1

S y s te m s a n d S a fe ty D o cu m e n ta t io n a n d I A & T

Pro g ra m M a n a g e rsT o d d M o s h er

C h ar les S w en s o nR ees e F u lm er

R an d y J o s t

EC E L EA DT y ler /J e f f

M A E L EA DQ u in n Yo u n g

B u s in e s s L EA D

C & D H

I n s tru m e n ts &S cie n ce

Po we r

O rbit A n a ly s is & O pe ra t io n s

A D A C S

Th e rm a l

M A ED e pu ty L e a d

EC ED e pu ty L e a d

S tru ctu re

M e ch a n is m sM a s s A n a ly s isG ro u n dS ta t io n

C o m m

S o f twa re

D o cu m e n ta t io nS ch e du le

C o s t

Program Managers oversee the Student Design of the satellite

System Engineers report to Program Managers about the status of the satellite

Each Subsystem has a Team Lead who oversees the subsystem design

The Team Lead reports back to the Systems Engineers

The team works cooperatively to achieve the best spacecraft design

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Objectives• Main

– Acquire scientific data on the fundamental parameters of ionospheric density irregularities.

– Validate a quick design-manufacture-fly approach to spacecraft production.

• Other– Flight test of a PDA-level avionics package.– Integrated command and data handling system, including direct software

access to the spacecraft.– Low power 3 axis attitude control system

• Torque coils for main control• Mini reaction wheels for stabilization.

– Use CMOS cameras for star and sun sensors.– Validate Utah State University ground station.– Flight test deployable communication antennas.

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• Orbit– Inclination: 28.5 – 55– Altitude: 350 – 425 km

• Size– Diameter < 18.7”– Height < 18.7”

• Mass– Total (including ½ SIP ring) < 25 kg.– Center of gravity less then +0.25” from center of ICU, and < 12” above separation plain.

• Separation: Light band system (see ICU User guide)– 2 micro switches to signal separation.– ½ SIP ring must remain attached.– 2 x 15 pin connectors for comm. while inside ICU.

• Structure– All fundamental frequencies above 100 Hz.– Acceletation load factors (X,Y,Z + 20g)– Depressurization < 0.76 psi/sec (2.0 safty factor)

• Thermal: TDB• Electrical Conductivity

– R < 2.5 m Ohm between SIP and nanosat– R < 1 M Ohm between return paths an structure.

Requirements

4

Mechanical / Structure

• 33 x 33 x 33 cm in stowed configuration

• Aluminum iso-grid structure

• Deployable honeycomb panels

• Spacecraft mass estimated to be ~ 20 kg

• Nadir pointing with sun tracking constraint YawX

Y

Z

Roll

Pitch

YawX

Y

Z

Roll

Pitch

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F l i g ht C o m pute r(C P U )

&I/O B o ar d

Po we r S u bs y s te m :P o w er Bo ar d sS o lar Ar r ay s

Batte r ies

S cie n ce I n s tru m e n ts :P I P

M ag n eto m eter

NR L D e v ice s :UHF R ec eiv er

G P S

C a m e ra s

Th e rm a l:S en s o r sHeater s

A D A C S :S u n an d E ar th S en s o r s

T o r q u er C o ilsM in i R eac tio n W h eels

Electrical Subsystems fo r USUSA TII

TT& C :T r an s m itte r & R ec e iv er

P a tc h An ten n as

A ddit io n a l M e m o ry

Many of the electrical components have been fabricated.

USUSAT I Boards: CPU boardI/O boardTelemetry boardCamera boardPower boardsPIP board

NRL Board:UHF Receiver & GPS

ADACS System to be the most challenging with a completely new design

NRL

Design Needed

Finalized

Testing

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Telemetry and Communications

165.6

160

19.29

347.3

38.84

2.4

.142

.439

.130

14.5

4

751.602

Overhead

PIP

UHF Receiver

Camera

Power Sys.

Thermal

CPU & MM

ADS

ACS

Magnetom.

