Project X pedition Spacecraft Senior Design – Spring 2009 .

Project Xpedition

Spacecraft Senior Design – Spring 2009https://engineering.purdue.edu/AAE/Academics/Courses/aae450/2009/spring

Motivation:Lunar Payload Delivery

Resupply Lunar BaseSmall Payload

Project Xpedition Requirements

•Land on the Moon

•Move 500 meters

•Transmit HD pictures and video to Earth

•Survive the Lunar Night

•Minimize cost with 90% success

Project Xpedition

Payloads

•100 g•10 kg•1700 kg

Mission Phases

•Earth Launch•Lunar Transfer•Lunar Descent•Locomotion

500m

Earth Launch

Dnepr-1

110 ft

160 ft

Falcon - 9

180 ft

Earth Launch• Site: Baikonur Cosmodrome, Kazakhstan• Cost: $5M

250 Mile Parking Orbit

Lunar LanderOrbital Transfer

Vehicle

880 lbs

8’

Solar Arrays unfold

InternalView

Hall Thruster produces 80

mN of thrust

Power

CommunicationS-Band Antenna

2 Solar ArraysLithium-Ion Battery

Attitude

Chemical Thrusters

Sun SensorStar Sensor

Reaction Wheels

Lunar Transfer

•16 mile parking orbit•2 hour orbital period

•Lander is self sufficient•350 lb Lander mass •Half of mass is propellant

Space Balls Housing

Communication Antenna and Motor

Solar Panel

Attitude Control Thrusters

RadiatorAttitude Sensors

CPU

H2O2 Tank

Helium Tank

Radial Flow Hybrid Engine

Camera

Surveyor 3

Apollo 12

25 miles

Landing Site: Mare Cognitum

Final Descent Attitude: 12 Control Thrusters Translation: Radial Flow Hybrid Engine

Mission Requirements Land on Moon Move Payload 500 m Survive Lunar Night

Lexan Shell

Camera

CPU

Dust Removal Vibration Motor

Battery

100g Payload

Main Axel and Motor Housing

Communications Transceiver

Mission Requirements

1. Land on Moon

2. Move 500m

3. Take Picture

4. Survive NightTaking Photo of Lander

Removing Dust

All Systems Are GO!

Avoiding Obstacle

Cruise Speed: 3.2 mph

Minimum Turning Radius: 2.5 in

-280 °F

10 kg Lunar Lander

230 lbs Lander270 lbs Propellant500 lbs Total


1. Land on Moon

2. Move 500m

3. Take Picture

4. Survive Night

Hybrid Engine

Thrust: 45 lbsBurn Time: 135 sec10 kg Payload

500m

Record Video


1. Land on Moon

2. Move 500m

3. Take Picture

4. Survive Night

Completed lunar descent• Full stop• Begin locomotion

Attitude Thrusters

16 ft

6 ft

300 ft

Main EngineAvg. Thrust: 230 lbsBurn time: 60 s

Large Payload

Mission Requirements:1. Move 500 meters2. Land on moon3. Resupply base

$27MCost

- $22M Prize

= $5MNet Mission Cost

100 g 10 kg 1743 kg0

50

100

150

200

250

Mission Cost, $M in 2009 Dollars

Mission Cost

$27 Million72% Success



Cost Per Kilogram$271Million

$3 Million

$130k

Payload Delivery:

1. Most economical payload: 2 tons2. Electric Propulsion for Lunar transfer3. Soft land on Lunar surface

Google Lunar X PRIZE:

1. Several viable locomotion methods2. Potential to open commercial market3. $27M mission accomplished for $5M

Project XpeditionResults

Question & Answer

https://engineering.purdue.edu/AAE/Academics/Courses/aae450/2009/spring

Project Xpedition

Backup Slide ListingPropulsionBrad AppelThaddaeus HalsmerRyan LehtoSaad Tanvir

AttitudeBrian ErsonKris EzraChristine TroyBrittany Waletzko

PowerTony CoferAdham FahkryJeff KnowltonIan Meginnis

Structures & ThermalKelly LeffelCaitlyn McKayRyan Nelson

CommunicationsMike ChristopherJohn DixonTrent Muller

Mission OperationsJohn AitchisonCory AlbanLevi BrownAndrew DamonAlex Whiteman

Solomon Westerman

Backup SlidesSaad Tanvir

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Propulsion System Mass Finals100 g Payload case (Ball)Propellant mass = 78.2 kgPropulsion System Inert mass = 29.9 kgTotal Prop System Mass = 108.1 kg

Arbitrary Payload case (Falcon 9)Propellant mass = 1783.62 kgPropulsion System Inert mass = 227 kgTotal Prop System Mass = 2010.62 kg

10 kg Payload case (Hopper)Propellant mass = 121.2 kgPropulsion System Inert mass = 45.4 kgTotal Prop System Mass = 166.6 kg

2Saad Tanvir

Propulsion GroupReturn to Listing

100 g – Hybrid Propulsion System Mass Breakdown

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4

10 kg – Hybrid Propulsion System Mass Breakdown

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Large payload – Hybrid Propulsion System Mass Breakdown

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Propellant Tank Specifications

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Pressurant Tank Specifications

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8

Hydrogen Peroxide Tanks - Thermodynamic Analysis

Assumptions:

Tank operating Temperature = 283 K (50 F) Surrounding Temperature = 2.73 K

Power Required ~ 35 W

ΔT = 280.3 K

Q: Rate of Heat transfer [W]A: Area of Cross section of the tank [m2]k: Thermal Conductivity [0.044 W/mK]ΔT: Temperature Difference [K]t: Thickness of the blanket [200 mm]

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Lunar Descent – Thermodynamic Analysis on Prop System

9

Temperature Drop < 5 K

No power required to heat the propulsion system during Lunar Descent

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Propellant Tank – Operating Pressure

Pchamber = 2.07 MPa

∆Pdynamic = ½𝜌v2 ~ 0.072 MPa

∆Pfeed (Upper bound) ~ 0.05 MPa

∆Pcool ~ 0.15pc = 0.31 MPa

∆Pinjector ~ 0.3pc = 0.62 Mpa

Ptank ~ 3.07 MPa

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Lunar Transfer: Chemical Alternative

Significant mass savings using the Electric Propulsion system

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Backup SlidesChristine Troy

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Lander Attitude Control

12 General Kinetics H2O2 thrusters

Lander Side view

Lander Top view

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Attitude Prop Mass Estimate

• Based on Rauschenbakh, Ovchinnikov, and McKenna-Lawlor

θ.

θ

+θ1

-θ1

No External Torque

θ

+θ1

-θ1

“Large” External Torque

θ.

gIspL

Mm b

Mb = external moment applied

g = gravitational acceleration

Isp = specific impulse of thrusters

L = distance from thruster to vehicle center of mass

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Spinning Lander Attitude Control Propellant and thrusters still needed for

spin up and axis reorientation– Estimate ~2.2 kg propellant savings for

100g/10kg cases Additional mass: spinning landing gear,

propulsion system redesigns, additional attitude sensing devices

Increased complexity: Liquid propellant feed while spinning, landing while spinning, reorientation of axis

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Compressed Gas Spring Energy Storage

• Some or all travel could be obtained from bouncing using stored descent energy

• Compressed gas not recommended – highly temperature sensitive, limited velocity and acceleration inputs

– Commercial gas springs limited to approx. -23° to 82°• Lunar surface temperature -153° to 107° C

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Backup SlidesBrittany Waletzko

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System Masses

Mass 100g 10 kg Large

Injected Mass to Low Earth Orbit (kg) 436 584 9953

Injected Mass to Low Lunar Orbit (kg) 156 228 4545

Mass on Lunar Surface (kg) 79 107 2325

Payload Delivered to Lunar Surface 100g 10kg 1743kg

Systems Overview

100g Payload

10kg Payload

Large Payload

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Mission Timelines (Backup)Elapsed Time (ddd:hh:mm) Event Vehicle

