CDF FLORAD 6 Technical Reportdma.ing.uniroma1.it/users/ls_sas/MATERIALE/PhA study Small EO s… · Spacecraft Attitude Determination and Control, Microcosm, Inc., Torrance 1978. 70

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CDF Study Report

FLORAD

MasterSat 2007-2008

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F R O N T C O V E R

The 3-axis stabilised FLORAD spacecraft is shown in its eliptic orbit at 600 ± 20 km altitude.

The FLOMIR radiometer is pointing towards the Earth. Accommodated inside the box-shaped

spacecraft are a tank, used to manoeuvres.

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STUDY TEAM

This FLORAD Assessment Study was performed in the MasterSat Concurrent Design Facility

(CDF) by the following interdisciplinary team:

TEAM LEADER R. Quartucci

AOCS M. Broglia Propulsion L. Visconti

S. Robbio

Configuration R. Quartucci Structures I. Sarandrea

Data Handling A. Cotichelli

G. De Donato

Systems R. Quartucci

Instruments Prof. F. Marzano Telecomms G. De Donato

A. Cotichelli

Mission analysis S. Robbio

L. Visconti

Thermal F. Capece

Power F. Capece

Study Manager Prof. P. GAUDENZI

Science Advisor Prof. F. MARZANO

Systems Advisor Ing. G. MORELLI

The team would like to thank: G. BELVEDERE

A. GOLKAR

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Further information and/or additional copies of this report are available from:

Università di Roma “SAPIENZA”

Prof. PAOLO GAUDENZI

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Table of contents

FLORAD..................................................................................................................1

STUDY TEAM........................................................................................................3

TABLE OF CONTENTS .......................................................................................5

1 INTRODUCTION ........................................................................................12

1.1 Background 12

1.2 Document Structure 12

1.3 Terminology 12

2 EXECUTIVE SUMMARY............................................................................13

2.1 Study Flow 13

2.2 Mission Objective 13

2.3 Requirements and Constraints 13

2.4 Scientific requirements 14

2.5 Mission Summary 1

3 MISSION OVERVIEW..................................................................................2

3.1 Mission analysis 2

3.1.1 Four different launches: 2

3.1.2 Launch with multiple satellites with orbit plane change: 2

3.1.3 Launch with multiple satellites via electric propulsion: 2

3.1.4 Launch with multiple satellites via orbit precession: 3

3.2 Mission baseline 6

3.3 Operational orbit 7

3.3.1 Ground Station coverage 7

3.3.2 Orbit injection 7

3.3.3 Hohmann Manoeuver 7

3.3.4 Station-keeping 8

3.3.5 End of Life De-orbit 8

3.4 Constellation design 9

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4 SYSTEM CONCEPT.....................................................................................10

4.1 Objectives 10

4.2 System requirements 10

4.3 Design drivers 10

4.4 Mass and power budgets design 13

4.4.1 S/C subsystems design 14

4.4.2 Preliminary mass and power budgets 15

4.5 Spacecraft modes of operation 16

4.6 Spacecraft mechanical states 19

4.7 Margin philosophy 19

4.8 Redundancy philosophy 19

4.9 System summary 20

4.9.1 Mass evolution 20

4.9.2 Final mass budget 21

4.9.3 Power budget 22

4.9.4 Spacecraft functional diagram 23

4.9.5 Spacecraft equipments 23

4.9.6 Spacecraft characteristics and performances 24

5 POWER SUBSYSTEM .................................................................................27

5.1 Requirements and design drivers 27

5.2 Baseline design 27

5.2.1 Architecture 27

5.2.2 Solar Array 28

5.2.3 Battery 28

5.3 Performances and budgets 29

5.4 Other options 30

5.5 Conclusions 31

6 AOCS SUBSYSTEM.....................................................................................32


6.2 Control modes specifications 32

6.3 S/C features 33

6.4 Assumptions and trade-offs 33

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6.5.1 Reaction wheels sizing 33

6.5.2 Desaturation 34

6.5.3 AOCS sensors 34

6.6 AOCS equipments 35

6.7 AOCS data rate 35

6.8 AOCS mass and power budgets 36

6.9 Conclusions 36

7 PROPULSION SUBSYSTEM.....................................................................37




7.4 Equipments 40

7.4.1 Propellant Tanks 40

7.4.2 Thruster 40

7.4.3 Latching Valve (LV) 41

7.4.4 Liquid Filter (LF) 41

7.4.5 Fill and Drain Valve (FDV) 41

7.4.6 Pressure Transducer (PT) 41

7.4.7 Feeding lines (Pipework) 41

7.5 Budgets 41

7.5.1 ∆V budget 42

7.5.2 Propellant budget 42

7.5.3 Dry mass budget 42

7.6 Conclusions 43

8 DATA HANDLING .....................................................................................44

8.1 Requirements and constraints 44

8.2 Data architecture design 45

8.3 Mass memory and throughput 46

8.4 Payload data processing and transmission 47

8.5 Hardware architecture and equipment overview 48

8.6 Conclusions and remarks 48

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9 TT&C SUBSYSTEM.....................................................................................50

9.1 Requirements and constraints 50



9.3.1 Data Transmission 51

9.3.2 Ground Station 51

9.3.3 Satellite antennas 51

9.4 Performances and budgets 52

9.4.1 Mass and power budgets 53

9.5 Hardware architecture and equipment overview 54

9.5.1 Hardware architecture 54

9.5.2 Antenna 55

9.5.3 Transponder 56

9.6 Conclusion and remarks 57

10 THERMAL CONTROL SUBSYSTEM..................................................58

10.1 Thermal environment 58

10.2 Thermal requirements and constraints 58

10.3 Thermal design assumptions 59


10.4.1 The platform 59

10.4.2 The Solar Array 59

10.5 Trade-off 60

10.6 Thermal equipment 60

10.7 Conclusions 60

11 CONFIGURATION ANALYSIS ............................................................62

11.1 General requirements and constraints 62

11.2 Configuration 62

11.2.1 Baseline design 62

11.2.2 Internal accommodation 62

11.2.3 External accommodation 63

12 STRUCTURES ...........................................................................................67

12.1 General requirements and constraints 67

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12.1.1 Requirements and design drivers 67

12.1.2 Baseline design 68

12.1.3 Assumptions and trade-offs 68

12.1.4 Mass budgets and baseline sizing 69

12.2 Conclusions 69

13 REFERENCES ............................................................................................70

a. Small satellites CNES programme, Myrade Platform, Paper 2008 70

b. D. Pereira, Spacecraft design project, 2004 FINAL DESIGN REPORT, Carleton University 70

c. STPSat-1 Data Sheet, Aeroastro, 2008 70

d. SSTL MicroSat-70 Documentation, Surrey 2008 70

e. SSTL MicroSat-100 Documentation, Surrey 2008 70

f. MicroStar Satellite Platform Documentation, Orbital 2008-04-10 70

g. SA-200B Data Sheet, Spectrum Astro 2008 70

h. J.R.Wertz. Mission Geometry; Orbit and Constellation Design and Management, Space Technology Library, 2001 Microcosm, Inc. 70

i. J.Rumbaugh, I.Jacobs and G.Booch. The Unified Modeling Language Reference Manual, Addison--Wesley. 70

j. G.T.French. Understanding the GPS, GeoResearch, Inc., 1996. 70

k. http://www.nasa.gov/ 70

l. J.R.Wertz. Spacecraft Attitude Determination and Control, Microcosm, Inc., Torrance 1978. 70

m. J.R.Wertz. Space Mission Analysis and Design, Microcosm, Inc., El Segundo 1999. 70

n. M. Guelman, R. Waller, A. Shiryaev, M. Psiaki.``Design and Testing of Magnetic Controllers for Satellites Stabilization''. 70

o. E. Perez, “Vega User’s Manual / Revision 0”, 2004 70

p. P. Fortesque, “Spacecraft System Engeneering”, Wiley 70

q. E.M.Silverman. “Product Development of Engineered Thermal Composites for Cooling Spacecraft Electronics”. 71

r. R.D.Karam. Satellite Thermal Control for Systems Engineers, Edited by P. Zarchan, American Institute of Aeronautics and Astronautics, Massachusetts. 71

s. Agneni, On Board Power Generation, Master in Satellite and orbiting platforms, Scuola di Ingegneria Aerospaziale 71

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t. L. Schirone, Space Solar Cells, Master in Satellite and orbiting platforms, Scuola di Ingegneria Aerospaziale 71

u. W.J.Larson and J.R.Wertz. Space Mission Analysis And Design, Space Technology Library, 2005 Kluwer Academic Publishers, London. 71

v. Fundamentals of Space Systems, Edited by V.L.Pisacane and R.C.Moore, 1994 Oxford University Press, New York Oxford. 71

w. Spacecraft Systems Engineering, Edited by P.Fortescue and J.Stark, 1995 John Wiley and Sons, Singapore. 71

x. Spacecraft Thermal Control Handbook, Edited by D.G.Gilmore, 2002 The Aerospace Press, El Segundo California. 71

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1 INTRODUCTION

1.1 Background In February 2008, ASI Board of Directors approved five missions related to “Bando

per Piccole Missioni”, which phase A study will be found to lead a preliminary feasibility study. In the context of this programme, FLORAD mission plays an important role.

1.2 Document Structure The layout of this report of the study results can be seen in the Table of Contents.

The Executive Summary provides an overview of the scientific aims, instruments, spacecraft, launcher, mission exploitation and programmatics. Details of each domain addressed in the study are contained in specific chapters.

1.3 Terminology Throughout the report, the following terminology applies:

• Flower Constellation (FCs) is a general class of elliptical orbits which can be optimized, through genetic algorithms in order to maximize the revisiting time and the orbital height, ensuring also a repeating ground – track.

• FLOMIR is a millimetre wave (MMW) passive radiometer used to observe the tropospheric parameters of Earth atmosphere

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2 Executive Summary

2.1 Study Flow The assessment study of a potential FLORAD mission using the ESA Concurrent

Design Facility (CDF) was initiated by MasterSat 6 Team work of Sapienza University of Rome.

The study was conducted from the kick-off on 08th March 2008 through to the Internal Final Presentation on 12th April 2008. It involved seven technical sessions of the interdisciplinary study team.

2.2 Mission Objective The FLORAD mission is a fundamental physics mission addressing tropospheric

termo-dynamical profile and hydrological content on Earth atmosphere by taking advantage of the superior potential of a constellation of microsatellites with a MMW passive payload.

The mission objectives of FLORAD are threefold:

• Application o Tropospheric monitoring (e.g. over sea) o Assimilation within mesoscale models o Meteo – hydrological nowcasting

• Technological o MMW compact radiometers o Exploitation of micro – satellites o Flower constellation of satellites

• Context o Eumetsat priorities (MRD doc. 2007) o Multi – satellites synergy

2.3 Requirements and Constraints The FLORAD assessment study was performed by an interdisciplinary team using

the CDF to support the selection of the mission. The study objectives were to demonstrate technical feasibility providing analysis leading to system and subsystem conceptual design and compliance with the mission requirements.

The mission requirements and constraints for FLORAD are driven by the science objectives and the constraints proposed for ASI Announcement of Opportunity :

• � mission lifetime about 2 years

• � launch in the period 2014 to 2015

• � hardware recurring or derived from earlier missions

• � cost < 50 M€

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2.4 Scientific requirements

Parameter Preliminary FLORAD Mission Requirements Target area at a regional scale • Mediterranean area and Southern

Europe

• Life cycle: > 2 years

• Launch (possible): 2012

Temperature vertical profile in presence of non-precipitating cloud

• -Accuracy: < 2 K

• -Vertical resolution: < 2 km

• -Horizontal resolution: < 25 km2

• -Temporal repeativity: < 3 h

• -Geolocation error: < Horizontal resolution

Integrated water vapor content • -Accuracy: < 10 kg/m2 (su mare)




Non-precipitating integrated cloud liquid and ice water

• -Accuracy: < 1 kg/m2 (su mare)




Stratiform precipitation intensity • -Accuracy: < 5 mm/h




Snowfall precipitation intensity • -Accuracy: < 1 mm/h (equivalente)




Table 2.1- FLORAD scientific requirements

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2.5 Mission Summary

Mission Objective Launch a constellation of micro – satellites with a MMW passive payload to retrieve trophospheric termo – dynamical profile and hydrological content

Payload

Milliwave Imaging Radiometer (FLOMIR)

• Sounding for temperature and water vapour

• Configuration of 19 channels (4 bands) o 50 – 57 GHz & 113 – 123 GHz & 178 – 188 GHz

• Antenna conical scanning mode

• Broadband corrugated horns

Launcher VEGA

• Performance: 1500 kg (Low Earth Orbit) Orbit Quasi circular Orbit

• hp = 600±20 km

• i = 63.4° Mission

Launch date 15 August 2015

Operations Ground station Fucino Baseline Design lifetime At least 2 years (including transfer) Platform Buy or build

Total dry mass 150 kg (including margin) Spacecraft main body dimensions

1m x 1m x 1m

Spacecraft

Power Consumption 200 W

Payload Mass 40 Kg Attitude control 3-axis stabilised (wheel, magneto –

torque) Attitude Determination Magnetometer, GPS

Solar panels 4 GaAs modules Thermal S/S Passive TT&C S – Band trasponder and data

transmission Battery Li - Ion

Table 2.2 – FLORAD mission specifications

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3 Mission Overview

The mission analysis activities focused on contributing input for the selection of an appropriate strategy for the constellation launch via a comparative analysis of the respective merits of different options.

