Asteroid Sample Return

AEM 4332Final Design Review

5/7/2008•Becky Wacker•Carla Bodensteiner•Ashley Chipman•John Edquist •Paul Krueger•Jessica Lattimer•Nick Meinhardt•Derek Steffes•Sam Zarovy

1

Becky Wacker

Team OverviewAttachment Method

2

Mission GoalDrill into the asteroid and return 100 g core sample

from a depth of 1 meter for study on Earth.

3

Our Mission Objectives

1. Leave orbit about the asteroid and position for landing

2. Land on the asteroid with drill and return vehicle intact

3. Attach the spacecraft to the asteroid for 24 hours withstanding a 10 N reaction force and a 0.1 N/m torque from the drill

4

Requirements• Spacecraft must weigh no more than 750kg including:

– 50kg payload package (drill and stereo cameras)– 200kg return vehicle

• The spacecraft must land and support drilling operations for 24 hours

• Landing must not exceed 15 g’s

• Spacecraft must support drilling operations:– 10 N reaction force– 0.1 N/m torque

5

Expectations• Explain the results of all system and hardware trade studies including

rationale for final selections

• Provide spacecraft layout drawing(s)

• Demonstrate compliance with all requirements

• Provide high-level block diagram of the spacecraft subsystems

• Define system and hardware requirements sufficiently to allow future subsystem architecture definition and hardware trade studies

• Demonstrate that the landing loads do not exceed requirements and can be supported by the structural/mechanical configuration

6

Assumptions• Orbiter will supply pictures of asteroid to give targeted landing area• Asteroid is only gravitational pull in area• Orbiter can place lander in preferred initial placement• Major Tasks 1 and 2 are most important• Key asteroid physical properties

a. Rotation period is 5.0 hoursb. Semimajor axis: 1.5 AUc. Radius: 10 km (assume spherical shape)d. Mass is 1x1016 kge. Surface rock distribution is Gaussian with 0.25% probability of a 0.5

m rock in a 5 m2 area. f. Surface gravity: 0.006 m/s2

7

Major Tasks1) Perform a trade study of approach, landing and anchoring options

which considers:

a. Control system architectureb. Structural impact loadsc. Anchoring mechanismd. Propulsion system options

2) Develop a spacecraft layout includinga. Mechanical/structural configurationb. GN&C sensor locations and field-of-view (FOV) clearancesc. Location and orientations of reaction control system thrustersd. Telecommunication component locations including antenna FOV clearances

8

Major Tasks Continued3) Identify major hardware components of each spacecraft subsystem

a. Guidance, Navigation, and Controlb. Command and Data Handlingc. Electrical Powerd. Telecommunicationse. Propulsionf. Structures and mechanismsg. Thermal control system

4) Define the key system and hardware requirements5) Determine landing loads and associated structure/absorber sizing

9

Team Organization

C a rla B o de n s te in er

D e rek S te ffes

P a u l K rue g er

Jo h n E dq u ist

S tru c tu re s + M e cha n ism s

Je ss ica L a tt im er

S a m Za ro vy

N ick M e in h a rd t

A sh ley C h ip m an

G N & C + P ro p u ls ion

B e cky W a cker

10

Structures Group Responsibilities

• Design Structure to hold all components and protect payload package and return vehicle

• Perform ANSYS analysis on structural components to ensure structural stability

• Provide solution for thermal control needs• Determine attachment method and prove

effectiveness• Determine Center of Mass of spacecraft

11

GN&C and Propulsion

• Determine proper trajectory for mission• Complete trade studies for GN&C components• Develop high level block diagrams for GN&C

systems• Provide necessary propulsion system to satisfy

mission requirements• Provide solution for telecommunication needs

12

Project Summary

• Approach asteroid from orbiter using a Hohmmann transfer and descend from 500m altitude

• Use pressure regulated tank with monoprop clusters

• Use spikes post-impact to attach to asteroid surface

13

Layout

14

Attachment Method

15

Initial Trade StudyIdea How works Pros Cons

Spikes on Impact

Drive stakes in when land from legs

Don’t need separate propulsion system (less weight)Highest probability for success

Harpoon Shoot something out and reel ourselves in

Uses more fuelMany mechanisms involvedPossibly get tangled to roll back up.

