CDF Technical Report

FFIINNGGEERRSSAATT PPrroojjeecctt

MasterSat 2005-2006

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Table of contents

1. MISSION OVERVIEW 1

1.1 Mission baseline 1

1.2 Launcher selection 1

1.3 Launch window 2

1.4 Launch date 2

1.5 Operational orbit 2

1.5.1 Ground Station coverage 2

1.5.2 Orbit injection 3

1.5.3 Station-keeping 3

1.5.4 End of Life De-orbit 3

1.6 Constellation design 4

2. SYSTEM CONCEPT 5

2.1 Objectives 5

2.2 System requirements 5

2.3 Design drivers 5

2.4 Mass and power budgets design 8

2.4.1 Payload design 8

2.4.2 Other subsystems design 9

2.4.3 Preliminary mass and power budgets 10

2.5 Spacecraft modes of operation 11

2.6 Spacecraft mechanical states 14

2.7 Margin philosophy 14

2.8 Redundancy philosophy 14

2.9 System summary 15

2.9.1 Mass evolution 15

2.9.2 Final mass budget 15

2.9.3 Power budget 17

2.9.4 Spacecraft functional diagram 18

2.9.5 Spacecraft equipments 18

2.9.6 Spacecraft characteristics and performances 20

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3. SAR PAYLOAD DESIGN 23

3.1 Requirements and constraints 23

3.2 Baseline design 23

3.3 Assumptions and trade-offs 24

3.4 Performances and budgets 25

3.5 Hardware architecture and equipment overview 27

3.6 Conclusions and remarks 30

3.7 References 31

4. POWER SUBSYSTEM 32

4.1 Requirements and design drivers 32

4.2 Baseline design 32

4.2.1 Architecture 32

4.2.2 Solar Array 33

4.2.3 Battery 33

4.3 Performances and budgets 34

4.4 Other options 35

4.5 Conclusions 36

5. AOCS SUBSYSTEM 37

5.1 Requirements and design drivers 37

5.2 Control modes specifications 37

5.3 S/C features 37

5.4 Assumptions and trade-offs 38

5.5 Baseline design 39

5.5.1 Reaction wheels sizing 39

5.5.2 Desaturation 40

5.5.3 AOCS sensors 40

5.6 AOCS equipments 40

5.6.1 CT-631 41

5.6.2 MT 140-2 41

5.6.3 FSS 42

5.7 AOCS data rate 42

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5.8 AOCS mass and power budgets 43

5.9 Conclusions 43

6. PROPULSION SUBSYSTEM 45

6.1 Requirements and design drivers 45

6.2 Assumptions and trade-offs 45

6.3 Baseline design 46

6.4 Equipments 49

6.4.1 Propellant Tanks 49

6.4.2 Thruster 49

6.4.3 Latching Valve (LV) 49

6.4.4 Liquid Filter (LF) 50

6.4.5 Fill and Drain Valve (FDV) 50

6.4.6 Pressure Transducer (PT) 50

6.4.7 Feeding lines (Pipework) 50

6.5 Budgets 50

6.5.1 ∆V budget 51

6.5.2 Propellant budget 51

6.5.3 Dry mass budget 51

6.6 Conclusions 52

7. DATA HANDLING 54

7.1 Requirements and constraints 54

7.2 Baseline design 55

7.3 Assumptions and trade-offs 57

7.4 Performances and budgets 59

7.5 Hardware architecture and equipment overview 62

7.6 Conclusions and remarks 63

8. TT&C SUBSYSTEM 65

8.1 Requirements and constraints 65

8.2 Baseline design 66

8.3 Assumptions and trade-offs 66

8.4 Performances and budgets 67

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8.5 Hardware architecture and equipment overview 69

8.6 Conclusion and remarks 71

9. THERMAL CONTROL SUBSYSTEM 73

9.1 Thermal environment 73

9.2 Thermal requirements and constraints 73

9.3 Thermal design assumptions 74

9.4 Baseline design 74

9.4.1 The platform 74

9.4.2 The payload 75

9.4.3 The Solar Array 75

9.5 Trade-off 76

9.6 Thermal equipment 76

9.7 Conclusions 77

10. STRUCTURES AND CONFIGURATION 78

10.1 General requirements and constraints 78

10.2 Configuration 78

10.2.1 Baseline design 78

10.3 Structure 83

10.3.1 Requirements and design drivers 83

10.3.2 Baseline design 83

10.3.3 Assumptions and trade-offs 84

10.3.4 Mass budgets and baseline sizing 85

10.4 Conclusions 85

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1. Mission Overview

The mission analysis activities focused on contributing input for the selection of an appropriate class of orbits for the FINGERSat mission via a comparative analysis of the respective merits of different options.

