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Taller de Diseño de Picosatélites (CUBESATS) Taller de Diseño de Picosatélites (CUBESATS) y Estaciones de Tierray Estaciones de Tierray Estaciones de Tierray Estaciones de Tierra

Session Session 44SubsystemsSubsystemsyy

J M l d l CJ M l d l C M d R iM d R iJuan Manuel del CuraJuan Manuel del CuraDirector de Director de ProyectoProyecto, SENER, SENER

DptoDpto. . VehículosVehículos AerospacialesAerospaciales, ,

Mercedes RuizMercedes RuizIngenieraIngeniera de de SistemasSistemas, SENER, SENER

[email protected]@sener.es

ETSIA. UPMETSIA. [email protected]@sener.es

1Taller de Diseño de Picosatélites (CUBESATS) y Estaciones de Tierra. J.M. del Cura

ContentContent

• Picosat=Picodesign?• Picosat subsystems

D t H dli g• System Engineering process• Main elements of a mission/spacecraft

– Data Handling – Communications– Power p

• System drivers • Picosat payloads

– Thermal– Structure

P l ip y

• Picosat subsystems– Attitude and Orbit Control

– Propulsion

– Data Handling – Communications– PowerPower– Thermal– Structure


– Propulsion

D t H dliD t H dliData HandlingData HandlingSubsystemSubsystemSubsystemSubsystem


Data Handling SubsystemData Handling Subsystem

Objectives:Objectives:Objectives:Objectives:• Two main functions of the Data Handling subsystem:

– Distribution of all commands to all S/C elements– Preparation of the information transferred to ground:

• Housekeeping• Mission dataMission data

• Additional functions:– S/C timekeeping– Computer health monitoring– Security interfacesSecurity interfaces

• All these functions can be implemented in a (set of) t ( )


computer(s)


Data Handling Design ProcessData Handling Design ProcessReliability

Command Rate

Number of

channelsOn board

Computer?Bus

constraints

Command Storage?

Mission Time

clock?Satellite Lifetime

Identification of functionsIdentification of

requirements and

gComputer Watchdog

?

ACSconstraintsH/K

RateNumber

of channels

P/L DataRate

Computer I/F

ACS functions

?

Schedule

Determination of the complexity of CDH

functions

Radiation Environmen

t Budget

functions

Determination ofEstimation of size,

mass and power for


Determination of overall CDH level of

complexity

mass and power for each component


System Complexity DefinitionSystem Complexity DefinitionSimple Typical Complex

System ComplexityRequirement or Constraint Simple Typical Complex

Processing CommandsCMD rates 50 cmds/s 50 cmds/s ≥50 cmds/sComputer interface none Computer or stored (not both) yesStored commannds none Computer or stored (not both) not needed

Requirement or Constraint

( )Number of channels <200 channels 300-500 channels > 500 channels

Processing of Telemetry DataTLM rates

H/K data 500-4kbps 4-64kbps 64-256kbpsPayload data none 1-200kbps 10kbps-10MbpsPayload data none 1 200kbps 10kbps 10Mbps

Computer interface none none yesNumber of channels <200 channels 400-700 channels > 500 channels

OtherMission time clock none included includedComputer watchdog none included if OBC includedComputer watchdog none included if OBC includedAOCS functions none none included

Bus Constraints Single Unit Single Unit or Multiple Units Integrated or DistributedReliability-Class B parts

Single String 0,8233 0,761 0,6983Redundant 0 9875 0 9736 0 9496Redundant 0,9875 0,9736 0,9496

Reliability-Class S partsSingle String 0,9394 0,9083 0,8285Redundant 0,9987 0,9964 0,9829

Radiation Environment /total dose) <2krads 2-50krads 50krads-1Mrads


Schedule (in months, after order)Class B parts 6-12 6-12 9-18Class S parts 9-18 9-18 9-24


Parametric Estimation of CDH Size, Mass and PowerParametric Estimation of CDH Size, Mass and PowerSimple Typical Complexp yp p

Command only 1500-3000 2000-4000 5000-6000Telemetry only 1500-3000 4000-6000 9000-10000Combined Systems 2500-6000 6000-9000 13000-15000

Size (cm3)

yCommand only 1,5-2,5 1,5-3,0 4,0-5,0Telemetry only 1,5-2,5 2,5-4,0 6,5-7,5Combined Systems 2 75-5 5 4 5-6 5 9 5-10 5

Mass (kg)

Combined Systems 2,75 5,5 4,5 6,5 9,5 10,5Command only 2 2 2Telemetry only 5-10 10-16 13-20Combined Systems 7 12 13 18 15 25

Power (nominal) (W) Combined Systems 7-12 13-18 15-25(W)



Computer architecture at system levelComputer architecture at system level



Elements of an onboard computerElements of an onboard computer



Definition of the onboard computer:• Identify the spacecraft bus and payload operational modes• Allocate top‐level requirements for the computer system• Define sub system interfaces• Define sub‐system interfaces• Specify baseline computer system

– Define computer systems operational modes and states– Functionally partition and allocate computational

requirements to • spacecraft sub‐systems, hardware, or software

d t ti• ground station– Analyze data flow– Evaluate candidate architectures

S l b i hi– Select basic architecture– Develop baseline system configuration

• Do we need a new computing system, or can we use an old h i l d ifi d


system that is already certified?


Computer functional partitioningComputer functional partitioning



Onboard computer architectureOnboard computer architecture



Computer Resources EstimationComputer Resources Estimation

B h kB h kBenchmark Benchmark ProgrammesProgrammes



Software Estimation ProcessSoftware Estimation Process

Identification of application functions

allocated to the

Breakdown of the function into basic

elementscomputer

elements

Definition of the real-time execution

frequency for each ofBottom

s-upSimilarit

y frequency for each of the basic elementsEstimation of the

SLOC and memory need for each

Bottoms-up

Similarity

function

Estimate throughput requirements

Similarity

Estimate the operating system and

overhead requirementsDetermination of the


requirementsDetermination of the margins for growth and on-orbit spare


Development Phase IssuesDevelopment Phase Issues



Computer system integration and testingComputer system integration and testing



Software EstimationsSoftware EstimationsSize (Kwords)Function Typical Typical

Code Data

CommunicationsCommand Processing 1,0 4,0 7,0 10,0Telemetry Processing 1,0 2,5 3,0 10,0

Attitude Sensor Processing

throughput (KIPS)

Execution Frequency

Attitude Sensor ProcessingRate Gyro 0,8 0,5 9,0 10,0Sun Sensor 0,5 0,1 1,0 1,0Earth Sensor 1,5 0,8 12,0 10,0Magnetometer 0,2 0,1 1,0 2,0Star Tracker 2,0 15,0 2,0 0,01

Attit d D t i ti & C t lAttitude Determination & ControlKinematic Integration 2,0 0,2 15,0 10,0Error Determination 1,0 0,1 12,0 10,0Precession Control 3,3 1,5 30,0 10,0Magnetic Control 1,0 0,2 1,0 2,0Thruster Control 0,6 0,4 1,2 2,0Reaction Wheel Control 1,0 0,3 5,0 2,0CMG Control 1,5 0,3 15,0 10,0Ephemerids Propagation 2,0 0,3 2,0 1,0Complex Ephemerids 3,5 2,5 4,0 0,5Orbit Propagation 13,0 4,0 20,0 1,0

AutonomyAutonomySimple Autonomy 2,0 1,0 1,0 1,0Complex Autonomy 15,0 10,0 20,0 10,0

Fault DetectionMonitors 4,0 1,0 15,0 5,0Fault Coreection 2,0 10,0 5,0 5,0

Other Functions


Other FunctionsPower Management 1,2 0,5 5,0 1,0Thermal Control 0,8 1,5 3,0 0,1Kalman Filter 8,0 1,0 80,0 0,01


Conversion of SLOC to Words of MemoryConversion of SLOC to Words of Memory

Language Assembly I t ti

Bytes per SLOC f 32 bitInstructions per

SLOCSLOC for 32-bit

ProcessorFORTRAN 6 36C 7 42PASCAL 6 36JOVIAL 4 24JOVIAL 4 24ADA 5 30



Operating SystemsOperating Systems

Function Size (Kwords) Typical Comments

Code Data

Executive 3,5 2 0,3n n is the number of tasks scheduled per secondT i l 200

throughput (KIPS)

Typical: n=200Run-Time Kernel 8 4 see

commentsThroughput is included in functions which use the features

I/O Device Handlers

2 0,7 0,05m m is the number of data words handled per secondHandlersBuilt-In Test and Diagnostics

0,7 0,4 0,5 Throughput estimated assuming 0,1Hz

Math Utilities 1,2 0,2 see comments

Throughput is included in estimate of application throughput



Flight ComputersFlight ComputersNo se puede mostrar la imagen. Puede que su equipo no tenga suficiente memoria para abrir la imagen o que ésta esté dañada. Reinicie el equipo y, a continuación, abra el archivo de nuevo. Si sigue apareciendo la x roja, puede que tenga que borrar la imagen e insertarla de nuevo.


DHS Examples DHS Examples –– Generic CubesatGeneric Cubesat


DHS Examples DHS Examples -- AAUAAU


DHS Examples DHS Examples –– XIXI--IVIV

CW-CtoERx-EtoC

Rx-TNCThermometer0 to 7

MPX

CW-TNCCW-EtoC

Rx-CtoERx-TNC

OBC MPX_SEL0 ~2

Reset Signal (Power Sub Sys.)

Tx-TNCTx-CtoE

Tx-EtoC

OBC Program&

(Power Sub Sys.)

SEL Detect

C-DCDC 5VTo Comm

E-DCDC 5V

ROM Read/WritePins

SCL LineROM0ROM0ROM0ROM0

Battery VoltageCharge Current

Battery Charger IC Reset Signal

Sub Sys.

