Satellite Engineering PROBA Familly - LTAS-SDRG · Title: PROBA-3 Kick-off Meeting Phase B2.2 Author: jtallineau Created Date: 12/9/2012 8:13:03 PM

© Copyright QinetiQ Limited 2012 QinetiQ Proprietary

Julien Tallineau

A presentation to: ULg

12/12/2012

1

Satellite Engineering PROBA Familly

QinetiQ Space nv © Copyright QinetiQ Limited 2010

0 TABLE OF CONTENT

1. Introduction

2. PROBA Approach

3. PROBA Satellites

4. PROBA Case Study

5. Conclusions

2


1 INTRODUCTION

3


1 INTRODUCTION

4

Mandatory

Usefull


1 INTRODUCTION

5


1 INTRODUCTION

6


1 INTRODUCTION

Satellite Engineering

1. Customer Need (Scientist)

2. Creation of System Requirements

3. Phase 0 (CDF Study)

4. Phase A (Feasibility Study)

5. Phase B (Preliminary Design)

6. Phase C (Final Design)

7. Phase D (Manufacturing / Testing)

8. Phase E (Launch / Commissionning/Operations)

9. Phase F (De-orbiting)

7


1 INTRODUCTION


1. Customer Need (Scientist)

2. Creation of System Requirements

3. Phase 0 (CDF Study)

4. Phase A (Feasibility Study)

8


1 INTRODUCTION


5. Phase B (Preliminary Design)

• Detailed System Analysis

• Preliminary Subsystem Analysis

• Trade-offs

6. Phase C (Final Design)

• Detailed Subsystem Analysis

• Procurement

• Qualification testing

9


1 INTRODUCTION


7. Phase D

• Manufacturing

• Acceptance Testing

• Requirement Verification

• Shipment to Launch site

10


1 INTRODUCTION


8. Phase E

• Launch

• Commissionning

• Operations

9. Phase F (De-orbiting/End of Life)

• None in this case

11


2 PROBA APPROACH

12


13

2 PROBA APPROACH

PROBA Philosophy

• PRoject for On-Board

Autonomy (PROBA)

• Missions

− In Orbit Demonstration

− Earth Observation

• Developed under ESA TEC

department

PROBA-1

PROBA-2

PROBA-V

PROBA-ALTIUS


2 PROBA APPROACH

14

LightSat Approach

• In-Orbit Demonstration (IOD) aiming at

- Demonstrating new techniques that could lead to new space system

- Reduced cost

• Drastic reduction of the number of requirements

• Keep the system as simple as possible with only limited inter-dependencies

• Results rather than paper oriented reviews

• Accept higher level of risk


2 PROBA APPROACH

Satellite Engineering Tools

1. Requirement Management Tool (DOORS)

2. Mission Analysis Tool (MATLAB)

• Orbit propagator (J2, Drag, Third Body, SRP)

• Attitude Control (Magnetic, Nadir, Sun, Target)

• Incoming fluxes

• Ground Stations Visibility

• In orbit Maintainance (To be developed)

• In orbit Manoeuvering (To be developed)

• Formation Flying/Rendez vous (To be developed)

15


2 PROBA APPROACH

Satellite Engineering Tools

1. Thermal Analysis Tool (ESATAN-TMS)

2. Configuration Tool (IRON-CAD/PRO-E)

• IRON-CAD allows for fast iteration

• PRO-E is similar to CATIA

3. Structural Analysis Tool (NASTRAN)

16


3 PROBA SATELLITES

• PROBA – 1

• PROBA – 2

• PROBA – V

17


3 PROBA - 1

PROBA – 1 Mission

18


3 PROBA - 1

Mission

1. In-orbit demonstration and evaluation of new hardware/software spacecraft

technologies and onboard operational autonomy

2. Obrital Parameter

3. Injection Parameter

19


3 PROBA - 1

Mission

1. RAAN selected for 10:00 Local Time Descending Node.

20


3 PROBA - 1

Mission

1. Launcher is PSLV

• Nominal separation rate of 2° per sec.

• Worst Case separation rate of 8° per sec.