After comparative merit analysis, a controlled Sun Synchronous Orbit (SSO) was selected as baseline and a more in-depth analysis was performed taking into account:

• Launch window and launcher options

• Global coverage and revisit time

• Orbital evolution and stability, mission lifetime

• Correcting the injection dispersion and station keeping

• Ground station visibility conditions

• Eclipses

• Thermal and radiation environment

1.1 Mission baseline

A polar orbit is mostly suitable for a remote sensing mission. A particular class of polar orbits are the Sun Synchronous Orbits (SSO). They offer a stable thermal environment (Sun incidence on spacecraft and Earth radiation nearly constant) and allow for a simple solar panel design. This is particularly true for the dawn-dusk orbit, closely over-flying the Earth terminator. In addition, eclipse time is relatively short and there is only one eclipse season, in the southern emisphere. Such an orbit is proposed for FINGERSat.

The characteristics of the selected orbit are listed here:

• Circular at 619.14 km altitude above equatorial Earth radius (6378 km)

• Inclination: 97.87°

• Period: 97.1 minutes

• Maximum eclipse duration: 20.5 minutes (21 % of period)

• Length of eclipse period: 90 days (24.6 % of year)

1.2 Launcher selection

The requirements call for a low-cost launch vehicle able to inject up to 1000 kg into low, circular, polar orbit. Furthermore, to simplify the early orbit operations, an ability to deploy the satellite in a de-spinned mode, and with an elevate accuracy in semi major axis and inclination.

A number of ’small’ launch vehicles, potentially meeting the above requirements, is either available or in an advanced development stage. The following table shows a comparison of the characteristics, based on the information available to date.

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VEGA LM 2C ROCKOT DNEPR SOYUZ - S

H [mm] 3515 3605 3711 1880 2364

h [mm] 2000 2200 2554 2960 3193

D [mm] 2380 3000 2380 2700 3480

d [mm] 1080 1823 666 1000 345

Maximum P/L weight [kg] 1600 1750 1150 1200 4600

First Longitudinal Frequency [Hz] > 35 > 35 > 33 > 20 > 35

First Lateral Frequency [Hz] > 15 > 12 > 15 > 10 > 15

Max Longitudinal Dyn Acc. [g] -5 ,0 and + 3 ,0 ± 3 ,0 -8 ,1 and + 1 ,8 ± 1 ,0 ± 0 ,7

Max Dyn + Static Lateral Acc. [g] ± 1 ,5 ± 1 ,0 ± 0 ,9 ± 1 ,0 ± 1 ,8

Max Longitudinal Static Acc. [g] -5 ,5 -6 ,7 -7 ,2 -7 ,8 -4 ,3

OASPL (20 - 2828 Hz) [dB] < 142 < 140 < 137 ,9 < 140 < 140

Table 1.1

For the purposes of this preliminary design exercise, the VEGA laucher is supposed to be used. The choise above ensures that a design compatible with VEGA will be readily adapted to other launchers.

Moreover the SOYUZ-S option is taken into account for a multiple launch.

1.3 Launch window

The orbital geometry is such that a launch opportunity exists once per day throughout the year, either the launch time has to be timed such that injection takes place precisely into the dawn-dusk orbit choosen.

1.4 Launch date

August the 15th 2015 is defined to be the launch date, taking into account that VEGA is still in a develop phase and that a low solar activity phase is foreseen.

1.5 Operational orbit

1.5.1 Ground Station coverage

The most suitable ground station for a SSO is Svalbard. The coverage characteristics are summarised in the following table.

Minimum elevation 10 deg

Longest pass duration 530 s

Mean pass duration 452 s

Mean number of passes per day 11.4

Ground contact per day 86.3 min/day

Table 1.2

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1.5.2 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 7.6 m/s.

1.5.3 Station-keeping

At an altitude of 619 km, the atmospheric air density is between 7.90×10-15 and

5.10×10-12 kg/m3 depending on the solar activity. This translates into a force up to

3.20×10-4 N per square meter cross sectional area and it drives to about 5 m/s for mantaining the orbit for the entire mission.

Moreover for controlling orbit inclination about 24 m/s are necessary during mission lifetime.

1.5.4 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). To de-orbit the spacecraft for an immediate re-entry, a 135 m/s velocity decrement would be needed.

Figure 1.1

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1.6 Constellation design

Different configurations are analized to reduce global coverage time, using equi-spaced satellite in a singular plane or in equi-spaced planes. Time to cover the 20°-60° latitude zone is evaluated. The following table shows the results. It is remarkable that the 0,31% of Earth surface not covered is localized close to the geographical poles.