-365:00:00 Launch Launch Vehicle/OTV000:00:00 Arrive in LLO OTV000:00:03 In lower orbit Lander000:00:04 Rotate and Land Lander000:00:04 Systems check Space Ball000:00:05 Deployment from Lander Space Ball000:00:06 Orientation Space Ball000:00:06 Travel 500m Space Ball

000:00:14 Braking maneuver, dust removal Space Ball

000:00:15 Take picture of Lander, Begin transmission to Lander Space Ball

000:00:23 End photo transmission Space Ball

000:00:23

Transmit arrival Mooncast (near real-time video, photos, HD video, XPF set asides, data uplink set) to Earth

Lander

001:33:56Transmit Mission Complete Mooncast (near real time video, photos, HD video)

Lander

002:08:04 Finished transmitting, prepare for night Lander

009:00:00 Standby for lunar night Lander025:00:00 Power up after night Lander026:00:00 Transmit telemetry and photo Lander026:00:14 Mission Complete

Elapsed Time(ddd:hh:mm)

Event

-365:00:00 Launch

0:00:00 Lunar Lander reaches LLO and separates from OTV

0:00:04 Lands on lunar surface and starts video taping

0:00:12 Finishes taping and begins transmission of video

0:03:44 Completes video transmission and takes panoramic pictures

0:03:45 Finishes panoramic pictures and begins transmission of pictures

0:03:59 Completes picture transmission and begins hop for locomotion

0:04:01 Locomotion phase complete and begins HD video taping

0:12:01 Begins transmission of HD video and takes panoramic pictures

2:06:24 Ends transmission of HD video and begins transmission of pictures

2:06:36 Ends transmission of pictures and shuts down for lunar night

15:23:24 Turns on and sends signal after lunar night.

Elapsed Time given in days, hours, and minutes

100g Payload Mission Timeline 10kg Payload Mission Timeline

100g and 10kg Payload Return to Listing

Mission Timelines—cont. (Backup)

Elapsed Time (ddd:hh:mm)

Event

-365:00:00 Launch000:00:00 Arrive in Low Lunar Orbit

Transfer to Lunar Descent Transfer OrbitBegin Final Lunar Descent burnCome to rest 100 m above surface/begin hover locomotionTouch down on lunar surface

Large Payload Mission Timeline

Elapsed Time given in days, hours, and minutes

Large Payload

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Trajectory Correction (backup)100g Payload Correction Maneuver Configuration

Parameter ValueIsp (s) 1952

mo (kg) 436.0Propellant for Correction (kg) 1.1Thrust per Engine (mN) 75Time for ΔV (hr) 80.7

m/s 50 =dt m

TV

10kg Payload Correction Maneuver Configuration


mo (kg) 585.6Propellant for Correction (kg) 1.5Thrust per Engine (mN) 75Time for ΔV (hr) 92.6

Large Payload Correction Maneuver Configuration


mo (kg) 9953Propellant for Correction (kg) 22.5Thrust per Engine (mN) (x4 engines)

424

Time for ΔV (hr) 54.2

T

Vmt

T = instantaneous thrust (assumed constant over interval)m = instantaneous mass (assumed constant over interval)

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Thruster Locations and Thrust Direction Vectors

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Hydrazine and Hydrogen Peroxide Thrusters

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Environmental Forces Codes

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Output (in Newtons, Kilograms)• “Environmental” :• Felec =• 1.5231e-005• Fref =• 2.8793e-022• Ftherm =• 1.9623e-022• Fscrad =• 4.0027e-006• Fswind =• 5.1750e-009• Fmag =• 2.1599e-013• Fexp =• 7.9937e-007• Ftotal =• 2.0039e-005

“Environmentalpropmass” : mm_cyl_month = 30.8571 mm_cyl_5000 = 0.0595 mm_cyl_50000 = 0.5952 mm_cube_month = 0.3086 mm_cube_5000 = 5.9524e-004 mm_cube_50000 = 0.0060

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Backup SlidesIan Meginnis

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OTV Power Subsystems

PPU(Electric Propulsion)

PCDU

Battery(LL)

100V

100V

DC-DC Converters

Individual OTV Components

Acronym Definitions:PCDU - Power Conditioning and Distribution UnitPPU - Power Processing UnitLL - Lunar LanderDC - Direct Current

Solar Array

<200V

Solar Array

Note: Not to scale

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Group Power (Watts)Propulsion 1529

Communication See “Lunar Lander”

Attitude 101.4Power 120Lunar Lander(during Lunar Transfer) 105

TOTAL 1959

100g Payload Case Power Budget



Attitude 145.4Power 120

Lunar Lander(during Lunar Transfer) 105

TOTAL 2534

10kg Payload Case Power Budget



Attitude 305.4Power 1731.5Lunar Lander(during Lunar Transfer) 105

TOTAL 42960

Large Payload Case Power Budget

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Power Distribution: 100g Payload OTV

Propulsion: 2029W

Attitude: 145W

Power: 120W

Lunar Lander: 105W

Propulsion:

1529W

Attitude : 101W

Power: 120W

Lunar Lander: 105W

Power Distribution: 10kg Payload OTV

Propulsion: 38773W

Attitude: 305WPower: 1731W

Lunar Lander: 105W

Power Distribution: Large Payload OTV

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Payload Size Component Variable Value

100gSolar Arrays(2 circular arrays)

Mass 13.06kgDeployed Area 6.54m2

Cost $1.96 Million

Battery Mass 12.17kgCost $22,000

10kgSolar Arrays(2 circular arrays)


Cost $2.53 Million


LargeSolar Arrays(2 rectangular arrays)


Cost $42.96 Million


OTV Power Dimensions

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Note: Not to scale

Acronym Definitions:PCDU - Power Conditioning and Distribution UnitPPU - Power Processing UnitBatt - BatteryDC - Direct Current

PPU PCDU BattDC/DC

Converter

Aluminum Heat Pipes with Ammonia

Aluminum Mount

2 Radiators

Electronics Board Thermal Control

Hall Thruster Thermal Control

Hall Thruster

Radiating Heat Shroud

Radiating Heat Shroud

(Exhaust)

OTV Thermal Control (all payloads)

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OTV Electronics Thermal Control Payload Size Component Mass (kg)

100gRadiators 1.2Ammonia NegligibleHeat Pipes 2.2TOTALS 3.4

10kgRadiators 1.6Ammonia NegligibleHeat Pipes 2.5TOTALS 4.1

LargeRadiators 23.4Ammonia NegligibleHeat Pipes 15.3TOTALS 38.7

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Note: Not to scale

• At least 1 of the OTV’s set of radiators will not be exposed to sun’s rays at any point during the trajectory

• Each radiator, alone, can provide thermal control for OTV electronics

Earth

Sun

Moon

Single, Simplified Orbit of OTV (Large Payload)

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Backup SlidesRyan Nelson

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Basic Frame Design• Drivers in frame mass

– Total Lunar Lander (LL) mass at lunar touchdown

– Volume of LL

• Shape: Conic Frustum– Stores all Lunar Lander subsystems while

minimizing volume

• All frame components hollow– Small leg diameter allows for storage

within side supports prior to lunar touchdown

Schematic of frame listing componentsReturn to Listing

Final Mass and Volumes

Lunar Lander Volume Height Top Diameter Bottom Diameter Mass

100g 1.05 m3 1.0 m 1.0 m 1.3 m 11.44 kg

10kg 1.15 m3 1.1 m 1.0 m 1.3 m 19.97 kg

Large 14.33 m3 2.0 m 2.4 m 3.6 m 104.86 kg

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Frame Design• Thickness of all frame components varies

– First mode of failure (factor of safety = 1.5)– Payload case

• Cross sectional shape is circular or rectangular for all components

• 0.5 mm magnesium skin place around Lunar Lander frame– Micrometeorite protection– Thermal protection

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Floor Supports

• Support a majority of landing loads• Thickness altered until load is supported• Hollow Rectangular Cross section

– Moment of Inertia

– Bending Stress acting on beam

I

My

I

My

12

)2)(2( 33 thtbbhI

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Side Supports/Legs

• First mode of failure is buckling

– K=0.5 for side supports (both ends fixed)– K=2.0 for legs (one end is free to move)– Hollowing the rod decreases moment of Inertia

and critical load

2(kl)