• Inclination accuracy of 0.2°

• Altitude accuracy of 35km

2. Radiations of 9 krad for 3 years mission

• Solar maximum is considered

• 3.5mm of Aluminium considered

21


3 PROBA - 1

Mission

1. Ground segment visibility (REDU)

• 30% of all orbits have contact with the ground station (duration > 0 minutes)

• 21.23 more than 6 minutes

• 10.88% more than 8 minutes

• The longest time between two passes

over the ground station is 11hr 44 min

22


3 PROBA - 1

Mission

1. Scenario

• Launch and Early Operation Phase (LEOP)

− Boot after separation

− De-tumbling

− First Ground Contact + AOCS/GPS switched ON

• Commissioning (two months)

− Low Autonomy

• Nominal Operations

− High Autonomy

23


3 PROBA - 1

Satellite Design - Configuration

1. Single H Structure

2. High Unit Density

3. Body Mounted Solar Panel

4. Cut out in panels:

• Star-Trackers

• Instruments

24


3 PROBA - 1

Satellite Design - Structure

1. Honeycomb panels

• Aluminium core

• Aluminium edge

• Alumiunium facesheet (Primary structure)

• CFRP facesheet (Secondary structure)

25


3 PROBA - 1

Satellite Design - Mechanics

1. Spacecraft Mass = 95 kg

2. Balance Mass of 2.5kg for

CoG Location Requirement

26


3 PROBA - 1

Satellite Design - Power

1. Power budget approach could be:

• Rely on Solar Array (SA) in Sun and on Battery in Eclipse

• Rely on both Battery and SA when available.

2. Power budget is positive, independently of the scenario.

27


3 PROBA - 1

Satellite Design – RF COM

1. Downlink (S-Band)

• Flux margin of 7dB (@max rate = 1Mbps)

• Flux margin of - 2dB (@min rate = 250kbps)

• Telemetry Recovery margin of 8dB

2. Uplink (S-Band)

• Carrier Recovery margin of 26dB

• Telecommand Recovery margin of 17dB

28


3 PROBA - 1

Satellite Design – Thermal

1. Passive thermal control

• Thermal Blankets + MLI

• Black paint (internal panel+electronic box)

29


3 PROBA - 1

Satellite Design – Data & memory

1. One Image sequence can be stored before next downlink

• Mass Memory size = 1Gbit

• Image sequence size = 5 image x 19 spectral lines x 742 spatial lines x 9 kbit/line

= 665 Mbit.

2. Data downlinked = 1140Mbit every 12 hours

• Max data rate

• Time ≈ 1300sec (Total Ground Visibility)

30


3 PROBA - 1

23 000 km

27 000 km

25 000 km

23 000 km23 000 km

27 000 km27 000 km

25 000 km25 000 km

Position determination

GPS RX

Position determination

Magnetometer (2)

Measure continuously the

Earth Magnetic field and

determine itselve where the North is.

N

Z

N

Z

N

Z

•Takes pictures of stars

• Compare it with its internat catalog.

• Compute the satellite orientation/position

Orientation/Position

Star-Tracker (3)

Orientation/Manoeuvre

Magnetotorquer (4)

N

Z

.

N

Z

N

ZZ

.

N

Z

N

Z

• Magnetic Coil

• Align itself to Magnetic lines

Orientation/Manoeuvre

Rection Wheel (4)

Accelerates or decelerates while momentum conservation implies the PROBA to rotate the other way.


3 PROBA - 1

Satellite Design – AOCS

1. Guidance

= determination of the desired path of

travel from the satellite's current

location to a designated target

2. Navigation

= determination of the satellite’s

location, velocity and attitude

3. Control

= manipulation of the forces needed to

track guidance commands while

maintaining satellite stability

32

AOCS Software - Navigation - Guidance - Control

Reaction Wheels (4)

Magneto- meters (2)

Magneto- torquers (4)

GPS Receiver

Star Tracker (2)

Head

Head


3 PROBA - 1

Satellite Design – AOCS ( Pointing Modes)

33

• Inertial pointing

• Sun Pointing

• Motion compensation

• Multiple scans • On-track

• Off-track

• On-track

• Off-track

Utilisation:

• Astronomy

• Solar Physics

Utilisation:

• Earth Observation

• Telecommunications

• Space Weather

Utilisation:

• High Resolution

Imaging

• Elevation Modeling

• Disaster Monitoring

Inertial Pointing Earth Pointing Point and Stare Scanning Modes


3 PROBA - 1


1. Specifications

2. AOCS SW Modes

• Quiscient

• Magnetic

• Celestial

• Terrestrial

34

Errors at 95% confidence level

Absolute Pointing Error (APE) 100 arcsec

Relative Pointing Error (RPE) 5 arcsec over 10 s

Absolute Measurement Error (AME) 10 arcsec

Quiscient Mode

- No AOCS

- Most units are OFF


3 PROBA - 1


1. Specifications

2. AOCS SW Modes

• Quiscient

• Magnetic

• Celestial

• Terrestrial

35





Magnetic Mode

- Sensor: magnetometer

- Actuator: magnetotorquers + 1 RW stby (momentum bias)

- Guidance: No

- Navigation: No

- Control: Bdot Algorithms (Kinetic Energy Dumping)


3 PROBA - 1


1. Specifications

2. AOCS SW Modes

• Quiscient

• Magnetic

• Celestial

• Terrestrial

36





Celestial Mode

- Sensor: Star-tracker + Magnetic mode sensors

- Actuator: Reaction Wheels + Magnetic mode actuatord

- Guidance: Sun pointing, Inertial pointing

- Navigation: Attitude and Orbit evaluator (Kalman Filter)

- Control: State feedback, PID, angular momentum control


3 PROBA - 1


1. Specifications

2. AOCS SW Modes

• Quiscient

• Magnetic

• Celestial

• Terrestrial

37





Terrestrial Mode

- Sensor: GPS + celestial mode sensors

- Actuator: celestial mode actuators

- Guidance: Nadir, Fixed-target pointing, Imaging scans

- Navigation: Same as Celestial mode

- Control: Same as Celestial mode


3 PROBA - 1

Launched on the 22/10/2001 from India

38


3 PROBA - 1

LEOP Activities

1. Rotating Energy is dissipitated progressively using the magneto-torquers.

39

0,0E+00

2,0E-03

4,0E-03

6,0E-03

8,0E-03

1,0E-02

1,2E-02

1,4E-02

0,00 1,00 2,00 3,00 4,00 5,00 6,00

Sq

ua

re a

ng

ula

r ra

te (

rad

/se

c)

2

Five hours

to detumble!


3 PROBA - 1

Normal Operations Activities

40

Ile de Ré,

France


3 PROBA - 1


41

Etna eruption,

Sicily, 20.10.2002


3 PROBA - 1


42

North Sentinel Island

India, 29.05.06


3 PROBA - 1


43

Palm Island

Dubai


3 PROBA - 2

PROBA – 2 Mission

44


3 PROBA - 2

Mission

1. In orbit Demonstration, PROBA-2 aimed at technological innovation.

Altogether, 17 new technological developments and four scientific

experiments are being flown on Proba-2.

2. Orbital Parameter

• RAAN selected for

6:00 AM Local Time

Ascending Node

45


3 PROBA - 2

Mission

1. RAAN selected for 6:00 AM Local Time Ascending Node

46


3 PROBA - 2

Mission

1. Launcher is Rockot

• Worst Case separation rate of 8° per sec.

• Inclination accuracy of 0.05°

• Altitude accuracy of 12km

• RAAN accuracy of 3.75° (≈15min LT)

2. Injected via the Breeze upper stage

47


3 PROBA - 2

Mission

48


3 PROBA - 2

Mission

1. Ground segment visibility

• REDU

• KIROUNA

49


3 PROBA - 2

Mission

1. Scenario

• LEOP

• Commissioning (three months)

• Nominal Operations

2. Spacecraft Modes

• Separation

• Safe

• Observation

• Stand-by

50


3 PROBA - 2

Mission

51


3 PROBA - 2

Satellite Design - Configuration

1. Single H Structure

2. Sun Shield

• Standard STR

• Bepi-Colombo STR

3. High Unit Density

4. Deployable Solar Panel (x2)

52


3 PROBA - 1

Satellite Design - Mechanics

1. Spacecraft Mass = 122.5 kg

2. CoG Choice

• Folded Configuration (LV requirement)

• Deployed Configuration (GNC requirement)

53

NB: LV I/F Ring Mass included


3 PROBA - 2


1. Power budget is positive, independently of the mode

• Observation (w/o TX

• Observation with TX

• Safe mode with TX (worst Beta-angle).