ThrustersPropulsion Only

Use thrusters to hold in place

Extra fuelTorque when drilling

Cork Screw Screws into ground like a cork screw into a cork

Not a high resulting force Complicated mechanisms and possibility of stirring soil instead of screwing in

Spikes Post-Impact

Two arms with small explosion that would rotate and come back while nailing into the ground

Would work…kind ofAble to retry

Even more complicated than harpoon

16

Initial Trade Study Cont.Thrusters Harpoon Spikes on

ImpactSpikes Post-Impact

Corkscrew

Weight Bad Bad Good Okay Bad

MechanicalComplexity

Good Bad Good Okay Bad

FuelConsumption

Bad Good Good Okay Good

Power Requirements

Good Bad Good Good Bad

Dependability Good Bad Okay Good Bad

Applied Loads

Good Good Bad Okay Good

Reusability Good Bad Bad Okay Okay

17

Attachment Method Results

18

Spikes Post-Impact Thruster Attachment

Mechanical Mass 8 kg (approx) 0kg

Fuel Mass 1.198 kg 590.4 kg

Total Mass 9.198 kg 590.4 kg

Attachment Method Overview• Assume built weight for guns to be same as hilti gun

(2kg ea.)• Use F=m(v/t) to determine force from anchor

deployment• t is assumed to be 0.01 s• Deploy two anchors simultaneously at 45° to reduce

thruster requirements• Determine required Thrust and Fuel based on Force• Assume 30s to deploy all 4 anchors• Use Anchor based on future research

19

Gun• Use Mechanical fastening

system based on Hilti Gun model DX E72

• .22 caliber powder actuated fastening tool

• Drives up to 7.1cm nails into concrete or steel

• 2 kg each

http://www.us.hilti.com/holus/modules/prcat/prca_navigation.jsp?OID=-16951

20

Reaction Force

Ballistics information found at: http://www.korabrno.cz/bal-22.htmlUsed: Remington .22 Short CB Cap

• Bullet Weight = 1.87 g• Velocity = 213.36 m/s• Force from Attachment =

(213.36/.01)*(1.87/1000)=39.898N per anchor

21

Spike deployment requires a 21 N force from the 4 thrusters pointing in the +Z direction using a safety factor of 1.5

(56/4=14*1.5(safety factor)=21N/per thruster)

Reaction Force Continued

SpacecraftSpacecraft

+28N+28N+28N+28N

-28N-28N+28N+28N

Gun #1Gun #1Gun #2Gun #2

22

Future Work

• Make gun space worthy-space worthy construction material

• Test force required to insert nails-soil testing experiments

• Determine spike specifications-force experiments

23

Structures and Mechanisms

Paul KruegerJohn EdquistDerek Steffes

Carla Bodensteiner24

Paul Krueger

Spike ExperimentPackaging

25

Force Required to hold Spacecraft

2

2

515

0.006 /

10

1.5

4 spikes

1.5 10 515 0.006 / 2.59

4 spikes

sc

ast

dri

dri sc astfoot

m kg

g m s

F N

n

s

n F m gF

s

N kg m sN

26

Shear Force on Spikes

1 4

2 3

1 1 2 2

1 4

2 3

1.346

2.358

0.1

1.5

2 2

0.018

0.010

sc

sc

R R m

R R m

N m

n

F R F R

F F N

F F N

27

Spike Force Experiment

• Force spike can withstand anchored in ground?

• Driven in at 45 degree angle

• 45 L x 3 D mm spike– 3.6 N normal force– 4.3 N shear force– Within Hilti gun driving

capacity

28

Spacecraft Frame

• Basic truss properties• Opening for return

vehicle and drill• Octagonal Shape• Dimensions based on

interior components– Each side 1m x 2m

• Constructed of aluminum and aluminum honeycomb panels

29

Packaging Trade StudyIdea How works Pros Cons

Dual Shear Plate Mounts electronics on flat plates that are mounted to shear plates that are then bolted to frame

Strong structure,Efficient packaging,Good heat dissipation

Needs custom electronics,Expensive

Shelf Shelf inside spacecraft where electronics are mounted

Uses standard electronics,

Inefficient packaging,Complicated heat dissipation

Skin Panel/Frame Electronics mounted to panels that are mounted to frame

Uses standard electronics,Good heat dissipation,Easy access

Less rigid than Dual Shear Plate configuration

30

Packaging Trade StudyDual Shear Plate Shelf Skin

Panel/Frame

Structure Strength

High Low Medium

Heat-transfer Good Okay Good

Volume needed Small Medium Small

Needs custom electronics

Yes No No

Cost High Low Low

31

Future Work

• Build basic mock up of feet• Add mounting brackets to frame

– Return vehicle, drill, fuel tank, etc

• Lighten spacecraft frame

32

John Edquist

Landing Gear

33

Landing Gear

• Basic Truss– Solid bar, D=20mm

• Low gravity allows smaller structure

• Multiple designs– ANSYS prevented

further testing

Landing Gear Ratcheting System• Gear is spring loaded• Must prevent it from bouncing off asteroid• Stopper prevents craft from bottoming out