We have considered four different options:

− Four different launches;

− Launch with multiple satellites with orbit plane change;

− Launch with multiple satellite with electric propulsion;

− Launch with multiple satellite via orbit precession; After comparative merit analysis was selected as baseline and a more in-depth

analysis was performed taking into account:

• Ground station coverage;

• ∆V budget

• Thermal and radiation environment

3.1 Mission analysis

3.1.1 Four different launches:

Each launch for a single FLORAD satellite to its specific orbit. It is very simple but costly. Possibility to exploit spare-room in big launcher dispensers (piggy-bag). For this kind of strategy we have problems with cost budget and disposal time of constellation. Alternative could be a possible future study about four different aircraft launch.

3.1.2 Launch with multiple satellites with orbit plane change:

Launch first to set up the RAAN of Sat1; than carry out the impulsive maneuver to change the orbital plane reach the Sat2 RAAN and release the Sat2…and so on. Rapid positioning, but need to provide a lot of propellant in the satellite to accomplish

maneuvers. In fact we estimated a ∆V budget to change orbit plane from polar orbit to

nominal orbit (∆i = 26,6°) of about 3000 m/s. To provide this ∆V we need for over 90% of satellite mass in propellant (using chemical propulsion system).

3.1.3 Launch with multiple satellites via electric propulsion:

The launch strategy is the same of the previous option. In this case we cannot use impulsive maneuver so we have a problem of slow positioning. But the really problem is the high power demanding for this kind of propulsion. In fact the lower power required by a electric thruster is about 600 W.

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3.1.4 Launch with multiple satellites via orbit precession:

We use a single launch to put in a parking orbit, higher than nominal orbit, all our satellites. Then carry out the impulsive Homann maneuver to reach the nominal orbit altitude for each satellite. The time between each maneuver is calculated from J2 effect study.

Figure 3.1

As showed by foregoing graphic, we can observe that the increasing parking orbit altitude have a relative precession velocity as regards the nominal one.

Increasing the altitude of parking orbit the ∆V budget augments.

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Figure 3.2

For a fixed satellite dry mass the total satellite mass increase with the altitude parking orbit. In fact we need for more propellant little by little that the altitude of parking orbit increases as showed in the following graphic:

Figure 3.3

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If we consider launcher payload for each altitude and the constraint of total mass of satellite, we will choose the optimum altitude of parking orbit. After some iteration of CDF, we fixed dry mass about 95 Kg and total mass 135 Kg for each satellite and we choose a parking orbit altitude about 2000 km.

Figure 3.4

Considering the ∆RAAN diagram we can estimate the time between the first and last maneuvers, (270° of RAAN).

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Figure 3.5

3.2 Mission baseline

Considering power and mass constraints imposed by mission requirements, the strategy selected is a launch with multiple satellite via orbit precession. After the analysis to obtain the total disposal of flower constellation we need for the following characteristics:

• Altitude for parking orbit: 2000 Km;

• Inclination of parking orbit must be the same of nominal orbit = 63.4°;

• ∆V disposal constellation ≈ 650 m/s;

• Time total disposal ≈ 6 months.

It is important to underline that the time disposal, about six months, is not the time of the begin operative life of the constellation, in fact just after the launch we have a satellite in operative status, after two months the second satellite and so on till six months where the constellation is totally deployed.

The characteristics of the selected orbit are listed here:

• Circular at 600 km altitude above equatorial Earth radius (6378 km)

• Inclination: 63.4°

• Period: 96.68 minutes

• Maximum eclipse duration: 35.25 minutes

They offer a stable thermal environment but solar panel design is not so simple (Sun incidence angle is variable very much during lifetime). In addition, eclipse time is

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relatively short and it doesn’t depend on seasons. Such an orbit is proposed for Florad mission.

3.3 Operational orbit

3.3.1 Ground Station coverage

The ground station at the center of target are is Fucino. The access area characteristics are summarised in the following table.

Minimum elevation 10 Deg

Longest pass duration 546 s

Mean pass duration 419 s

Mean number of passes per day 20

Table 3.1

As showed in the following diagram we can observe that the distribution of contacts with ground

station is homogeneous during the day. In particular the range time between two successive

contacts is about 4 min ÷ 2 h 30 min.

Figure 3.6

3.3.2 Orbit injection

Typically after separation satellites are injected not in the nominal orbit but in a similar one with some errors in the keplerian elements. Given these errors, the propulsion system provide for corrections for an amount of about 12 m/s in worst case. To consider the de-spin of each satellite the propulsion system provide 18 m/s.

3.3.3 Hohmann Manoeuver

To reach the operative orbit with Homann transfer we need for a V∆ of about 650 m/s.

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3.3.4 Station-keeping

At an altitude of 620 km, the atmospheric air density is between 7.90T10-15 and 5.10T10-12 kg/m3 depending on the solar activity. This translates into a force up to 3.20T10-4 N per square meter cross sectional area and it drives to about 1 m/s for mantaining the orbit for the entire mission. Moreover for controlling orbit inclination about 7,5 m/s are necessary during mission lifetime.

3.3.5 End of Life De-orbit

Orbit decay rate is about 0.2 to 200 m per day depending on solar activity. Natural decay will take about 40 years (to a 150 km orbit).

∆V for launcher dispersion 12 m/s

∆V for inclination control 9,4 m/s

∆V for de-spin 18 m/s

∆V for station keeping 0,5 m/s

∆V for manoeuver 688 m/s

∆∆∆∆V total budget : 727 m/s

Table 3.2

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3.4 Constellation design

Different configurations are analyzed to optimize the revisit time and global coverage. For FC we use 4 satellite in 4 different orbit plane ( i=63.4° , RAAN= 0° , 90°, 180°, 270°).

Figure 3.7 Flower constellation

Figure 3.8 Flower constellation ground track

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4 System Concept

4.1 Objectives

The FLORAD system design report is issued to support the formal preliminary design Pre-Phase A of FLORAD programme, with the goal of explaining all the technical choices at system and sub-system level performed during the CDF sessions.

The objectives of this document are to:

• Describe FLORAD spacecraft design

• Demonstrate the full and optimal compatibility of the described system with respect to the requirements specification.

• Develop a secondary configuration in order to minimize costs and mass budgets

4.2 System requirements In the following, the system requirements are summarised and related to their

consequences on the spacecraft system design. The scientific goals of the FLORAD mission require:

• Payload: Radiometer (RAD) ,millimetric-wave band radiometer; 19 channels (4 bands).

• Pointing Accuracy: 0,5 deg

• Altitude: 600 ± 20 km

• Launcher: VEGA

• Stabilization Type: 3-axis

• Bus Voltage: Unregulated 22V-34V

• OBDH: Integrated

• Propulsion type: Monopropellant

• TT&C: S band bit rate 2Mbit

• Operative LifeTime: 2 years

• Autonomy: 2 hours

4.3 Design drivers

In the following, the main system design driving requirements, resulting from the FLORAD mission profile are identified, and listed w.r.t. the relevant system design areas. The overall system design requirements are:

• to minimize spacecraft weight (maximum wet mass: 150 Kg, including adapter, µ-satellite class)

• system reliability shall be maximised by an “as simple as possible“ system design

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• subsystem reliability shall be maximised by use of technologies used in past missions (commercial technologies “space qualified”)

• Vega is the preferred launch vehicle but other launcer could be studied

• to consider with attention heritage from previous analogous missions According to the design drivers described above, the system engineer has produced,

in accordance with subsystem engineers, a preliminary baseline design for each subsystem, and, where useful, a set of alternatives to be considered. A summarized table is presented below:

Name: Baseline

No.: 1 System Option

Key: A

System Approach Option

Redundancy 2 Functional

Use of Existing Platform 2 no

Existing Platforms Identified

Mission

Altitude 1 600 ± 20

Life time Duration (yrs) 1 2

Launch

Launch year 1 15-Aug-2015

Launcher 2 VEGA

Number of satellites 1 4

Wet mass 2 150

Propulsion

Type of Propusion 1 Chemical Monopropellant

Specific Description 3 Hydrazine

No. and position of thrusters 2 8 in the corner

Operations

No. Ground Station 1 1

Ground Station Operational 2 Fucino

Configuration

Stabilisation 1 3-Axis

S/C Modules 1 (P/L-S) + (SEP)

Payload Accommodation 2 Internal/external

Satellite Platform Shape 1 Box

AOCS

Desaturation Time 1 3orbit

Wheels Desaturation 2 Magnetic Torquers

Actuation System 1 Momentum Wheel

Power

Solar Array Technology 2 4 GaAs

Battery type 1 Li-Ion

SEP SA Configuration 1 2 wings

SEP SA Movement 2 fixed

Data Handling

Data Retrieval Process 2 Store & dump

Data Rate Philosophy 1 Fixed

Comms TT&C

LGA Antenna Type 1 helix conical

Thermal & Mechanisms

Thermal Control 3 Passive

Table 4.1: preliminary baseline design

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In the following the sets of sub-system options is presented:

System Approach 1 2 3

Redundancy Full Functional None

Use of Existing Platform yes no

Existing Platforms Identified

CNES - Myriade SSTL – MicroSat100

Mission

Altitude 600 <600 >600

Life time Duration (yrs) 2

Launch

Launch year 15-Aug-2015

Launcher VEGA EFA-2000

Number of satellites 4

Wet mass 150 100

Propulsion

Type of Propusion Chemical Ionic

Specific Description T5 SPT Hydrazine

No. and position of thrusters

4, -X face 4, -X face, 2 –Y face 8 in the corner

Operations

No. Ground Station 1 2

Ground Station Operational

Svalbard Fucino

Configuration

Stabilisation 3-Axis

S/C Modules (P/L-S) + (SEP) Single Module

Payload Accommodation embedded external Mixed

Satellite Platform Shape Box

AOCS

Desaturation Time 2orbit 4orbit 10 orbit

Wheels Desaturation Thrusters Magnetic Torquers

Actuation System Reaction Wheels Momentum Wheels

Power

Solar Array Technology Silicon GaAs – Triple GaAs – Double

Battery type Li-Ion Nich-Hi Nich-Cadmium

SEP SA Configuration 2 wings 4 wings

SEP SA Movement fixed tiltable 1-axis tiltable 2-axis

Data Handling

Data Retrieval Process Real Time

Elaboration Store & dump

Data Rate Philosophy fixed Adjustable

Comms TT&C

LGA Antenna Type helix conical quadrifilar

Thermal & Mechanisms

Thermal Control Active Passive Active&Passive

Table 4.2: Sub-system options

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4.4 Mass and power budgets design In this section is described the method utilized by the system engineer for the

preliminary estimation of the S/C Mass and Power Budget. The method developed provides a jointly design of Mass and Power budget because it isn’t possible to obtain a good design considering the two budget separately. The design diagram is represented in the figure below. The concept is to start from payload mission requirements and orbit characteristic, that drive the payload design. Then the mass of electrical power system is determined, according to payload power consumption.

Figure 4.1: Mass and Power budget design scheme

MISSION REQ.

PAYLOAD DESIGN

MASS BUDGET POWER BUDGET

PROP MASS

PAYLOAD POWER

S/S 1 POWER

S/S N POWER

S/C TOTAL POWER

EPS MASS

STR MASS

TCS MASS

Other S/S MASS

S/C TOTAL MASS

CHECK

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4.4.1 S/C subsystems design

Considering the baseline S/S design previous presented in Table 4.2 and according to Tables 4.3-4.4 (presented below) of typical mass and power budget for mini satellites and micro satellites, the mass and power budget for each subsystem are generated.

Typical Mass Budgets for MICRO SATELLITES

Typ Mass (Kg) Myriade Platform (Min)

Min % SSTL MicroSat100 (Max)

Max %

Payload Payload 25 23,29 20 27,1

Platform Structure (inclusive of mechanism) 16 14,91

AOCS 11.2 10,44

Electrical Power 23.5 21,90

Data Handling 5.3 4,94

TT&C 8.6 8,01

Propulsion 0 0

Thermal Control 2 1,86

total PLATFORM 66,6 76,71 100 67,5

Miscell Harness, Supports and Miscellanea 2 1,86 10 6,7

Total S/C Dry Mass 107,32 100 143 94,2

Propellant 0 0 5 5,8

Total S/C Launch Mass 107,32 100,00 148 100,00

Table 4.3: Typical Mass Budget for micro satellites

Typical Power Budgets for MICRO SATELLITE

Typ Power (W)

Min Max

Payload Payload 20 50

Platform Structure 0 0

AOCS 0 15

Electrical Power 10 30

Data Handling 5 5

TT&C 5 15

Propulsion 0 5

Thermal Control 0 5

total BUS 40 125

Harness (ohmic losses) 1 5

Total S/C Power Consumption 41 130

Solar Array available power (typical range) 80 120

Power provided by battery 48 100

Table 4.4: Typical peak power budget for radar observation satellite

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4.4.2 Preliminary mass and power budgets

The preliminary mass and power budget are developed according to method diagram above presented. The system baseline was described in Table 4.1.