EI

2(kl)

EIFcr =

)( 21

22 rrI

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Side Supports/Legs

• Compression failure occurs after buckling for both side supports and legs– Despite small cross sectional area– Compression failure

• Top, Bottom, and Engine Support rings all designed to have same cross sectional dimensions

A

F

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Cross Sectional Dimensions

100g Payload CaseStructural Component Cross sectional shape Outer Dimensions Thickness

Outer Ring Circular 10 cm Diameter 4mm

Engine Support Ring Circular 10 cm Diameter 4mm

Rectangular Floor Supports Rectangular 10 cm Height, 6 cm Width 6mm

Side Supports Circular 10 cm Diameter 2mm

Top Ring Circular 10 cm Diameter 4mm

Legs Circular 6 cm Diameter 3mm

10kg Payload CaseStructural Component Cross sectional shape Outer Dimensions Thickness

Outer Ring Circular 10 cm Diameter 5mm

Engine support Ring Circular 10 cm Diameter 5mm

Rectangular Floor supports Rectangular 10 cm Height, 6 cm Width 7mm

Side supports Circular 10 cm Diameter 2mm

Top Ring Circular 10 cm Diameter 5mm

Legs Circular 5 cm Diameter 3mm

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Placement of Hop Engines

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Backup SlidesBrian Erson

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Backup Slide 1

Calculation of Thrust Misalignment Torque• Estimate of Thrust at 1 kW ~100 mN• Estimate of Thrust misalignment ~ 0.05 m

• Conservative Max Misalignment Torque ~ 5 mNm

Calculation of Drift Error• Tracking error from Reaction wheel spec sheet <1rpm• Operating speed 3000 rpm• Max Wheel Torque 12 mNm

• Drift Error = 1/3000 * 12 = 0.004 mNm

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Backup Slide 2

Calculation of Mass Requirement (Ref, Smart-1 Lunar Probe)• Reaction Wheel assembly ~ 12 kg• Sun Sensors ~ 4 kg• Angular rate sensors ~ 0.3 kg• Star Tracker ~ 3 kg• Total ~ 19.3 kg

• ACS/Launch mass: 19.3/380 = .05 = 5%• Conservative estimate of IMTLI: 700 kg * 5%• Conservative estimate of ACS mass < 35 kg

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Backup Slide 3

Calculation of Pointing Accuracy• Max pointing error of SMART-1: 60 arcminutes• 1 arcminute = 1/60*deg • 1 deg = .017 rad

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Backup Slide 4

Attitude Control Mass Calculations:3-axis

Sun Sensors 0.7Star Sensors 6.4Reaction Wheels 6H2O2 Thrusters 1Propellant(H2O2)* 26Total: 40.1 kg

SpinConical Scanner 6Doppler Device 1H2O2 Thrusters 1Propellant(H202)* * 19.1Total: 27.1

*Prop Mass Includes Lunar Descent

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Backup Slide 5

**H2O2 De-Saturation(DS) Mass Calculation:• DS of reaction wheels:

• Estimate of DS maneuvers/day: 6• Reaction wheel max torque: 0.03 N-m• H2O2 Thrust: 9.5 N/kg• Max Mission Length: 365 days

Total Mission DS H2O2 mass:(365)(6)(1/9.5)(0.03) = 6.9 kg

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Backup Slide 6

Other Mass Calculations:• Power:

• Masses based on posted Power Group Data• Mass of battery without solar cells based on assumption of >1.2 kW needed to

power OTV• Communication:

• 3 kg mass based on posted Com Group Data• Thermal:

• 3 axis mass based on posted data• Spin stabilized Thermal Protection:

Assume 15 rpmMass = [(1/15(17.1))+4*] =~ 5 kg

*Estimated mass of standard thermal protection

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Backup Slide 7

Cost tradeoff:Xe thruster system cost: $100,000Current Earth to LEO cost/kg: $4400Economical mass savings 22.7 kgXe system mass savings < 5.0 kgXe system Earth to LEO cost: ~$20,000/kg

Note:Unless Xe system saves upward of 22.7 kg,or the cost decreases, it is not economically feasible to install the system. Further analysis will be done to improve mass and dollar cost numbers of both systems.

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Backup Slide 8

Xe DS propellant calculation

Total DS force needed:Max torque of Rxn Wheel: 0.03 NmDS per day: 6Mission length: 365 daysMoment arm 1.0 mTotal: 65.7 NTotal number of thrusts @ .015 N 4380 thrustsMarotta Cold Xe gas thruster Specs:Mass: 0.075 kgIsp: 68 secThrust: 0.015 NTime per thrust: 0.04 secMass Flow Calculation:Isp = Force/massflow * gravityMass flow: 0.0022 kg/secTotal Mass:

0.0022 kg/sec * 0.04 sec/thrust * 4380 thrusts = 0.385 kg

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Thruster Analysis

100g 10kg Arbitrary

Added Inert mass (kg) 1.924 3.02 13.84

Added Volume (m^3) 4.3x10-6 7.88x10-5 5.97x10-4

Cost savings($) 7000 6000 38000

•Consultation with Purdue Hybrid(H202) Rocket Team led to development of an alternate OTV attitude control system

•System consists of 4 small H202 tanks enclosed within OTV

•Each system is independent

•All payload cases can be developed in-house for a fraction of purchase cost

Backup Slide 9

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Reaction Wheel Update

Payload Device Manufacturer Mass (kg) Size (cm) Power Required peak (W) Max Torque (mNm)

100g VF MR 4.0 (4) Valley Forge Composites 2.6 (each) 20 x10 (each) 76 (total) 20 (each)

10kg VF MR 10.0 (4) Valley Forge Composites 5.0 (each) 25 x15 (each) 120 (total) 30 (each)

Arbitrary VF MR 19.6 (4) Valley Forge Composites 10.5 (each) 39 x17 (each) 280 (total) 260 (each)

•Each Reaction Wheel had to be upgraded within each payload to account for increases in system mass

•Relevant changes to note:

100g 10kg Arbitrary

Mass Increase (kg) 4.4 9.6 22

Power Increase (W) 20 44 160

Backup Slide 9

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Backup Slide 10

Cost Savings Calculation:100gGeneral Kinetics Cost for 4 – 1N thrusters: $12,000In-house Manufacturing cost: $5,000Cost Savings: $7,000

10kgGeneral Kinetics Cost for 4 – 1N thrusters: $12,000In-house Manufacturing cost: $6,000Cost Savings: $6,000

ArbitraryGeneral Kinetics Cost for 4 – 13N thrusters: $48,000In-house Manufacturing cost: $10,000Cost Savings: $38,000

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Backup Slide 11

Inert Mass CalculationsDensity of H202: 1.11 kg/LMass of aluminum tank per .001 m^3: 3.68 kgKg(prop) = massflow*(sec/thrust)*thrustsKg(tank) = (3.68/.001)*volumeH202100g4 – 0.0048kg H202 Tanks 0.064 kg4 – 0.02N H202 Thrusters 0.36 kgFeed Lines, Valves 1.5 kgTotal Inert Mass 1.924 kg10kg4 – 0.315kg H202 Tanks 1.16 kg4 – 0.03N H202 Thrusters 0.36 kgFeed Lines, Valves 1.5 kgTotal Inert Mass 3.02 kgArbitrary4 – 2.65kg H202 Tanks 8.8 kg4 – 0.26N H202 Thrusters 0.36 kgFeed Lines, Valves 1.5 kgTotal Inert Mass 13.84 kg

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Backup SlidesCaitlyn McKay

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DeploymentLinear Shaped Charge

System Mass (kg) / Space Ball

Charge 0.580

Foam 0.040

Total 0.620

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Accordion Landing

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Impulse Momentum

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Kamikaze RoverSolar Panels 0 kgBatteries 0.422 kgPower (extra) 1.92 kgCommunications 1.51 kgDrive System 0.298 kgStructure (frame) 0.40 kgSpace Blankets 0.58 kgWheels 0.58 kgCooling System 0 kg

System Mass 5.7074kgBallast Mass 10kg Total 15.7074kg

Length 0.23mWidth 0.21m

Height 0.21m

* Life of 13 minutes.

