing Scientific Instruments for SPRITE
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 Completed steps Detail the mass and volume requirements for scientific hardware used in previous J-Class missions. Ryan Livingston

Instruments - Crew Experiments
Original Mass (kg) Returned Stored Volume (m3) Soil Mechanics Investigation** 15.7 TBD Solar Wind Composition Experiment 0.46 0.385 1.3e-3 Lunar Portable Magnetometer 1.18e-2 Far Ultraviolet Camera/Spectrograph 22 0.25 * Hand Tools to assist experiments = approx 50 kg ** includes ALSD (drill) Ryan Livingston

Instruments - Crew Experiments
Original Mass (kg) Returned Stored Volume (m3) Cosmic Ray Detector 0.163 0.13e-3 Transverse Gravimeter Experiment 14.6 0.0351 Lunar Neutron Probe 2.27 0.4 0.38e-3 Surface Electrical Properties 16 1 0.024 Ryan Livingston

Instruments - Deployed
Experiments Original Mass (kg) Returned Volume (m3) Passive Seismic Experiment 11.5 0.012* Heat Flow Experiment 9.9 0.023 Lunar Surface Magnetometer 8.6 0.044 Laser Ranging Retroreflector 36.2 0.135 Cold Cathode Gauge 5.7 0.012 Suprathermal Ion Detector Experiment 8.8 0.014 Solar Wind Spectrometer 5.3 0.007 * does not include foldable skirt Ryan Livingston

Instruments - Deployed
Experiments Original Mass (kg) Returned Mass (kg) Volume (m3) Lunar Dust Detector 0.27 TBD Active Seismic Experiment 11.2 Lunar Seismic Profiling Experiment 25.1 Lunar Atmospheric Composition Experiment 9.1 0.018 Lunar Ejecta and Meteorites Experiment 7.4 0.02 Lunar Surface Gravimeter 12.7 0.027 Ryan Livingston

Choosing Scientific Instruments for SPRITE
Future Steps 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 Ryan Livingston

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 - 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

REACH Configuration Several configurations considered Single arm
Two arms (one on each end of rover) Single arm on track David Gruntz David Gruntz

REACH Material Carbon/Epoxy resin ideal choice Lightweight and strong
Density (kg/m3) Yield Stress (MPa) Elasticity (GPa) Yield Stress/Density Ratio Comments Aluminum, wrought, 2024-T4 2800 325 73 0.12 Easy to machine Titanium alloy, annealed 4460 1230 TBD 0.28 Expensive, Too strong Carbon/Epoxy resin 1600 800 125 0.50 Extremely lightweight Beer, Ferdinand. Mechanics of Materials Werelety, Norman. ENAE423 Lectures - Composite Materials Carbon/Epoxy resin ideal choice Lightweight and strong David Gruntz

Initial Sizing & Mass Static analysis performed to determine size and mass 100 kg payload in Martian gravity (3.7 m/s2) Configuration Length Mass (per arm) (kg) Max Stress (MPa) MOS SF = 2 Material 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

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

Contingencies

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 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

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 Samuel Schreiber

PANIC - Habitat Habitat composed of space suit-like material for insulation and pressurization ~ .4 kg/m2 Habitat is inflated to 3.5 psi of 100% oxygen Provides an opportunity for astronauts to remove space suits, eat, rest, and discard waste 10 m3 minimal habitable volume for two 95th percentile American male astronauts with space suits. Reusable - Only one needed throughout return to base Samuel Schreiber

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 hrs; Resting – 78 hrs 3.2 kg oxygen needed to pressurize habitat at each stop (only 0.6 kg needed for respiration) Each cache: 7.6 kg water for traverse 7.9 kg Oxygen Tank 5.7 kg water for rest 0.8 kg oxygen for traverse 1.3 kg Water Tank 3.2 kg oxygen for rest Samuel Schreiber

PANIC - Mass Estimates Estimated Total Masses*:
26.4 kg in each cache + food + habitat 344 kg Total + food + habitat Habitat Mass: 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

PANIC - Concerns and Questions
Overall reliability and probability of failure Astronaut exhaustion, malnutrition and overheating Probability of excessive radiation dosage due to solar flare Amount of time spent on return – upwards of 8 days Carbon dioxide build up in habitat Heating Oxygen leaks in habitat Different rover paths provide differences in difficulty of a walk return Samuel Schreiber

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 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

TERROR - Trajectory 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

TERROR - Mass and Volume Estimates
Fuel 56 kg Oxidizer 103 kg Tank (Fuel) 4 kg Tank (Oxidizer) 5 kg Pressure Tank Wiring 10 kg Engine 16 kg Thrust Structure 1 kg Avionics Seats 25 kg Total Mass 224 kg Volume Fuel Tank 0.065 m3 Oxidizer Tank Engine 0.016 m3 Platform Total Volume 1.14 m3 Timothy Wasserman Dan Zelman