Estimated Mass Budget

(kg) % (on dry

mass)

Preliminary Budget

Payload Florad Payload 30,000 33,708 27,000

total payload 30,000 23,256 27,000

Platform Structure (incl. Mechanism) 17,000 19,101 15,300

Electrical Power 17,000 19,101 15,300

Integrated Control ( DH) 4,000 4,494 3,600

AOCS 8,000 8,989 7,200

TT&C 4,000 4,494 3,600

Propulsion 2,000 2,247 1,800

Thermal Control 1,500 1,685 1,350

total BUS 53,500 60,112 48,150

Miscell Harness 3,000 3,371 2,700

Supports and Miscellanea 1,500 1,685 1,350

Balancing Masses 1,000 1,124 0,900

Total S/C Dry Mass 89,000 68,992 4,950

Other 0,000 0,000 0,000

Propellant 40,000 31,008 36,000

Total S/C Launch Mass 129,000 100,000 116,100

Estimated Volume Budget

Density kg/m3 Volume

Platform Structure (incl. Mechanism) 89,000 1,000

Table 4.5: Preliminary system mass budget

Typ Power Nominal

(W)

Payload Florad Payload 30,000000

total payload 30,000

Platform Structure 0,000

Electrical Power 30,000

Integrated Control (DH) 30,000

AOCS 15,000

TT&C 30,000

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Propulsion 0,000

Thermal Control 0,000

total BUS 105,000

Harness (ohmic losses) 3,000

Total S/C Power Consumption 138,000

Solar Array available power (typical range) 200,000

Power provided by battery -62,000

Table 4.6: Preliminary system power budget

4.5 Spacecraft modes of operation

The basic system modes of operation are defined below. Table 2.7 gives a general summary of the system modes.

Mode Name Definition

Launch Mode

Onboard launcher:: If we are using Cone Adapter, all sub-systems are off, except essential equipment (e.g.RX).Satellite capable of receiving and executing Telecommands.The radiometer is off. An automatic switch is used at separation to activate the equipment start-up sequence.

Initialisation Mode

Initial Deployment and Attitude acquisition: SA deployed and operational Attitude acquisition with SUN pointing Transmission of Telemetry Data;failure detected and corrected by the ground. Contingency situation possible

Nominal Mode

Nominal Earth Atmosphere Observation with Radiometer The spacecraft is kept pointing. Accuracy of pointing is determined by AOCS Acquisition of optical image data . Acquisition, Storage and trasmission to ground RX stations of data Contingency Situation possible

Eclipse Mode

Thermal Control passive: Switch Power generation from SA to battery. Accuracy determined by AOCS Payload and all subsystem are on Acquisition, Storage and trasmission to ground RX stations of data Contingency Situation possible

Calibration Mode

Calibration of radiometer The spacecraft is kept pointing. Accuracy determined by AOCS. TT&C activated. Payload Activated Contingency Situation possible

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Safe Mode

Hibernation and Failure Recovery mode: The spacecraft is kept SUN pointing. Accuracy determined by AOCS. Payload is put on standby or switched off. Non-essential functions are halted. TM/TC access to DHS is guaranteed to enable failure detection and reconfiguration. Failure detection and recovery are executed by the ground. Contingency Situation possible

Table 4.7: Modes of operation at system level

Figure 4.2: Modes of Operation transition diagram

LAUNCH

INITIALISATION

CALIBRATION

ECLIPSE

Mode Transitions

Diagram

NOMINAL

SAFE

18

Modes of operation at subsystem level are described in the next table:

Mode Sub-system Requirements THERMAL Passive

COMMS functions switched,only RX.

DATA HANDLING All functions - OFF

AOCS All functions - OFF

PROPULSION All functions - OFF

POWER ON.

1 LM

PAYLOAD All functions - OFF

THERMAL Passive

COMMS All functions - ON

DATA HANDLING Initialisation functions, service module and payload module commissioning.ON

AOCS All functions - ON

PROPULSION All functions - ON

POWER Activate SA.

2 IM

PAYLOAD All functions - ON

THERMAL Passive


DATA HANDLING All Functions- ON


PROPULSION OFF

POWER Battery provide power for Payload and S/S.ON

3 NM


THERMAL Passive


DATA HANDLING All Functions- ON



POWER SAW & Battery provide power for Payload and S/S.

4 EM


THERMAL Passive


DATA HANDLING OFF




5 CM

PAYLOAD OFF

THERMAL Passive


DATA HANDLING OFF




6 SM

PAYLOAD OFF

Table 4.8: Modes of operation at subsystem level

19

4.6 Spacecraft mechanical states

The principal mechanical states are presented in the next table:

Mode Name Definition

Stowed State All mechanisms stowed for launch.

Fuel tanks full

Deployed with Fuel State (Nominal)

All mechanisms deployed: -Solar Arrays

Fuel tanks almost empty (Fuel Decreasing)

Table 4.9: Mechanical states of S/C

4.7 Margin philosophy

At sub-system level, margins are given depending on the confidence of results, such as those associated with equations used, the values of input parameters, and general uncertainties about a given design solution. Then, depending on the maturity of the items, contingency is applied on unit/item level. For each equipment a mass margin is applied in relation to its level of development:

• 5% Off-the-Shelf Items

• 10% Items to be modified

• 20% Items to be developed A System level margin of 10% is placed on the spacecraft dry mass (dry mass

including sub-system margins). In the first CDF session a margin of 10% for each subsystem was applied. This

margin was managed by system engineer during all project phases.

4.8 Redundancy philosophy

The drive to reduce mass/volume/cost led to the philosophy that redundancy would be kept to a minimum. For a few sub-systems some level of redundancy has been introduced because of design necessities and in some cases when cost/mass/volume constraints were not a large factor. Further studies could be done to identify where higher reliability can be gained with only a small increase in mass/volume/cost.

20

4.9 System summary

In this section an overall spacecraft summary is presented. The mass evolution during all project phases, final mass budget, final power budget, S/S principal characteristics, system architecture and S/S equipment are analyzed.

4.9.1 Mass evolution

In the following the spacecraft mass variation during the principal project steps of CDF sessions is presented.

120,000

125,000

130,000

135,000

140,000

145,000

150,000

155,000

160,000

1 2 3 4 5 6 7

Total Launch Mass

Target Spacecraft Mass

Figure 4.3: Mass evolution of S/C during CDF sessions

120,000

125,000

130,000

135,000

140,000

145,000

150,000

155,000

160,000

1 2 3 4 5 6 7

Total Launch Mass

Target Spacecraft Mass

Figure 4.4: Mass evolution of S/C during CDF sessions

21

4.9.2 Final mass budget

The mass identified in the system budget is based on the specified values of the individual units and subsystems. Depending on the maturity of the items, contingency is applied on unit/item level. The applied mass margin was 10%.

S/C Subsystems S/S Mass

(kg)

S/S Mass

Margin (%)

S/S Mass Margin (Kg)

S/S MASS with Margin

Structure 21,42 5 1,06 22,48 kg

Thermal Control 0,596 20 0,119 0,715 kg

Communications 6,548 10 0,655 7,203 kg

Data Handling 2,005 10 0,212 2,217 kg

AOCS 5,145 10 0,515 5,66 kg

Propulsion 8,897 5 0,88 9,777 kg

Power 8,505 10 0,86 9,365 kg

Harness 1,093 20 0,187 1,28 kg

Payload 32 10 3,2 35,2 kg

Supports & Miscellanea 1,5 10 0,15 1,65 kg

Balancing Mass 1 10 0,1 1,1 kg

Total S/C Dry Mass 88,7098 7,93776 96,6476 kg

Propellant (main) 37,4 kg

Total S/C Wet Mass 134,048 kg

System Margin 10% --

S/C Wet Mass with Margin 147,452

kg

Adapter Mass 2,026 kg

Total S/C Launch Mass 149,478 kg

Table 4.10: S/C mass budget

Balancing Mass

1%

Harness

1%

Power

9%

AOCS

5%

Propulsion

10%

Thermal Control

1%

Communications

8%

Structure

24%

Data Handling

3%

2%

Payload

36%

Structure

Thermal Control

Communications

Data Handling

AOCS

Propulsion

Power

Harness

Payload

Supports &

MiscellaneaBalancing Mass

Figure 4.5: S/C dry mass breakdown

22

4.9.3 Power budget

In this paragraph is listed the maximum spacecraft power consumption for each operative mode and for each subsystem. Six operational modes have been identified for which for the power subsystem has been dimensioned. The corresponding S/C power demand is given in the following table.

Table 4.11: S/C power budget for different modes of operation

In the next table is presented the mean power demand for orbit of total spacecraft for the nominal mode.

S/S Mean Orbit Nominal Mode

Power(W)

THERMAL 0,00

COMMS 26,00

DATA HANDLING 2,22

AOCS 10,66

PROPULSION 8,32

POWER 20,00

HARNESS 0,97

PAYLOAD 37,00

Total S/C Power 105,157

Table 4.12: Mean/Orbit S/C power

Power Levels: Absolute maximum (at Sub-system Level) for each mode.

Power (W)

Mode 1 Mode 2 Mode 3 Mode 4 Mode 5 Mode 6

LM IM NM EM CM SM

THERMAL 0,000 0,000 0,000 0,000 0,000 0,000 COMMS 26,000 26,000 26,000 26,000 26,000 26,000 DATA HANDLING 0,000 2,216 2,216 2,216 2,216 0,000 AOCS 0,000 10,656 10,656 10,656 10,656 10,656 PROPULSION 0,000 13,010 13,010 0,000 13,010 13,010 POWER 20,000 20,000 20,000 20,000 20,000 20,000 HARNESS 0,969 0,969 0,969 0,969 0,969 0,969 PAYLOAD 0,000 37,000 37,000 37,000 37,000 0,000

TOTAL POWER 46,969 109,851 109,851 96,841 109,851 70,635

23

4.9.4 Spacecraft functional diagram

Figure 4.6: S/C Functional diagram

4.9.5 Spacecraft equipments

The list of system equipment is shown in Table 2.13 for the baseline spacecraft:

FUNCTIONAL SUBSYSTEM nr Total Mass (kg) Margin (%) Margin (kg) Mass with Margin

Mass (kg) per unit

FLORAD

Structure 21,42 4,93 1,06 22,48

Closure Panel 4 3,22 12,87 5,00 0,64 13,511

Skeleton 1 0,29 0,29 0,00 0,00 0,292

Platform (top) 1 2,75 2,75 5,00 0,14 2,892

Platform (main) 1 2,75 2,75 5,00 0,14 2,892

Platform (bottom) 1 2,75 2,75 5,00 0,14 2,892

Thermal Control 0,5962 20,0000 0,1192 0,7154

MLI 1 0,5051 0,5051 20,0000 0,1010 0,6061

Radiator 1 0,0878 0,0878 20,0000 0,0176 0,1054

Communications 6,548 10,000 0,655 7,203

Alcatel TRC 2 3,000 6,000 10,000 0,600 6,600

model 2 1 0,036 0,036 10,000 0,004 0,040

SAAB Helix Quadrifilar Antenna 2 0,24 0,480 10,000 0,048 0,528

Data Handling 2,005 10,000 0,212 2,262

RAD6000 1 2,000 2,000 10,000 0,200 2,200

RAM 150X XPRESS 1 0,05 0,05 10,000 0,12 0,062

TCS S/S

SMU DH S/S

AOCS S/S

MW

MTR STR

ES

MGT

TT&C

XPN

XPNSAW

PCDU

BAT

EPS S/S

P/L

OMT

Antenna

RF

Data Lines TM Lines PROP S/S

DH Lines Power Lines

PCU

24

AOCS 5,1456 10,000 0,51456 5,66

MicroWheel-10SP-S 1 1,1 1,100 10,000 0,11 1,21

M.Torquer MT2-1 6 0,3 1,8 10,000 0,18 1,98

Magnetometer HMC6352 4 0,00014 0,0056 10,000 0,00056 0,001

Star Tracker Aero Astro MST 2 0,3 0,6 10,000 0,06 0,66

Earth Sensor MMS 13 410 2 0,8 1,6 10,000 0,16 1,76

GPS SGR-05 2 0,02 0,04 10,00 0,004 0,044

Propulsion 8,897 5,00 0,88 9,374

Pressurant 1 0,12 0,0976 5,00 0,44 0,126

Tanks (propellant) 1 6,40 6,4000 5,00 0,32 6,720

Fill Drain Valves 5 0,02 0,0180 5,00 0,00 0,095

Latching Valves 8 0,00 0,0040 5,00 0,00 0,034

Filters 4 0,08 0,0800 5,00 0,00 0,336

Lines and fittings 1 0,02 0,0200 5,00 0,00 0,021

Temp. Transducers 10 0,00 0,010 5,00 0,00 0,010

Pres. Transducers 5 0,08 0,0750 5,00 0,03 0,394

Thrusters 8 0,20 0,1950 5,00 0,09 1,638

Power 8,505 10,000 0,860 9,365

Solar Array 3 1,208 3,625 10,000 0,362 3,987

Battery (cell) 280 0,016 4,480 10,000 0,448 4,928

PDU 1 0,250 0,250 10,000 0,025 0,275

PCU 1 0,250 0,250 10,000 0,025 0,275

Harness 1,093 20 0,187 1,28

DH Cable 1 0,780 0,780 20,000 0,156 0,936

Power Harness 1 0,313 0,313 20,000 0,031 0,344

Payload 32,000 10,000 3,200 35,200

Radiometer 1 32,000 32,000 10,000 3,200 35,200

Table 4.13: Equipment list

4.9.6 Spacecraft characteristics and performances

In this section a final overview of spacecraft baseline and s/s technical specification are presented. The proposed spacecraft satisfies all the technical requirements. The spacecraft is lightweight and presents high reliability. The stowed S/C dimensions are 0,9x0,9x0,9 m and the structure is composed by aluminium. The payload dimensions are 0,7x0,7x0,13 m, is composed by 9 channels wide band radiometer. The spacecraft is 3 -axis stabilized using star trackers for attitude determination and momentum wheel for attitude control, with magnetic torquers to desaturate the wheel (magnetometers measures the earth’s magnetic field). The spacecraft propulsion system is a monopropellant hydrazine system. Approximately a total of 200 m/s is required to correct the launch dispersion, to maintain the orbit over the lifetime of the mission. This is achieved using eight 6-N thrusters. The spacecraft is powered by a 1,3 m2 two-wing, fixed GaAs triple junction solar array and a 0,81 m2 “body – mounted” on bottom panel sized to accommodate the maximum duration of data taking. Eclipse operation is ensured by a 105 Ah Li-Ion battery, sized by the higher charge/discharge rate. Data is processed in a commercial CPU, and stored in a RAM of 16 Gb capacity. The data handling architecture is integrated for telemetry data, command data and AOCS sensors data. Telemetry, commands and science data are handled by an S-band telecom system with redundant transponder. The thermal subsystem uses passive elements: multilayer insulation (MLI), radiator, and temperature sensors.

25

Mission

Operational Orbit Description Flower Orbit

Proposed Launch Date 15-ago-2015

Launch Vehicle VEGA

Design Life Time 2,00 yrs

Total Mission Time 2 yrs

Orbit Injection type Hohmann Manoeuvre

Orbit Type

LEO elliptical orbit (Flower)

Nominal Orbit Altitude 600 ± 20 Km

Orbit Inclination 63,4 deg

Eccentricity 0,02

Ground Station for Operations Fucino

Spacecraft Configuration

Heritage no

Spacecraft Dimension 1 0,9 m



Max Dim. With deployed SAW 3,5 m

Structure material aluminium

Stabilisation Stabilisation Type 3-axis

Pointing Accuracy 0,5 deg

Actuation system Momentum Wheel

Wheel desaturation Magnetic torquers

Actuation system SAFE Momentum Wheel

Power Distribution type Unregulated

Bus Voltage 22-34 V

Solar Array Type GaAs MJ

Peak Power 110 W

SAW (EOL) average power 214.5 W

Battery type Li-Ion

Battery capacity 93.23 Wh

Total SAW Area 2.8323 m^2

Propulsion

Propulsion type Chemical

Propellant Hydrazine

Number of thrusters 8

Position Thrusters 2-X face ,2 X face; 4 corner

Data Handling Architecture Centralized

Data Storage 16 Gbit

DH CPU throughput 2,04 MIPS

Comms TT&C frequency uplink 2,05 GHz

Telemetry data rate 2 Mbps

Telemetry BER 10^-6

Ratio fup/fdown 1,02

Modulation type

BPSK

26

Thermal Thermal Control type Passive

Propulsion control type MLI (Passive)

Battery thermal control type MLI (Passive)

Table 4.14: S/S summary

Figure 4.7: FLORAD Spacecraft

27

5 Power Subsystem

5.1 Requirements and design drivers

The main design drivers for the Power sub-system are the following:

• Payload duty cycle of 100%

• Main bus must be unregulated with voltage between 22 V and 38 V

• Life Time : 2 years

• Maximum duration of eclipse is 35.25 minutes with 4782 eclipses each year

• Elliptic Low Earth Orbit with inclination of 63.4° and semi axis of 6970 Km

• For configuration reasons, dimensions of each sub-panel of Solar Array must satisfy this constraints: Length 0.9 m, Width 0.9 m

• Minimization of subsystem mass

• Payload, TT&C and DH subsystems has an associated power profile, generally characterized by three values:

o Peak power o Standby power o Duty Cycle

• Instead the other subsystems, that don’t need peak power, request power for operative modes in which they operate

• During launch phase battery supplies power for spacecraft essential functions. Initialization phase is characterized by Solar Array deployment while in Safe mode power subsystem must guarantee the safety of all subsystems. In normal mode solar array supply all subsystem and recharge the battery. In eclipse mode the battery must supply all power request.

• The worst case is considered: during eclipse battery must supply peak power of Payload, TT&C and DH and AOCS subsystems for the entire instruments duty cycle

5.2 Baseline design

5.2.1 Architecture

The main techniques to control power generated by Solar Array are Peak Power Tracking (PPT) and Direct Energy Transfer (DET). PPT is non dissipative system while DET dissipates power not used by loads through shunt resistor. Power Control Unit (PCU) provides both these choices and the possibility to regulate Battery charge. Power System feeds all other subsystems by way of a distribution unit. Figure 5.1 illustrates Electronic Power System (EPS) architecture. Each Solar Array Wing (SA/W) is deployed by a solar array mechanics.

28

Figure 5.1: Power Subsystem block diagram

5.2.2 Solar Array

The main solar array design drivers are the following:

• Solar array sizing has been designed considering power required by all subsystems and power necessary to charge battery. The result of this kind of analysis is that solar array must supply a mean power of 214 W

• The technology selected for Solar Array is second choice GaAs triple junction cells with an efficiency of 19 % at 28 °C, which is the maximum efficiency reachable with the constraint on solar cell (constraint about costs)

• Operational temperatures vary from -150°C to 110 °C

• Solar array design includes also a mechanism for deployment.

The resulting solar array area is 2.833 m2, with 1029 cells in every side. Solar array is characterized by an incident power per unit area of 260 W/m2 at BOL and 232 W/m2 at EOL

5.2.3 Battery

The main battery design drivers are the following:

• Battery has been dimensioned to generate the power required in eclipse mode

• The battery design have been based on Li-ion technology, with a DOD of 70% and an efficiency of 96%

• The required capacity is 93.24 Wh, with 29 shunt battery and 10 series battery

• Battery mass is 4.480 kg

• Operational temperature is 20 °C

29

5.3 Performances and budgets

A summary of Power sub-system performances is given below: Main bus characteristic Bus Type Unregulated Max MB Voltage 38V Min MB Voltage 22V Distribution concept Centralised PDU Power subsystem characteristic and units Power s/s Power s/s mass 8.145 Kg (10% mass margin) Harness mass 0.313 Kg (20% margin) Solar Array (SA) SA type Flat SA Technology GaAs (TJ) second choice Area 2.833 m2

Number of wings 2 Number of sub-panel per wing 2 Sub-panel dimensions 0.7 x 0.9 m2 Number of body-mounted panel 1 Body-mounted panel dimension 0.9 x 0.9 m2 BOL power 241.3 W EOL power 214.6 W SA mass 4.359 Kg Power Control Unit (PCU) SA regulator modules DET (Direct Energy Transfer) Battery charge BCR (Battery Charge Regulator) PCU mass 0.5 Kg (10 % margin) PCU dimensions 0,2 x 0,2 x 0,1 m Batteries (BTRs) BTR technology Li-Ion Number of BTRs 1 Nominal cell capacity 0.84 Ah BTR capacity 243 Ah Number of series cells 10 Number of parallel cells 29 DOD 70 % BTR dimensions 0,35 x 0,038 x 0,062 m BTR mass 4.480 Kg (10 % margin)

30

Power Distribution Unit (PDU) PDU mass 0.25 Kg (10 % margin) Power budget at the last iteration of CDF is summarized below:

• TT&C request 6 W in standby and 26 W when it transmits data (mean contact time with ground station of 7.55 minutes)

• Power requested by AOCS subsystem is 14.31 W

• Thermal Control needs 0 W ( Totally passive thermal control )

• Propulsion subsystem requests 20 W with a very small duty cycle

• DH subsystem requests 2.22 W

5.4 Other options

In this section two alternative solutions for the EPS are described. Both this options foresee the reduction of payload duty cycle.

The first alternative assumes the radiometer used only in daylight, independently on the world area covered. Moreover DH and TT&C duty cycles are scheduled during illumination time. This choice leads to a smaller solar array area and battery dimensions (with the same technological choices even if battery performances are not completely exploited).

The second solution assumes that radiometer works only on Mediterranean area, independently on time passage, either in daylight or in eclipse. In this solution is considered that the TT&C and DH subsystem work with the payload. Obviously the worst case must be evaluated, i.e. when the satellite collect and download data during eclipse. This case implies a smaller a solar array area and battery dimension with respect to the selected solution, but a smaller solar array area and a bigger battery dimension with respect to first alternative. This second alternative leads to a reduction of solar array such as to permit body-mounted ones.

Table 5.1 shows a comparison between the selected solution and the two alternatives exposed above (each case is analysed with the same technological choices assumed previously).

Selected

solution

First

alternative

Second

alternative

Power supplied by Solar array [W] 214.5 117.22 90.11

Solar array area [m2] 2.823 2.569 2.336

Solar array mass [kg] 3.625 3.288 2.990

Battery dimension [cm] 35x62x110,2 35x62x38 35x62x60.8

Battery mass [kg] 4.480 1.600 2.560

Table 5.1: Comparison between different solutions for EPS design. Data are compared for the same technological choices (second choice GaAs TJ cells and Li-Ion battery).

31

5.5 Conclusions At the end of CDF iterations, design choices satisfied all requirements and

constraints. The use of second choice GaAs TJ solar cells and Li-ion battery guarantee minimization of subsystem mass and the fulfilment of the worst power request. Power subsystem accommodation is shown in Figure 5.2.

Figure 5.2: Power subsystem accommodation in FLORAD.

32

6 AOCS Subsystem

This section presents a preliminary AOCS architecture. The assumptions made for the preliminary sizing calculations are given together with their justification.

6.1 Requirements and design drivers

Design drivers for the FLORAD mission are listed below:

• Elliptical orbit (inclination = 63,4 deg, e= 0,02 )

• Two years life

• Pointing accuracy = 0,5 deg

• Fixed nominal nadir pointing

• Launcher : Vega

• Altitude ≈ 600 ± 20 km

6.2 Control modes specifications

Initialization: Vega leaves the S/C with a tip-off rate. In particular Vega ([8]) leaves

a three-axis stabilized satellite with following accuracy:

• Longitudinal axis depointing = 1 deg; • Trasversal axis depointing = 1,5 deg; • Angular tip-off rates along longitudinal axis = 0,6 deg/s; • Angular tip-off rates along trasversal axis = 1 deg/s.

For our analisys the worst case has been studied.

De-spin manoeuvre can use both trusters and magnetometers. Using magnetometers we can save propellant and they usually work also if it’s impossible to use trusters (for example because of some problems in communicating with ground during de-spin manoeuvres). The disadvantage is that the de-spin manoeuvre takes a longer time respect to the use of trusters.

Nominal: We consider the Weather acquisition mode as the nominal mode. Eclipse: Also during Eclipse periods, weather acquisition should be provided. Safe: The Satellite should be oriented toward the sun.

33

6.3 S/C features

S/C properties are provided below

Vers. S/C

mass

Ixx Iyy Izz CMx CMy CMz Worst

Mom.

Arm

FLORAD ~138 19,83 18,24 11,59 0,008 0 0,034 0,45

Table 6.1

6.4 Assumptions and trade-offs

In this preliminary study many simplifying and conservative assumptions are used.

All internal perturbation torques are neglected and the whole S/C is considered as a rigid body; all this torques can be compensated using a closed-loop control law. A margin factor is considered in the computation of the perturbation torques and momentum, to put internal torques into account.

Four sources of external perturbations are analyzed: atmospheric drag, solar radiation pressure, gravity gradient and earth magnetic field acting on the S/C residual dipole. The modules of the perturbation torques are taken into consideration, disregarding their direction. The total external torque is evaluated as a sum of scalar values. Each perturbation is calculated considering the worst case. To size reaction wheels, an estimation of the angular momentum to be stored is necessary; torques produced by drag, solar radiation pressure and gravity gradient are considered, as secular perturbations producing a growing with time momentum, whereas earth magnetic field produces a cyclic perturbation and the resulting angular momentum is null over one orbit period.

To evaluate the required momentum capability and torque authority a margin of 20% is applied.

The AOCS hardware sizing leads to two important trade-offs. The first one concerns with the number of orbits between each momentum damping. If the wheel desaturation is frequent, smaller wheel will be required but a much more time is needed to desaturate with magnetic torquers. So if the momentum dumping is actuated by magnetic torquers is better a frequent desaturation, while if thrusters are used, larger wheels will be preferred and desaturation manoeuvres will be infrequent.