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Rover Deployment

Linear Shaped Charge

System to lower Rover from Lander to surface.

Item Linear Shaped Charge

SOLIMIDE Foam

Steel Cable

Platform Motor Support Beams

Total

Mass (kg) 0.580 0.030 0.82 0.025 0.025 0.13 1.610

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Backup SlidesTrent Muller

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1. Mt. Pleasant Radio Observatory. Hobart, Tasmania, Australia. A 26 meter dish.

2. Hartebeesthoek Radio Astronomy Observatory (HRAO). Johannesburg, South Africa. A 26 meter dish.

3. Pisgah Astronomical Research Institute (PARI). Rosman, North Carolina. USA. One of the 26 meter dishes.

4. James Clark Maxwell Telescope. Mauna Kea Observatory, Hawaii, USA. A 15 meter dish. Return to Listing

Ground Stations Altitude (km) of Non-Tracking Zone

1-2 6241.59

2-3 4831.64

3-4 1527.09

4-1 896.85

Ground Station

Latitude (o) Longitude (o)

1 42.81 S 147.44 E

2 25.55 S 27.68 E

3 35.20 N 82.87 W

4 19.82 N 155.48 W

Communications Coverage

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Equipment Model Manufacturer Mass (kg) Power Usage (W) Price (2009 $)

Lander-Earth Antenna (2)

Patch Antenna SSTL 0.16 -- 40,000

Lander-Earth Receiver

RX-200S SpaceQuest 0.2 1.5 30,000

Lander-Earth Transmitter

TX-2400 SpaceQuest 0.2 34 24,000

Lander-Rover Antenna

ANT-100 SpaceQuest 0.1 -- 500

Lander-Rover Transceiver

TR-400 SpaceQuest 0.21 6 20,000

Computer Board RAD6000 BAE 0.85 13 200,000

Video Camera HF10 Canon 0.38 3.9 1,000

Antenna Pivot (2) -- -- 0.38 2.13 168

Totals 2.48 60.53 315,668

Communications Equipment Onboard Lander for 100 g Payload

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Communications Equipment Onboard Lander for 10kg and Large Payload

Equipment Model Manufacturer Mass (kg) Power Usage (W) Price (2009 $)

Lander-Earth Antenna (2)

Patch Antenna SSTL 0.16 -- 40,000

Lander-Earth Receiver

RX-200S SpaceQuest 0.2 1.5 30,000

Lander-Earth Transmitter

TX-2400 SpaceQuest 0.2 34 24,000

Computer Board RAD6000 BAE 0.85 13 200,000

Video Camera HF10 Canon 0.38 3.9 1,000

Antenna Pivot (2) -- -- 0.38 2.13 168

Totals 2.17 54.53 295,167

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Backup SlidesTony Cofer

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Hydrazine Heater for 100g and 10kg Payloads

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Nocturnal Power Controller

• Save 11.5 kg of batteries• Controllable• Size 2”X2”X1/2”• Weight~20 g• Power diss. 0.1mW• Requires 0.23 g battery for 14 days

ControllerInterface

Solar Source

ComparatorWith

Hysteresis

Actuator

Command Computer

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Backup SlidesMike Christopher

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Mooncast ScheduleLunar Arrival Mooncast

Item Link Direction Size [MB] Transmission Time [hr]Photos Down 5 0.24XPF Set Asides Down 10 0.47Data Uplink Set Up 10 0.47Data Uplink Set Down 10 0.47

Totals 35 MB 1.65 hrs

Locomotion Mooncast Item Link Direction Size [MB] Transmission Time [hr]

8 min Near Real Time Video Down 75 3.538 min High Definition Video Down 900 42.37Photos Down 5 0.24


Survival Mooncast (BONUS PRIZE) Item Link Direction Size [MB] Transmission Time [hr]

8 min Near Real Time Video Down 75 3.53Photos Down 5 0.24


Michael Christopher – Backup Slide Return to Listing

Patch Antenna and Pivot System

AdvantagesRedundancy (2 pivots and antennae and 2 motors on each pivot.Reduces the need for more antennae on the Orbital Transfer Vehicle (OTV)Low cost pivot: $83.50 Low mass pivot: 0.2 kg

System Mounted on OTV


Antenna MountBase Plate

Stepper Gear Motor



• Mass:2*(0.0454 kg/motor) + 0.1kg = 0.1908 kg

• Power Consumption:2.128 Watts

• Cost:2*($16.75 /motor) + ~$50 Al = $83.50



Michael Christopher – Backup Slide


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Backup Slides John Aitchison

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Lunar Descent OverviewNote: Not to Scale

Lunar Parking Orbit

Lunar Descent Transfer Orbit

Final Descent

Moon

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Final Descent Overview

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Final Descent Trajectory

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100 g Payload Descent Overview

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10 kg Payload Descent Overview

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1743 kg Payload Descent Overview

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Descent Validity Check• ∆V ~ 2,000 m/s to move from LPO to zero velocity on lunar surface

Isp = 320 s

g0 = 9.8 m/s2

Mi = Total Lander Mass in LPO = 157 kg

.

• Mf = 83 kg

• Propellant Used = Mi – Mf = 74 kg

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Equations of Motion

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Altitude & Range

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Surface Clearance: Worst Case Scenario

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Lander Mass vs. Time

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Lunar Descent Transfer Orbit

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Sample Descent Code Output

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Backup SlidesJohn Dixon

Thermal Considerations• Assumptions:

– Solar Panels Reflect Unused Solar Energy Completely– Thermal Blanket keeps Energy Transfer through body to 0

J/s• Above includes MLI comprised of Kapton (or Teflon) /

Silver Lined Reflective Surface, Kapton Insulation (with scrim separation)

– Thermal Heat Sinks radiate to Coldest Possible Surface– Steady State Conduction

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Insulation/Heat Sink

• Copper Heat Emission– q/t = 730.432 J/s (from emissivity of Copper)– Cu mass = 6.08 kg

• One heat vane traveling to each side of the rover• @max CPU Operating Temp

• Multi-Layer Insulation (MLI)– Insulation mass= 0.898 kg

• Total Thermal Control Mass: 6.98 kg

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Copper Sink Properties

• Copper Slab– 0.03m thick X 0.065m wide X 0.08m long– Volume: 0.000156 m^3

• Copper Vein– 0.02m height X 0.065m wide X [0.001:0.372]m

thick– Volume(max distance) = 0.000677 m^3

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System Description

• N2 gas @ 1 atm inside Toy Ball Enclosure– Mass of N2 gas = 0.01023 kg

– Temp of N2 gas = 0 oC (273.15K)

• Total Heat Dissipation– Z-93 White Paint Coating (α = 0.17)

• Qsun = 5384.435 J

• Qelectronics = 1060 J

• Qtotal = 6444.435 J• Total Energy Rate Into System = 13.426 W

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Thermal Transport Over Time

10 20 30 40 50 60273

273.05

273.1

273.15

273.2

273.25

273.3

273.35

Temperature of N2/Al vs Time

Time, t

Tem

pera

ture

, K

Tgas

TAl

10 20 30 40 50 60 70

-10

-8

-6

-4

-2

0

2

4

6

Heat Transfer of Al vs Time

Time, t

Hea

t T

rans

fer,

J/s

qdotAlin

qdotAlout

•Steady State Equilibrium Occurs at ~50 seconds•Total Temperature Rise Over 8 min = 0.7K