Ratchet Spring Calculations• Total mass of craft: 519 kg• Force from gravity = Mass*Gravity

= 1.5*519 * 0.006 m/s2 = 5 N• Force on impact: F=SF*Mass*Velocity/Δt

=1.5*(519 kg)*(2.45 m/s)/(1 sec) = 1912 N• Force per pad = (1912+5)/4 = 479.25 N• Force from weight negligible due to very low gravity• Spring constant k=Force*Distance/Δθ

= (479.3 N)*(0.385 m)/(π/6 Rad) = 352 N*m/rad (per pad)= 6.2 N*m/degree (Δθ=30⁰)

Divide by 4 because 4 springs in parallel per pad:Individual spring constant: 1.5 N*m/degree

Source: http://en.wikipedia.org/wiki/Torsion_spring

Recommendations

• Optimize landing gear to reduce weight and areas of high stress concentration

• Continue design of ratchet system and work with Derek to perform ANSYS testing

• Perform more detailed analysis of internal components and structure for the SolidWorks model

Derek Steffes

Structural Impact LoadsMass Budget

38

Structural Impact Loads

39

Structural Shape

40

Mass Budget• From Design

– GN&C: 21 kg– Propulsion: 26.8 kg– Telecommunication: 10 kg– Structure: 110.4 kg– Thermal protection: 18.4 kg– Attachment guns/spikes: 8 kg– Propellant: 4.4 kg

• Given– Drill: 50 kg– Return vehicle: 200 kg

• Allocated– Electrical: 50 kg– Data handling: 20 kg

Total mass: 519 kgMass margin: 231 kg

41

Center of Mass

42

Structural Analysis• Material Type

– Aluminum Alloy 6061-T6• Source: Alcoa Engineered Products

– Aluminum Honeycomb Paneling 3003• Source: Portafab

• FEA element type– 10-node tetragon

43

Structural Analysis• Impact Load

– Linear momentum-Impulse• mvi+(FImp/SF)Δt=mvf

– Gravitational Force• Fg = SF(mg)

– Total impact load• F = Fimp + Fg

• Valuesm = 519 kgvi = 2.45 m/s

vf = 0 m/s

SF = 1.5Δt = 1.0 sg = -0.006 m/s2

FImp= -1907 N

Fg = -5 N

F = -1912 N44

Landing Assembly (Stress)

45

Landing Assembly (Stress)

46

Landing Assembly (Strain)

47

Landing Assembly (Strain)

48

Octagonal Structure (Stress)

49

Octagonal Structure (Stress)

50

Octagonal Structure (Strain)

51

Octagonal Structure (Strain)

52

Hilti Gun Recoil (Stress)

53

Hilti Gun Recoil (Stress)

54

Hilti Gun Recoil (Strain)

55

Hilti Gun Recoil (Strain)

56

Future Work

• Single-structure analysis– Requires 256,000+ elements

• Dynamic load analysis• Spring-Ratchet system analysis• Non-ideal landing conditions analysis

– Single landing assembly impact

57

Carla Bodensteiner

Thermal Control

58

Requirements

• Define high level diagram of thermal control• Identify major components required• Define system and hardware requirements

sufficiently to allow future thermal control architecture definition and hardware trade studies

59

• Passive Thermal Control– Multilayer Insulation (MLI)– Surface Finishes

• Active Thermal Control– Panel (strip) heaters

• Closed switch controlled by PRTs (Platinum Resistance Thermometers) monitoring optimal temperature for spacecraft components

• Used to heat hydrazine thrusters before initial burns

Basic Hardware Required

60

Active Thermal Control

Electronics and Thrusters

PRT monitor equipment

Command and Data Handling

Panel Heating ON/OFF

61

Plume Protection

62

Plume Protection Trade StudyMaterial Basic

InformationPros Cons

Molybdenum Alloys

Used in turbine blades

Meets temperature requirements

Unsure of space worthiness

PICA Used as heat shield for entry into Earth/Mars atmospheres

Light weight, used in previous space missions

Expensive materials

63

Plume Protection Trade Study Cont.