54

Definition

Beta Angle = angle between the Sun-Earth vector and the orbital plane.


3 PROBA - 2


55


3 PROBA - 2


56


3 PROBA - 2

Satellite Design – Power

1. Battery DoD (Ah)

in function of time

2. Non Regulated bus 28V

57


3 PROBA - 2


1. Power budget while de-tumbling !

2. Trade-off between:

• Performance (GNC)

• Time (LEOP schedule)

• Battery Discharge (Higher DoD)

58


3 PROBA - 2

Satellite Design – RF COM

59


3 PROBA - 2

Satellite Design – Thermal

1. Passive thermal control

• Thermal Blankets + MLI + Chotherm

• Black paint (internal panel+electronic box)

60


3 PROBA - 2

Satellite Design – Data & Processing

1. Processing budget shows how busy is the processor with all the units +

instruments.

• On Software Verification Facility

• On S/C during System Validation Test 5

61


3 PROBA - 2


1. Low power resistojet (Xenon)

• 15W for heater (x2)

• 50s Isp (min)

• 20mN Thrust

• Total ∆V = 2m/s

62


3 PROBA - 2


1. Sensors

• 2 Star-tracker

• 2 GPS RX

• 2 Magnetor-Meter

2. Actuators

• 4 Reaction Wheels

• 3 dual-coil magneto-Torquer

63


4 CASE STUDY: PROBA-3

• The Mission

• Phase A

• Phase B

64


3 PROBA - 3

Mission

1. The PROBA-3 mission will provide an opportunity to validate and develop

• The Metrology and Actuation techniques / technologies

• The Guidance strategies and Navigation and Control algorithms necessary for

formation flying.

2. One mission, two spacecrafts:

• Coronagraph S/C (CSC)

• Occulter SC (OSC)

65


3 PROBA - 3

Mission

1. High Elliptical Orbit (HEO) with 20 hours period

• Low perturbation in Apogee

• Low AoP drift (fixed above REDU)

• Limited Eclipse duration

• Radition Issue

• Stringent Constraint on RF Link

• Delta-V issue.

66

Parameter Value

Orbit type HEO

Perigee altitude 600km

Apogee altitude 60530km

Inclination 59°

Eccentricity 0.8062

AoP 188°

RAAN 173°

Epoch Jan 2017


3 PROBA - 3

Possible Launchers

1. PSLV Launcher

• Low Cost

• Heritage from PROBA-1

• Reduced Volume

2. Falcon 9 Launcher

• High Cost

• Higher Performance

• Large Volume

• No Heritage

67

PSLV

Falcon 9


3 PROBA - 3

Which one of the two shall you design your Spacecraft for ?

68


3 PROBA - 3

Which one of the two shall you design your Spacecraft for ?

ANSWER = The Two !

But it is up to ESA + Delegates to decide the mission budget

69


3 PROBA - 3

What would you put in your spacecraft ?

70


3 PROBA - 3

What would you put in your spacecraft ?

71

Subsystem Design

Structure Aluminium/CFRP/Invar/Titanium ?

Thermal Passive/Active ?

Mechanism Body Mounted/Deployable SA?

Power Large/Small SA ?

Large/Small Battery?

GNC Sensor?

Actuator?

RF High/Low Gain COM Antenna?

High/Low Gain FF Antenna?

Payload What do you need to perform the mission?


3 PROBA - 3

Welcome in Phase A !

