Justin Treptow Alex MorreseAlexis MendezCasselle Russell

John KlausRobert CooperThilina FernandoZoe Morozko

Faculty Advisors: Dr. Dan KirkGreg Peebles

Paul Martin Ben BurnettShriman ShivaDamian Harasiuk

1

Develop “Mission Plan” with in 4 years, minimum expense, with no military or heritage hardware

Establish a program that continues beyond our graduation

Bring useful talent into the program such as: Universities, Professionals, Graduates

2

Land in GLXP approved landing site near a Historical Artifact

Traverse 500m in a deliberate manner around a Historic Artifact

Transmit minimum dataset to earth On arrival Mooncast > 500MB Mission Complete Mooncast > 500MB

3

On Arrival Mooncast

• Descent Video > 2min. Near Real Time (NRT) transmitted ASAP• Camera shot vertically or horizontally• Arrival Video (Landing Site and Environment) > 30sec NRT transmitted ASAP• Post Arrival Images: Panorama, 3 self portraits depicting 40% of craft, GLXP Logo Cluster and Lunar surface• Transmit video message, email, text message• NRT video > 6 min, shot from various angles• HD video > 6 min, shot from various angles

4

Mission Complete Mooncast

NRT departure video > 30 sec, transmitted ASAP, shot of transition onto lunar surface

NRT looking back video > 30 sec, transmitted ASAP, shot of original landing site as vehicle moves away

Looking back photo depicting landing location NRT Mid-Journey video > 30 sec, transmitted ASAP, shot while

transversing Mid-Journey photo at 250 m from landing site depicting > 40% of rover

surface Journey’s End Panoramic at 500m with lander/landing sight in view Journey’s End Self Portrait portraying > 40% of transversing vehicle 3 Journey’s End photos of logo cluster and moonscape NRT video > 5min, shot from various angles HD video > 5 min, shot from various angles

5

Design Requirements

System

System Drivers

(Metrics)Analysi

s

Trade Study

Design Conflict

s

Big Picture

System

Level

Subsystem Level

Interface Level

System Drivers

(Metrics)Analysi

s

Trade Study

Sub -System

System Drivers

(Metrics)Analysi

s

Trade Study

Interface

6

Requirements Personnel AllocationLanding in GLXP approved landing site Zoe, Rob, John, Alex, Justin, Damian, Paul, Ben

Traverse 500m Thilina, Alexis, Shriman

Transmit Minimum Dataset Casselle

Note: There are no walls in our group

Project Start: 5 Seniors

Junior Design: 8 Seniors

Senior Design: 11 Members ( 8 seniors & 3 graduate level members)

Senior Design Current: 12 members (8 seniors & 4 graduate level members)

7

Launch Vehicle Trade Study

8

Properties Delta II Atlas V - 401 Zenit 3SL Falcon 1 Pegasus XL Dnepr Payload to LEO (kg) ~6,096 9,750 ~7,000 723 454 3700Cost (millions of dollars) 50 75 70 7 16 9.5Reliability (success rate, %) 99% 100% 97% 0% 87% 97%

Launch Location (distance to equator, km) ~3165 km (Cape)

~3165 km (Cape)

Equator (0 km)

~3165 km (Cape) The sky Baikonur,

Kazakhstan

Weight% Scores

Payload to LEO (kg) 0.4 4 10 8 3 1 6Cost (millions of dollars) 0.35 5 1 2 10 8 9Reliability (success rate, %) 0.2 9 10 8 1 7 8Launch Location (dist. to equator, km) 0.05 8 8 10 8 10 7

1.6 4 3.2 1.2 0.4 2.41.75 0.35 0.7 3.5 2.8 3.151.8 2 1.6 0.2 1.4 1.60.4 0.4 0.5 0.4 0.5 0.35

Overall 5.55 6.75 6 5.3 5.1 7.5

Converted Soviet ICBM

97 % success rate as ICBM

3 Stages

Payload to LEO: 3700 kg

Cost: $9.5 million

Launches out of Kazakhstan

at different inclinations

http://snebulos.mit.edu/projects/reference/launch_vehicles/DNEPR/Dnepr_User_Guide.pdf

9

Inclination = 50.5Inclination =64.5Inclination = 87.3Inclination = 98 10

Inclination (i) Orbit inclination with respect to earth’s equator

92% Drop in Payload Mass over Orbit Altitude Increase of 600 km (i = 50.5°)

Dnepr Users Guide pg 1911

Hohmann Transfer is the preferred method

Ion propulsion’s time scale is outside of mission requirements

SMART -1 ion propelled mission took 13 months

12

PurposeTo develop a analytic tool that accelerated analysis or

Hohmann Transfers

Input (Design Specifications)Orbit altitudeISPMass to be at final location

Output (System Drivers)Total & Incremental ΔVTotal Propellant Mass RequiredMass Ratio

13

Δv: velocity change required to alter the orbit of a space craft

From these results:▪ Velocity can be determined at a specific point

along the trajectory ▪ Mass of the fuel required will aid in tank

design requirements▪ Mass available for landing will allow for

payload mass estimates

14

Acquired Satellite Tool Kit (STK) Using the Astrogator Suite $100,000.00 value

Utilized by: Lockheed Martin, Northrop Grumman, NORAD, Florida Institute of Technology (Dana Carmody), ect.