SDA Line

O 0ROM0ROM0ROM0ROM0ROM0ROM0

y g g

(Structure Mother Board)

Solar Cell Current1 t 6MPX


1 to 6MPX


OBCCRNT /ROND /SOLATEMP /VOLT

Uplink CommandFixed length = 17 bytes

Tx-TNC CW-TNC

ANTD /CRNT /DCDCMTQC /POWR/ROMDSOLA /TEMP /VOLTTx TNC CW TNC

Ground Station in UT



Components of ElectronicsComponents of ElectronicsFor Thermometer

ROM READ/WRITE Pin

For ThermometerJumperPin For ROM


ROM ModuleFor CameraXI-II model


Components of ElectronicsComponents of Electronics--(2)(2)OPA M d lOPAmp Module Program Pin


Thermometer Module XI-II model


Components of ElectronicsComponents of Electronics--(3)(3)

PIC 16F877• Clock :4MHz• Memory :8kword• RAM :368bytes• EEPROM :256bytes

ROM (24LC256)

EEPROM :256bytes• Operative Voltage:2.0~5.5V

ROM (24LC256)• I2C serial EEPROM• Memory :256Kbit(32Kbyte)• Memory :256Kbit(32Kbyte)• Max erase/write cycles:100,000• Max write-cycle time :5ms

M l k f 400kH


• Max clock frequency :400kHz

DHS Examples DHS Examples –– QuakesatQuakesat

• Pros Linux– Drivers (baypac & ax25) built‐in– <10k loc+linux = flight software

• 3k loc for low level A/D timers– Utilities already written

• Md5sums ( errror checking)• Bzip2 ( file compression )• Shell utilities• Shell utilities

• Pros Prometheus– 16 channel/16bit A/D built‐in

H d ti /i t t– Hardware timers/interrupts– Multitasking 66 MHz– 32 Meg RAM/128 Meg Flash

• Cons• Cons– Power hog 2.5 W– Flexibility require more testing!!


C i tiC i tiCommunicationsCommunicationsSubsystemSubsystemSubsystemSubsystem


Communications SubsystemCommunications Subsystem

ObjectivesObjectives

• Main functions of the Communication subsystem:– Interface between the spacecraft and the ground system

• Transfer of P/L data• Transfer of H/K dataT f f t d• Transfer of operator commands

– Carrier tracking– Command reception and detectionCommand reception and detection– Telemetry modulation and transmission– Rangingg g– Subsystem operations



Definition of the communications architectureDefinition of the communications architectureUse of relay satellites and

Mission data flow diagram

Data sources,

end users and

locationsQuantity of

data per unit time

Identification of links and ground station locations

relay ground stations? Data processing

location

Identification of Communicationrequirements

diagram unit timeSelection of alternative

communications architectures

q

Transmission delay

Access time

Availability,

reliabilityEvaluate

Design & Size Each

Sampling rates

alternatives and compare

Determination of Data Rates for Each

Link

Design & Size Each Link

Quantization levels


Documentation reasons for selection

Bits per sample


Definition of the communications architectureDefinition of the communications architecture

f l k• Types of links:– Ground station‐to‐satellite uplink

Satellite to ground station do nlink– Satellite‐to‐ground station downlink– Satellite‐to‐satellite crosslink– Intersatellite linkIntersatellite link

• Constraints:– Direct viewDirect view– Frequencies high enough– Satellite‐Ground station geometry


Communications SubsystemCommunications SubsystemArchitectureArchitecture AdvantagesAdvantages DisadvantagesDisadvantagesStore and forward •Low‐cost launch

•Low‐cost satellite•Long message access time and transmission delay (up to several hours)

•Polar coverage with inclined orbit

GEO •No switching between satellites•Ground station antenna tracking often not required

•High‐cost launch•High‐cost satellite•Need for stationkeepingp g•Propagation delay•No coverage of polar regions

Molniya •Provides coverage of polar regions•Low‐cost launch per satellite

•Requires several satellites for continuous coverage of one hemisphere•Need for ground station antenna pointing and satellite handover•Network control more complex•Need for stationkeeping

GEO with crosslink •Communication over greater distance without •Higher satellite complexity and costGEO with crosslink •Communication over greater distance without intermediate ground‐station relay•Reduced propagation delay•No ground stations in foreign territory:

–Increased securityR d d t

•Higher satellite complexity and cost•Need for stationkeeping•Relay satellite and launch costs•No coverage of polar regions

–Reduced cost

Low‐Altitude Multiple satellites with crosslink

•Highly survivable‐multiple paths•Reduced jamming susceptibility due to limited Earth view area•Reduced transmitter power due to low altitude

•Complex link acquisition ground station (antenna pointing, frequency, time)•Complex dynamic network control•Many satellites required for high link availability


p•Low‐cost launch per satellite•Polar coverage with inclined orbit

y q g y


Communications architecture functionsCommunications architecture functions

• TTC:– TrackingTracking– TelemetryC di

••Point to pointPoint to point– Commanding

• Data collection••BroadcastBroadcast

• Data relay



Criteria for selecting the communications architectureCriteria for selecting the communications architecture•Earth coverage

OrbitOrbit

RF S tRF S t

•Range•View times

•Carrier frequencyRF SpectrumRF Spectrum

Data rateData rate

•Legal assignment

•Direct impact on sizeO b d i

CriteriaCriteriaDuty factorDuty factor

•On board processing

•Mission and orbit

E i li bili

LinkLinkAvailabilityAvailability

•Equipment reliability•Redundancies•Time required for reparation•Outgage•Use of alternative links

Link Access TimeLink Access Time •Mission dependant


ThreatThreat •Mission dependant


Main Elements of the on board Communication S/SMain Elements of the on board Communication S/S

Power

Data

OBDHCmds Tlm

TransmitterTransponder A

Receiver

Storage Tlm

OBDH Tlm

OBDH Cmds

LowPassFilter

BandRejectFilter

Transmit

Diplexer

GNC

Antenna A

Gimbal/AntennaControlReceiver

Transmitter

DataStorage Tlm

OBDH Tlm

LowPass

BandReject

RF Switch2P2T

Di l

Electronics

Antenna B

Transponder BReceiver

OBDH Tlm

OBDH CmdsFilter Filter Diplexer

GNC

Gimbal/AntennaControl

ElectronicsTransmitRF Switch

LowPass

Power 2P2T

Cmds Tlm

Filter

LowPass


Cmds TlmOBDH

Filter


TTC Design ProcessTTC Design ProcessData rate

Existing, assigned

Range, orbit and S/C

geometry

and volume

Minimum

elevation angle

Select frequency

Data rate

Definition of requirements

Determination of the required bandwidth

Data rate

W t q

Subsystem trades based on link budget

Worst case rain

conditionsBit error

rate

Receiver

Calculation of performance

Receiver noise temperature

gainpparameters

Estimation of

EIRP

System trades between the different

subsystems

Transmitter

gain

Transmitter

Estimation of subsystem weight

and powerG/T

M i


power MarginDocumentation and

iteration


Main RequirementsMain RequirementsRequirementRequirement Alternative/considerationsAlternative/considerationsqq

Data rates:Command

Health & status TM4000bps typical, 8‐64 bps deep space8000bps is common

Mission/Science Low <32bps; Medium: 32bps‐1Mbps; High>1Mbps‐1Gbps

Data volume Record data, compress data and transmit during longer windows

Data storage Solid‐state recorders 128x106 bitsg

Frequency Using existing assigned frequencies and channels

Bandwidths Use Shannon’s theorem to calculate channel capacity

Po er Use larger antennas higher efficienc amplifiersPower Use larger antennas, higher efficiency amplifiers

Mass Use TWTAs for higher RF power output to reduce antenna size

Beamwidth Different antenna types, beam shapes and beamwidths

EIRP (Eff ti I t i A t i ( i ) i th t itt i t EIRP (Effective Isotropic Radiated Power)

As antenna size (gain) increases, the transmitter power requirement decreases

G/T (Antenna gain to system noise

)

Various communication system temperatures and G/Ts


temperature)


TTC Requirements on other subsystemsTTC Requirements on other subsystemsAOCSAOCS:: PayloadPayload::AOCSAOCS::•• Antenna Pointing (Gimballed)Antenna Pointing (Gimballed)••Pointing reqs of the lesser of 1/10 Pointing reqs of the lesser of 1/10 of antenna beamwidth or 0.3degof antenna beamwidth or 0.3deg

PayloadPayload::••Storing mission dataStoring mission data••RF and EMC interface reqsRF and EMC interface reqs••Special reqs for modulation, coding and decodingSpecial reqs for modulation, coding and decodinggg

•• ClosedClosed--loop pointing reqsloop pointing reqsSpecial reqs for modulation, coding and decodingSpecial reqs for modulation, coding and decoding

OBDHOBDH::••Command and telemetry data ratesCommand and telemetry data rates

TTCTTC ••Clock, bit sync and timing reqsClock, bit sync and timing reqs••22--way comm reqsway comm reqs••Autonomous fault detection and Autonomous fault detection and recovery reqsrecovery reqsSt t /Th lSt t /Th l recovery reqsrecovery reqs••Command and telemetry electrical Command and telemetry electrical I/FI/F

Structure/Thermal:Structure/Thermal:••Heat sinks for travelling wave tube Heat sinks for travelling wave tube amplifiersamplifiers••Heat dissipation of all active boxesHeat dissipation of all active boxes

Propulsion:Propulsion:••NoneNone

Power:Power:••Distribution reqsDistribution reqs

Heat dissipation of all active boxesHeat dissipation of all active boxes••Location of TTC electronics and Location of TTC electronics and antennasantennas•• A clear FoV and movement for all A clear FoV and movement for all

i b ll d ti b ll d t


gimballed antennasgimballed antennas


TTC Constraints on other subsystemsTTC Constraints on other subsystems

AOCSAOCS::PayloadPayload::••Maximum data ratesMaximum data rates

•• Pointing for fixed antennasPointing for fixed antennas•• Pointing lossPointing loss

•• Maximum data volumeMaximum data volume

TTCTTC

OBDHOBDH::••On board storage and On board storage and processingprocessing

Propulsion:Propulsion:Structure/Thermal:Structure/Thermal: Propulsion:Propulsion:••NoneNone

Power:Power:A t d litA t d lit

••Temperature uncertainty Temperature uncertainty leading to frequency leading to frequency uncertaintyuncertainty


••Amount and quality Amount and quality of powerof power


Design ParametersDesign Parameters Antenna sidelobeLevels

Design to minimise. Sidelobes degrade the antenna’s directionalityLevels

Polarization

directionality.