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Backup Slide 1

10 20 30 40 50 60 70

-20

-15

-10

-5

0

5

10

Total Heat Transfer vs Time

Time, t

Hea

t T

rans

fer,

J/s

qdottot

N2

qdottot

Al

0 50 100 150 200 250 300 350 400 450 500-25

-20

-15

-10

-5

0

5

10

15Total Heat Transfer vs Time

Time, tH

eat

Tra

nsfe

r, J

/s

qdottot

N2

qdottot

Al

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Backup Slide 2

0 50 100 150 200 250 300 350 400 450 500273

273.1

273.2

273.3

273.4

273.5

273.6

273.7

273.8

273.9

274

Temperature of N2/Al vs Time

Time, t

Tem

pera

ture

, K

Tgas

TAl

0 50 100 150 200 250 300 350 400 450 500-12

-10

-8

-6

-4

-2

0

2

4

6

8Heat Transfer of Al vs Time

Time, tH

eat

Tra

nsfe

r, J

/s

qdotAlin

qdotAlout

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Lander to Earth Transmission

20 25 30 35 40 45 50 55 60 65 700

1

2

3

4

5

6

7

X: 33Y: 2.989

Power, Watts

Mar

gin

of S

ucce

ss, d

B

Margin of Success vs. Power, Beamwidth = 50o

Distance from 200 km Parking Orbit to 440,000 km of Moon at Apogee

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5

x 105

0

10

20

30

40

50

60

70

80

X: 4.057e+005Y: 4.671

Distance, km

Ma

rgin

of S

ucc

ess

, dB

Margin of Success vs. Distance, Beamwidth = 50o, Power = 50W

Transmit Satellite: 0.191 mReceiver Satellite: DSN 26 m

Minimum Power: 33 WFrequency: 2.2 GHz

(S-Band Range)Data Rate: 51.2 kbps

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Rover to Lander Transmission

0 50 100 150 200 250 300 350 400 450 50080

90

100

110

120

130

140

Distance, m

Ma

rgin

of S

ucc

ess

, dB

Margin of Success vs. Distance, Beamwidth = 25o, Power = 5W

0 5 10 15 20 25 30 35 40 45 5076

78

80

82

84

86

88

90

92

94

Power, Watts

Ma

rgin

of S

ucc

ess

, dB

Margin of Success vs. Power, Beamwidth = 25o

Distance from 0 m Lander to 500 m Maximum Travel

Transmit Satellite: 0.381 mReceiver Satellite: Lander 0.2 m

Minimum Power: Open ConditionFrequency: 2.2 GHz

(S-Band)Data Rate: 51.2 kbps

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Beamwidth Optimization (Backup)

20 21 22 23 24 25 26 27 28 29 3097.1

97.2

97.3

97.4

97.5

97.6

97.7

X: 25Y: 97.69

Beamwidth, meters

Ma

rgin

of S

ucc

ess

, dB

ROVER: Margin of Success vs. Beamwidth, Power = 5W

30 35 40 45 50 55 60 65 701.5

2

2.5

3

3.5

4

4.5

5

X: 50Y: 4.794

Beamwidth, metersM

arg

in o

f Su

cce

ss, d

B

LANDER: Margin of Success vs. Beamwidth, Power = 50W

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Backup SlidesJeff Knowlton

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Overview•1 Minute Deploy•2 Status Relays •8 Minutes Travel•1 minute Prep/ Photograph•8 Minutes Transmitting

Drive Motors6%

Trans-mis-sion55%

Camera4%

Re-serve34%

Battery Distribution

Space Ball Power

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Ball Power SystemBattery (using three)Lithium Manganese Dioxide Coin

(CR2330 )3 volts.26ampere-hrCylinder Dimensions

23mm diameter 3mm height

0.004 kg each5% loss per month(self discharge 1 year)

Total• 2.34watt-hr at Liftoff• 45.96% loss over 1 year• 1.26 Watt-hrs after 1 year• 0.112kg including housing

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Backup SlidesThaddaeus Halsmer

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(2)

(3)

(4)

(1)

Table 1 Engine performance parametersEngine No. Payload case/Description F_max/min [N] tb [s]

1 10 kg/hop engine 2x 192 (avg.) 134.52 100 g/main engine 1100/110 198.63 10 kg/main engine 1650/165 190.44 Arbitrary/main engine 27000/2700 250.2

Stick is 6.5 feet high, same as a standard doorway

Lunar Lander Propulsion – Engine Specifications

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SV01

SV02

High PressureHelium Tank

HV01

REG

CK01

CK02

MOV

F01

H2O2

Tank

HV02

RV01

Lunar Lander Propulsion –fluid system diagrams

SV01

SV02

High PressureHelium TankHV01

REG

CK01

CK02

MOV

F01

H2O2

Tank

HV02

RV01

SV04SV03 SV05

100g and Large payload cases 10kg payload case

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150 200 250 300 350 400 450 500 550 60020

40

60

80

100

120

140

160

180

200

220

240

Hybrid

Mono-Prop

Bi-Prop

Isp [s]

Pro

pella

nt m

ass

[kg]

Figure X: Propellant mass vs. Isp trade

Lunar Lander Propulsion - Propellant/Propulsion system selection

Selection Criteria:

1. Thrusta. min/maxb. throttling

2. Dimensionsa. Short and fat

3. Mass – minimize

4. Propellant storability

5. Purchase/development costs

6. High Reliability

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• As area ratio, ε, increases Mnozzle increases, but Isp increases also• As Isp increases Mprop decreases for a given thrust and burn time

Wrote Matlab script that used Matlab CEA interface to compute multiple Isp’s for different area ratio’s and the corresponding Mprop and Mnozzle for a given thrust, and burn time

Results: Area ratio for minimum mass occurred at ~150, however this nozzle would be very large and little is gained above ~100

41

32

105400125

propnozzle

MM

Lunar Lander Propulsion - Nozzle area ratio and mass optimization

50 100 150 200126

127

128

129

130

nozzle area ratio

Noz

zle

+ P

rop

mas

s [k

g]

Pc = 0.862 MPa, 2000N thrust

Pc = 1.72 MPa, 2000N thrustPC = 0.345 MPai, 2000N thrust

Used CEA to compute Isp for given nozzle area ratio

• All other inputs constant

Empirical nozzle mass equation

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Fuel grain dimension definitions

Burn time [s] Isp, ave [s]210 329400 322 0 5 10 15 20

260

280

300

320

340

O/F ratio

Isp

[s]

Lunar Lander Propulsion – Isp analysis approach

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Lunar Lander Propulsion – fuel grain and chamber sizing approach

1. Choosea. Empirical value for initial fuel regression rateb. Initial O/F ratio for optimum Isp

c. Initial propellant mass flow rate

Compute required burn surface area

2. Dimensions of fuel grainsa. Diameter is derived from burn surface area found from values in step #1 and chosen

fuel grain geometryb. Thickness is function of burn time and regression rate

3. Compute Chamber dimensionsa. Chamber dimensions approximated from fuel grain size and additional room for

insulating materials

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Backup SlidesAlex Whiteman

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-1000 -900 -800 -700 -600 -500 -400 -300 -200 -100 0 100-100

-50

0

50

100

150

200

range (m)

altitu

de (m

)

MoonHop Trajectory

LiftoffTouchdown

10kg Hop Trajectory

Trajectory Timeline• First, throttle up and then

throttle down engine while pitching Lander in clockwise direction.

• Next, Lander remains at constant pitch angle and altitude while thrusting in direction opposite of hop

• Finally, Lander pitches in counter-clockwise direction in order to land in a vertical orientation.

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Large Payload Hover Trajectory

-700 -600 -500 -400 -300 -200 -100 0 100-50

0

50

100

150

range (m)

altitu

de (m

)

MoonHover TrajectoryTrajectory Timeline

• First, Attitude control system moves Lander horizontally while slowly descending.

• Next, Attitude control system thrusts in opposite direction to cancel horizontal velocity.