Paul Martin is taking point with the simulation with oversight from senior design team members

Capacity to model complete mission simulation Launch – Mission Complete

Images provided by AGI.com

15

What orbit altitude corresponds to most efficient use of Dnepr’s lifting capabilities and propellant in spacecraft?

ΔV variations occur when transferring from a circular earth orbit to an elliptical earth orbit with apoapsis at L1 point.

300km

600km

900km

• L1: location where gravitational force from earth and moon are equal

L1

16

Orbit Altitude (km)

Circular Earth Orbit Velocity (km/s)

Required Velocity at Periapsis of Elliptic Orbit (km/s)

ΔV (km/s)

Dnepr Mass Delivered

(kg)

300 7.73 10.81 3.58 3700

400 7.67 10.73 3.06 3300

500 7.61 10.65 3.04 2800

600 7.56 10.57 3.01 1900

700 7.50 10.50 3.00 1200

800 7.45 10.42 2.97 700

900 7.40 10.35 2.95 300

• Maximum ΔV difference of 0.63km/s, 17.6% change

17

• Large change in orbiting mass for small change in ΔV.• Best to use Dnepr to launch between 3300kg and 3700kg mass into orbit 18

Fuel

Specific Volume (m3/kg) Oxidizer

Specific Volume (m3/kg)

Mixture Ratio (Ox:Fuel) Isp (s) MR

Liquid Bi- Propellants

Hydrogen 1.41E-02 Oxygen 8.77E-04 5 390 4.12

Kerosene 1.34E-03 Oxygen 8.77E-04 2.29 301 6.25

Hydrazine 9.96E-04 Oxygen 8.77E-04 0.74 313 5.83

MMH 1.15E-03 Nitrogen Tetroxide 6.90E-04 1.73 280 7.18

UDMH 1.26E-03 Nitrogen Tetroxide 6.90E-04 2.1 277 7.33

Liquid Mono-propellants

Hydrazine 9.96E-04 -  - - 199 16.00

Hydrogen Peroxide 6.94E-04 -  - - 165 28.34

Solid Propellants

Aluminum+HTPB  Ammonium

Perchlorate 5.12E-04 2.12 267 7.90

Aluminum+PBAN  Ammonium

Perchlorate 5.12E-04 2.33 266 7.96

• Mass Ratio based on total Δv of 5.42km/s

Spacecraft System Engineering. 3rd ed. Fortescue, P. pg 183http://www.braeunig.us/space/propel.htmhttp://www.astronautix.com/ROCKET PROPELLANTS, Warren FA, Reinhold Pubising 1958

19

Initial Tank Volume Required

(m3)

% Volume Required

Fuel OxidizerMixture Ratio

(Oxidizer:Fuel)

Specific Impulse (s) MR FUEL OXIDIZER TOTAL

Standard Payload Module

Hydrogen Oxygen 5 390 4.12 6.75 2.10 8.86 67.6

Kerosene Oxygen 2.29 301 6.25 1.25 6.18 7.43 56.57

Hydrazine Oxygen 0.74 313 5.83 1.24 4.08 5.32 40.6

MMHNitrogen Tetroxide 1.73 280 7.18 1.35 3.81 5.17 39.4

UDMHNitrogen Tetroxide 2.1 277 7.33 1.32 4.68 6.00 45.8

Standard Payload Module Volume: ~13m3

20

For equipment operating on For equipment operating on Earth's surface under ideal Earth's surface under ideal operating conditions.operating conditions.Specific cells cited are not Specific cells cited are not space rated.space rated.Only hydrogen reactant Only hydrogen reactant weight included, oxygen weight included, oxygen requirement has been requirement has been neglected.neglected.All cells use manufacturer's All cells use manufacturer's stated output.stated output.