Circular or linear. For reducing losses, compatibility is needed.

FrequencySt bilit

, p y

For quick acquisition: known and stable. Short-term, temperature

DesignDesignParametersParameters

Stability

Capture&

, pand ageing.

Capture: range of frequencies for locking the signal. Tracking:

Tracking Rangelocking the signal. Tracking: range with the signal locked

Diplexer Same antenna for Rx and Tx. Low isolation requires a band rejectIsolation

Coupling between

isolation requires a band-reject filter.

T b d d i b th h l


p gAntennas

To be reduced in both channels


Antenna typesAntenna types



Selection criteriaSelection criteria •Mass•Volume

P f

Volume•Power•Bit error rateN i fiPerformance •Noise figure

•Frequency stability•Insertion loss

SelectionSelectionCriteriaCriteria

•Reliability•Efficiency

Compatibility•With existing systems•SGLS

Oth

•TDRSS

•Technology risk


Other Technology risk•Heritage


Carrier frequenciesCarrier frequencies



Detected powerDetected power



Receiver noiseReceiver noise



Other noise sourcesOther noise sources



Signal to noise and information contentSignal to noise and information content



Signal to noise ratio per bitSignal to noise ratio per bit



Typical command and telemetry characteristicsTypical command and telemetry characteristics


Communications SubsystemCommunications SubsystemTypical Communication Satellite Transponder CharacteristicsTypical Communication Satellite Transponder Characteristics


Communications Examples Communications Examples -- AAUAAU


Communications Examples Communications Examples –– XIXI--IVIV

OBCOBCS

Tx TNCC16C622

Rx TNCC16C 11

Telemetry data Beacon data Up-link command

AD Convert

SensorsSensors

NegotiationMorse encoderPIC16C622 PIC16C711

AX25 Coded datawith Parity

AX25 Coded command

M C d d d

Morse encoderPIC16C716

PLL PLLModulator

MX614Demodulator

MX614

FSK modulated command

Morse Coded dataPTT Control

FSK modulated data

PLLControl

PLLControl

Control

Nishi RF Lab.Custom made

FM transmitter

Nishi RF Lab.Custom made

CW transmitter

Nishi RF Lab.Custom madeFM receiver

FSK modulated data Control

Half wave lengthdipole antenna

Half wave lengthmonopole antenna

FM transmitter CW transmitter FM receiver

Antenna SW

switching


dipole antenna monopole antennaAntenna SW


Tx TNC (AX.25 encoder)Tx TNC (AX.25 encoder)

■Tx TNC：Micro controller PIC16C622－program memory(EPROM) : 2 kbyte－data memory(RAM) : 128 bytedata memory(RAM) : 128 byte－clock : 4 MHz－I/O port : 13 (4 AD Converters)－power consumption : 2 0 mA @ 5V－power consumption : 2.0 mA @ 5V

■Tx TNC receives telemetry data from OBC ■Puts Parity byte for error detection■E d th t l t d t ith AX 25 t l PIC16C622■Encodes the telemetry data with AX.25 protocol■Sends encoded data to FSK modulator

PIC16C622

AX.25 Protocol■This protocol is mainly used for data transmission by HAM■Every Amateur Radio Station all around the world can decode our telemetry data!!!


our telemetry data!!! Flag Destination Source Control PID parity data parity data FCS Flag

AX.25 frame structure(with Parity)

Communications Examples Communications Examples –– XIXI--IVIVTx TNC ProgramTx TNC Program

Start & InitializationStart & Initialization

data from OBC ?No

Receive data from OBCYes

Packetize into AX25 format

Send packet to FSK modulator


p


FM TransmitterFM Transmitter

■FM Transmitter is used to transmit telemetry data■Nishi RF Laboratory custom made transmitter

f 437 490MH－frequency:－band width:－RF output power:

437.490MHz20kHz1W

－input power:－operative temp.:－volume:

under 6W-30 ～+6090×60×10cm FM transmitter

(including CW transmitter)


FM transmitter System Diagram


CW Generator ProgramCW Generator ProgramStart & InitializationStart & Initialization

No

YOBC ready OBC ready

to send data?Data Sampling

Receive data from OBCYes

N

Yesto send data?to send data?

Counter < 10secCounter < 10sec

Data sensing (AD Convert)

UT1 www space t u tokyo ac jp

No

UT1 www.space.t.u-tokyo.ac.jpUT2 AA BB CC UT3 DD EE FF Data SendingUT4 GG HH II

UT5 JK LM NO

Data Sending


UT6 PQ RS TU


Rx TNC (AX.25 decoder)Rx TNC (AX.25 decoder)

■Rx TNC：Micro controller PIC16C711program memory(EPROM) : 1 kbyte－program memory(EPROM) : 1 kbyte

－data memory(RAM) : 64 byte－clock : 4 MHz

4 AD C t (8bit)－4 AD Converters (8bit)－power consumption : 2.0 mA @ 5V

■Rx TNC receives AX.25 encoded command from FSK demodulator

■Decodes it and sends command to OBC PIC16C711

OBC Reset System■If the command is “Reset Command”, resets OBC■Monitors OBC’s current and resets OBC in case of SEL


■Monitors OBC s current and resets OBC in case of SEL(Countermeasure of OBC’s SEL)


Rx TNC ProgramRx TNC Program

Interruption Routine

Start & InitializationMain Routine

Interruption Routine

set ‘Receiving’ flag

Receive Uplink command A/D convert ‘Total I’A/D convert ‘Total I’

g gset Receiving flag

Command = “rset”Command = “rset”or flag rst 1 ?

Yes No

‘Total I’ > Threshold ?‘Total I’ > Threshold ?

Yes

Reset OBCReset OBC

or flag_rst = 1 ?or flag_rst = 1 ?

Wait 10 [ms]

OBC ready to receive?OBC ready to receive?No

Yes

g_flag_rst = 1

[ ][ ]Send serial data to OBCSend serial data to OBCflag_rst = 0flag_rst = 0


clear ‘Receiving’ flag


FM ReceiverFM Receiver

■FM Receiver is used to receive up-link command■FM Receiver is used to receive up link command■Nishi RF Laboratory custom made receiver

－frequency:－input power:

145.835MHzunder 100mW

－receive sensitivity:－receive output:

ti t

under -16dBμ16dBV typ.30 +60－operative temp.:

－volume:-30 ～+6050×60×10cm FM receiver



Antenna ConfigurationAntenna Configuration

Antenna for Transmitters430MHz band Half wavelength dipole antenna

Antenna for Receiver144MHz Half wavelength monopole antenna144MHz Half wavelength monopole antenna



Antenna Pattern (Transmitter)Antenna Pattern (Transmitter)

Antenna Absolute GainTransmitters' Half wavelength dipole Antenna

(dBm)(dBm)

-5 00

0.00

5.00 The gain which we can decode th d t i

-20.00

-15.00

-10.00

5.00 the data in our ground station

-25.00

Gt


Gt,req


Antenna Pattern (Receiver)Antenna Pattern (Receiver)

Antenna GainReceiver's Half wavelength monopole antenna

(dBm)

-30.00

-25.00

-20.00

-45.00

-40.00

-35.00

30.00

-55.00

-50.00



Link Budget (Telemetry Tx)Link Budget (Telemetry Tx)Link B udgetT l t (TD M A )

Sym bol U nit Telem etry R em ark

Frequency f M H z 437.400Transm it P W 0.600 Param eterTransm it P dB W -2.218

Telem etry (TD M A )

Transm itter Line Loss Ll dB -3.000 U sually -1dB～-3dBTransm it A ntenna H alf-Pow er B eam w id θt deg 110.000 Ideal dipl ePeak Transm it A ntenna G ain G pt dB 2.148 Ideal dipl eTransm it A ntenna Pointing O ffset et deg 90.000 U ncontrolledTransm it A ntenna Pointing Loss Lpt dB -8.033

CUBESATComm. System

Transm it A ntenna G ain G t dB -5.885Equiv. Isotropic R adiated Pow er EIR P dB W -11.103Propagation Path Length S km 1439.940 50kbyte/1passSpace Loss Ls dB -148.434Propagation & Polarization Loss La dB -0.470 Polarization (-0.3dB )

Comm. System

Peak R eceive A ntenna G ain G rp dB 12.500 G S 435H S20R eceive A ntenna H alf-Pow er B eam w idtθr deg 29.000 G S 435H S20R eceive A ntenna Pointing Error er deg 15.000 A ssum ptionR eceive A ntenna Pointing Loss Lpr dB -3.210R eceive A ntenna G ain G r dB 9.290

UT’sGround Station

System N oise Tem perature Ts dB K 25.700D ata R ate R bps 1200.000 M X614Eb Eb0 dB 21.390B it Er B ER 0.000

R equired Eb/N 0 R eq Eb/N 0dB -H z 13.000 FSK, B ER =10-5


Im plem entation Loss dB -5.000M argine dB 3.390


Link Budget (Command Rx)Link Budget (Command Rx)Link B udgetU li k C d

Sym bol U nit U plink R em ark

Frequency f M H z 145.835Transm it P W 20.000 Param eterTransm it P dB W 13.010

U plink C om m and

Transm itter Line Loss Ll dB -3.000 U sually -1dB～-3BTransm it A ntenna H alf-Pow er B eam w id θt deg 33.000 G S 144H S12Peak Transm it A ntenna G ain G pt dB 10.000 G S 144H S12Transm it A ntenna Pointing O f fe tet deg 15.000 A ssum ptionTransm it A ntenna Pointing Loss Lpt dB -2.479