• Main engine fires to cancel vertical velocity

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Hopper Trajectory Results

0 20 40 60 80 100 120 140-6

-4

-2

0

2

4

6

time (sec)

vertica

l velo

city (

m/s

)

Vertical Hopper Velocity vs Time

0 20 40 60 80 100 120 140-50

0

50

100

150

200

time (sec)

altitu

de (m

)

Hopper Altitude vs Time

0 20 40 60 80 100 120 140-10

-5

0

5

10

15

20

time (sec)

horiz

ontal v

elocit

y (m/s

)

Horizontal Hopper Velocity vs Time

Backup Slide 1

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Hover Trajectory Results

Backup Slide 2

0 10 20 30 40 50 60 70-3.5

-3

-2.5

-2

-1.5

-1

-0.5

0

time (sec)

vert

ica

l ve

loci

ty (

m/s

)

Vertical Hover Velocity vs Time

0 10 20 30 40 50 60 70-20

0

20

40

60

80

100

time (sec)

alti

tud

e (

m)

Hover Altitude vs Time

0 10 20 30 40 50 60 700

5

10

15

20

time (sec)

ho

rizo

nta

l ve

loci

ty (

m/s

)

Horizontal Hover Velocity vs Time

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m

T

rrr radial

2

2

r

r

mr

Ttheta 2

EOM’s r = distance of the Lander from the center of the moon

θ = angular displacement along the surface of the moon measured from the start of the trajectory

μ = gravitational parameter of the moon equal to 4902.8 km3/s2

Tradial = thrust in the radial (r) directionTtheta = thrust in the angular (θ) directionm = mass of Lander

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Hover Trajectory Assumptions and Constraints

• Initially, Lander comes to complete stop 100m above lunar surface

• Lander remains in upright position throughout trajectory• Lander touches down with near zero horizontal and

vertical velocity• Main lunar descent engine responsible for all vertical

movement• Attitude control system responsible for all horizontal

movement• Lander must cover 500m distance in greater than 60

seconds• Horizontal velocity limited by maximum thrust provided

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-1000 -900 -800 -700 -600 -500 -400 -300 -200 -100 0 100-100

-50

0

50

100

150

200

range (m)

alt

itu

de (

m)

Moon

Hop Trajectory

10kg Hop Trajectory

LiftoffTouchdown

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Hop Trajectory Assumptions and Constraints

• 2-D trajectory in plane normal to Moon’s surface

• Instantaneous throttling of hybrid engine

• Lander takes off and touches down with near zero vertical and horizontal velocity and upright orientation

• Rotation rate of Lander limited by torque provided by attitude control system

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Hop Trajectory Design

• In order to maintain 90% chance of success, cannot relight main lunar descent engine to perform hop.

• Instead use pair of redundant thrusters to perform hop.

• Unusual trajectory shape due to thruster configuration: one thruster firing at 32° from vertical.

• Must offset thrust direction by having Lander velocity in opposite direction to ensure no horizontal velocity upon landing.

• With out this trajectory shape, Lander would crash and/or land on its side.

• This trajectory adds only 2.5kg of propellant compared to a trajectory using a vertical thruster.

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Backup SlidesCory Alban

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Completion of Mission RequirementsStep Time (min) Tasks to be completed

1 0 Space Ball performs a system diagnosis.2 1 Deployment from Lander.

3 2 Direction of travel received from mission control. Space Ball orients to path of travel.

4 2-10 Accelerate to cruising speed of 1.04m/s. Travel for 8 minutes until 500m objective achieved.

5 11Braking maneuver with a 90 degree orientation change to point camera toward Lander. Shake off dust if necessary.

6 12 Snap photo of Lander from ball and begin transmission.7 20 Finish Photo Transmission.

Requirement Steps to Completion

Travel 500m in a controlled manner 1-4

Carry 100g payload 500m 1-4

Transmit Mission Complete Mooncast 6-7

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Lunar Surface Hazard AnalysisPotential Hazard Solution

Lunar regolithVery fine dry powderSticks to everything

Using gradual acceleration, the space ball avoids peeling out and digging into the regolith

Vibration Motor shakes off any collected regolith

Impact Craters2cm to several meters in diameter

Choose path to avoid large craters Built up momentum reduces chance of getting

caught in a crater

Debris/RocksDebris size: 0.0005m to 0.50m

Lexan shell will withstand a full speed collision At cruising speed, momentum carries ball over small

rocks and retains stability (similar to a rolling wheel)

TemperatureAverage day temperature 107CHighest day temperature 123C

Temperatures are within tolerances for Lexan 1atm of N2 inside Lexan shell to control

temperature rise within the space ball Temperatures are within thermal range for Lexan

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Space Ball Structure AnalysisBending Moment in Drive Axel

• Model as a thin circular rod• R = 0.125m

• Aluminum 2024 Alloy• σ= 220 MPa• ρ= 2730 kg/m3

• Maximum loading conditions (8.3g)• g = 8.3 * 9.80665m/s2 = 75.25m/s2

• Mpay = 1.529 kg• Minimum required radius: 1.17*10-8m

Torsion Stress in Drive Axel• Maximum Torque, T = 0.31 Nm• Minimum required radius: 1.10*10-4m

Design radius: 0.003mFactor of Safety: 27

RMpay*g

T

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Space Ball Structure AnalysisSphere Impact Analysis

• Assume all kinetic energy converted to impact energy• Cruise Speed, v = 1.04m/s• Ball Mass, m = 2.435kg• Total Kinetic Energy, K = 1.317J• Impact Strength of Lexan, σ = 600 – 850 J/m• Minimum wall thickness: 1.55*10-3m

Pressure Vessel Analysis• Pressure, P = 101325 Pa (1atm) • Radius of sphere, R = 0.125 m• Maximum Stress of Lexan, σ = 75 Mpa• Minimum wall thickness: 8.4*10-5m

Design wall thickness: 3.82*10-3mFactor of safety: 2.5

R

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Backup SlidesAdham Fakhry

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Final Power Systems 100 gram Lander Mass (kg) Dimensions (m) Cost ($)

Solar Cells 2.0 0.785 m2 250,000

Batteries 0.422 0.1016 X 0.0252 X 0.0709 1500

DC-DC Converters 0.725 0.06 X 0.05 X 0.04 51,000

PCDU (Power Conditions and Distribution unit)

1.9 0.033 X 0.033 X 0.033 12,000

10 kg Lander Mass (kg) Dimensions (m) Cost ($)Solar Cells 2.0 0.785 m2 250,000

Batteries 0.645 0.142 X 0.0534 X 0.1502 2000



1.9 0.033 X 0.033 X 0.033 12,000

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Final Power Systems for Arbitrary

Arbitrary Lander Mass (kg) Dimensions (m) Cost ($)

Solar Cells 2.0 0.785 m2 250,000

Batteries 0.89 0.142 X 0.0276 X 0.095 2000



1.9 0.033 X 0.033 X 0.033 12,000

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Backup Slide 1: Power Available to the Lander

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Battery Design (1)

• Battery is designed for meet three power goals for 100 g Lander:– Delivers 124 W for 250 seconds for operating the

Lander engine– Delivers 30 W for 576 seconds of attitude– Delivers 60.4 W for 30 minutes for all

communication gear

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Battery Design (2)

• Battery is designed for meet three power goals for 10 kg Lander:– Delivers 150 W for 450 seconds for operating the


communication gear

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Battery Design (3)

• Battery is designed for meet three power goals for Large Lander:– Delivers 275 W for 500 seconds for operating the


communication gear

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Solar Array sizing• Solar array Calculations:• Dimensions of Solar cells:

– Area of Lander roof = π(1/2)2 = 0.785 m2

– Solar efficiency = 300 W/m2

– Potential max power = 235.6 W

• Cost of Solar Cells:– Cost of cells per watt = 1000 $/W– Cost of Cells = 235,619.45 = $235,600– Total cost = $235,600 + 4,400 (for additional costs) =

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Hydrazine Tanks• 100 g

– 3.9 kg Hydrazine + 0.3 kg Tank = 4.2 kg– 0.2 m diameter tanks, V= 0.00133 m3

• 10 kg– 4.13 kg Hydrazine + 0.31 kg Tank = 4.41 kg– 0.21 m diameter tanks, V= 0.0015 m3

• Large Payload– 42.6 kg Hydrazine + 2.25 kg Tank = 44.85 kg– 0.43 m diameter tanks, V= 0.0133 m3

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Battery Specifications

• 3.6 V, 20 Ah Lithium Ion Cell• Gives 72 W-hr only need 44 W-hr• Energy Density = 140 W-hr/kg• Dimensions = 0.142 m X 0.0534 m X 0.1502 m• Cost $2000 per cell• From Yardney - Lithion