Source Power Output Mass Volume Reactant usage Op. Temp. Reactant Pressure

PEM fuel cell 1.2-2 kW 44-110kg 0.128 – 0.192 m3 2.14 – 4.69 m3/hr 2 – 46 C 125 - 142 Kpa

alkaline fuel cell 6kW 200kg 0.553 m3 4.3 m3/hr -20 - 40 C 400-600 Kpa

solar panels 205W 17.96kg .0734 m3 n/a n/a

Li-ion batteries 300-1500W/kg 576 J/g 972 J/m3 n/a n/a

RTG 330W 55.5 kg none stated 4.5kg 4400W n/a

21

Source reactant/unit #units total mass req. total volume req.

PEM fuel cell 25.798752 5 678.99376 0.96

alkaline fuel cell 90.93984 1 290.93984 0.553

solar panels n/a 29.26829 525.6585366 2.148292683

li-ion batteries 3600 6.666667 24006.66667 2133333.333

1. Alkaline Fuel Cell Used on Apollo and STS missions High performance compared Rarely sees modern design Prone to carbon dioxide saturation

2. Polymer Electrolyte Membrane Fuel Cell Current industry standard Has never been space rated Lower performance than alkaline cells Highly sensitive to fuel impurities

3. Photo Voltaic Cells (Solar Panels) Widely used on previous missions Power generation is time independent Requires deployment system Large volume requirement

22

Site Latitude Longitude Comments Apollo 11 Landing 1N 24E First manned moon mission

Apollo 12 Landing 3S 24W

Apollo 12 LEM impact 4S 21W

Apollo 13 Saturn IVB impact 3S 19W Only mission success during Apollo 13

Apollo 14 Landing 4S 18W Allan Shepherd plays golf

Apollo 14 LEM impact 3S 20W

Apollo 14 S-IVB impact 8S 26W

Apollo 15 Landing 26N 5E Used modular equipment transporter

Apollo 15 LEM impact 26N 1E

Apollo 15 S-IVB impact 2S 11W

Apollo 16 Landing 9S 16E First mission to use the lunar roving vehicle

Apollo 16 S-IVB impact 1N 24W

Apollo 17 Landing 21N 31E Last manned moon mission

Apollo 17 LEM impact 21N 31E Used lunar roving vehicle

Apollo 17 S-IVB impact 5S 13W

23

The goal has been made; land at the location of the first Apollo manned Moon landing, Apollo 11

Luna 20

Luna 16

24

• Qualifies as a historical landing spot which accomplishes one of the extra goals presented by the X-Prize competition

• Location of “Sea of Tranquility”:▪ Proven successful landing location▪ relatively rock and obstruction free

• Equatorial location requires less fuel for orbital corrections and an easier trajectory

• Equatorial landing trajectory between ±10º Latitude provides the most possible landing sites in case we “miss”

25

Motors Manufacture Model Power S/RDrive motors Maxon Motors RE-max 29 15W 48V 0.283A

1.59A (Starting )Upon Request

Steering motors Maxon Motors RE-max 21 6W 48V 0.151A 0.598A(starting)

Upon Request

Cameras

HD 11.16 W (max estimation)

No

SD =16.8 W (max estimation)

No

•Tentative power budget based on selection of components. •Research into cameras and other onboard electronics is on going.•Given this, a total consumption of 150 watts will be estimated.•This is based on the power consumption of previous rovers that have been designed.

26

Comparable Systems:•Spirit & Opportunity both operated at 150W

SpectroLab’s Triple Junction Solar Cell• 316 W/m2

Design Recommendation:•Power consumption <=150W

27

NASA

Ideal Power Generated by Solar Panels• Triple Junction Solar Panels = 0.474 m2 surface area

Lunar Conditions ≠ Ideal• Direct solar Radiation • Panels above normal operating conditions,• Loose efficiency.

Assumptions: Solar Radiation Sc = 1353 W/m2

Emissivity of Solar Panels 0.85 Temp Lunar Environment 107⁰C

Analytical Analysis (Radiation Transfer) Panel Surface Temp Ts = 197°C Voltage drop 1.13v/cm2 Negligible Current Increase 1.24*10-4 Amp

These losses will have to be compensated by including extra Solar Cells

28

Parameters that determine performance:

Radiation SusceptibilityPower ConsumptionOperation TemperaturesMassResolution Frame Rate

29

30

Requirements for HD camera 720p resolution 1280x720 pixels 15 frames a sec minimum

Requirements for SD camera Near real time transmission 15 frames a sec minimum Always on

CAMERA COST ($) POWER OPERATING TEMP MASS

BODY ADAPT BATTCanon EOS Digital Rebel XS

Standard DefinitionCMOS Sensor

800 TYPE: Rechargeable Lithium Ion

VOLTAGE: 7.4V DCCAPACITY: 1080mAhLIFE: 500-600 shots{NO FLASH]400-500[50% FLASH]

0-40˚C(32-104 ˚F)