UT’sGround Stationg p

Transm it A ntenna G ain G t dB 7.521Equiv. Isotropic R adiated Pow er EIR P dB W 17.531Propagation Path Length S km 1439.940Space Loss Ls dB -138.894Propagation & Polarization Loss La dB -0.470 Polarization (-0.3dB )p gPeak R eceive A nteG rp dB -2.521 M onopoleR eceive Antenna H alf-Pow er B eam w idtθr deg 100.000 M onopoleR eceive Antenna Pointing Error er deg 90.000 U ncontrolledR eceive Antenna Pointing Loss Lpr dB -9.720R eceive Antenna G ain G r dB -12.241

CUBESATComm. System

System N oise Tem perature Ts dB K 31.100D ata R ate R bps 1200.000Eb Eb0 dB 32.634B it Er B ER 0.000

R equired Eb/N 0 R eq Eb/N 0dB -H z 13.000 FSK,B ER =10-5

y


R equired Eb/N 0 R eq Eb/N 0dB H z 13.000 FSK, B ER 10Im plem ention Loss dB -5.000M argine dB 14.634

Communications Examples Communications Examples –– QuakesatQuakesat

QuakeSat Tasking & Data Flow ConceptQuakeSat Tasking & Data Flow Concept

Research TasksNORAD Tracking

Health Files

Mission Data FilesAX.25 Protocol

RequestsResults

NORAD Tracking2 line element sets

Uplink

Additional Ground Stanford

G d Q k Fi d

Tasking Files,New Software

FTP FilesStations

(Fairbanks)Ground Station

(unmanned)

QuakeFinderMission Control

Center (MCC)

FTP Files

Internet Control


MCS Components

Communication Examples Communication Examples –– QuakesatQuakesat

DeployableRadio Antennas

M t tMagnetometerDeployablesolar panels Deployablep y

2-section boom

1 foot 1 foot 2 feet

Total weight = 9 9 lbs (4 5 kg)


Total weight = 9.9 lbs (4.5 kg)

Communications Examples Communications Examples –– QuakesatQuakesat

CommunicationCommunication

• 9600 baud, AX.25 packet system• Stanford developed a customized version with PFR/PFS to

handle packet control of long files (fill holes)handle packet control of long files (fill holes)• Typical magnetometer and housekeeping file length is 100‐

300kB L t fil i kB– Longest file in one pass: 700kB

– Avg. 8 magnetometer collects per day (1 MB)• Beacon every 10 sec. (disabled w/ mag. collects)y g

– 33 data points plus time and date• Stanford Ground Station (SGS)

Access via Internet remote controlled standardized I/F– Access via Internet, remote controlled, standardized I/F– 15 db Yagi, auto antenna control using El Sets– New features being added, (polarity control, signal strength)


PPPowerPowerSubsystemSubsystemSubsystemSubsystem


Power SubsystemPower Subsystem

ObjectivesObjectivesProvide a stable and reliable energy supply to all the S/C subsystems andProvide a stable and reliable energy supply to all the S/C subsystems andpayloads during all the mission life. For this to be done the Power subsystemshall:

• Generate and store electric energy to be supplied to other S/C subsystems• Control the electric current flow:

– To the secondary energy source (batteries)– To be distributed to the S/C subsystems

• Distribute the electric power.• Adapt the current to the different equipments requirements.• Autonomous power management during Sun-Eclipse transitions.• Protect all the electric and electronic equipments against power failures or

system degradation.



Power subsystem components (I/II)Power subsystem components (I/II)

Primary energy source Power distribution, control and main bus protection

Power source

Power conversion

Charge control Discharge bypass

Source control

Load

s

Secondary energy source

Shunt voltage limiter Regulation

wer

con

ditio

ning

Energy storage


Pow

Energy storage control



Primary energy source

Power subsystem components (II/II)Power subsystem components (II/II)

Power source

Primary energy source

• Solar radiation• Chemical energy• Nuclear energy

Power conversion

Source control

gy

• Solar cell arrays• Fuel cells• RTG in

g


Pow

er c

ondi

tion

• DC-DC Converters• DC-AC Converters• Current/Voltage

adaptors

Energy storage • Batteries• Fuel cells

P adaptors

Energy storage control



Average versus peak powerAverage versus peak power

Power (W)

300300 W Peak Power

200

100 90 W Average Power

1 2 3 4 5 6 7 8 9 10 T (Hours)



Typical power requirementsTypical power requirements

T Average power P k (kW)Type g p(kW) Peak power (kW)

Pico satellites ~10‐3 ~10‐3

Micro satellites 10‐3 – 10‐1 0.1 – 0.2

Small satellites 0.1 – 0.3 0.2 – 0.43

Comm. Satellites (GEO) 1.5 – 5.5 2.0 – 6.5

C S t llit (LEO) 8 Comm. Satellites (LEO) 0.5 ‐ 0.8 0.7 – 1.2

Remote sensing 2.0 – 6.5 2.8 – 8.7

I l b Interplanetary probes 0.3 – 0.5 0.8 – 1.0

Space shuttle 10 ‐ 15 13 ‐ 17


Space platforms 25 ‐ 110 50 ‐ 150

Manned Mars mission 2000 ‐ 4000


Power subsystems evolutionPower subsystems evolution



Design requirementsDesign requirements

• In-orbit autonomous and continous power supply to S/C equipmentsand payloads.

• A t t d i S E li t iti• Autonomous power management during Sun-Eclipse transitions.• Simplicity in power interface with the loads.• High reliability applying modular design and redundanciesHigh reliability, applying modular design and redundancies.• Protection against failures and degradation.• Minimum mass to optimise the charge capacity.p g p y• Minimum recurring cost.• Bus voltage compatible with existing equipments and payloads.



Design drivers (I/II)Design drivers (I/II)• Mission

– Client / Final user– Distance to Sun– Manouvering

• Orbital parameters– Altitude– InclinationManouvering

• Vehicle configuration– Mass restrictions– Size

– Eclipse cycles• Payload requirements

– Power, voltage and currentD t l k– Launch vehicle imposed restrictions

– Thermal dissipation capacity• Duty cycle

T t l i i lif ti

– Duty cycle, power peaks– Protection against failures

– Total mission lifetime– Power levels in different modes– Power levels during different mission phases

• Attitude controlAttitude control– Spinning S/C– 3-Axis stabilisation– Pointing requirements


– Thrusters position


Design drivers (II/II)Design drivers (II/II)



Preliminary design process for the Power SubsystemPreliminary design process for the Power SubsystemStep Information Required Derived Requirements

1. Identify requirements

Top‐level requirements, mission type (LEO, GEO), spacecraft configuration, mission life,

Design requirements, spacecraft electrical power profile (average

d k)configuration, mission life, payload definition and peak)

S l t d i

Mission type, spacecraft configuration, average load

EOL power requirement, type of solar cell, mass and area of solar

l fi ti (2. Select and size power source g , grequirements for electrical requirements

array, solar array configuration (2‐axis tracking panel, body‐mounted)

Eclipse and load‐leveling enerfy

3. Select and size energy storageMission orbital parameters, average and peak load requirements for electrical power

Eclipse and load‐leveling, enerfy storage requirement (battery capacity requirement), battery mass and volume, battery type

4. Identify power regulation and control

Power source selection, mission life, requirements for regulating mission load, thermal control requirements

Peak‐power tracker or direct‐energy‐transfer system, thermal‐control requirements. Bus‐voltage quality power control algorithms


requirements quality, power control algorithms

Power SubsystemPower SubsystemEffects of systemEffects of system--level parameters on the Power Subsystemlevel parameters on the Power Subsystem

Parameter Effects on designg

Average electrical power requirement

Sizes the power generation system (e.g., number of solar cells, primary battery size) and possibly the energy storage system given the eclipse period and depth of discharge

Peak electrical power required

Sizes the energy storage system (e.g., number of batteries, capacitor bank size) and the power processing and distribution equipment

Mission lifeLonger mission life (>7 yr) implies extra redundancy design, independent battery charging, larger capacity batteries and larger arrays

Orbital parameters Defines incident solar energy, eclipse/Sun periods and radiation environment

Spacecraft configurationSpinner typically implies body‐mounted solar cells; 3‐axis stabilised typically implies body‐fixed and deployable solar panels


p


Common spacecraft power sources comparisonCommon spacecraft power sources comparisonEPS Design P t Solar photovoltaic Radio‐isotope Fuel cellParameters p p

Power range [kW] 0.2 ‐ 300 0.2 ‐ 10 0.2 ‐ 50

Specific power [W/kg] 25 ‐ 200 5 ‐ 20 275

Specific cost [$/W] 800 – 3000 16K – 200K Insufficient data

Low‐orbit drag High Low Low

Degradation over life Medium Low Low

Storage required for solar eclipse Yes No No

Sensitivity to Sun angle Medium None None

Sensitivity to S/C shadowing Low (with bypass diodes) None None

Obstruction of S/C Hi h L N/viewing High Low None

IR signature Low Medium Medium

Principal applications Earth‐orbiting spacecrafts Inter‐planetary Manned missions


Principal applications Earth orbiting spacecrafts Inter planetary Manned missions


Issues in designing the energy storage capacityIssues in designing the energy storage capacity

Issue Effects on design

Physical Size, weight, configuration, operating position, static and dynamic environments.

Electrical Voltage, current loading, duty cycles, activation time and storage time and limits on depth of discharge.

Programmatic Cost, shell and cycle life, mission, reliability, maintainability and produceability.