72


3 PROBA – 3 Phase A

Outcome of Phase A is our STARTING POINT

1. Configuration

2. Mass

3. Power

4. Avionics

5. Thermal

6. Propulsion

7. Link & Data

8. Payload

73



Coronagraph Satellite Overview

1. Twelve subsystems

2. Five entities

74




1. From Top

2. From Bottom

75




1. From Back Left

2. From Back Right

76




1. From Front

2. From Inside

77




1. Mass Budget

• “Total CSC dry mass including I/F ring with L/V shall be less than 360kg (with

margins)”.

• Current estimate: 376.44kg

2. Power Budget

• “ Total CSC maximum power consumption shall be less than 294W (with margins)”

• Current estimate: 353.15W

78




1. Avionics

• Centrilized around Advance Data & Power Management System (ADPMS)

• Support of Interface Electronics

See block diagram

79




1. Thermal

• Solar array temperature range (-169°C to +79°C).

• Battery operational temperature range:

− -10°C minimum operational temperature

− +40°C maximum operational temperature

• Star Tracker detector maximum operational temperature

− Current prediction: 40°C

− Required: 15°C

80




1. Propulsion System (HPGP)

• 2x4 thrusters to raise orbit

• Constrains

− Pre-warming needs 64W (=2x4x8W) during 30 min

− Propellant needs to be kept

above 10°C (Always)

• Performance

− Isp = 202 EoL

− Thrust = 1N

81




1. Propulsion System (Cold Gas)

• Siwteen thruster for GNC and FF

• Performance

82




1. GNC Sensors

• 6 Sun Acquisition Sensors

• 2x3 Gyros

• 2x3 Acceleromter

• 3 Star Trackers Head + 2x1 Electronics

• 2x1 GPS

2. GNC Actuators

• 3+1 Reaction Wheels

• Cold Gas thrusters

83

3. GNC FF

• Omni-directional RF

Sensor (FFRF)

• Coarse Lateral Sensor

• Fine Formation Sensor




1. Link Budget

• “The CSC shall be able to downlink data at a symbol rates of 256ksps and 2Msps

using REDU-3 Ground Station”

• “The CSC shall be able to receive TC data at a symbol rates of 64ksps

using REDU-3 Ground Station”

• Current estimation:

− Downlink not closed for 2Msps (@apogee)

− Uplink not closed for 64ksps (@apogee)

84




1. Data Budget

• “The CSC shall be able to downlink 8.5Gbit of data per day”

• Current estimate

− Min rate (256ksps)

− Max rate (2Msps)

85

0,00%

10,00%

20,00%

30,00%

40,00%

50,00%

60,00%

70,00%

80,00%

90,00%

100,00%

0 10 20 30 40 50 60 70 80D

aily C

overa

ge P

erc

enta

ge

COR Altitude (1000 km)

Coverage Percentage Needed for Downlink of CORONAGRAPH data (8.5GBit/Daily)

Redu Coverage

Redu+Santiago+Perth Coverage

Needed Coverage Percentage Low Rate DL (8.5GBit Data/24h)

Needed Coverage Percentage High Rate DL (8.5GBit Data/24h)



SUMMARY

1. Satellite is too heavy

2. Satellite is too power consuming

3. Satellite is too hot or too cold

4. Satellite is too far to close its downlink and uplink

5. Satellite is too far to downlink its science data

6. Satellite is too un-protected wrt radiations (20krad under 2mm Al)

86


3 PROBA - 3

What can we do?

87



ANSWER

1. Mass & Power reduction exercise

2. Perform a detailed thermal analysis

• Active control

• Electronics Location

• Radiator size

3. Review Operation concept for U/D link

4. Increase Satellite Shielding

88

+ MINIMIZE THE COST !


3 PROBA - 3

Welcome in Phase B !

89


3 PROBA – 3 Phase B

Mass reduction

1. Who are the biggest contributors ?

• Structure

• Formation Flying

• Propulsion System

• Payload

90


3 PROBA - 3

What can we do?

91



Mass reduction

1. Structure

• Shield the spacecraft with outer panels !

• Remove materials in the Bottom Board !