Another important trade-off is the selection of the desaturation actuator. If magnetic torquers are used instead of thrusters, the whole subsystem will require a larger amount of power but a lot of fuel will be saved up.

6.5 Baseline design

6.5.1 Reaction wheels sizing

The consideration and requirements presented in the previous paragraph lead to select the simpliest control techniques.

34

The choice is toward a momentum bias system. It has just one wheel with its spin axis mounted along the pitch axis, normal to the orbit plane. The wheel runs at a nearly constant high speed to provide gyroscopic stiffness to the vehicle. Around the pitch axis, however, the spacecraft can control attitude by torquing the wheel, slightly increasing or decreasing its speed. Periodically, the pitch wheel must be desaturated (brought back to its nominal speed), as in zero momentum system, using trusters or magnets. So for the AOCS control of the FLORAD satellites, it has been chosen a single pitch wheel for momentum and magnetic torquers for momentum dumping and roll and yaw control. These torquers use magnetic coils to generate magnetic dipole moments. A magnetic torquer produces torque proportional (and perpendicular) to the Earth’s varying magnetic field. As said, they are used to desaturate the momentum wheel and to counteract secular perturbation for controlling the yaw and roll axis. Indeed for a momentum bias system, we typically design the stored angular momentum, determined by the inertia of the spinning body, to be large enough to keep the cyclic motion within the pointing specification without active control during an orbit. The more angular momentum in the body, the more resistant it is to external torques. The stored angular momentum depends on the orbit period and the accuracy requested in one orbit for Roll and Yaw (see Theory AOCS report for further information). With this technique the cost is limited and the requested accuracy is achieved (usually this control can assure an accuracy between 0,1 and 0,5 degrees)

6.5.2 Desaturation

To save up fuel, the desaturation process is accomplished using three redundant magnetic torquers. The torque produced by each coil is estimated by calculating the worst case magnetic field, i.e. at the highest altitude within the orbit box. The required torque authority is estimated to be the same as that of the wheel.

6.5.3 AOCS sensors

For FLORAD mission, the following sensors have been selected. For the 0,5 degrees Earth-relative pointing requirement, horizon sensors are the most

obvious choice since they directly measure two axis we need to control. Also a star sensor has been used to initially acquire vehicle attitude from unknown

orientation, for coarse attitude data and for fine data for Yaw. Two magnetometers (plus the redundant ones) are also adopted for attitude

measurement. To determine the position of the satellite has been used a GPS sensor. It’s true that

there are some studies ([8]) that show how it’s possible to use Earth sensors and Star tracker to calculate the satellite’s orbit, but these algorithms are under investigation and nowadays many improvements have been made in the realization of GPS sensors. So high accuracy and reliability are achieved using light and not expensive (in terms of costs and power) GPS sensors.

35

6.6 AOCS equipments In Table 6.1 all the components of the AOCS subsystem are presented.

Class Company Model Number Mass

(kg)

Power

(W)

Whell Surrey MicroWheel-

10SP-S 1 1,1 5

M.Torquer Zarm Technik MT2-1 6 0,3 0,77

Magnetometer Honeywell HMC6352 4 0,0014 0,052

Star Tracker Aero Astro MST 2 0,3 1

Earth Sensor Goodrich MMS 13 410 2 0,8 3

GPS Surrey SGR-05 2 0,02 0,8

Table 6.2: AOCS Equipment for the first version

The total mass and peak power (including a safety margin of 10%) for the AOCS system are

respectively 5,65 kg and 13,43 W.

6.7 AOCS data rate

A preliminary estimation of the AOCS data rate is provided. Attitude sensor

software handle data from instruments and produces internal variables. Earth sensors, and magnetometers involves decoding and calibrating sensed data. The selected star trackers are capable to produce S/C quaternions. Attitude determination and control require several computing functions; in this analysis these functions are supposed to be performed every 0,25 sec (4 Hz). Star trackers are supposed to transmit data with a delay of 1 second, whereas all the other sensors are supposed to work at a frequency of 4 Hz.

Attitude Sensor

Processing Frequency (Hz) words/frame bits/word Bit Rate (Kbit/s)

Earth Sensor 4 32 16 2048

Magnetometer 4 3 16 192

Star Tracker 1 4 32 128

Table 6.3 : Sensors bit rate

Errore. L'origine riferimento non è stata trovata. presents the

estimated bit rate for each sensor. Errore. L'origine riferimento non è stata trovata. provides an estimation of the required functions and their operational frequency. Using error determination, it is possible to find how far the spacecraft’s orientation and position is from that desired.

Determination & Control Frequency (Hz)

Kinematic Integration 4

36

Error Determination 4

Magnetic Control 4

Reaction Wheel Control 4

Ephemeridi propagation 4

Table 6.4 :AOCS Functions

6.8 AOCS mass and power budgets

In this preliminary analysis an estimation of the AOCS budget is produced and margin factors are taken into consideration. In Errore. L'origine riferimento non è stata

trovata. preliminary results are presented. Mass and power budget are evaluated for both nominal and safe modes. A 10% margin factor is applied.

FLORAD

Mass Budget (kg) 5,14

Power Budget-Nominal (W) 12,21

Power Budget-Safe (W) 12,21

Mass Margin (%) 10

Power Margin (%) 10

Table 6.5: AOCS Budgets

6.9 Conclusions The architecture of the AOCS for the FLORAD mission has been presented. It has

been shown that the AOCS units employed for the design of the AOCS satisfy the requirements and leave ample margins.

Further analysis of the entire AOCS should be performed to consolidate the preliminary architecture proposed in this study. Of particular importance are initialisation and acquisition manoeuvres.

Figure 6.1:AOCS subsystem accommodation in FLORAD

37

7 Propulsion Subsystem

7.1 Requirements and design drivers Target of this chapter is to provide the main propulsion subsystem characteristics for

the FLORAD mission as regards the FC strategy, where is necessary a propulsion subsystem. The propulsion subsystem is designed to meet the mission requirements of VEGA launcher scenarios.

The purpose of the FLORAD propulsion subsystem is to provide adequate forces during the mission lifetime (2 years) and to complete the following manoeuvres:

• To correct launcher dispersions

• Orbit maintenance manœuvres

The propulsion system shall be as simple as possible and consume as little propellant as possible; besides the propulsion system used shall be able to re-orbit and correct the inclination of the spacecraft.

In terms of performance, Ion thrusters have far better specific impulse compared to chemical propulsion systems. The complexity of the system suggest the selection of a monopropellant/bipropellant system. Cold gas systems have the lowest performance.

Ion thrusters requires a specific study and correction strategy. A monopropellant solution seems to be feasible. The bipropellant system is not convenient if compared to the monopropellant one.

7.2 Assumptions and trade-offs According to the discussion in section 6.1, a monopropellant propulsion system

could be selected for the mission. In any case, a trade-off between an ion propulsion system and a monopropellant one shows that the latter seems more suitable compared to the first one. Several characteristics of the two systems has been marked with values ranging from 1 to 5 where 5 is the best score (see Table 6.1). Marking of the propulsion features of the two different systems are discussed below.

The performance of an ion engine is superior compared to a monopropellant system. The specific impulse for a typical monopropellant system using hydrazine gas as propellant is between 200 – 230 seconds depending on the thruster design, temperature and duty cycle, while the specific impulse for a typical ion engine is ~3500 seconds.

The maximum thrust of the ion engine, QuinetiQ T5, is approximately 20 mN while a typical monopropellant system can provide a thrust between 15 mN and 20 N.

In terms of system complexity it is clear that the ion propulsion system is much more complex compared to a simple monopropellant system. Therefore, the monopropellant marking is superior compared to the marking for the ion engine.

Maturity level: The technology readiness level for monopropellant system is believed to be 8 or 9. Therefore, the marking is equal to 5. The technology readiness level for ion engines is a little lower.

38

The dry mass of a propulsion system comprising 2 ion engines is higher compared to the total dry mass of a equivalent monopropellant system. But the total mass, comprising propellant, is not so different. Therefore, the marking of the ion propulsion system is lower compared to monopropellant system.

The monopropellant system is superior to the ion engine in terms of required power. The monopropellant system requires only a few watts. However, an ion engine of type T5 requires 600W.

Monopropellant Ion engine (T5)

Specific impulse 2 5

Thrust level 3 1

System complexity 5 1

Maturity level 5 3

Mass 3 4

Power demand 5 1

Total 23 15

Table 7.1

7.3 Baseline design

The propulsive subsystem planned for FLORAD is a monopropellant system using hydrazine (N2H4) in blow down mode.

The propellant is stored in a single tanks, of 0.054 3m . The propellant is supplied to

thruster branches, constituted by two thrusters each. Each of the thruster branches can be isolated by a Latch Valve.

The system is designed to work in blow down mode: this means that the pressure will vary from the MEOP (22 bar) down to 5.5 bar. Consequently the supplied thrust level of the single thruster varies from 6 N down to 1.85 N. The system can operate in Steady State Mode as well as in Pulse Modulation.

The subsystem shall be loaded up to latching valves, because the liquid lines must be completely wet to avoid the undesired phenomenon of the detonation due to the hydrazine adiabatic compression. The Subsystem architecture is shown in Figure 6.1.

The tank is loaded by separated liquid lines. Also the pressurising gas is loaded by means of two independent lines: this is necessary to keep the ullage volumes separated, avoiding the propellant migration during launch. In the tank will be stored about 40 Kg of propellant. The propellant can be evacuated from the tank via the Fill and Drain Valve. The Fill and Drain Valve has an internal soft seal and an external double seal included in the valve cap. Totally, three independent seals are provided to satisfy the safety requirement.

Each thrusters pair is located in the opposite corners of the panel of satellite and the thrust developed is heading perpendicularly to the panel plane; such a way to dispose the thrusters is necessary to have symmetric forces to realize Homann and station keeping manoeuvres. Each branch has a latch valve that will be kept closed during ground operations and during Pre-Launch and Launch phase, in order to have three mechanical barriers from the tank to the thruster outlet. The Latch Valve has a reverse relief capability to prevent over pressurization of the downstream lines due to temperature increase.

39

The propellant is fed to the thruster through a filter with a filtration capability of 15 micron, which prevents the components to be contaminated by particles contained in the liquid.

Figure 7.1

The pressure inside the tank is monitored by one Pressure Transducer (P), while a pressure transducer monitors the pressure in the thruster branches. A thermocouple (temperature transducer) is mounted on each thruster decomposition chamber, to monitor the catalyst bed temperature and the status of the thruster at the same time. Thermistors (temperature transducer) are instead placed on the tank to monitor the temperature during loading operations and in orbit for gauging reasons. A thermistor will be used to reduce the thermal drift on the tank pressure transducer reading, improving the accuracy of the final pressure value.

Each of the eight branches enables the subsystem to perform all the manoeuvres required by the mission requirement. The thrusters are the CHT 0.5, manufactured by EADS space. Hydrazine must be prevented from freezing; the freezing point is about 2°C. If freezing should occur, the hydrazine shrinks. Line rupture will occur during defrost if liquid fills the volume behind frozen hydrazine and is trapped. The solution is to provide heather and thermostats on the lines, tank and thrust chamber valves. Catalyst bed are also heated to increase performance and bed life.

The propulsion system provides the following telemetry that can be used for failures identification:

• temperature transducer

• downstream Pressure Transducer

• Latch Valve switching status

40

In Table 7.3 are described the equipments:

FUNCTION NUMBER

Thruster 8

Catalyst Bed Heaters Internally redundant

Temperature transducer � One on each thruster chamber � Two on the Propellant Tank � Total of 10 thermistors on pipeworks,…

Heaters All heaters are internally redundant

Fill/Drain Valves 3 independent seals

Down-stream 3 independent seals (1 seal of valve 3, 2 seals for each Thruster Control Valve)

Pressure Transducer 1 in the tank and 1 after each filter.

Table 7.2

7.4 Equipments

In the following there is a description of the single equipment selected for the FLORAD propulsion subsystem in the FC strategy.

7.4.1 Propellant Tanks

The Propellant Tank proposed is EADS OST 31/0 tank. Propellant Management Device tanks are used in monopropellant systems for the control of fluid and separation of the pressurant gas from the fuel to provide gas-free propellant to the thrusters through the spacecraft life. The tank is a spherical pressure vessel; it comprises two hemispherical shells linked by a cylindrical centre section and is joined by identical equatorial tungsten-inert-gas (TIG) girth welds. The shell material is forged titanium. The tank is assembled to the spacecraft by means of three out of four trunnions equally spaced around the circumference of the centre section.

7.4.2 Thruster

The thrusters used for FLORAD spacecraft are CHT 0.5 thruster units manufactured by EADS Space; they were designed and tested for blow-down applications for a max inlet pressure of 22 bar.