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 Higher Clearance 0.5 m Requirement Faster and More Powerful Newer engines Laurie Knorr

FEAR - Mass and Material
Aluminum Alloy 2219 Carbon Epoxy Density 2.84 g/cm3 1.6 g/cm3 Tensile Strength 359 MPa 600 MPa Yield Strength 248 MPa Modulus of Elasticity 73.1 GPa 70 GPa Shear Modulus 27 GPa 5 GPa Shear Strength 230 MPa 90 MPa Aerospace Specification Metals Inc - < Goodfellow - < Laurie Knorr

FEAR - Height Change Increase the size of the wheels
Mass of new wheel would be 1.69 times the mass of old wheel if the diameter is increased by 20 cm Change the suspension Mass increase minuscule Small loss in strength Chassis Fitting Lower Arm Upper Arm Damper LRV Operations Handbook, 1973 Contract NASA-25145 Lunar Rover Operations Handbook - < Laurie Knorr

FEAR - Motors Four motors: One on each wheel Old motors New motors TBD
36 V Input 0.25 hp Power 10,000 rpm New motors TBD Lighter More powerful Lunar Rover Operations Handbook - < Laurie Knorr

FEAR - Special Features
Drive back to base in less than 10 hours Folds up into 0.9 m3 space Attaches to the outside of SPRITE Lunar Rover Operations Handbook - < Laurie Knorr

Emergency Return Operations
ERV Mass Mass Launched Volume PANIC 344 kg* 13 kg TBD TERROR 224 kg 76 kg 1.14 m3 FEAR 240 kg 210 kg 0.9 m3 * Does not include food Laurie Knorr

Safety Time Advantages Disadvantages PANIC 182 hr Fairly simple
Time Advantages Disadvantages PANIC 182 hr Fairly simple Can be used in conjunction with other safety procedures Takes time to walk Very tiring on crew Increased probability of solar flare exposure TERROR 13 min Fast return to base Can work if one crew member is injured Very unsafe Complicated system FEAR 10 hr Crew exerts little energy Complicated detachment procedures Laurie Knorr

Program Timeline and Costs

Program Timeline Development/Production: 2005-2015 Launch: 2016
Requirement M8: The SPRITE shall be ready for initial lunar operations by 2016 Development/Production: 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

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 *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

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 Charles Bacon

Cost Analysis No specified limitations for cost budget
Heuristics from NASA Cost Estimation site: C(FY04 $M)= ami[kg]b* Manned Spacecraft (SPRITE) a = , b = .556 Liquid Rocket Engine (TERROR, landing engine) a = , b = .551 Other system cost estimates derived uniquely for each system *Derived from NASA Cost Models Charles Bacon

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

Cost Totals Cost (FY 04 $M) System 1940 SPRITE Vehicle (1) 1460
Landing Engine (1-stage LOX/LH2) 1460 Landing Structure TBD FEAR 144 PANIC TERROR 320 Robotic Arm 185 Launch 254 Charles Bacon

Cost Analysis Current total – $4.1 Billion
Requirement I6: The SPRITE design shall be designed to minimize life cycle costs Current total – $4.1 Billion Estimated final cost to launch: 1 SPRITE + 1 ERV Worst Case Scenario – TERROR: most expensive Cost will increase for another SPRITE, but not significantly (production is only 2-6% of total cost) Other costs include consumables and fuels (relatively low cost) Charles Bacon

Mission Operation and Data Analysis Cost
Mission Operational Costs - $154M/yr* Includes maintaining and upgrading ground systems, mission control; tracking; telemetry; command functions; mission planning; data reduction and analysis; crew training and related activities *assume investment price - $3.9B Charles Bacon

Cost Spreading Development and Production would occur from 2005-2015
Requirement M8: The SPRITE shall be ready for initial lunar operations by 2016 Development and Production would occur from Launch in 2016 Beta Function Non-Recurring costs over 11 years Recurring Costs take over in 2016. Charles Bacon

Cost Spreading Charles Bacon

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

Systems Integration Future Work…
Thorough itemized analysis for SPRITE to result in a reasonable projected cost Work breakdown timeline ( ) to illustrate key systems, milestones, and deliverables with projected due dates Costs of major systems still unknown Create System Block Diagrams Charles Bacon

The End

Small Pressurized Rover for Independent Transport and Exploration

Similar presentations

Presentation on theme: "Small Pressurized Rover for Independent Transport and Exploration"— Presentation transcript:

Similar presentations

About project

Feedback

Log in

Auth with social network:

Small Pressurized Rover for Independent Transport and Exploration

Similar presentations

Presentation on theme: "Small Pressurized Rover for Independent Transport and Exploration"— Presentation transcript:

Similar presentations

About project

Feedback