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Heats of Reaction Calculations

• 10 W 14 days =10W 14 days 24 hrs/day.60 min/s.6 ∙ ∙secs= 12096000 Joules

• Hrxn = -112093 J/mol = 3502916 J/Kg• Mass of Hydrazine = 3.45 kg

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Backup SlidesKelly Leffel

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Schematic of Heat Transfer

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Thermal Control Total

100 gram payload 10 kg payload Arbitrary payload

MLI blanket 2.35 kg 2.38 kg 21.4 kg

Heaters 0.5 kg 0.45 kg 34.1 kg

Cooling System 6.72 kg 6.73 kg 1.03 kg

TOTAL 9.57 kg 9.56 kg 56.53 kg

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Component Mass (kg) Dimensions (m)

MLI blanket 2.35 lays on equip

Al plate 1.4 0.005 x 0.1 m2

Heat pipe 2.6 5 m, Ø 0.0560

Radiators 2.70.005 x 0.311

x0.311

Ammonia 0.021 -

Heaters0.5

0.005 thick

100g

Component Mass (kg)Dimensions

(m)

MLI blanket 2.38 lays on equip

Al plate 1.4 0.005 x 0.1 m2

Heat pipe 2.51 5 m, Ø 0.0575

Radiators 2.80.005 x 0.38 x0.38

Ammonia 0.0215 -

Heaters 0.5 0.005 thick

10kg

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Large Payload

Component Mass (kg) Dimensions (m)

MLI blanket21.4

lays on equip

Al plate1.4 0.5 x 0.1 m^2

Heat pipe10.53 12 m, Ø 0.1039

Radiators22 0.5 x 0.81 x 0.81

m

Ammonia0.1727

-

Heaters 1.03 0.005 thick

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MLI BlanketLander surface, propulsion system, and space balls’ compartments (100 g)

• 30 layers• Aluminized Mylar (0.007 g/cm^2)•Effective emissivity= 0.005

•Q = e*(A)*sb*(Th^4-Tc^4)e = Effective emissivity = 0.005A = Surface area (changes for each lander)sb = Stefan-Boltzmann constant = 5.67 *10^-8 J/K^4.m^2.sTh = Hot temperature (temperature in the sun) = 393 KTc = Cold temperature (temperature in the lander) = 293 K

•Additional 0.4 kg on the 100 g case for the ball storage box

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Heat needed to be removed

Assume 70% efficient equipment

With 40 Watts required, 12 Watts of heat released

Communication Equipment Heat

100g – 49 Watts

10 kg – 38 Watts

Arbitrary – 282 Watts

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• Communication Equipment has a Max Temperature of 313 K, keep at 303 K as a factor of safety

• Keep Lander Operating Temperature around 293 K

• Similar Thermal Control as the OTV– Area of Plate : 0.1 m^2– Aluminum (Al) thermal conductivity : 236 W/(m*K)– Al density: 2700 kg/m^3– Thickness < AK(T1- T2)/q < 3.8 m (for both cases)

• Choose 0.5 cm ( 0.005 m)

– Mass of plate = density * thickness * area = 1.4 kg

Aluminum Plate

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•Ammonia•Latent heat of vaporization of Ammonia: 1371 kJ/kg•Mass (100 g) = 0.061 kW * 450 sec /(1371 kJ/kg) = 0.02 kg•Mass (10 kg) = 0.050 kW * 450 sec /(1371 kJ/kg) = 0.017 kg

•Aluminum Heat Pipes (100g)• Volume needed to simulate P=1 atm : 0.02313 m^3• Choose pipe of 5 m long• 0.00463 m^2 cross sectional area•pi*ri^2 = 0.00463 m^2 : ri = 0.0384 m , ro = 0.0394 m

•Mass = 2700 * pi * (ro^2 – ri^2) * length = 3.3 kg

Heat Pipes

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Heat Pipe Continued

• Aluminum (10 kg)– Volume : 0.01532 m^3, choose length = 5 m– 0.00306 m^2 cross sectional area– pi*ri^2 = 0.00306 m^2 : ri = 0.0312 m , ro = 0.0322 m– Mass = 2700*pi*(0.0322^2-0.0312^2)*5 = 2.7 kg

• Radiators– Dissipate 61 and 50 Watts– Emissivity of 0.92 for white paint– Area of the radiators:0.1762 m^2(100 g) and 0.1444 m^2 (10kg)– Mass = 2.38kg(100 g) , 1.95kg (10 kg)

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Backup SlidesRyan Lehto

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Space Ball Propulsion System Performance•Average Velocity: 1.04 m/s (2.33 mph)•Max Inclination: 14.42°•Acceleration: 0.0043 m/s2

•Power Usage: 0.543 W•Turning Radius: .0625 m (2.46 in)•Largest Boulder Traversable: 0.325m (12.79 in)•Propulsion System Mass: 0.172 kg (0.379 lbm)

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Ball Movement

• Forward/Back Movement • Left/Right Movement

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Motor DataMotor

Power Input (W) Efficiency Power Nominal Output (W) Mass (g)

No-load Velocity (RPM)

Dia (mm) Length (mm) Cost

349190 - RE 6 Ø6 mm, Precious Metal Brushes,

0.3 Watt 0.534 55.2% 0.3 2.3 18500 6 22.9 $58.71 (45.87 Eur)

Gearing Ratio Efficiency Mass (g) Diameter mm Length mm Cost

304181- Planetary Gearhead GP 6 A Ø6 mm 221:1 60% 2.9 6 25.8 $94.55 (73.87 Eur)

Sources: http://shop.maxonmotor.com/ishop and http://motion-controls.globalspec.com

Stepper MotorHolding Torque (Nm) Step Angle Voltage (V) Mass (g)

No-load Velocity (RPM) Cost

ARSAPE Two Phase Stepper Motor -- AM2224-R3AV-4.8 0.045 15° 3 2 18500 $58.71 (45.87 Eur)

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http://shop.maxonmotor.com/ishop

http://motion-controls.globalspec.com/

Alternative Propulsion ComparisonSpace Ball• Motors: 2 (One Stepper & One

Continuous D/C Motor)• Additional Mass: Drive Shaft & Swing

Arms 0.172 kg (0.379 lbm)• Largest Boulder Traversable: 0.325 m

(12.79 in)

Rover• Motors: 4 Continuous D/C• Additional Mass: 4 Wheels and Motor

Mounts 2.513 kg (5.54 lbm)• Largest Boulder Traversable: 0.113 m

(4.45 in)

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Backup SlidesKris Ezra

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Gimbaled Main Engine AlternativeGimbal Mount Specifications:1. Approximate mass of 6 kg2. Angular maximum motion of 20º

– 3 Axis Gimbal

20º

1.06 m

0.3858 m

Mission Length (days) 365Desaturation Maneuvers (#/day) 6Max Reation Wheel Torque (Nm) 0.03H202 Specific Thrust (N/kg) 9.5Attitude Moment Arm (m) 1.1Engine Moment Arm (m) 0.3858

Total mass (kg) (Attitude DS) 6.287081Total mass (kg) (Engine DS) 17.92584

Gimbal Alternative Discarded based on Mass Cost

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Spinning Mass Tether Alternative

Rationale for Discarding Momentum Transfer Concept:The momentum transfer concept was analyzed just using work/energy relationships subject to the conditions that the Lander could not experience an acceleration greater than 10g and that the Lander would initially be traveling at an orbital speed of 1.7 km/s. Because the constraint on the system is an acceleration and the frame of the moving Lander is not inertial, the system was analyzed using work/energy but in an inertial frame. This approach has obvious limitations; however, it also should provide a more conservative analysis meaning that, if the results are unfeasible for this simplified model, the addition of a gravitational component by the moon will only make exacerbate the outcome. Shown below is a plot of the acceleration felt by the Lander versus collision/spring distance through which some force must act to slow the Lander to zero. A reasonable distance for this “collision” would be between 1 and 2 meters since a spring of this relaxed length must be carried on the OTV with a mass less than that of the Lander descent propellant. From the graph, it can be seen that, at this distance, the accelerations are on the order of 1x10 5 Earth g’s. This is four orders of magnitude higher than that sustainable by the communications equipment (10g) and is probably higher than what is able to be withstood by the molecular bonds in the vehicular structure. Additionally, to maintain an acceleration less than 10g during a deceleration from 1.7 km/s it would be necessary to have a collision distance of approximately 150 km. For these reasons among others, the momentum transfer concept is infeasible.