Bits/sUSUSAT II

GPS

Total

• Uplink: Based on USUSAT I• 450 MHz

• TEKK KS960-L Transceiver & TNC/Modem

• Patch Antenna Array (x3) + Combiner

• Downlink Based on USUSAT I• 2.3 GHz

• L3 Comm. ST-802 Transmitter

• Patch Antenna Array (x3) + Splitter

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C&DH Overall Capabilities

- Fault tolerant hardware and software- 32-bit 80-MIPS processing power- 64 ADC channels- 71 Digital inputs/outputs- Two RS232 and one RS422 port- A SPI bus able to connect up to 32

additional devices- 1-Wire® bus- DMA oriented telemetry and camera

image buffers

Mass Startup * Standby Peak Average

CPU Board 88.86 g 1188 mW 1010 mW 1300 mW (burn flash) 1100 mWIO Board 89.52 g 0 W 274mW 330 mW 300 mW

Telemetry Board 73.1 g 0 W 70mW 346 mW (downlink) 120 mWCamera Board 78 g 0 W 71mW 1074 mW (image capture) 88 mW

Total 1608 mW

Measured Power

* Only CPU board on, other boards switched off.

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C&DH Overall Capabilities

Mass Startup * Standby Peak Average

CPU Board 88.86 g 1188 mW 1010 mW 1300 mW (burn flash) 1100 mWIO Board 89.52 g 0 W 274mW 330 mW 300 mW

Telemetry Board 73.1 g 0 W 70mW 346 mW (downlink) 120 mWCamera Board 78 g 0 W 71mW 1074 mW (image capture) 88 mW

Total 1608 mW

Measured Power * Only CPU board on, other boards switched off.

CPU IO Board

Telemetry Camera

Command And Control

SEU Current Monitoring

Telemetry formatter mass storage

Downlink

Serial Communications

Additional device interface

Temperature Sensing

Command On/Off

Magnetometer

1 MB Camera image buffer

Horizon/Sun Sensors

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SoftwareWind River Systems’ VxWorks RTOS version 5.4.1

– Truly multitasking

– Highly deterministic

– Market leader

Complete software in C/C++

OS image – 600 Kb

75% of software – 500 Kb

VxWorks on the CPU board can be manipulated

and controlled using a simple Hyperterminal

connection. That could be connected to a

receiver when in space.

in

Both shells come

Power On

BootROMs

TSFS Boot Virtual IO console

VxWorks image download through Port 0 (38400 bps)

Reboot

Both shells come up

Default boot through image

flash

up

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Risk assessment/MitigationRisk element Description Proposed Mitigation

Battery leakage

•Incapacity of assuring battery range temperatures (on the Orbiter)

•Batteries leak due to mechanical and thermal stresses

•Perform local thermal FEA

•Increase the confidence of the design by performing thermo-vac tests with temperature monitoring

•Reliable containment of the battery box and mechanical test validation

Solar panel release mechanism

•Deployable mechanisms are prematurely released due to mechanical stresses or temperature impact on material (SMA, paraffin)

•Cold weld developed on elements from long storage time in un-deployed position

•Premature electrical signal to release mechanism

•Increase the confidence of the design by performing thermo-vac tests followed by visual inspection

•Perform functional test of the mechanism prior and after mechanical test

•Perform thermal test on mechanism with functional test

•Verify from analysis that thermal sensitive release mechanisms are in a safe zone

•Kick off components installation (flex, spring) to prevent cold weld components

•Electrical inhibits on the control board for the release

ADACS •Influence of the torquer-coil on the magnetometer

•Increase the understanding of the magnetometer environment by performing test and calibration

•Locate the magnetometer in a controlled zone of influence (boom, end of solar panel)

Thermal Control

•Incapacity of the nanosatellite to maintain and control temperatures

•Maintain an accurate FEA model of the satellite (on-orbit)

•Provide active or passive thermal control on sensitive components with down linked temperatures

•Perform thermo-vac test

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Safety considerations

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Launch Vehicle Interface

• Mechanical interface: Lightband

• Electrical interface: TBD

• Thermal properties: TBD

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Thermal• Thermal analysis will predict max and min

temperature and heater power predictions for:• on-orbit conditions• launch conditions (in the shuttle)

• USUSat-1 temperature extremes were predicted to be -40 to 85 °C

• Improvements in the thermal design are expected to result in equipment temperatures inside the satellite that are on the order of -20 to +50 °C

• Survival heaters will be used to ensure safe operating temperatures for electronics

USUSat-1 steady state temperatures of the CEE, battery enclosure, and transmitter

USUSat-1 Equipment temperature requirements

ComponentOperating

Temperature Range (°C)

Survival Temperature Range (°C)