The thruster consists mainly of two parts:

• one Thrust Chamber Assembly (TCA)

• one single-coil, dual seat Thruster Control Valve (TCV) The TCA head end consists of a structural support acting as thermal barrier between

thrust chamber and TCV, propellant feed tube and injector head plate. The TCA

41

contains also four redundant cartridge heater elements for the catalyst bed (to prevent propellant freezing) and a thermocouple, that indicates the thruster temperature condition prior and during firing.

7.4.3 Latching Valve (LV)

The Latch Valve (LV) is an electrically pulse-actuated device which is used to control the propellant flow into the thruster branches. When commanded to open, the valve opens and remains open; when commanded to close, the valve closes and remains closed. The Latch Valves are utilised to isolate the thruster branches in case of failure, then a very low number of switch-over cycles are foreseen. The latch valves are characterised by a back relief capability to avoid overpressurization due to temperature increasing. They are equipped with a position indicator (micro-switch) to enable monitoring of Latch Valve status (open or closed).

7.4.4 Liquid Filter (LF)

Liquid filters are required up-stream of the thruster control valves in order to protect the valve seats from fine contaminants. They employ filtration elements which consist of a stack of chemically etched titanium alloy disks.

7.4.5 Fill and Drain Valve (FDV)

The Fill and Drain Valve (FDV) used for propellant loading and draining is characterised by an internal soft seal and an external double seal provided by the valve cap. In total, the FDV fulfils the safety requirement of three independent seals.

7.4.6 Pressure Transducer (PT)

The low pressure transducers are strain gauge analogue devices. They use a titanium diaphragm to sense pressure and a thin film strain gauge bridge network to monitor the deflection of the diaphragm. The units are hermetically sealed (all welded design).

7.4.7 Feeding lines (Pipework)

The propulsion system lines are manufactured from titanium alloy. The tubing are of 1/4 inch diameter tube. In order to increase reliability by reducing leakage risks the pipework is entirely welded with only exception of the thrusters final stretches, both sides.

7.5 Budgets This section outlines the various budgets for the FLORAD propulsion subsystem in

the FC strategy. The budgets presented are the ∆V budget, propellant budget, dry mass budget.

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7.5.1 ∆V budget

The following ∆V budget has been considered for the FLORAD. Table 6.4 shows a summary of the manoeuvres required by the propulsion subsystems.

∆V for launcher dispersion 12 m/s

∆V for inclination control 9,4 m/s

∆V for de-spin 18 m/s

∆V for station keeping 0,5 m/s

∆V for manoeuver 688 m/s

∆∆∆∆V total budget : 727 m/s

Table 7.3

7.5.2 Propellant budget

The propellant budget is presented in this section.. The major part of hydrazine gas required comes from the Homann manoeuvres (92%), while the remaining represents the necessary propellant for the orbit maintenance. In total, including a 5% propellant reserve, 37,4 kg of hydrazine are required to complete the mission.

7.5.3 Dry mass budget

The total dry mass for the system is 9,374 kg. The propellant tank dry mass adds up to 6,4 kg which is equivalent to 68% of the total, dry mass.

The propulsion system dry mass budget is shown in Table 5.5:

Unit Element 1 Unit Name Quantity Mass per quantity

excl. margin Margin

Total Mass incl. margin

1 Pressurant 1 0,0976 5 0,126

2 Propellant tank (OTS 31/0) 1 6,4000 5 6,720

3 Fill Drain Valves (Vacco MRS) 5 0,0180 5 0,095

4 Latching Valves (LEE VHS-M/P) 8 0,0040 5 0,034

5 Filters (Vacco LP filter) 4 0,0800 5 0,336

6 Lines and fittings 1 0,0200 5 0,021

7 Temp. Transducers (National LM335Z) 10 0,010 5 0,010

8 Pres. Transducers 5 0,0750 5 0,394

9 Thrusters (EADS CHT-0.5) 8 0,1950 5 1,638

SUBSYSTEM TOTAL 8,897 5,0 9,374

Table 7.4

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7.6 Conclusions

The design process showed in this chapter has been performed under the key constraints of design simplicity. The initial trade-off study is necessary to meet exactly the mission requirement. In different conditions the ion propulsion would be convenient. Further analysis of the subsystem should be performed to strengthen the preliminary design identified in this chapter.

Figure 7.2: Propulsion subsystem accommodation in FLORAD.

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8 Data Handling

8.1 Requirements and constraints

DH subsystem has the main task of managing AOCS Data Handling and possibly compressing Payload scientific data and transmit them.

Automated functions that are performed on the satellite and which DH has to manage are:

Handling of AOCS sensors Data and performing AOCS function Commanding of AOCS Actuators, SADAs and thruster Handling of telemetry and sensors data

Sensors and functions for AOCS are described in the table below:

Sensors Frequency (Hz) Data rate (bps)

Earth Sensor frequency 4 2048

Magnetometer frequency 4 192

Star Tracker frequency 1 128,00

Telemetry (Kbit) 1 0,0444

Command (Kbit) 1 1,6

Table 6.8.1

Function Frequency (Hz)

Error Determination frequency 4

Magnetic Control frequency 4

Reaction Wheel Control frequency 4

Ephemeris propagation 4

Table 6.8.2

Moreover data handling equipment will manage other units such as thermal and power components, however these are not considered in this work because the requirements for AOCS are preponderant.

Telemetry and telecommand data handling requirements are taken into account with the following requirements:

Telemetry Word Length (bit) 10

Telemetry frequency (Hz) 1/45

Telemetries number 200

Table 6.8.3

Command Word Length (bit) 16

Command frequency (Hz) 1

Commands number 100

Table 6.4

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Telemetries parameters are taken from [Fortescue, “Spacecraft system engineering”] while command parameters are taken from [Larson; Wertz; “Space Mission Analysis and Design”], and adapted to the characteristics of a small satellite. We had to had a 100% as margin of uncertain (as recommended By Wertz) and for on orbit spare because this is only the first part of the design and many functions are not well defined.

8.2 Data architecture design

The presence of only one payload makes preferable a centralized architecture for the presence of a low number of connections. So the better flexibility of ring architecture isn’t so favourable respect the reliability given by a centralized architecture. The advantage in terms of less weight due to smaller wiring harnesses isn’t so much to justify ring architecture. It is remarkable that Myriade small satellite has this kind of data architecture. In the figure below is illustrated the data architecture.

Figure 6.1

The core of the system is the box named SMU (Satellite Management Unit). It contains:

- Elaboration Units (both for AOCS and eventual payload data processing) - RAM memory boards - PROM memory units - Connections boards

TT&C Thermal

SMU

AOCS

Power

Payload Propulsion

46

It has been used a 1553 bus to connect payload interface equipment with the processor while with other subsystems a serial link. This is due to the fact that AOCS sensors and actuators don’t support a bus architecture.

8.3 Mass memory and throughput

The sizing of spacecraft computer analyzing primarily AOCS functions gave the following results:

Requirements Component Chosen

Code Memory – ROM (Kbit) 1542,4 2048

Data Memory – RAM (Kbit) 1059

Telemetry Memory – RAM (Kbit) 1600 16000000

Throughput (MIPS) 2,04 4

Table 6.5

The telemetry needs are calculated supposing that for six consecutive orbits earth station is not visible from the spacecraft, however the memory provided can support more orbits without earth station visibility and gives flexibility to TT&C designers to select among different downlink strategies. The component has a computational power of 4 MIPS quite greater than requirements, we accepted this choice because the chosen DSP has integrated 2048 Kbit of PROM memory sufficient for our data managing, so we don’t have to buy and insert a PROM memory apart with increase of power consumption. Instead using a DSP we could use a Field Gate Programmable Array (FPGA). This kind of device can develop more functions in parallel while DSP is programmed with a sequential stream of instruction to do. It could be an advantage to use an FPGA as processor because this device must be connected with a great number of connections. But FPGAs are preferred for easy and repetitive algorithms, while DSPs are used in more irregular algorithms. So in managing many signals, like those coming from TT&C and from the payload, and with much instruction to do, we need a great quantity of ROM. FPGA hasn’t this quantity of non-volatile memory on the contrary of DSP. For this reason DSP utilization is preferable respect FPGA.

For what concerns RAM we have selected a larger memory but, as it will be

explained successively, we will need for storing data acquisitions from payload. The box that contains spacecraft computer is connected to other equipment with

cables. We gave an approximate calculation of data harness in the table below.

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Number Average Length (m) Mass (Kg/m)

Sensors 22 0,60 0,02

Bus 1 1,00 0,5

Actuators 4 0,6 0,02

TL sensors 200 0,6 0,02

TOTAL 0,63

Table 6.6

Sensors number has been chosen giving a number of them for each subsystem. This distribution is showed in next table:

Number sensors

Actuators 4

Propulsion 2

Attitude control 7

Thermal control 5

Power 2

Table 6.7

To these sensors number we have added a 10% of margin.

8.4 Payload data processing and transmission

The task of the mission is the acquisition of scientific data, made by a radiometer, about the atmosphere of the Mediterranean area. The ground station is placed in Fucino center. For this cause the strategy chosen implicates the contemporary acquisition of scientific data and their downlink at each passage over the Mediterranean sea. No processing is done aboard because it could not permit an immediate downlink for limit of time. So it’s preferable downlink immediately all scientific data and store them in a RAM memory the data acquired in the last passage. In this way, in case of error in downlinking, it can be repeated in the successive passage over the ground station. Each acquisition has a total size of 100 Mbit, so for this application we need a RAM memory of about 200 Mbit to conserve all scientific data of the last two passage. However it’s better oversize the RAM memory. In fact the radiometer could acquire not only data of Mediterranean sea, but also of other earth places. In this way is possible storing many scientific data from other areas. This justifies the greater quantity of RAM memory chosen. However their low mass and low power consumption isn’t a problem. To spare weight and power the transmission of the data to the ground station is made by the TT&C subsystem. It can be done thanks to the low data rate to transmit scientific data acquired.

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8.5 Hardware architecture and equipment overview

The architecture used is the one described in “Data architecture design”, while the following table summarize the equipment list, and principal component properties.

Unit Unit Name # Mass (Kg) Mass

Margin (%) Power (W)

Features

1 DSP

RAD6000 1 2,2 10 2,1

MIPS: 4 RAM: 8 Mbit

ROM: 2,48 Mbit

2 Cable - 0,75 20 - Data harness

3 RAM

150X XPRESS 1 0,05 10 0,12 16Gb

Table 6.8

Units 1 and 3 are assembled in a single box: Spacecraft Management Unit (SMU).

8.6 Conclusions and remarks

The subsystem sized for data handling is designed to be integrated and reliable, besides it is based on real space compliant components.

In a deeper analysis it should be important to study the following subjects:

• Updating reference table taken from [Larson; Wertz]

• Analysis of bus throughput, bus technology and communication protocols

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Figure 6.8.1: Data Handling subsystem accommodation in FLORAD.

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9 TT&C Subsystem

9.1 Requirements and constraints

Basic requirement for TT&C subsystem is to provide a telecommunication link in S-band between FLORAD space and ground segments.

As regards design constraints it is required to be able to guarantee link availability with ground station even in emergency situations. Downlink design has to guarantee a BER for telemetry rate up to 10-6. Main driver in component choice is not to overcome system mass and power budgets.

A compendium of all inputs from System and Mission subsystems is reported in the following table.

MACRO INPUT INPUT VARIABLE NAME

VARIABLE VALUE

MEASURE UNIT

Minimum Voltage

Vmin 22 V

Voltage Maximum

Voltage Power Vmax 34 V

Mass Budget MTT&C 7,21 kg

System Budgets

Power Budget WTT&C 26 W

Bit rate (Data rate + Telemetry

bit rate) Bit Rate Rb 2000,00 Kb/sec

Telemetry BER Bit Error Rate BER 1,00E-06 adim

MACRO INPUT INPUT VARIABLE NAMES

VARIABLE VALUE MEASURE UNIT

Orbit Orbit

Height H 591,86 Km

Frequency TT&C up-

link Frequency

flin 2050 MHz

Table 9.1

9.2 Baseline design

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The communication subsystem consists of the following elements:

• Two Low Gain Antennas (LGA);

• RF Distribution Unit equipped with four microwaves switches and associated cabling;

• Two S-Band Standard transponders with coherency and ranging capabilities, including: – Diplexer – Transmitter – Receiver

The transmitters shall operate in cold stand-by and the receivers in hot stand-by.

9.3 Assumptions and trade-offs

9.3.1 Data Transmission

Data rate for uplink telecommand is estimated in 4 Kbit/sec. Telemetry data from

different satellite equipments are stored in the On Board Data Handling memory and downloaded when link communications with ground station are established. We made a rough estimate of telemetry generated on board. Also it has been chosen to transmit the scientific data by same TT&C subsystem (this involved a high bit rate for downlinking about 2Mbps).

9.3.2 Ground Station

Ground operations are supported by Fucino Station. According to the link margins,

we are able to select the antenna with the minimum allowable diameter in order to limit operative costs. It has been chosen the 11m antenna because this guarantees the communication link between the spacecraft and the ground station also in the first phase of the mission with the satellite height is 2000Km.

9.3.3 Satellite antennas

Link communication has to be available from any aspect angle. In order to obtain a

global coverage we have decided to use omni-directional LGA antennas, located on the two opposite satellite platforms (zenith/nadir).