• Remove two panels while re-arranging structure

• Shorten the satellite

• Consider LV I/F ring as part of system mass margin

2. GNC & Propulsion

• Remove the Accelerometer

• Remove Cold Gas Propulsion System (Transfer to OSC)

• Include 2x4 additional HPGP thruster for 6DoF

92

3. Payload

• Replace the FFRF

• Shorten Payload

4. Other

• Direct Injection !



Mass reduction results

1. Structure has been reinforced (corner bracket)

• Honeycomb Al-Al-Al (Primary & Secondary Structure )

• Honeycomb CFRP-Al-CFRP (Solar Cells accommodation)

2. Satellite has better shielding

3. Mass has decreased to 318kg

See Mass Budget

93



Power reduction


• Thermal control = 91W (due to propellant thermal constraint)

• Propulsion (HPGP) = 128W (16x8W after the mass reduction)

• Payload = 40W (pre-warming)

• Reaction Wheel = 42W (when 3 accelerating + 1 constant speed)

2. Still to be included within the 294W

• STR / Gyro / GPS / RF systems / On-Board Computer / Propulsion Electronics / …

94

30%

43%

13%

13%

99%



Power reduction


• Thermal control = 91W (due to propellant thermal constraint)

• Propulsion (HPGP) = 128W (16x8W after the mass reduction)

• Payload = 40W (pre-warming)

• Reaction Wheel = 42W (when 3 accelerating + 1 constant speed)

2. Still to be included within the 294W

• STR / Gyro / GPS / RF systems / On-Board Computer / Propulsion Electronics / …

95

30%

43%

13%

13%

99%

What would do a good system engineer ?



Power reduction

1. Thermal Control

• Satellite wrapped into MLI

• Electronics as close as possible to cold points

2. Propulsion

• Only half of thruster pre-warmed + duty cycle

• Only less than half the power is given at the same time (longer pre-warming!)

3. Payload

• Longer pre-warming

4. GNC

• FFRF removed

96



Power reduction

1. Thermal Control reduced to 75W (instead of 91W)

2. Propulsion reduced to 33,6W (instead of 128W)

3. Payload reduced to 10W (instead of 40W)

97



Power reduction

1. Thermal Control reduced to 75W (instead of 91W)

2. Propulsion reduced to 33,6W (instead of 128W)

3. Payload reduced to 10W (instead of 40W)

98

What is the next

step ?



Electrical Architecture

1. Power Generation & Storage

• Solar Array

• Battery

2. Power Conditioning

Distribution Unit

• ADPMS

3. Connections

• Safe & Arm

• Umbilical Connection

99



Solar Array Design

1. Main components

• Cells Series (TBD)

• String Parallel (TBD)

• 6xSection (TBD strings)

2. Secondary components

• Shunt Selection

• Dump Resistor

100



Solar Array Design

1. Main components



• 6xSection (TBD strings)


• Shunt Selection

• Dump Resistor

101

How can we

calculate this?



Solar Array Design (3G28%)

1. Evaluate the degradation?

• Coverglass thickness

• Degradation of Electrical Parameters

• Degradation of Temp. Coefficient

102



Solar Array Design (3G28%)

1. Evaluate the degradation?

• Coverglass thickness

• Degradation of Electrical Parameters

• Degradation of Temp. Coefficient

103



Solar Array Design

2. Evaluate the number of cells ?

• Max battery voltage to be provided (29.4V)

• Compute all voltage drop

• Compute number of cells (MPP)

104

18 Cells / Strings



Solar Array Design

3. Evaluate the number of strings ?

• Max current allowed by ADPMS (12A)

• Compute string current

− EoL ( di/dt < 0)

− Minimum Solar Cste

(di/dC > 0)

− Operating temperature

(di/dT>0 but dU/dT<<0)

− Operating point

105



Solar Array Design


• Max current allowed by ADPMS (12A)

• Compute string current

− EoL ( di/dt < 0)

− Minimum Solar Cste

(di/dC > 0)

− Operating temperature

(di/dT>0 but dU/dT<<0)

− Operating point

106

23 (+1) Strings



Solar Array Design

4. Evaluate the Power available?

• Power = Current * Voltage

• 1 String Failure Tolerance

107



Solar Array Design

4. Evaluate the Power available?

• Power = Current * Voltage

• 1 String Failure Tolerance

108


3 PROBA - 3

What if GNC has a failure?