Molybdenum PICA

Density Okay Good

Space worthiness Unknown Good

Temperature Appropriate

Good Good

Costs Okay Bad

64

Basic Material Properties

PICA (ablative material) (Phenolic Impregnated Carbon Ablators)

• Used on heat shield of Stardust Sample Return Capsule

• Designed by Lockheed Martin• Density of 0.224- 0.321 g/cm3

• Withstand at least 2700 degrees C

65

Propulsion

Nick Meinhardt

66

Propulsion Trade Study for Translational Motion

Main Thrusters Isp (s) Fuel OxidizerPower

Required ThrustFlight

History Complexity

Controlled shutoff/ restart

Monoprop Rocket medium storable N/A N/Amiddle range extensive simple yes

Storable Biprop medium storable storable N/Amiddle range extensive moderate yes

Solid Rocket medium storable N/A N/A high extensive simple no

Cryogenic Biprop highnot

storablenot

storable N/Aextremely

high limited complex yes

Ion Enginevery high storable N/A high

extremely small limited complex yes

Hall Thrustersvery high storable N/A high small limited complex yes

Electrothermal (Arcjet) high storable N/A high small extensive complex yes

Electrothermal (Resistojet) high storable N/A

extremely high small extensive complex yes

1

2

3

67

Propulsion Options for Attitude Control

Attitude Control Adjustment Translation

Power Required

3 Direction Rotation

Rapid Change in Direction

Combined Rot/Trans

Professional's Opinion (Steve Lee)

Monoprop Clusters fine-rough yes no yes yes

yes, off modulation great idea!

Momentum Wheel very fine none yes yes no

yes, with main thrust

probably would use a reaction wheel instead

Reaction Wheel very fine none yes yes no

yes, with main thrust

overkill for this application, usually used for photos

1

68

Thruster Requirements

• Attachment Method Frequired = 21N

• Provide attitude control capability• Provide translational capability• Exit mass flow rate

spe Ig

Fm

0

69

Flow Component RequirementsDesign Drivers:• Minimal pressure drop for rated flow• Monitor fuel conditions• Isolate systems: multi-fault tolerance against thruster fire• Provide parallel & series valve redundancy for reliability

Additional affecting factors:• Low power consumption (affects sizing of power supply)• Low response time (affects minimum turn angle)

70

Fuel Tank Requirements

•

•

- assume incompressible flow

- assume conservation of mass systemtank PPF

attach man error loading

attach man trapped

error loadingtrapped

attachmentmaneuvers

fuelfuel

fuelfuel

fuelfuel

fuelfuelfuel

005.0

03.0

VV

VV

VV

VVV

2

0sppiping

required

HNsystemtank

422

1

gIA

FPP

2pipingHNtank ))()(2

1(42UP

pipingpipingHN )(42

AUm

0 , gIUUmF spee

Burn ΔV or F Burn Time (s)

Fuel(kg)

Fuel(in3)

ΔV1 0.736 m/s 5.07 0.185

ΔV2 -8.07 m/s 55.7 2.03

Fattach 22.5 N 30.0 1.20

Total 4.32 261

71

Pressure Regulated vs. Blowdown Tank Pressurization

Comparison Pressure Regulated Blowdown

Component Complexity Complex Simple

Controls Complexity Simple Complex

Pressure loss None Function of propellant consumed

Loss in thruster force None Function of tank pressure

Mass Additional mass from pressurant monitoring components

Additional mass from added tank thickness to withstand high pressures

Pressurization Required Low 299 psi High 721 psi to complete mission

#tanks needed One Potentially two if not enough fuel volume available in high pressure tank

Change in performance between ΔV and Attitude maneuvers

None Isentropic expansion of pressurant results in temperature drop within the tank temporary lull in pressure.

Fits mission profile Yes No72

Propulsion Isolation Assembly &Thruster Orientation

Parallel

Series

73

Fuel Lines• Same length and number of 90 degree bends

in lines going to each thruster.