52

9.4 Performances and budgets

Link budgets

Main terms of uplink and downlink budgets are summarized below. We have considered:

• An S-Band 11 meters antenna presents an EIRP of 67,50 dBW and a G/T of 22,9 dB/K at 5,5 degrees of elevation angle;

• Slant range is calculated for a circular orbit of 600 Km and at 5,5 degrees of elevation angle;

• Spacecraft transponder with 5 W RF power (performances refer to Alcatel TCR transponder, for other characteristics see “Hardware architecture and equipment overview” section);

• LGA antennas with semispherical coverage and -2,5 dB gain for +/- 95 deg around the boresight (performances refer to SAAB Ericsson Space, for other characteristics see “Hardware architecture and equipment overview” section);

• An RF Distribution Unit with 3,3 dB losses, including cables;

• Uplink carrier frequency has been fixed at 2050 MHz and a coherent turnaround ratio of 1,02 has been utilized;

• Uplink modulation (Telecommand Data Rate: 4Kbps) selected is NRZ-L/BPSK/PM with 0,7 rad pp modulation index;

• Downlink modulation (Telemetry Data Rate: 4Kbps) selected NRZ-L/BPSK/PM with 0,7 rad pp modulation index.

Ground Station Modulation Indices

EIRP G/S 73,50 dBW TELECOMMAND 0,70 rad pk

RANGING (RNG) 1,00 rad pk

Propagation Losses Carrier Recovery

FREQUENCY 2050,00 GHz CARRIER SUPPRESSION 3,42 dB

SLANT RANGE 2284,80 Km PLL-BDW 2Bl0 2000,00 Hz

TOTAL PROP. LOSS 165,86 dB IMPLEMENTATION LOSS 3,00 dB

POW. FLUX at S/C -34,67 dBm/m^2 REQ C/N in PLL BDW 25,00 dB

Spacecraft Receiver CARRIER MARGIN 32,17 dB

RX ANT GAIN -2,46 dBi Telecommand Recovery TOTAL CIRCUIT&CABLE LOSS 4,50 dB MODULATION LOSS 8,97 dB

S/C RX G/T -36,38 dB/K IMPLEMENTATION LOSS 1,50 dB

RX POWER -72,38 dBm BIT RATE 4,00 kbps

CAR ACQ THRSH -128,00 dBm REQ Eb/N0 12, dB

TC RX THRSH -118,00 dBm

TELECOMMAND MARGIN 38,11 dB

REQ RX POWER -118,00 dBm

RX POWER MARGIN 45,62 dB

RX S/N0 96,60 dBHz

Uplink

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S/C Transmitter Carrier Recovery

S/C TX POWER 6,99 dBW CARRIER SUPPRESSION 3,39 dB

TOTAL LOSS 3,32 dB PLL BANDWIDTH 2*Bl 3000,00 Hz

S/C TX ANT GAIN -2,46 dBi PLL BANDWIDTH 34,77 dBHz

EIRP S/C 1,21 dBW REQ LOOP S/N 25,00 dB

Propagation Losses CARRIER MARGIN 21,25 dB

FREQUENCY 2091,00 GHz Telemetry Recovery SLANT RANGE 2284,80 Km TLM MODULATION LOSS 4,88 dB

TOTAL PROP. LOSS 169,29 dB DEMODULATOR TECH LOSS 1,50 dB

POW. FLUX at G/S -106,96 dBm/m^2 BIT RATE 2000,00 kbps

Ground Station MODULATION TYPE DE – BPSK dB

RX G/T 23,90 dB/K REQ Eb/N0 12,30 dB

RX S/N0 84,42 dBHz

Modulation Indices TELEMETRY MARGIN 4,23 dB

TELEMETRY(TM) 0,70 rad pk Tone Recovery

RANGING (RNG) 0,70 rad pk TONE MODULATION LOSS 31,79 dB

IMPLEMENTATION LOSS 3,00 dB

REQ S(Tone)/N 25,00 dB

MAX REQ LOOP-BDW 578869,25 mHz

Table 9.2

9.4.1 Mass and power budgets

Mass and power budgets are shown in the section below. The chosen equipments guarantee best performances and are the most efficient available ones. A 10% maturity margin has been applied to all elements mass, according to ESA standard (fully developed product).

TT&C MASS BUDGET

Unit No. of Units

Unit Mass (kg)

Raw Mass (kg)

Maturity Margin to be applied (%)

Predicted Mass (kg)

S-Band Transponder 2 3,00 6,00 10 6,60

Radio Frequency Distribution Unit 1 0,04 0,04 10 0,04

S/S

S-Band Helix Low Gain Antenna 2 0,24 0,48 10 0,53

TOTAL MASS 7,17

TARGET MASS 7,21

DownLink

54

DIFF. 0,04

TT&C POWER BUDGET

Unit No. of Units

Unit Power (W)

Raw Power (W)

Maturity Margin to be applied (%)

Predicted Power (W)

S-Band Transponder 2 26,00 52,00 0 26,00

Radio Frequency Distribution Unit 1 0,00 0,00 0 0,00

S/S

S-Band Helix Low Gain Antenna 2 0,00 0,00 0 0,00

TOTAL POWER 26,00

TARGET POWER 26,00

DIFF. 0,

Table 9.3

9.5 Hardware architecture and equipment overview

9.5.1 Hardware architecture

In the uplink chain, command data received by the antenna are sent to the

transponder (XPNDA-B) after passing through the RFDU, and then are routed to the DHU units. In the downlink sequence telemetry data reach the transponder and after the opportune elaboration and, through the RFDU, are routed to one of the two antennas, which are usually activated by a telecommand switch. Subsystem architecture is finally shown.

D

T

RXPND

XPND

CMDS STATUS

MONITOR

TELECOMMAND

RFDU

ANT A

ANT B

ON BOARD

DATA

HANDLING

TELEMETRY

TELECOMMAND

TELEMETRY

Figure 9.1

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9.5.2 Antenna

Conical Helix LGA antennas are produced by SAAB Ericsson Space. The S-band

conical helix antenna operates between 2000 – 2150 MHz and 2200 – 2300 MHz. It is a circularly polarized antenna. The radiation performance is optimised for hemispherical coverage with a maximum radiation at 95°. A great number of satellites (both LEO and GEO) have mounted this antenna type. The antenna pattern and its main characteristics are reported respectively in Figure 7.2 and in Table 7.4.

Frequency range (MHZ)

TC 2000 – 2150 TM 2200 – 2300

Diameter 65 mm

Polarisation LHCP or RHCP Height 285 mm

Coverage 0°<θ<95° Temperature

range -145°C to +140°C

Gain G95° > -2.5 dBi Mass < 240 g

Axial ratio 3,3 dB Electrical I/F 1 port, SMA

female

Table 9.4

Figure 9.2

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9.5.3 Transponder

In Florad project the Alcatel Transponder (showed in figure 7.3) is used to establish the communication link ; the transponder characteristics are reported in the table 7.5:

Figure 9.2: Alcatel Transponder

Model Alcatel TRC

Receiver Frequency (MHz)

2025 to 2120

Carrier Acquisition Threshold (dBm)

-128,00

Tc. Threshold (dBm)

-118,00

Noise Figure (dB) 5,00

Tc Modulation Index (rad)

0,70

Rg. Mod. Index (rad)

0,5 to 1,5

Transmit Frequency

2200 to 2300

Coherent Turn-around Ratio

1,09

RF Power (W) 5,00

Mass (Kg) 3,00

Mass margin (%) 10,00

Dim Length (mm) 275,00

Dim Width (mm) 110,00

Dim Height (mm) 197,00

Op. max Temperature (°C)

61,00

Op. min. Temperature (°C)

-24,00

Pon (W) 26,00

Pstby (W) 6,00

Table 9.4: Characteristic of the Alcatel Transponder

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9.6 Conclusion and remarks A fully off-the-shelf equipped TT&C subsystem was designed. Thus the subsystem

is lightweight, low-cost, high reliable. Moreover all technical requirements are achieved completely and performances make it possible to manage a higher amount of TM/TC e scientific data than strictly requested by mission.

Figure 9.3: TT&C subsystem accommodation in FLORAD

.

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10 Thermal control subsystem

10.1 Thermal environment

Spacecraft thermal control is a process of energy management in which environmental heating plays a major role. The principal forms of environmental heating on Earth orbit are direct sunlight, sunlight reflected by Earth (albedo), and infrared (IR) energy emitted by Earth.

The overall thermal control of an orbiting satellite is usually achieved by balancing the energy emitted by the spacecraft as IR radiation against the energy dissipated by its internal electrical components plus the energy absorbed from the environment.

10.2 Thermal requirements and constraints

The task of the thermal control subsystem is to maintain all spacecraft and payload components and subsystems within their required temperature ranges during the mission. Two limits are frequently defined: operational limits that the component must remain within while operating and survival limits that the component must remain within at all times, even when not powered. Table 10.1 gives typical component temperature ranges for representative spacecraft components.

Operating Temperature Non Operating Temperature Subsystem

Tmin (°C) Tmax (°C) Tmin(°C) Tmax(°C)

Electronics -10 50 -20 60 Batteries 11 21 10 30

Hydrazine Tanks and Lines

15 40 5 50

Antennas -90 90 -110 110 Earth Sensors -40 40 / /

Sun Sensors -30 65 -30 70 Star Trackers -30 50 -40 60

Gyros -50 90 -60 100 Reaction wheels -25 60 -35 70

Table 10.1 :Spacecraft subsystems typical temperature requirements

The fundamental design drivers and constraints for the FLORAD spacecraft’s thermal control are:

• Elliptic Low Earth Orbit (altitude: 600 ± 20 Km);

• An inclination of 63.4 deg;

• A payload duty cycle of 100% for every orbit;

• Eclipse duration about thirty-five minutes for each orbit;

• A payload operative temperature respectively between 0 °C and 50 °C.

• A solar array maximum operative temperature of +110 °C, and minimum non operative temperature of –200 °C.

59

10.3 Thermal design assumptions

The following assumptions have been used in the design process:

• The platform’s components are supposed to be at the same temperature between 0 °C and 50 °C, in the nominal phase, and between –10 °C and 60 °C, in the safe phase;

• Only the external view factors between the platform and the solar array are taken into account: possible interaction between the internal units are neglected.

• The worst hot and cold case are considered in the nominal and safe phase for sizing the satellite thermal control;

• The satellite configuration is supposed to be a cube (size: 0.9 × 0.9 × 0.9 m3). The hydrazine propulsion system is prevented from freezing (2 °C) by a dedicated

Multilayer insulation blankets (MLI), whereas the batteries are maintained within their operative range of temperatures (usually between 0 °C and 20 °C) using MLI, radiators and doublers.

10.4 Baseline design

10.4.1 The platform

The FLORAD thermal design is based on passive thermal control techniques. The appropriate radiating area is designed for the maximum dissipation of the

spacecraft. MLI blankets (21 layers) cover all other external surfaces of the spacecraft wall

panels; the external side consists of Aluminized foil tape with 2 mil adhesive (2 mil), whereas the internal side consists of Y9630-3M aluminized Mylar (1 mil).

The usage of radiators with a little area (0.22 m2) allow to maintain the temperature within the limit.

The thermal analysis has demonstrated that the platform’s temperatures can be maintained within the limits specified in Table 10.1.

10.4.2 The Solar Array

Solar Arrays are thermally isolated from the spacecraft structure and are treated independently. Again the initial temperatures are calculated in the worst hot case and during the eclipse (when the solar arrays does not work) and result within the specified limits. However a thermal control based on the presence of coatings, radiators and heaters is regarded.

60

10.5 Trade-off

Many factors influence the design and development of the Thermal Control subsystem: mission constraints, mission objectives, and the physical design of a spacecraft determines the inputs and the outputs of the thermal system’s interface. Trade-off studies are conducted on improving temperatures at the expense of added weight , specialized hardware and heater power.

In the FLORAD mission semi-active, or active, thermal components are not taken in account for the mass and cost budget. However, louvers are not taken in account because they constitute a single point of failure.

10.6 Thermal equipment

In Tables 10.2 the platform and payload thermal equipment are enumerated: for each component class, the used component, the number of items, the principal propriety, the mass, the mass margin, the total area and eventually the requested power are presented.

0.22

-

Area

[m2]

0200.505

0.210.030.210.03

Y9360-3M aluminized

Mylar

MLI

(21 layers)

0.00270.00070.0030.0007Silverized

fused silicaoptical

0200.088

0.150.040.170.04

Aluminiumfoil tape

with 2 miladhesive

Radiator

Power

(W)

Mass

Margi

n (%)

Mas

s

(kg)

ααααBOLεεεεBOLααααEOLεεεεEOL

Used

Component

Compone

nt Class

0.22

-

Area

[m2]

0200.505

0.210.030.210.03

Y9360-3M aluminized

Mylar

MLI

(21 layers)

0.00270.00070.0030.0007Silverized

fused silicaoptical

0200.088

0.150.040.170.04

Aluminiumfoil tape

with 2 miladhesive

Radiator

Power

(W)

Mass

Margi

n (%)

Mas

s

(kg)

ααααBOLεεεεBOLααααEOLεεεεEOL

Used

Component

Compone

nt Class

Table 10.2 :Platform thermal equipment

10.7 Conclusions

The total preliminary and current mass budget for the FLORAD thermal control mass budget is provided in Table 10.3 below.