6 19 la t 2 0 °-6 0 °

n° planes n° s/c Coverage [%] Time Time

1 1 99 ,69 3 d 15 h 52m 2 d 8 h 52m

1 2 99 ,69 1 d 17 h 52m 1 d 14 h 12m

1 3 99 ,69 1 d 11 h 51m 1 d 11 h 51m

1 4 99 ,69 13 h 20m 12 h 20m

2 2 99 ,69 17 h 59m 17 h 10m

2 4 99 ,69 7 h 56m 6 h 30m

Glo ba l

Configura

tion

Table 1.3

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2. System Concept

2.1 Objectives

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

The objectives of this document are to:

• Describe FINGERSAT spacecraft design

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

2.2 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 FINGERSAT mission require:

• Payload: SAR (Synthetic Aperture Radar) , STRIPMAP mode, 3m ground resolution, >40Km Swath, Incidence Angle 20°-50°, X band

• Pointing Accuracy: 0,028°

• Altitude: 619Km

• Duty Cycle: 10%

• # Images/day: >50

• Launcher: VEGA

• Stabilization Type: 3-axis

• Bus Voltage: Unregulated 23V-38V

• PDHT: downlink data-rate > 155Mbps, X band

• OBDH: Integrated

• Propulsion type: Monopropellant

• TT&C: S band

• Operative LifeTime: 5 years

• Autonomy: 24 hours

2.3 Design drivers

In the following, the main system design driving requirements, resulting from the FINGERSAT 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: 1050 Kg, including adapter)

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

• system performances shall be maximised by use of new technologies

• subsystem reliability shall be maximised by use of technologies “space qualified”

• subsystem reliability shall be maximised by cross strapping of redundant units

• 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 1 619

Life time Duration (yrs) 1 5

Local hour ascending node 1 6 a.m.

Launch

Launch year 1 ago-13

Launcher 2 VEGA

Number of satellites 1 1

Wet mass 2 1050

Propulsion

Type of Propusion 1 Chemical Monopropellant

Specific Description 3 Hydrazine

No. and position of thrusters 2 2, -X face, 4 -Y face

Operations

No. Ground Station 1 1

Ground Station Operational 1 Svalbard

Configuration

Stabilisation 1 3-Axis

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

Payload Accommodation 2 external

Satellite Platform Shape 1 Box

AOCS

Desaturation Time 1 3orbit

Wheels Desaturation 2 Magnetic Torquers

Actuation System 1 Reaction Wheels

Power

Solar Array Technology 2 MJ GaAs

Battery type 1 Li-Ion

SEP SA Configuration 1 2 wings

SEP SA Movement 2 tiltable 1-axis

Data Handling

Data Retrieval Process 2 Store & dump

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Data Rate Philosophy 1 fixed

Comms TT&C

LGA Antenna Type 1 helix conical

Thermal & Mechanisms

Thermal Control 3 Active&Passive

Table 2.1: prelimimary baseline design

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

PRIMA TERRASAR

Mission

Altitude 619 <600 <500

Life time Duration (yrs) 5

Local hour Qscendine node

6 a.m. 12 a.m.

Launch

Launch year ago-13

Launcher Soyuz-S VEGA ROCKOT

Number of satellites 1 2 4

Wet mass 1000 1050 950

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 2, -X face

Operations

No. Ground Station 1 2

Ground Station Operational

Svalbard Thune McMurdo-

Antarctica

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 3orbit 6orbit 10 orbit

Wheels Desaturation Thrusters Magnetic Torquers

Actuation System Reaction Wheels Magnetic Torquers

Power

Solar Array Technology Silicon GaAs – Triple GaAs – Double

Battery type Li-Ion Nich-Hi Nich-Cadmium

SEP SA Configuration 2 wings 1 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

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LGA Antenna Type helix conical quadrifilar

Thermal & Mechanisms

Thermal Control Active Passive Active&Passive

Table 2.2: Sub-system options

2.4 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, in fact, 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 2.2.1: Mass and Power budget design scheme

2.4.1 Payload design

The payload mass is calculated according to performance requirements and using typical formulas.

The estimated length is calculated by:

)2

( azest

RL =

Where Raz is the required azimuth resolution. Similarly the estimated width:

MISSION REQ.

PAYLOAD DESIGN

MASS BUDGET POWER BUDGET

PAYLOAD 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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f

c

SW

hWest

)cos(ϑ=

Where h is the orbit altitude, SW is the minimum swath, θ is the off-nadir angle, c is

the light speed and f is the operation frequency. According to technical requirements the estimated antenna dimensions result:

Lest=6 m West = 0,97 m then the total area: A = 5,84 m2. Using a Terra-SAR technology for SAR antenna (about 48 Kg/m2) the estimated

antenna weight is about 280 kg. Assuming a 10% for SEI (SAR electrical interface), the total payload estimated mass is 310 Kg. For the selected technology the mean power consumption is 460 W/m2, then the total payload power budget is about 2690 W.