Acceleration sustainable by Communication Equipment: 10gRequired Tether Length to Match Orbital Velocity: ~50 kmAdditional mass cost at this Length: 325 kg (Total mass of 400 kg)Orbital Height: ~100 km

Result: Weight of tether exceeds propellant mass and tether length is nearly half the orbital height. Completely infeasible.

0 10 20 30 40 50 600

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

Tether Length (km)

Mag

nitu

de o

f Li

near

Vel

ocity

(km

/s)

Linear Velocity vs Tether Length

0 10 20 30 40 50 60-400

-350

-300

-250

-200

-150

-100

-50

0

50

100

Tether Length (km)

Mas

s S

avin

gs (

kg)

Mass Savings

v

w2v

v=0

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Backup SlidesAndrew Damon

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X

Y

x

y

Barycenter

Circular-Restricted Three-Body Problem

*2

3 3

(1 )( ) ( 1 )2 x

OTV

Tx xx ny n x

d r m

*2

3 3

(1 )2 y

OTV

Ty yy nx n y

d r m

Two coordinate frames:• and fixed inertially• x and y rotate with Earth-Moon

system

X Y

Equations of Motion (including thrust):

Variable Descriptionxyd

Component of OTV position in x-directionComponent of OTV position in y-direction

Distance from the Earth’s center to the OTV

R Distance from the Moon’s center to the OTVDistance from the Earth’s center to the barycenter

Gravitational parameter

N Mean motion of the system, normalized to 1.0Tx*Ty*m*Mo*

Thrust in the direction of x velocity componentThrust in the direction of y velocity component

Current mass of the OTVInitial mass of the OTV in Earth parking orbit

Mass flow rate of the EP system

*m

G

*Much more accurate than patched two body model

*Gravity effects of Earth and moon are always taken into account

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Recommend: Parking Orbit of 400 km – Drag drops to less than 5% of Thrust, Within capabilities of Dnepr Launch Vehicle

• Assume Thrust of 110 mN

• Assume CD = 1.0

Analysis based on cross section area of:

Solar Panels ~ 8 m2

OTV ~ 4 m2

Total Area ~ 12 m2

Circular Parking Orbit Altitude

(km)Drag (mN)

T/DAssume Thrust of

110 mN

200 76.4 1.44

300 17.8 6.18

400 4.1 26.83

500 0.96 114.6

Atmospheric Drag for Circular Parking Orbits

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Drag Calculations

FD ~ Newtons

ρ ~ kg/m3 CD ~ dimensionless

v ~ m/sA ~ m2

21

2D DF C v A

Backup Slides

Altitude (km) Circular Velocity (km/s)

200 7.78

300 7.73

400 7.67

500 7.61

Curve fit for density based on altitude:

Where h is altitude in km and ρ is in ng/m3

( 570.6)/(-69.3) he

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Backup SlidesLevi Brown

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Correction Maneuver:

50 m/s Burn:Additional Propellant Requirements

100 g – 1.1 kg10 kg – 1.5 kg

Large – 22.5 kg

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Method of Matching Spirals has Errors

Position <6000 km (1.5 % Earth-Moon Distance)

Velocity ≈ 425 m/sRequires ≈ 13 kg Propellant

Better trajectory matching requires more accurate model but

Nothing to indicate infeasibility

Trajectory Mismatch

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Parking Orbit SelectionLower Orbit?

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Parking Orbit SelectionHigher Orbit?

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Backup SlidesSolomon Westerman

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-$15

-$10

-$5

$0

$5

$10

$15

$20

$25

100g Cash Flow Diagram

PurseLaunchR&DIntegrationPurchase CostOverhead

End-of-Year

Mill

ion

$

20102009 2011 2012

Total Cost ($M USD)Launch 4.8R&D 2.3Integration 4.5Purchase 7.2Overhead 9.1Total 27.9

GLXP Prize ($M USD)Grand Prize 20.0Lunar Night 5.0Total 25.0

Lose $2.9 M in 2012 USD!

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Costing Model Differences

1. Overhead– 100g, 10kg

• 15 engineers @ 3 years @ 150k each• 3 STK license, 15 MATLAB license

– Arbitrary• 100 engineers @ 3 years @ 150k each• 10 STK license, 75 MATLAB license

2. R&D – 100g, 10kg

• 20 Engineers @ 150k salary each + 50k per month equipment increases reliability by 2% per month– Arbitrary

• 40 Engineers @ 150k salary each + 50k per month equipment increases reliability by 2% per month

3. Integration– 100g, 10kg

• $10k / kilogram– Arbitrary

• $10k / kilogram

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Backup SlidesBrad Appel

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Electric Propulsion System Setup

# Component Mass [kg] Power [W] Price [$ US]

1 Tank Heating Coil 0.25 5 --

2 Xenon Storage Tank 21 -- 130000*

3 Tank Multi-Layer Insulation -- -- --

4 Solenoid Latch Valve 0.3 4 35,000

5 Pressure Regulator 0.55 -- 20,000

6 High Purity Filter 0.2 -- 5,000

7 Sintered Flow Restrictors 0.4 -- 600

8 Feed System Heating Coil 0.25 5 --

9 Thruster Radiator Sleeve 2 -- 500

10 Hall Thruster 5.7 see PPU 230,000

11 Power Processing Unit 10 1244 535,000

12 Feed Lines 1 --- --

13 Power / Intercomm Harness 2 -- --

14 Xenon Propellant 139 -- 204,000

TOTAL PROP SYSTEM 169 1539 1,030,100

TOTAL IMLEO 418

XENON TANK

S

P

HALL THRUSTER

P

T

PPU T 1

2

3

45

6

7

8

10

11

From PCDUS/C Communication

9

Xenon System

Thermal System

Power / Intercomm

0.2 m

• No redundancies, no integration costs

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Electric Propulsion System Specifications

Specifications for the Hall Thruster – 100g Mission

Variable Value UnitsThrust 78.5 mNSpecific Impulse 1950 sMass Flow Rate 4.1 mg/sPower Input 1526 WEfficiency 0.53 --Input Voltage 350 VDCMass 5.7 kg

Propulsion System Totals – 10kg Mission

Variable Value UnitsWet Mass 215 kgDry Mass 30 kgRequired Power 2,043 WattsBurn time 365 daysThrust 104 mNSpecific Impulse 1964 sMass flow Rate 5.4 mg/s

Specifications for the BHT-8000 Hall Thruster – Large Mission

Variable Value UnitsThrust 424 mNSpecific Impulse 2250 sMass Flow Rate 19.2 mg/sPower Input 7,600 WEfficiency 0.64 --Mass 25 kg

Propulsion System Totals - Large Mission

Variable Value UnitsWet Mass 3,810 kgDry Mass 520 kgRequired Power 38,773 WattsBurn time 365 daysPayload Capability 4,545 kg

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• LOx/LH2 would require an extra 600 kg, costing an extra $2.6M

• An ion thruster could accomplish the mission, but would require much more power than the HET

• Current technology places HET lifetime over 1 year

Other Propulsion Options

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Xenon Storage Thermal Analysis

SupercriticalStorageRegion

Gas

Liquid

Allowed temperature path of propellant

• Maximize storage pressure for volume efficiency (~ 150 bar)• Maintain tank temperature for gaseous Xenon phase: Balance heat due to radiation and pressure drop with a 5 watt resistance heater

•Curve data from National Institute of Standards and Technology

Temperature (K)

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Project X pedition Spacecraft Senior Design – Spring 2009 .

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Transcript of Project X pedition Spacecraft Senior Design – Spring 2009 .