Nevertheless the mass and power budgets will follow the natural evolution of the project with some updating as soon as the various parts of the spacecraft and the operation modes will be frozen.

61

00200.5961

Safe mode

Nominal

and Eclipse

modes

Total TCS Power Budget

(W)Total TCS

Mass

Margin

(%)

Total

Current

TCS Mass

Budget (no

margin) (kg)

Total

Preliminary

TCS Mass

Budget (kg)

00200.5961

Safe mode

Nominal

and Eclipse

modes

Total TCS Power Budget

(W)Total TCS

Mass

Margin

(%)

Total

Current

TCS Mass

Budget (no

margin) (kg)

Total

Preliminary

TCS Mass

Budget (kg)

Table 10.3: Total thermal mass and power budgets

62

11 Configuration Analysis

11.1 General requirements and constraints Structures and Configuration are two closely connected subsystems. The choices

adopted for one subsystem influence the analysis of the other one. The main configuration and structure requirement is the accommodation of the

spacecraft in the chosen launcher. Design drivers for the configuration are:

• Available volume in the chosen fairing: - Baseline Vega long fairing (cylinder part d= 2.2 m, h=5.5 m) - Option launchers : Rockot, Long March 2C, DNEPR, Soyuz-S. All these

launchers have available dimensions larger than Vega. So the latter represents the stronger constraint for the stowed configuration.

• Single Launch

• Structural - mechanical requirements of the spacecraft during mission lifetime

• Deployment mechanisms

• Thermal requirement of the spacecraft elements

• Pointing direction and field of view of solar panels

• Pointing direction and field of view of SAR Antenna.

11.2 Configuration

11.2.1 Baseline design

The Vega launcher is chosen to be the baseline launcher for this mission. A type-long fairing of the launcher will be used to accommodate the spacecraft.

Reference System centred in the geometric centre of the spacecraft has:

• Z axis pointing to hearth (yaw);

• X axis in the orbital plane along flight direction (roll);

• Y axis in order to have an anticlockwise reference system (pitch).

11.2.2 Internal accommodation

The spacecraft can be divided into two modules: the service module (SVM) and the

payload module (PLM). The SVM provides all necessary services to the PLM. In the first one power, propulsion, some AOCS and TT&C subsystems units are

accommodated. The second one contains the payload and some AOCS units. In particular in the SVM there are:

• Battery, PCU, PDU, on the inner side of the bottom closure panels;

• A propellant tanks in the middle of the satellite bottom panel.

• Momentum wheels, mounted on the main panel;

63

• magnetometers, accommodated on panel; In the PLM there are:

• RF electronics subsystem and digital electronics subsystem (payload), mounted on the closure panel top –X;

• the harness connecting all the instruments units.

11.2.3 External accommodation

• Sun sensors are accommodated on top closure panels +Y;

• Star trackers, on the top closure panels –Y;

• Earth sensors, on the upper side of top closure panels +Y and –Y;

• 2 monopropellant thrusters are located in pairs approximately in the central position of panel –X and other 4 thrusters on top closure panels -Y;

• Antennas of the TT&C subsystem are accommodated on the top and bottom platform in order to allow zenith/nadir pointing;

• Solar panels (two wings, three panels per each wings, and another mounted on the bottom panel) are hinged on the edges of the top closure panels in order to point the Sun in the deployed configuration;

• Radiometer

• Radiators are on all closure panels except the panel pointing toward the Sun.

In the stowed configuration solar arrays are folded up on the closure panels. For this reason they increase one dimension of the cross section of 1 m. So the final stowed satellite has cross section of 1.1 x 0.9 m and is 0.9 m high.

In the deployed configuration solar panels open symmetrically by means of a deployment mechanism. In this way, the deployed configuration will be balanced by an inertial point of view. The centre of gravity in X and Y direction will be not influenced by the presence of these appendixes but will depend only by the internal arrangement of the elements.

Balancing masses can be accommodated on the bottom platform, if required. Stowed and deployed configurations are shown in Figures 11.1 and 11.2. Internal

arrangement of the elements is shown in Figures 11.3 and 11.4.

64

Figure 11.1: FLORAD stowed configuration

Figure 11.2: FLORAD deployed configuration

65

In Figure 11.3 a X-Y cross section of the spacecraft is shown.

Figure 11.3: X-Y cross section

Figure 11.4: Z – Y section

66

Figure 11.5: X – Z section

67

12 Structures

12.1 General requirements and constraints Structures and Configuration are two closely connected subsystems. The choices

adopted for one subsystem influence the analysis of the other one. The main configuration and structure requirement is the accommodation of the

spacecraft in the chosen launcher. Design drivers for the configuration are:

• Available volume in the chosen fairing: - Baseline Vega long fairing (cylinder part d= 2.2 m, h=5.5 m) - Option launchers : Rockot, Long March 2C, DNEPR, Soyuz-S. All these

launchers have available dimensions larger than Vega. So the latter represents the stronger constraint for the stowed configuration.

• Single Launch

• Structural - mechanical requirements of the spacecraft during mission lifetime

• Deployment mechanisms

• Thermal requirement of the spacecraft elements

• Pointing direction and field of view of solar panels

• Pointing direction and field of view of SAR Antenna.

12.1.1 Requirements and design drivers

The main requirements and drivers for the structural design of the spacecraft derive from the compatibility with the chosen launcher, i.e. the payload has to fit inside the fairing and it has to be compatible statically and dynamically with the structural characteristics of the launcher. The first requirement is the maximum value of the spacecraft diameter, which has to be less than 2.2 m to fit into the Vega fairing. The chosen value for the three axis is 90 cm. On the 2 closure panels along x axis and on the bottom platform the stowed solar arrays are accommodated. Then the static compatibility with the launcher is ensured by means of an adapter. Vega is equipped with the 937B adapter for multiple payload configuration with a dispenser based on the experience and technologies developed through Ariane4 e Ariane5 program (Sylda, Spelda, Speltra). It is a structure in the form of a truncated cone, with a diameter of 937 mm at the level of the spacecraft separation plane. It is attached to the reference plane by a bolted connector frame, and also provides for spacecraft separation.

The dynamic compatibility is reached when the spacecraft stiffness is sufficiently higher than the launcher one. The first two structural frequencies (axial and lateral) of the spacecraft therefore have to be higher than the launcher ones. For the Vega, these are 35 Hz and 15 Hz, respectively.

Moreover Vega has a high longitudinal load factor, about 6.2 g, which impose a structure with high resistance to the axial compression, besides the resistance to the

68

bending caused by lateral acceleration, 1.5 g. For these reason the design of the spacecraft must satisfy not only dynamic requirements but also strength and stability ones. So the spacecraft structure shall be sized under the most severe combination of loads that can be encountered at any given instant of flight assuming the lateral loads may act in any direction simultaneously with longitudinal loads. Frequency analysis, Von Mises criteria, stability analysis have been used to size any structural elements.

From a structural point of view, the main consideration is the design of the inner structure that support the load of the whole spacecraft and in particular the payload. The structural stability of these items is very important since an incorrect inclination of the antennas may jeopardize the mission.

12.1.2 Baseline design

The spacecraft shape is a cube with a square side of 0.9m of side.There are 4 closure panels which withstand to the axial and lateral loads. These items are made with sandwich panels manufactured with honeycomb core in Aluminum. So the skins withstand all the axial load while the core supply enough resistance to the lateral loads.

On the top and bottom, panels close the spacecraft to avoid space contamination and

to supply a mounting surface for the radiometer. These items are made with sandwich panels manufactured with aluminum Alloy. They have a core thickness larger than the other elements since their main load is longitudinal one. So not only the skins withstand the load but especially the core. Moreover the bottom platform has to have a bigger thickness because of the attachment to the adapter. The latter is mounted to the spacecraft by means bolted attachments which form a “load paths” to the top of the spacecraft for the required stiffness during launch.

Payload module and service module are divided by a platform which supply a thermal insulation between the two environments and provide lateral stiffness to the structure. It is manufactured with sandwich panel of Aluminum alloy.

Four panels close the payload module and the service module. Also these items are made with honeycomb sandwich panels in Aluminum alloy.

Honeycomb core withstands the load since the main stress comes from the lateral acceleration of the launcher which acts perpendicularly to the plane of the panels.

In addition to these elements it’s important to underline the requirements to gain good joints between the items in order to assure stability and prevent dangerous gaps. So bolted and welded joints add a strong contribute to the baseline design.

12.1.3 Assumptions and trade-offs

Preliminary design has been performed considering a Safety Factor equal 1.8. Moreover Margin of Safety has been fixed to 10% in order to supply an efficient

preliminary design. Drivers in the structural materials selection have been the strength capability, the

stress corrosion resistance and CTE values for those parts which must insure high mechanical stability. The used material is:

• 7075 Al-alloy

69

Sandwich panels with honeycomb core in Aluminum have been utilized and above-mentioned material has been used for the skins.

12.1.4 Mass budgets and baseline sizing

A software tool has been realized in order to allow a baseline design. This tool allow to size thickness of the items in order to warrant their resistance during launch phases and mission lifetime. Inputs of this procedure are dynamic and static characteristics of the launcher, main dimensions of the items, material chosen for manufacturing. Several iterative processes have been performed in order to arrive to a final sizing and consequently to estimate the mass budget of the structure subsystem. In the table below the items and the main characteristics are listed.

MASS BUDGET AND BASELINE SIZING

Item #

items Materia

l

Mass per each Item (kg)

Mass with

margin (kg)

Mass Margin (%)

Height (mm)

Length (mm)

Width (mm)

Thickness Skin or Bulk (mm)

Closure Panel 4 Al 3,58 3,7585 5 900 900 0 0,5

Platform (top) 1 Al 2,754 2,8917 5 0 900 900 0,250

Platform (main)

1 Al 2,754 2,8917 5 0 900 900 0,250

Platform (bottom)

1 Al 2,754 2,8917 5 0 900 900 0,250

Total Mass of Structure (kg) 22,57 23,7 5

Figure 12.1: Mass budget and sizing for baseline design

12.2 Conclusions For this configuration has been stated that the minimum mass value to guarantee

requirements of safety is 23,7 kg considering a mass margin of 5%. However it’s possible to perform a trade-off on the feasible kind of configurations

both internal and external ones: configuration with a cruciform assy structure. Moreover it’s possible to evaluate use of other kind of materials in order to maximize performance-weight ratio.

70

13 References

a. Small satellites CNES programme, Myrade Platform, Paper 2008

b. D. Pereira, Spacecraft design project, 2004 FINAL DESIGN REPORT, Carleton University

c. STPSat-1 Data Sheet, Aeroastro, 2008

d. SSTL MicroSat-70 Documentation, Surrey 2008

e. SSTL MicroSat-100 Documentation, Surrey 2008

f. MicroStar Satellite Platform Documentation, Orbital 2008-04-10

g. SA-200B Data Sheet, Spectrum Astro 2008

h. J.R.Wertz. Mission Geometry; Orbit and Constellation Design and Management, Space Technology Library, 2001 Microcosm, Inc.

i. J.Rumbaugh, I.Jacobs and G.Booch. The Unified Modeling Language Reference Manual, Addison--Wesley.

j. G.T.French. Understanding the GPS, GeoResearch, Inc., 1996.

k. http://www.nasa.gov/

l. J.R.Wertz. Spacecraft Attitude Determination and Control, Microcosm, Inc., Torrance 1978.

m. J.R.Wertz. Space Mission Analysis and Design, Microcosm, Inc., El Segundo 1999.

n. M. Guelman, R. Waller, A. Shiryaev, M. Psiaki.``Design and Testing of Magnetic Controllers for Satellites Stabilization''.

o. E. Perez, “Vega User’s Manual / Revision 0”, 2004

p. P. Fortesque, “Spacecraft System Engeneering”, Wiley

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q. E.M.Silverman. “Product Development of Engineered Thermal Composites for Cooling Spacecraft Electronics”.

r. R.D.Karam. Satellite Thermal Control for Systems Engineers, Edited by P. Zarchan, American Institute of Aeronautics and Astronautics, Massachusetts.

s. Agneni, On Board Power Generation, Master in Satellite and orbiting platforms, Scuola di Ingegneria Aerospaziale

t. L. Schirone, Space Solar Cells, Master in Satellite and orbiting platforms, Scuola di Ingegneria Aerospaziale

u. W.J.Larson and J.R.Wertz. Space Mission Analysis And Design, Space Technology Library, 2005 Kluwer Academic Publishers, London.

v. Fundamentals of Space Systems, Edited by V.L.Pisacane and R.C.Moore, 1994 Oxford University Press, New York Oxford.

w. Spacecraft Systems Engineering, Edited by P.Fortescue and J.Stark, 1995 John Wiley and Sons, Singapore.

x. Spacecraft Thermal Control Handbook, Edited by D.G.Gilmore, 2002 The Aerospace Press, El Segundo California.

Documents

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