Solar photovoltaicSolar photovoltaic

Pout = Pin· ·cos ()

• Pout : Solar cell’s output power density (W/m2)

• Pin : Incoming solar power density (W/m2)

• : Solar cell’s energy conversion efficiency

• : Incidence angle (deg or rad)



Solar cell efficiencySolar cell efficiencyCell type Theoretical efficiency Achieved efficiencyCell type Theoretical efficiency Achieved efficiency

Thin sheet Amorphus Si 12 % 5 %

14 8 %Silicon (Si) 20.8 % 14.8 %

Gallium Arsenide (GaAs) 23.5 % 18.5 %

GaAs/Ge 19 % N/A

Indium Phosphide 22.8 % 18 %

Multijunction (GaInP/GaAs) 25.8 % 22 %

• Energy to solar array area for Si: ≈ 120 – 210 W/m2

• Energy to solar array area for GaAs: ≈ 170 – 260 W/m2



1. Determine requirements and constrains for power subsystem solar array

Solar array design processSolar array design process1. Determine requirements and constrains for power subsystem solar array

design:• Average power required during daylight and eclipses• Orbit altitude and duration• Design lifetime

2. Calculate amount of power that must be produced by the solar arrays3. Select type of solar cell and estimate power output with the Sun normal to the3. Select type of solar cell and estimate power output with the Sun normal to the

surface of the cells4. Determine the beginning of life (BOL) power production capability per unit area

of the array5. Determine the end of life (EOL) power production capability for the solar array6. Estimate the solar array area required to produce the necessary power based

on EOL power and alternate approach7. Estimate the mass of the solar array8. Document assumptions


Taller de Diseño Preliminar de Satélites, Escuela Técnica Superior de Ingenieros Aeronáuticos, 2009

© SENER Ingeniería y Sistemas S.A. Proprietary Information


Solar array design processSolar array design process



BatteriesBatteries

• Batteries store electrical energy in the form of chemical energy.• Spacecraft use two kind of batteries:

Primary batteries: Designed for single use for short missions (less than one month)and firing of pyrotechnic devices. Can’t be recharged by reversing the dischargecurrent flow (e.g. Silver Zinc).

Secondary batteries: Can be discharged and recharged several times. Use as partof the main power system to supply the load during eclipse and whenever the loadexceeds the solar array capability (e g NiCd NiH2)exceeds the solar array capability (e.g. NiCd, NiH2).

• Eclipse time, design life and average power requirements drive storage capacityrequirement.q

• Depth-of-discharge and number of discharge cycles determine available capacity.





BatteriesBatteries



Characteristics of primary batteriesCharacteristics of primary batteries

Primary battery couple

Specific energy density

[W h /k ]Typical applicationp [W∙hr/kg]

Silver Zinc 60 ‐ 130 High rate, short life (minutes)

Lithium Thiorryl Medium rate moderate life (< 4 Lithium Thiorryl Chloride 175 – 440 Medium rate, moderate life (< 4

hours)

Lithium Sulfur Dioxide 130 350 Low/medium rate long life (days)Dioxide 130 – 350 Low/medium rate, long life (days)

Lithium M fl id 130 – 350 Low rate, long life (months)Monofluoride 3 35 , g ( )

Thermal 90 ‐ 200 High rate, very short life (minutes)



Characteristics of secondary batteriesCharacteristics of secondary batteries

Secondary battery couple Specific energy density[W∙hr/kg] Status

Nickel‐Cadmium 25 ‐ 30 Space‐qualified, extensive database

Nickel‐Hydrogen (individual pressure vessel design)

35 – 43 Space‐qualified, good database

Nickel‐Hydrogen (common pressure vessel design)

40 – 56 Space‐qualified for GEO and planetary

Nickel‐Hydrogen (single S lifi dNickel Hydrogen (single pressure vessel design) 43 – 57 Space‐qualified

Lithium‐Ion (LiSO2, LiCF, LiSOCl2)

70 ‐ 110 Under development

Sodium‐Sulfur 140 ‐ 210 Under development



Characteristics of secondary batteriesCharacteristics of secondary batteries



Steps in the energy storage subsystem designSteps in the energy storage subsystem design

1. Determine the energy storage requirements. Consider:• Mission length• Primary and secondary power storage• Primary and secondary power storage• Orbital parameters (eclipse frequency, eclipse length)• Power use profile (voltage and current, DOD, duty cycles)• B tt h /di h l li it• Battery charge/discharge cycle limits

2. Select the type of secondary batteries3. Determine the size of the batteries (batteries capacity):

• N b f b i• Number of batteries• Transmission efficiency between the battery and the load

Battery capacity (for battery capacity in Amp-hr, divide by bus voltage)

hrWTPC ee ··


hrWnNDOD

Cr ··


Steps in the energy storage subsystem designSteps in the energy storage subsystem design



1 Determine the electrical load profile Consider:

Steps in the power distribution subsystem designSteps in the power distribution subsystem design1. Determine the electrical load profile. Consider:

• All spacecraft loads, their duty cycles and special operation modes• Inverters and arc requirements• Transient beha ior ithin each load• Transient behavior within each load• Load-failure isolation

2 D id t li d d t li d t l2. Decide on centralised or decentralised control:• Transient behaviour within each load• Total system mass

3. Determine the fault protection subsystem:• Detection (active or passive)• Isolation• Correction (change devices, reset fuses, work around lost subsystem)



Steps in the power regulation and control subsystem designSteps in the power regulation and control subsystem design

1. Determine the power source. Consider:• All spacecraft loads, their duty cycles and special operation modes (primary

batteries, photovoltaic, static power, dynamic power)2. Decide the electrical control subsystem:

• Power source• Battery charging (peak power tracking, direct energy transfer)• Spacecraft heating

3. Develop the electrical bus voltage control:• How much control does each load require? (unregulated, quasi-regulated, fully

regulated)• Battery voltage variation from charge to discharge• Battery recharge subsystem (parallel or individual charging, < 5yrs – parallel

charge, >5yrs – independent charge)• Battery cycle life


• Total system mass

Power Examples Power Examples –– Generic CubesatsGeneric Cubesats



2-section boom


Power Examples Power Examples –– XIXI--IVIV

ChargeCi i


AAAA

CircuitA A

M t t

AAAAAAAAAAAAAA TNC OBC OBC

MagnetometerDeployablesolar panels Deployable

Batteries SwitchingRegulator

SwitchingRegulator

DCDCConverterp y

2-section boomRegulator Regulator Converter

Electronics CommunicatiS b

Tx


Subsystem on Subsytem


Power Regulation & ControlPower Regulation & Control

• Bus voltage: main 5[V]• Regulated to 5V using switching

regulators and DCDC converter regulators and DCDC converter • Elect. subsystem power line &

Comm. subsystem power lines y pare independent so that they monitor each other and shutdown in case of SELshutdown in case of SEL


SourceSource

• Power is supplied by body mounted solar cells.• Cells are arranged on all 6 CubeSat surfaces• Cells are arranged on all 6 CubeSat surfaces.• Average power 1228 [mW] (typ @ 80 )



Solar PanelSolar Panel

Bass bar

■Cell type : Si Crystal (SHARP)■Efficiency : 16%■10 cells in series / panel■Cell size:■Cell size:

+X :28.25x13.8mm-X,+Y,-Y:47.75x13.8mm+Z Z 47 75 15 8+Z,-Z :47.75x15.8mm


Photo:3 cells in series


Solar Array Layout (+X panel)Solar Array Layout (+X panel)

+X panel:

4.5V x 172mA = 774mW(typ @ 25 )(typ. @ 25 )

4.5V x 162mA = 727mW(typ. @ 80 )



Solar Array Layout (Solar Array Layout (--X,+Y,X,+Y,--Y panel)Y panel)

-X,+Y,-Y panels:

4.5V x 297mA = 1336mW(typ ＠ 25 )(typ. ＠ 25 )

4.5V x 279mA = 1256mW(typ. ＠ 80 )



Solar Array Layout (+Z,Solar Array Layout (+Z,--Z panel)Z panel)

+Z,-Z panels:, p

4.5V x 340mA = 1530mW(t @ 25 )(typ. @ 25 )

4.5V x 319mA = 1438mW(typ. @ 80 )( yp @ )



Energy StorageEnergy Storage

• Batteries will be used during eclipse and downlinkLii d b i • Liion secondary batteries are selected.

• 8 batteries are set in parallel8 batteries are set in parallel.• DOD is 3%• Batteries only lifetime is 38 hrsBatteries only lifetime is 38 hrs



Liion batteryLiion battery

Cathode Material Lithium Manganate

A d M i l C bAnode Material Carbon

Operating Voltage 3 8[V]Operating Voltage 3.8[V]

Discharge Capacity 780 [mAhr]Discharge Capacity 780 [mAhr]

Single Cell Spec


Single Cell Spec.