Batteries 0 to 30 -20 to 50

Solar Cells -100 to 100 -100 to 170

Electronics Enclosure

-40 to 85 -55 to 125

Transmitter -40 to 85 -40 to 85

Camera -35 to 65 -40 to 100

Temperature Sensors

-55 to 125 -55 to 125

Magnetometer -40 to 85 -55 to 125

ComponentOperating

Temperature Range (°C)

Survival Temperature Range (°C)

Batteries 0 to 30 -20 to 50

Solar Cells -100 to 100 -100 to 170

Electronics Enclosure

-40 to 85 -55 to 125

Transmitter -40 to 85 -40 to 85

Camera -35 to 65 -40 to 100

Temperature Sensors

-55 to 125 -55 to 125

Magnetometer -40 to 85 -55 to 125

ComponentComponentComponentOperating

Temperature Range (°C)

Operating Temperature Range (°C)

Operating Temperature Range (°C)

Survival Temperature Range (°C)

Survival Temperature Range (°C)

Survival Temperature Range (°C)

BatteriesBatteries 0 to 300 to 30 -20 to 50-20 to 50

Solar CellsSolar Cells -100 to 100-100 to 100 -100 to 170-100 to 170

Electronics Enclosure

Electronics Enclosure

-40 to 85-40 to 85 -55 to 125-55 to 125

TransmitterTransmitter -40 to 85-40 to 85 -40 to 85-40 to 85

CameraCamera -35 to 65-35 to 65 -40 to 100-40 to 100

Temperature Sensors

Temperature Sensors

-55 to 125-55 to 125 -55 to 125-55 to 125

MagnetometerMagnetometer -40 to 85-40 to 85 -55 to 125-55 to 125

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USU Ground Station

• Helical array: 2.3 GHz

• Ground station rack

• Cross yagi for uplink of 450 MHz

• Gimbaled System

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Ground Support Equipment• Ground Support Equipment

• Facilities:• USU location of ground station• SDL participation with additional

support equipment• Communications equipment

• Transmitters, receivers, electronics

• Data display and storage• Special Test Equipment

• Electronics for testing satellite electronics

• Electronics to for controlling and monitoring satellite during test

• Test chamber equipment, existing at SDL, for thermal vacuum tests

• Shaker tables for vibration tests located at SDL

Ground station towers located at USU

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USUSat-II Systems Engineering Approach

USUSat-IIRequirements

NASASpace Shuttle Requirements

NASASpace Shuttle Requirements

Ground Station

Requirements

Ground Station

Requirements

Sub-system Detailed Design

•Thermal

•Ground Station

•ADACS

•Orbital Analysis and Operation

•C&DH

•Communication

•Structure

•Instruments and Science

•Software

•Power

NASA ICU Canister

Requirements

Navy Research Laboratory

Payload Requirements

Navy Research Laboratory

Payload Requirements

USUSat-II

Interface Control

Document

USUSat-II

Interface Control

Document

USUSat-II

Model Specification

•Top-level Design•Tracking Values / Metrics•Test Plan•Manufacturing Plan

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Test Plan• Component level:

• Electronics functional tests• Mechanism functional tests

• System level:• Functional tests in ambient

conditions• Functional tests in vacuum• Thermal vacuum tests for thermal

cycling• Thermal balance• Environmental functional tests• Vibration testing

ThermaVac chamber

Average pressure : 10 -6 torr

Temperature range: -30 to 60 °C

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Constraints• Power Constraints:

• USUSAT II requires ~ 25 watts of average power• This is dominated by science instruments, comm, torquer coils, and

large eclipse periods compared to the orbit.• In order to achieve this power deployable panels are needed

• The ability to do nadir pointing with sun constraint will allow a greater amount of power to be achieved

• Mass Constraints:• With the addition of the NRL UHF receiver @ 5 kg and the

deployable solar panels this is a large percentage of the mass • Torquer coil system will have a sizeable mass consumption due to

aggressive spacecraft maneuvers• Communications Constraints:

• No major problems seen in the communications subsystem• Thermal Constraints:

• Problems may arise, but lessons have been learned from USUSAT I

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Education Outreach and Focus

• Elementary School Outreach:• Convenient elementary school located on campus

• Organize field trips to educate students about the space program and its applications

• Make in class appearances to present information about space systems to students

• Focus on freshman and sophomores who are interested in aerospace sciences:• Involve the students with the NANOSAT 3 program by

inviting them to the weekly meeting to present the updates of the spacecraft design

• Small SAT Conference: • A good place to expand the knowledge of anyone who

attends the conference about small satellite programs including NANOSAT 3 and other possible future missions