109


3 PROBA - 3

What if GNC has a failure?

ANSWER = Need to be taken into account if pointing accuracy of SUN

POINTING is decreased!

110



Battery Design

1. Main components




• Internat heaters

• Thermistors

111



Battery Design

1. Main components




• Internat heaters

• Thermistors

112

How can we

calculate this?



Battery Design

1. Evaluate the number of cells ?

• Nominal non regulated bus voltage (28V)

• Cells characteristics (4.2V EoC)

113

2.5

2.6

2.7

2.8

2.9

3.0

3.1

3.2

3.3

3.4

3.5

3.6

3.7

3.8

3.9

4.0

4.1

4.2

0 10 20 30 40 50 60 70 80 90 100

State of Charge (%)

Ce

ll L

ev

el E

MF

(V

)

Discharge EMF Charge EMF

7 Cells / String



Battery Design


• Approach decision

− Battery for both Sun & Eclipse

− Battery for Eclipse only

• Check Power Budget (Wst Case)

− Detumbling (200W / 1hr)

− Eclipse (240W / 30min)

− Long Eclipse (190W / 3.5hrs)

114



Battery Design


• Capacityused = Power * Time

• Capacityrequired = CapacityUsed*(1+DoD)

• Capacity of 1 string = NbCells*CapacityCell

• One String Failure Tolerance

115

Scenario Used Capacity (Wh)

Detumbling 200

Eclipse 120

Long Eclipse 665

Scenario DoD choice Required

Capacity (Wh)

Detumbling 20% 240

Eclipse 20% 148

Long Eclipse 60% 1064

Parameter Value

NbCells 7

CapacityCell 5,4 Wh

Capacity 1 string 37,8 Wh

29 (+1) Strings



U/D Operational Concept Problems

1. Downlink not closed at apogee at Max Rate

2. Uplink not closed at apogee at Max Rate

3. Data not able to be downlinked below 50kkm at max rate (coverage)

4. Data not able to be downlinked at min rate (coverage)

See Link Budget

116


3 PROBA - 3

What can we do?

117



U/D Operational Concept Solutions

1. Use more power @ Ground Segment

2. Use more power @ Space Segment

3. Use one/several high gain antenna @ Space Segment

4. Use variable data rate

• Low Rate (Uplink & Downlink) when at apogee

• High rate (Uplink & Downlink) when at perigee

118



U/D Operational Concept Solutions

1. Use more power @ Ground Segment

2. Use more power @ Space Segment

3. Use one/several high gain antenna @ Space Segment

4. Use variable data rate

• Low Rate (Uplink & Downlink) when at apogee

• High rate (Uplink & Downlink) when at perigee

119

What is the consequence ?

What is the consequence ?


3 PROBA - 3

Results !

120



Updated Block Diagram

121



Updated Configuration

122




123




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Final Stack Configuration

1. Occulter Spacecraft

2. Coronagraph Spacecraft

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3 PROBA - 3

How would you now further reduce the cost ?

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3 PROBA - 3

How would you now reduce the cost ?

ANSWER = Increase the risk !

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4 CONCLUSION

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4 Conclusion

Satellite System Engineering

1. Follows the Project Life Cycle

• Starts with Mission Concept

• Prepares System Requirements

• Proposes System Designs based on Trade-Offs (Technical + Programmatics)

• Manufacture the S/C

• Verify Requirements (Review of Design / Analysis / Test)

• Launch it !

2. Iterative Multi-disciplinary approach + Massive Communication

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4 Conclusion

QinetiQ Space

1. Proposes Fast Learning Curve on Satellite Systems

2. Offers possibility to built international network quickly

3. Provides opportunity to be known at the European Space Agency

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www.QinetiQ.be

© Copyright QinetiQ Limited 2010

QinetiQ Space nv

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Documents

Satellite Engineering PROBA Familly - LTAS-SDRG · Title: PROBA-3 Kick-off Meeting Phase B2.2 Author: jtallineau Created Date: 12/9/2012 8:13:03 PM