Battery ChargerBattery Charger• 3 candidates for Battery Charge Circuit3 candidates for Battery Charge Circuit

MAX1679 MM1333 MM1485MAX1679 MM1333 MM1485

•Small package (8 pins), •Small package (8 pins), •Small power dissipation•Const. Voltage &

small power dissipation•Voltage&Temperature protection

small power dissipation•Const. Voltage & Current Charge Mode

gCurrent Charge Mode•Pre-charge Temperature protectionp

•Pre-charge, Timeout

•Need constant reset

g

•No pre-charge func or temperature protection

Temperature protection

•Large package (16 pins) and may be difficult to


Need constant reset before IC’s timeout

temperature protection and may be difficult to assembly


Energy ConsumptionEnergy Consumption

Components Power[mW] Frequency in use

OBC 20 All timessensors 20 All timesTx TNC 20 During downlinkTx 6000 During downlinkCW 300/125 All times (ON / OFF)CW TNC 20 All timesR 125 All iRx 125 All timesRx TNC 20 All timesCamera 150 Sometimes


Camera 150 SometimesMagnetic Plg. 800 Antennae deployment


Power BalancePower Balance

P i t• Points– Beacon can be sent by solar panels direct drive

Source and consumption must be balanced– Source and consumption must be balanced

• Solar cell average output 1228[mW] > Consumption at beacon use 900[mW]beacon use 900[mW]

• Maximum average supply power: 669[mW] > Average • Maximum average supply power: 669[mW] > Average consumption 616[mW] OK


OK

Power Examples Power Examples –– QuakesatQuakesat



2-section boom

1 foot 1 foot 2 feet

Total weight = 9 9 lbs (4 5 kg)


Total weight = 9.9 lbs (4.5 kg)

St tSt tStructureStructureSubsystemSubsystemSubsystemSubsystem


Structure SubsystemStructure Subsystem

ObjectivesObjectives

• Main functions of the structure subsystem:– Mechanically support all other spacecraft subsystems– Attaches the spacecraft to the launcher– Provides for ordnance‐activated separation

All tiff d t th i t– All stiffness and strength reuirements– Interfaces with booster

• T o elements• Two elements:– Primary structure, supporting major loads– Secondary structure for auxiliary elements weighting less – Secondary structure, for auxiliary elements weighting less

than 5kg



Structure Design ProcessStructure Design ProcessEnvelope

Accessibility

MissionLaunchVehicle

Environments

SubsystemRequirements

EnvelopeProducibility

Define Load Paths

Definition of requirementsDevelopment of configurations

Definition of Design OptionsDefinition of Test/Analysis

TestCriteria

DesignC i i

Construction Material

CriteriaCriteria

Options Options

Sizing and checkingi t


requirements Detailed Design


Main RequirementsMain Requirements

All Ph f f h d Mission PhaseMission Phase Source of RequirementsSource of Requirements• All Phases, from manufacture to the end of the mission

Mission PhaseMission Phase Source of RequirementsSource of Requirements

Manufacture and Assembly

•Handling fixture or container reactions• Stressess induced by manufacturing processes (welding)

Transport and dli

• Crane or dolly reactionsHandling • Land, sea or air transport environments

Testing • Environments from vibration or acoustic tests• Test fixture raction loads

Prelaunch •Handling during stacking sequence and pre‐flight checksPrelaunch Handling during stacking sequence and pre flight checks

Launch and Ascent

• Steady‐state booster accelerations• Vibro‐acoustic noise during launch and transonic phase• Propulsion system engine vibrationsT i l d i hi l ll l h d l d f i i • Transient loads, stage separations, vehicle manoeuvres, propellant slosh and payload fairing

separation• Pyrotechnic shock from separation events, deployments• Thermal environments

Mi i S d h l iMission operations

• Steady‐state thruster accelerations• Transient loads during attitude manoeuvres and attitude control burns or docking events• Pyrotechnic shocks from separation events, deployments• Thermal environments


Reentry and Landing

• Aerodynamic heating• Transient wind and landing loads


Main RequirementsMain RequirementsS/CS/C

Allowable Weight driven by:Allowable Weight driven by:•• LauncherLauncher

O bitO bitS/CS/CWeightWeight

•• OrbitOrbit•• Upper StageUpper Stage•• Weight growthWeight growth

S/CS/CSizeSize

Allowable Size driven by:Allowable Size driven by:•• Launcher fairingLauncher fairing

LaunchLaunchVehicleVehicle

S/CS/CS/CS/CRigidityRigidity

Required Rigidity driven by:Required Rigidity driven by:•• Avoiding natural frequenciesAvoiding natural frequencies

S/CS/CStrengthStrength

Required Strength driven by:Required Strength driven by:•• SteadySteady--state accelerations state accelerations (load factors)(load factors)•• Random/Acoustic vibrationRandom/Acoustic vibration


StrengthStrength •• Random/Acoustic vibrationRandom/Acoustic vibration•• Shock levelsShock levels


Development of the Configuration Development of the Configuration -- ConstraintsConstraints

AOCS:AOCS:•• StabilizationStabilization

Payload:Payload:•• AccomodationAccomodation

TTC Antennae:TTC Antennae:•• RigidityRigidity

•• SensorsSensors •• SizeSize •• Thermoelastic stabilityThermoelastic stability•• Clear FoVClear FoV

StructureStructure

OBDH:OBDH:•• Radiation shieldingRadiation shielding

Wi i l thWi i l th

Concurrent design:Concurrent design:•• WiringWiring•• PipingPiping

•• Wiring lengthWiring length

Propulsion:Propulsion:

•• Components distributionComponents distribution

Propulsion:Propulsion:•• Engine modulesEngine modules•• PeripheryPeriphery

N t i tiN t i ti

Power:Power:•• Stowed S/AStowed S/A

Thermal:Thermal:•• ComponentsComponents


•• No contaminationNo contamination•• Batteries accessBatteries access

Components Components distributiondistribution


Design optionsDesign options

Methods of Methods of constructionconstruction

MaterialsMaterials Type of Type of structurestructure

Trade Trade

structurestructure

StudiesStudies

WeightWeight CostCost RiskRisk


gg


Design optionsDesign options Methods of Methods of t tit ti

Methods of construction:Methods of construction:constructionconstruction

Type ofType ofMaterials:Materials:Type of structure:Type of structure:

•• FoldingFolding•• MachiningMachining•• AdhesivesAdhesives•• WeldingWelding

MaterialsMaterials Type of Type of structurestructure

•• StrengthStrength•• StiffnessStiffness•• Density (Weight)Density (Weight)

Th l d ti itTh l d ti it

Type of structure:Type of structure:•• Skin panel assembliesSkin panel assemblies•• TrussesTrusses•• Ring framesRing frames

•• WeldingWelding•• FastenersFasteners

•• Thermal conductivityThermal conductivity•• Thermal expansionThermal expansion•• Corrosion resistanceCorrosion resistance•• CostCost

•• Ring framesRing frames•• Pressure vesselsPressure vessels•• FittingsFittings•• BracketsBracketsCostCost

•• Ductility (can prevent Ductility (can prevent cracks)cracks)•• Fracture toughnessFracture toughness

•• Equipment boxesEquipment boxes

•• MonocoqueMonocoque•• Ease of fabricationEase of fabrication•• Versatility of attachment Versatility of attachment options (welding)options (welding)•• AvailabilityAvailability

•• SemimonocoqueSemimonocoque•• SkinSkin--stringerstringer•• SandwichSandwich


AvailabilityAvailability


Material selection (I/II)Material selection (I/II)• MaterialMaterial AdvantagesAdvantages DisadvantagesDisadvantages

Aluminium •High strength vs weight•Ductie; tolerant of concentrated stresses• Easy to machine• Low density; efficient in compression

• Relatively low strength vs volume• Low hardness•High coefficient of thermal expansion

Steel •High strength•Wide range of strength, hardness and ductility obtained by treatment

• Not efficient for stability (high density)•Most are hard to machine•Magnetic

Heat‐i t t

•High strength vs volume • Not efficient for stability (high density)resistant • Strength retained at high temperatures

•Ductile• Not as hard as some steels

Magnesium • Low density, very efficient for stability • Susceptible to corrosion• Low strength vs volume

Titatium • High strength vs weight• Low coefficient of thermal expansion

•Hard to machine• Poor fracture toughness if solution treated and aged

Beryllium •High stiffness vs density • Low ductility & fracture toughness• Low short transverse properties• Toxic

Composite • Tailored for high stiffness, high strength and extremely low coefficient of thermal expansion• Low density• Good in tension (eg pressurised tanks)

• Costly for low production volume, requires development programme• Strength depends on workmanship, ussually requires individual proof testing


• Good in tension (eg, pressurised tanks) proof testing• Laminated composites are not as strong in compression• Brittle, can be hard to attach


Material selection (II/II)Material selection (II/II)•

E/ E1/2/ E1/3/ f/Al 6061 T6 2700 68 276 24 2 9 1 5 98 6 23 6 186 97

Material Density(kg/m3)

Young's ModulusE(GPa)

Yield Strengthf(Mpa)

Thermal expansion(m/m K-1)

Fracture toughness

(MPa m)

Fatigue Strength

(MPa)

Selection criteria

Al 6061 T6 2700 68 276 24 2,9 1,5 98,6 23,6 186 97Al 7075 T6 2800 71 503 26 3,1 1,5 186,3 23,4 24 159Mg A2 31B 1700 45 220 26 3,9 2,1 129,4 26Mg ZK 60 A.T5 1700 45 234 26 3,9 2,1 137,6 26 124Ti 6Al 4V 4400 110 825 25 2 4 1 1 187 5 9 75 500Ti-6Al-4V 4400 110 825 25 2,4 1,1 187,5 9 75 500Be S 65 A 2000 304 207 151 8,7 3,4 103,5 11,5Be SR 200 E 2000 304 345Fe INVAR 150 275/415 1,66Fe AM 350 (SCT850) 7700 200 1034 26 1 84 0 8 134 3 11 9 40/60 550Fe AM 350 (SCT850) 7700 200 1034 26 1,84 0,8 134,3 11,9 40/60 550Fe 304L Ann 7800 193 170 25 1,8 0,7 21,8 17,2KEVLAR 49 0º 1380 76 1379 55 6,3 3,1 999,3 -4Aramid fibre 90º 1380 5,5 29,6 4 1,7 1,3 21,4 57Graphite epoxy 1620 282 586 174 10 4 4 361 7 -11 7/29 7Graphite epoxy 1620 282 586 174 10,4 4 361,7 -11,7/29,7

MIL-HDBK-5, “Metallic Materials and Elements for Aerospace Vehicle Structures”



Design philosophyDesign philosophy

LightnessLightness AffordabilityAffordability

ReliabilityReliability

Material strength:Material strength: Loads:Loads:

UncertaintiesUncertainties

•• RandomRandom•• Undetectable flawsUndetectable flaws•• Process variationsProcess variations

•• AcousticsAcoustics•• Engine vibrationEngine vibration•• Air turbulenceAir turbulence


UncertaintiesUncertainties



TermTerm DefinitionDefinition

Load Factor • A multiple of weight on Earth, representing the force of inertia that resists acceleration