• Ltotal = 3.78m

• Bends = 9

74

Propulsion System Losses

20 N HYDRAZINE THRUSTER Model CHT 20

Max Tank Operating Pressure: 385.3 psia

Filter Pressure Drop: 5 psid at rated flow

Latch Valve Pressure Drop: <1 psid at rated flow

Solenoid Valve Pressure Drop: 25 psid (max) at rated flow

Total System Pressure Drop: 31.6 psidThruster Operating Pressure for Required Force: 267 psiaRegulate Tank Pressure Near: 299 psia

In – line pressure drop: 0.625 psid at 283 K, fully developed laminar flow

75

Component PropertiesComponent Product Max Operating

Pressure∆P, Flow rate Power

ConsumptionMax Response Time

Fuel Tank 385.3 psiaN/A N/A N/A

Fuel Filter 585.3 psia 5 psid 0.039 kg/s N/A N/A

Torque-Driven Latch Valve

405.3 psia 1 psid0.00998 kg/s

11.1 watts 50 ms

Dual-Seat Solenoid Valve

385.3 psia 25 psid (max)0.00998 kg/s

27 watts 15 ms

Service Valve 650 psia 30 psid0.0680 kg/s N/A N/A

20 N Thruster 320 psia0.00998 kg/s

27.5 watts(heater: 10

Solenoid: 17.5)-76

Propulsion Mass BudgetComponent Type

Manufacturer/Model Number Qty Unit Mass (kg) Total Mass (kg) Heritage

Propellant Tank AKT-PSI/80222-1 1 1.293 1.293 I.U.E.

Filter VACCO/F1D10638-01 1 0.18 0.180 Not Listed

Pressure Transducer Gulton-Statham/PA4089-450 5 0.26 1.820 DS1

Service Valve (gas) VACCO/V1E10483-01 FDV 1 0.11 0.110 DS1

Service Valve (liq) MOOG/Ref Model 50-856 5 0.209 1.045 MOOG IR & D, PMA, NSTAR

Latch Valve MOOG/Ref Model 52-266 8 0.65 5.200

Koreasat, LM 700, Globalstar, Iridium®, ETS-8, Smart-1, MT-

SAT2, Cosmo Sky Med

Solenoid Valve MOOG/Ref Model 51-281 16 0.299 4.784

GPS, ACE, CSII, Centaur, Skyret 4, COBE, Topex,

SFV,DMSP, USERS, SERVIS

Engines (15N req) EADS Astrium CH20 16 0.395 6.320

EURECA, HAPS, XMM, Integral, METOP, Herschel,

Plank

Tubing 1/4" 0.028" wall tubing  N/A  5.00 5.000 DS1

Pressurant GHe  N/A N/A  0.010 DS1/MER

Total Dry N/A N/A N/A 25.24 N/A

Propellant loaded N2H4 N/A N/A 4.420 Requirements

Total Wet N/A N/A N/A 29.67 N/A77

Off-Modulation

System Impact:

• Operating up to 4 thrusters at a given time to produce translational motion

• Max required flow rate through the latch & solenoid valves

= exit flow rate = F/ve = 0.00998 kg/s

• Max required flow rate through the filter

= 4 * exit flow rate = 0.0399 kg/s

• Max power consumed at once = 262 Watts

Def: turn off thrusters during a ∆V maneuver to compensate for unwanted rotations that are caused by an offset in CM.

78

Future Work• Model pressure control assembly & find

components.• Complete off-modulation dynamic model to

find more accurate fuel consumption information.

• CFD analysis using Monte Carlo methods to determine plume impingement effects on thrust.

79

GN&C

Jessica LattimerSam Zarovy

80

Jessica Lattimer

ApproachGN&C Hardware

81

GN&C Assumptions• There exists space qualified versions of all devices• No redundancy in system devices – single string• Probability of landing on large rocks is so small that hazard avoidance

control system is not needed• Do not need pin point landing• Asteroid is the only gravity force acting on the spacecraft• Only disturbance torque acting on spacecraft is gravity• Cruise stage has cameras and GN&C system to navigate and determine

landing spot• Cruise stage will put spacecraft in any initial orbit we desire• Current project will use radar, next team will do trade study between

LIDAR and radar• Flight computer has enough processing power for all calculation and

command execution• The impulse force during landing only lasts t = 0.1 sec

82

Approach

Hohmman TransferNon-Hohmman Transfer

Time Okay Good

Fuel Efficiency Good Okay

Burn Complexity Good Bad

83

Free Fall Distance

Farther Away Closer

Time Okay Good

Crash Probability Good Bad

Reposition Chance Good Bad

Used a MatLab script to quantitatively analyze free fall distance

84

Approach

85

Attitude Determination

Star tracker Sun Sensors Magnometers GyroscopesEarth-Horizon Scanners

Range of Use Good Okay Bad Good Bad

Reliability Good Okay Bad Okay Bad

Errors Good Good Good Good Good

Weight Good Good Good Good Good

Size Good Good Good Good Good

Power Good Good Good Okay Good

86

Star TrackerGalileo Avionica A-STR Autonomous Star

TrackerFOV: 16.4 X 16.4˚Power Consumption: 8.9 W at 20˚C

13.5 W at 60˚CSize: 195 (L) X 175 (W) X 288 (H) mmMass: 3 kg

87

IMU

• Honeywell Miniature IMU• Mass: 5 kg• Size: 25 (D) X 20 (H) cm• Power Consumption: 30 W