475g (16.8oz) 80g (2.8oz)

Canon VIXIA HF10High Definition 1080i

810 SUPPLY: 100-240V AC,50/60HzOUTPUT: 8.4V DC,1.5A

0-40˚C(32-104 ˚F)

380g (13.4oz) 135g (4.8oz)

Nikon D40Standard Definition

CCD

480 TYPE: Lithium IonVOLTAGE:7.4V DCCAPACITY:1000mAh

0-40˚C(32-104 ˚F)

475g (17oz) 51g (1.8oz)

Panasonic HDC SD9High Definition 1080i

CCD

650 SUPPLY:110-240V AC, 50/60HzOUTPUT: 9.3V DC, 1.2A

0-40˚C(32-104 ˚F)

275g (9.7oz) 115g (4.0oz)

Sony Alpha DSLR-A200Standard Definition

CCD

500 TYPE: Lithium IonVOLTAGE: 8.4V DC, 2.0ACAPACITY:1650mAh

0-40˚C(32-104 ˚F)

532g (18.8oz) 78g (2.8oz)

Sony Alpha DSLR-A350Standard Definition

CCD

800 TYPE: Lithium IonVOLTAGE: 8.4V DC, 2.0ACAPACITY: 1650mAh

0-40˚C(32-104 ˚F)

582g (20.5oz) 78g (2.8oz)

Sony HDR-HC9High Definition 1080i

CMOS

980 SUPPLY:100-240V AC, 50/60HzOUTPUT: 8.4V DC, 0.35-0.18A

0-40˚C(32-104 ˚F)

550g (1lb 3oz) 170g (6.0oz)

Sony HDR-SR12High Definition 1080i

CMOS

1150 SUPPLY: 100-240V AV, 50/60HzOUTPUT:8.4V DC, 0.35-0.18A

0-40˚C(32-104 ˚F)

570g (1lb 4oz) 170g (6.0oz)

Camera Analysis

31

Selene(Kaguya) took HD images of lunar surface and

earth setting on its horizon 16.5kg,50W, CCD sensor

ISS (Imaging Science Subsystem) Cassini Orbiter- took images of Saturn 57.83kg,55.9W, CCD sensor

ITOS (NOAA) Satellites used for meteorology AVCS (Advanced Vidicon Camera

System)- used before CCD32

1. Total Ionizing Dose (TID)• Changes in threshold voltage

2. Displacement Damage (DD)• Movement in Si lattice• Effect functions (ex. Power)

3. Single Event Effects (SEE) • Single Event Upset• Single Event Latch up• Single Event induced Burnout

33

CCD camera w/o RH CCD camera w/ RH

t=0 (100Gy/hr)

t=1hr (100Gy/hr)

34

Radiation Hardening Techniques Physical-

▪ Microchips that are made from insulating substances (Silicone Oxide & Sapphire)

▪ Shielding the package or the chips themselves (depleted Boron)

▪ Using components with a wide band gap Logical-

▪ Error correction (parity check)▪ Redundant elements

35

Option 1 Design Electronics with hardening

in mind Option 2 Shield commercial electronics

Case study will be performed to select

36

System (with 4 wheels)

total # of Joints

# springs and dampers

Wheel travel

Rocker Bogie 2 differential joints

0 Depends on design

Independent suspension

24 (6 per wheel)

4 springs 4 dampers

Limited by the spring or damper

Source: Heiken et al, Lunar Source Book Cambridge university press 1991

System Drivers: Complexity, Mass, Vehicle Velocity

Current Conclusion•Rocker Bogie System

4 wheel Rocker Bogie system selected

No springs or dampers Has the greatest wheel travelLeast number of jointsProven history

37

Wheel DesignLiterature review of mechanical properties of lunar regolith Recommended ground pressure of 7-10 Kpa for wheeled vehicles

38

Complete trafficability calculations Wheel sinkage Gross pull per Wheel Soil resistance per wheel

Perform kinematic analysis on the Rocker Bogie and fine tune the design to fit our particular needs.

39

40

Launch Vehicle

Orbital Dynamics

Transfer Vehicle

Lander

End Fall 08

Big Picture

System

Sub System

Interface

End Spring 09

2 Iterations

3 Iteration >2

Iterations

0 Iterations

41

Big Picture

System

Sub System

Interface

Rover Power Imaging Communications

2 Iterations

2 Iteration

1 Iterations

End Fall 08

End Spring 09

Roving Vehicle

1 Iterations

Document will contain Requirements System level Analysis

▪ System Drivers (performance determining equations)▪ Evaluation of System Drivers▪ Sub-System level (limited)

Document will contain Requirements System level Analysis Sub-System level Analysis Interface Analysis

42

43