Limit Load • The maximum load expected during the mission at a specified or selected statistical probability

Allowable Load or • The highest load or stress a structure or material can withstand Stress without failure, based on statistical probability

Factor of Safety • A factor applied to the limit load to obtain the design load for the purpose of decreasing the chance of failure

Design Load • Limit load multiplied by the yield or ultimate factor of safety. Thisvalue must be no greater then the corresponding allowable load

Design Stress • Predicted stress caused by the design load This value must not Design Stress • Predicted stress caused by the design load. This value must not exceed the corresponding allowable stress

Margin of Safety • A measure of reserve strength: Allowable load/Design load – 1 ≥0




OptionOption Design Factors of Design Factors of SafetySafety

YieldYield UltimateUltimate

Ultimate test of dedicated qualification article (1.25xlimit) 1.0 1.25

f f ll fl h ( l )Proof test of all flight structures (1.1xlimit) 1.1 1.25

Proof test of one flight unit of a fleet (1.25xlimit) 1.25 1.4

No structural test 1.6 2.0



Preliminary SizingPreliminary Sizing

StrengthStrengthSt e gtSt e gt

StiffnessStiffness WeightWeight

Preliminary Preliminary SizingSizinggg


Structure Examples Structure Examples –– Generic CubesatsGeneric Cubesats


Th lTh lThermalThermalSubsystemSubsystemSubsystemSubsystem


Thermal SubsystemThermal Subsystem

Objectives• The role of the Thermal Control Subsystem (TCS) is to maintain all• The role of the Thermal Control Subsystem (TCS) is to maintain all

spacecraft and payload components and subsystems within theirrequired temperature limits for each mission phase:• Operational limits• Operational limits• Survival limits

• TCS is also used to ensure that temperature gradient requirements aremet in order to avoid any structural deformation and avoid pointingerrors.

The biggest problem is getting rid of excess heat



Design drivers (I/II)Design drivers (I/II)• The environment in which the spacecraft

Heat fluxes and not temperatures

• The environment in which the spacecrafthas to operate

• Th t t l t f h t di i t d are the subject of control• The total amount of heat dissipated onboard the spacecraft

• The distribution of the thermaldissipation inside the spacecraft

• Th t t i t f th• The temperature requirements of thevarious equipment items

• The config ration of the spacecraft and• The configuration of the spacecraft, andits reliability/verification requirements



Design drivers (II/II)Design drivers (II/II)



Thermal environment for a LEO SpacecraftThermal environment for a LEO SpacecraftR di i S

Direct solar radiation

Radiation to Space

(1358±5 W/m2)

Reflected sunlight

(30±5 W/m2)

Radiation from Earth

(237±21 W/m2)

Power generated in the S/C



Typical thermal requirements for spacecraft componentsTypical thermal requirements for spacecraft components

ComponentTypical temperature ranges [ºC]

Operational Survival

Batteries 0 to 15 ‐10 to 25

Power box baseplates ‐10 to 50 ‐20 to 60

R ti h l t t Reaction wheels ‐10 to 40 ‐20 to 50

Gyros/IMUs 0 to 40 ‐10 to 50

Star trackers 0 to 30 ‐10 to 40Star trackers 0 to 30 10 to 40

Hydrazine tanks and lines 15 to 40 5 to 50

Antenna Gimbals ‐40 to 80 ‐50 to 90

Antennas ‐100 to 100 ‐120 to 120

Solar panels ‐150 to 110 ‐200 to 130



Typical installed powers for various kinds of spacecraft Typical installed powers for various kinds of spacecraft



Radiation heat transfer: influence of surface propertyRadiation heat transfer: influence of surface property

Incident radiationRefelected radiation

Absorved radiation

Transmitted radiation

1

The steady state temperatures withdifferent surface coatings:

41

2··4····4·

RPST



Radiation equilibrium temperatureRadiation equilibrium temperature• The spacecraft skin temperature or radiation equilibrium temperature can be

calculated from the basic heat balance equation.• However:

• Orbital data (altitude and orientation relative to external heat sources) andOrbital data (altitude and orientation relative to external heat sources) and• Detailed spacecraft configuration need to be known

Heat In + Internal Heat = Heat Out Radiation equilibrium temperature

• Varying orbit conditions and power dissipation together with complex• Varying orbit conditions and power dissipation together with complexconfigurations require thermal analyses codes:• ESATAN• ESARAD• SINDA• RadCAD



Thermal control system design process (I/II)Thermal control system design process (I/II)

Step Inputs Outputs Key issues

1. Identify thermal C t th l ‐ System thermal requirements ‐ Identify payload thermal requirementsyrequirements and constrains

Component thermal requirements

y q‐ Specialised requirements for specific equipment

y p y q‐ Identify major elements that may present thermal challenges (see step 3)

2. Determine ‐ Orbit/attitude history ‐Total energy input to the S/C‐Max & min distances to Sun‐Max & min distance to Earth or other central thermal

environment‐ S/C size and shape‐ Internal heat sources

Total energy input to the S/C‐ Profile of energy input vs. time

Max & min distance to Earth or other central body‐ Chemical or nuclear internal heat sources

3 Identify thermal ‐ Thermal requirements List of specific thermal problem

Identify major elements that:‐ Generate large amounts of heat‐ Need cryo operating temperatures3. Identify thermal

challenges or problem areas

‐ Heat sources‐ Equipment placement and attitude history

List of specific thermal problem areas or problem times or events (hot, cold or stability)

Need cryo operating temperatures‐ Have boiling or freezing problems‐ Require a narrow temperature rangeIdentify extraordinary thermal events or actions

‐ Prefer passive over active means

4. Identify applicable thermal control techniques

‐ Thermal requirements and energy profile from above‐ Additional constrains

‐ Preliminary list of thermal control mechanisms for mission duration and principal S/C components, areas or times

Prefer passive over active means‐ Component placement often key‐ Pay particular attention to problem areas or severe thermal constrains‐Watch for mission critical issues (freezing propellant or hinge; fluid boiling or potential

l i )


explosions)


Thermal control system design process (II/II)Thermal control system design process (II/II)

Step Inputs Outputs Key issues

‐List of thermal T k i

5. Determine radiator and heater

List of thermal environments & events‐ Thermal control approach

‐ Radiator sizes and temperatures to manage hot case with marginH f ld

Take into account:‐Degradation of thermal surfaces over mission life‐ Longest eclipse furthest from a

l b dheater requirements

approach‐ Components temperature requirements

‐ Heater power for cold case thermal control

warm central body‐‐ Extraordinary thermal events or circumstances

6 E ti t T i ll % t % f d 6. Estimate TCS mass and power

‐ List of TCS methods and components

‐ TCS mass‐ TCS power

‐ Typically 2% to 10 % of dry mass‐May impact mass & power of other subsystems

7 Document Thermal robustness can be key to 7. Document and iterate

ysystem design flexibility and reducing operations cost





Thermal design development process IThermal design development process ILIST DESIGN REQUIREMENTSLIST DESIGN REQUIREMENTSPayload requirements usually dominates the design

BASELINE DESIGNMass

ESTABLISH HEAT INPUT SOURCESSun, Earth, electronics…

COMPUTE WORST-CASE HOT AND

SizePower

NO YES

COLD TEMPERATURES

SELECT THERMAL CONTROL TECHNIQUETECHNIQUE

DETERMINE HEATER POWER REQUIRED

MEET REQS?

COMPARE WORST-CASE VALUES WITH DESIGN TEMPERATURE LIMITS



Design requirementsDesign requirements• Temperature limits and reliability requirements for each component• Equipment power dissipations and operating modes• Range of mission orbit parameters• Operationl satellite attitudes• Attitudes during stressed or failure modesg• Launch phase configurations and attitudes• Ground cooling needs• Autonomy requirements• Thermal-distrotion budgetsThermal distrotion budgets• Launch-system interfaces• Contamination control• Special thermal control requirements for components such as batteries, crystal oscillators and sensors• Interfaces with other subsystems such as:Interfaces with other subsystems, such as:

• Payloads• Propulsion• Attitude conrtol• Electrical powerElectrical power• Structures• Telemetry, tracking and command• Computer and data handling



Thermal design development processThermal design development process

Input- S/C Configuration- Orbit data- Preliminary equipment temperature limits

Calculate:- TCS Wieight

- Preliminary equipment power profiles

Assume passive thermal design

TCS Wieight- TCS Power

Is TCS MP?YES

Establish:- Tompartment

I lt i i tYES

Calculate environmental heat loads on S/C

Calculate:T Are Ts

Is TCS MP?

- Insultaion requirements- Coatings

Establish:- Heater power

Heat ppe requirements

YES

NO

Assume modified passive TCS

- Tequipment

- Tompartment

Are Ts acceptable?

Is TCS MP?Is TCS SA?NO

- Heat ppe requirements

Establish:- Louver characteristics- Heat ppe requirements

YES

NO

Assume semi-active TCS

A ti TCS

Is TCS SA?Is TCS A?

NO

NO

Establish:- Radiator size- Transport loop characteristics

YES

NO


Assume active TCS

Assume spacial TCS

Is TCS A?NO

YESEstablish special requirementsNO

P l iP l iPropulsionPropulsionSubsystemSubsystemSubsystemSubsystem


Propulsion SubsystemPropulsion Subsystem

Basic definitionsBasic definitions• Propulsion system: system to provide thrust autonomously. It comprises every component

necessary for the fulfilment of the mission e g :necessary for the fulfilment of the mission, e.g.:

– Thrusters– Propellants

– Pressurisation subsystem– Tanks

– Valves– Filters– Pyrotechnic devices

– Electrical components such as power sources incase of electrical propulsion

– Sensors

• Propellant: material or materials that constitute a mass which, often modified from its originalstate, is ejected at high speed from a rocket engine to produce thrust.– In cold gas engines, the gas is accelerated due to the difference between storage and ambient

pressurepressure.– In chemical rocket engines, either a combustion reaction between two kind of propellants, fuel

and oxidiser, o a decomposition reaction of amonopropellant provides the energy to acceleratethemass.