88

Radar

• 3 Antennas– Mass : 1 kg each– Size : 20 cm diameter x 1 cm high each

• Central control box– Mass: 10 kg – Size: 30 cm x 20 cm x 15 cm

• Based off of MSL radar design

89

Radar Field of View

• Antennas only objects on bottom of spacecraft

• Landing gear is at 45˚ angle form vertical in descent

45˚

90

Sam Zarovy

Control System

91

Thruster Logic Vehicle Model

SensorUpdateGuidance

Algorithm+

-

Error

Vehicle State

Control Inputs

Estimated Vehicle State

Control Block

Control Update

Control System Architecture

Desired Vehicle State

92

Trajectory Input

93

Control Block

Attitude Error

Kp_att

+Attitude Update

Ki_att

Kd_att

1/S

S

++

Angular Rate Error

Kp_AR

+

Angular Rate Update

Ki_AR

Kd_AR

1/S

S

++

Velocity ErrorKp_vel

+Velocity Update

Ki_vel

Kd_vel

1/S

S

++

Altitude Error

Kp_alt

+

Altitude Update

Ki_alt

Kd_alt

1/S

S

++

94

Vehicle Model

• Use thruster firing as input to force and moments equations

• Estimate how thrusters will change vehicle state

95

Thruster Logic

• Torque produced by one thruster couple firing = 48 N*m• Minimum angle achievable: X-Axis = 0.0021 deg

Y-Axis = 0.0018 deg Z-Axis = 0.0023 deg

96

Control System Architecture: Sensor Update Block

Star tracker

Attitude Measurement

IMU - Gyros

Angular Rate Measurement

Filter: blend slow update rate of star tracker with fast update rate of gyros. Use highly accurate star tracker to remove gyro bias.

Estimated State

Filter: Use measurement to update estimated attitude and angular rate.

Estimated Attitude

IMU - Accels Acceleration Measurement

Estimated Velocity

Filter: Use measurement to update estimated velocity.

Radar

Altitude Measurement

Estimated Altitude Filter: Use measurement to update estimated

altitude.

Attitude State

Velocity State

Altitude State

Vehicle State

Filter: blend Radar ground speed measurements with acceleration measurements to calculate velocity measurement.

Ground Speed Measurement

Angular Rate StateEstimated Angular Rate

97

Flight Computer

98

Future Work

• LIDAR trade study• Further refine control blocks

99

Telecommunications

Ashley Chipman

100

Telecommunications

Requirement: Report back to orbiter

• Antenna: Conical Log Spiral (from Spacecraft Mission Analysis

and Design) 7.2 GHz receive band (X-band) Transmit 8.5 GHz

• Solid State Amp vs TWTA (traveling wave tube amp): Requires more power input SSA has a lower mass (wt limitations) more reliable (require lower voltages)

101

Low Gain Antenna

RF Switch Assembly

Telemetry Cond Unit

Comm. Det. Unit (Same as main command unit)

Transponder -- solid state amp

Telecommunications

102

Telecommunications

103

Component Mass (kg) Power (W)

Dimensions (m)

Placement

Med Gain Antenna(Conical Log Spiral)

1.36 10 Cone Diam 0.14Tube Diam 0.03Cone Depth 0.16

Ideally on top15 degree FOV

Transponder-- receiver-- transmitter

6.8712.540.0

0.14 x 0.33 x 0.07

Where fit** includes solid state amp

Comm. Detector Unit

N/A N/A N/A Same as main command unit

RF switches/cables(est. 10% of total)

0.908 N/AN/A

TBDTBD

Where needed

Telemetry Conditioner Unit

0.85 3.8 0.067 x 0.238 x 0.158

Where fit

Total Mass: 9.988 kg Total Power: 66.3WTotal Mass: 9.988 kg Total Power: 66.3W

Field of View

Star Tracker AntennaStar Tracker Antenna

1.82 meters

104

Thank you!

• Steve Lee – JPL• Professor Garrard• Professor Flaten• Professor Ketema

105

Questions?

106