– In electric engines, either an electromagnetic or an electrostatic field accelerates the mass,g gwhich in some cases has been heated to high temperatures or electric heating providesadditional energy to accelerate the mass (the latter in the case of power augmented thrustersand resistojets).

– Combinations of the above are possible.



Spacecraft pSpacecraft propulsionropulsion functionsfunctionsTask Descriptionp

Mission design (Translational velocity change)

Orbit changes Convert one orbit to another

Plane changes

Orbit trim Remove launch vehicle errors

Stationkeeping Maintain constellation position

Repositioning Change constellation position

A i d l ( l l h )Attitude control (Rotational velocity change)

Thrust vector control Remove error vectors

Attitude control Maintain an attitudeAttitude control Maintain an attitude

Attitude changes Change attitudes

Reaction wheel unloading Remove stored momentum


Manoeuvring Repositioning the spacecraft axes


Different types of space propulsion systemsDifferent types of space propulsion systems

Chemical rocketsLiquid

Cold gas

Monopropellant

Bipropellant

Dual mode

N

Solid

Hybrid

OPU

LSIO

ElectrothermalResistojets

Arcjets

PAC

E PR

O

Non-chemical rockets Electrostatic

j

Ion Thrusters

Hall Effect Thrusters

SP

Non-rocket propulsion

Electromagnetic Pulsed Plasma Thrusters


p p


Main components of a Space Propulsion SubsystemMain components of a Space Propulsion Subsystem

Tank(s)Control

PropulsionPower

ElectronicsPowerUnit

Propellant Th ( ) ThrustPropellant

Di ib iPropellant Thruster(s) ThrustDistributionElements



Propulsion subsystem selection and sizing processPropulsion subsystem selection and sizing processfEstablishment of

propulsion system requirements

Definition of propulsion system options

Iterations with other subsystems

system options

Trade-off among the candidate propulsion

systems

Functions and tasks

Propulsive requirements

I l t ti i t Propulsion options propellantsImplementation requirements Propulsion options propellants

Propulsion system configuration

Propulsion system components

Preliminary system performance

SWOT analysis

Compliance matrix

Selection of the most suitable system

y y p

Mass budget

Propulsion system definition and design

Power budget

System size

Configuration

t


etc.


Propulsion subsystem selection and sizing processPropulsion subsystem selection and sizing process1. List applicable spacecraft propulsion functions, e.g., orbit insertion, orbit

d l ll d d bmaintainance, attitude control, controlled de‐orbit...2. Determine V budget and thrust level constrains for orbit insertion and

maintenancel l f d l h l l f l h3. Determine total impulse for attitude control, thrust levels for control authority,

duty cycles (% on/off, total number of cycles) and mission life requirements4. Define propulsion system options:

C bi d t l i t f bit d ttit d t l• Combined or separate propulsion systems for orbit and attitude control• High vs. low thrust• Cold gas vs. liquid vs. solid vs. electric propulsion technology

5 Estimate key parameters for each option5. Estimate key parameters for each option• Effective Isp for orbit and attitude control• Thrust and MIB• Propellant mass• Propellant and pressurant volume• Configure the subsystem and create equipment list

6. Estimate total mass and power for each option


7. Establish baseline propulsion subsystem8. Document results and iterate as required


General subsystem requirementsGeneral subsystem requirements• Thrust levels torque levels and linear impulse levels are required as a• Thrust levels, torque levels, and linear impulse levels are required as a

function of mission phase• Total impulse required by all maneuvers• Layout envelopes or constraints, centre of mass profile throughout the

mission• Allowable weight, mass properties, power and TT&C channel budgets as a

function of mission phase• Environments which will be imposed on the subsystem components• Reliability and redundancy requirementsy y q• Cost and schedule constraints• Subsystem safety – proof and burst to operating pressure

A ibili f li i h d l di ll l h i• Accessibility – for aligning thrusters and loading propellants at launch site• Cleanliness – both internal and external• Life



General subsystem interfacesGeneral subsystem interfaces• Propellant Tanks are located for s/c mass properties reasons, not RCS

convenience• Thrusters must be located to avoid or minimize plume impingement

effects (forces, thermal, contamination) on solar arrays, antennas andother appendages. Additionally they must be located so that neededthrusters are not covered by these appendages when stowed.

• Locate thrusters to point through s/c centre of mass or to have equalmoment arms, and to have alignment capability.

• Thermal interfaces are generally quite complicated, thrust chambersreach high temperatures (1500°C) and must be isolated from their valvesand s/c surfaces. Also since thrusters protrude through s/c exteriorsurfaces, they form heat leaks which must be insulated. Also, manypropellants have a more narrow acceptable temperature range than most/ h d d i i l tis/c hardware and require special precautions.

• The power subsystem may be called upon to provide high power andpulsed loads. Transient suppression is required to protect the s/c againstEMI hi h ld ff t it h t t d l i i it


EMI which could affect switch states and logic circuitry.

Propulsion SubsystemPropulsion SubsystemPropulsion subsystem interfaces with other subsystemsPropulsion subsystem interfaces with other subsystems

Structure

Inserts

GSE

Tanks filling

Tanks support estructure

Vibration levels

…

Testing

…

Thermal control

Radiation levels

DHS and TM/TC

Health monitoring

Conduction

Thruster and line thermal control

…

Failure detection

Valve drivers

…

AOCS

Thrust levels

Impulse levels

Power supply

Heaters

Sensors


Impulse levels

Definition of firing modes

…

Sensors

Valves

…

Propulsion SubsystemPropulsion SubsystemPropulsion system design processPropulsion system design process

List aplicablespacacraft Functions

Orbit InsertionOrbit mainteneceAttitude Control

Determine TotalImpulse for Attitude

control.

Determine delta Vand Thrust level

constraints for orbitinsertion andMaintenace

Determine Thrustlevels for control

authority, duty cyclesand mission life

requirements

Determine propulsionsubsystem options

bi dDeterime level of

d d d

Estimate Key parameters foreach option

Eff i I d Th f Estimate total mass- combined or separate- Low or High thrust- Liquid, solid, electricor plasma

redundancy andoverall configuration

for each option

- Effective Isp and Thrust fororbit and attitude control- Propellant mass andPressurant Volume

Estimate total massand power for each

option

Qualify hardware atcomponent andsubsytem level

Finalize design andprocure/manufacture

equipment

Inetgrate intospacecraft system

level AssemblyIntegration and test

Are requirementsmet? Yes

No

subsytem levelequipment Integration and testProgram

NoNo


Propulsion SubsystemPropulsion SubsystemPrincipal options for spacecraft propulsion systems (I/II)Principal options for spacecraft propulsion systems (I/II)

Propulsion Technology

Orbit Insertion Orbit Maintenance and Manoeuvering

Attitude Control

Typical Steady State Isp [s]Perigee ApogeePerigee Apogee

Cold Gas X X 30 – 70

S lid X X 8 Solid X X 280 – 300

Liquid

Monopropellant X X 220 –240

Biproprellant X X X X 305 –310

Hybrid X X X 250 – 340

Electric X X 300 ‐ 3000



Principal options for spacecraft propulsion systems (II/II)Principal options for spacecraft propulsion systems (II/II)Effective exhaust Thrust Maximum Delta v Propulsion methods in current use velocity

[km/s]

Thrust(N) Firing Duration Maximum Delta‐v

[km/s]

Solid rocket 1 ‐ 4 103 ‐ 107 minutes ~ 7

Hybrid rocket 1.5 ‐ 4.2 <0.1 ‐ 107 minutes > 3

Monopropellant rocket 1 ‐ 3 0.1 ‐ 100 milliseconds ‐minutes ~ 3

Bipropellant rocket 1 ‐ 4.7 0.1 ‐ 107 minutes ~ 9

D l d k t 7 i t Dual mode rocket 2.5 ‐ 5.3 0.1 ‐ 107 minutes ~ 9

Resitojet 2 ‐ 6 10‐2 ‐ 10 minutes

Arcjet 4 ‐ 16 10‐2 ‐ 10 minutes

Hall Effect Thruster (HET) 8 ‐ 50 10‐3 ‐ 10 months/years > 100

Electrostatic Ion Thruster 15 ‐ 80 10‐3 ‐ 10 months/years > 100

Field Emission Electric Propulsion (FEEP) 100 ‐ 130 10‐6 ‐ 10‐3 months/years(FEEP) 3 y

Pulsed Plasma Thruster (PPT) ~ 20 ~ 0.1 ~ 2,000 ‐ ~ 10,000 hours

Pulse Inductive Thruster (PIT) 50 20 months


REFERENCESREFERENCES

1. Fortescue, “Spacecraft Systems Engineering”,2 Larson and Wertz “Space Mission Analysis and Design”2. Larson and Wertz, Space Mission Analysis and Design3. ECSS Standards4. Cubesat Standard5. Sarafin, “Spacecraft Structures and Mechanisms”6. Bruhn, “Analysis and Design of flight Vehicle Structures”7. Gilmore D.G, “Spacecraft Thermal Control Handbook Volume 1: Fundamental

Technologies8 P Sanz-Aranguez J S llorente J J Piñeiro “Vehículos espaciales II”8. P. Sanz Aranguez, J.S. llorente, J.J. Piñeiro, Vehículos espaciales II 9. Sutton and Biblarz, “Rocket Propulsion Elements”10.Brown, “Spacecraft Propulsion”11.http://ssdl.stanford.edu/cubesat12.http://www.cubesat.org (= http://www.cubesat.net)13.http://www.amsat.org14.http://www.ssel.montana.edu/merope15.http://ssdl.stanford.edu/cubesat15.http://ssdl.stanford.edu/cubesat16.http://www.cubesat.org (= http://www.cubesat.net)17.http://www.amsat.org18.http://www.ssel.montana.edu/merope


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