2.4.2 Other subsystems design

Considering the baseline S/S design previous presented in Table 2.2 and according to Tables 2.3-2.4 (presented below) of typical mass and power budget for radar observation satellite, the mass and power budget for each subsystem are generated.

Typical Mass Budgets for RADAR OBSERVATION

SATELLITE

Typ Mass (Kg)

Min Min % Max Max %

Payload SAR Payload (active antenna) 400,00 31,1 800,00 35,5

Data Handling & Transmission Payload

50,00 3,8 120,00 5,4

total payload 450,00 35,1 920,00 40,9

Platform Structure (inclusive of mechanism) 290,00 22,6 410,00 18,2

Electrical Power 250,00 19,4 380,00 16,8

Integrated Control (AOCS, DH) 80,00 6,2 120,00 5,3

TT&C 10,00 0,7 10,00 0,5

Propulsion 12,00 1,0 25,00 1,1

Thermal Control 45,00 3,5 60,00 2,6

total BUS 687,00 53,6 1005,00 44,66

Miscell Harness 55,00 4,3 85,00 3,7

Supports and Miscellanea 15,00 1,1 30,00 1,3

Balancing Masses 10,00 0,8 80,00 3,5

Total S/C Dry Mass 1217,00 94,9 2120,00 94,2

Propellant 65,00 5,1 130,00 5,8

Total S/C Launch Mass 1282,50 100,00 2250,50 100,00

Table 2.2: Typical Mass Budget for radar observation satellite

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Typical Power Budgets for RADAR OBSERVATION SATELLITE

Typ Power (W)

Min Max

Payload SAR Payload (active antenna) 3000,00 12000,00

Data Handling & Transmission Payload 200,00 550,00 total payload 3200,00 12550,00

Platform Structure

Electrical Power 80,00 150,00

Integrated Control (AOCS, DH, sensors) 200,00 450,00

TT&C 20,00 40,00

Propulsion 10,00 10,00

Thermal Control 50,00 350,00

total BUS 360,00 1000,00

Harness (ohmic losses) 80,00 550,00

Total S/C Power Consumption 3640,00 14100,00

Solar Array available power (typical range) 2800,00 4100,00

Power provided by battery 840,00 10000,00

Table 2.3: Typical peak power budget for radar observation satellite

2.4.3 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 2.1.

Estimated Mass Budget

(Kg) %

Payload SAR Payload (active antenna) 315,00 31,8

Data Handling & Transmission Payload

55,00 5,5

total payload 370,00 37,3

Platform Structure (incl. Mechanism) 190,00 19,1

Electrical Power 135,00 13,6

Integrated Control ( DH) 40,00 4,0

AOCS 40,00 4,0

TT&C 10,00 1,0

Propulsion 15,00 1,5

Thermal Control 40,00 4,0

total BUS 470,00 47,2

Miscell Harness 45,00 4,5

Supports and Miscellanea 20,00 2,0

Balancing Masses 10,00 1,0

Total S/C Dry Mass 915,00 92,0

Other 0,0000

Propellant 75,00 8,0

Total S/C Launch Mass 990,00 100,00

Table 2.5: Preliminary system mass budget

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Est. Power

(W)

Payload SAR Payload (active antenna) 2690

Data Handling & Transmission Payload 200

total payload 2890

Platform Structure (incl. Mech.) 0

Electrical Power 80

Integrated Control (DH) 80

AOCS 130

TT&C 30

Propulsion 10

Thermal Control 0,00

total BUS 330

Harness (ohmic losses) 80

Total S/C Power Consumption 3300

Solar Array available power (typical range) 2400

Power provided by battery 900

Table 2.6: Preliminary system power budget

2.5 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:: this mode is used until deployment of the solar array. Satellite capable of receiving and executing Telecommands (e.g. SA deployment).All sub-systems are off, except essential equipment (e.g.RX). 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 TT&C by LGA. Contingency situation possible

Nominal Mode

Nominal Earth Surface Observation: The spacecraft is kept SUN pointing. Accuracy determined by AOCS Acquisition of radar data with 10% orbit duty cicle. Storage and trasmission to ground RX stations of raw SAR data. Contingency Situation possible

Eclipse Mode

Nominal Earth Surface Observation: Switch Power generation from SA to battery. Accuracy determined by AOCS Collection of science data from all instruments. 100% use. Storage and transmission to ground RX stations of raw SAR data. Contingency Situation possible

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Safe Mode

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

Table 2.7: Modes of operation at system level

Figure 2.2: Modes of Operation transition diagram

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

LAUNCH

INITIALISATION

SAFE

NOMINAL ECLIPSE

Mode Transitions Diagram

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Mode Sub-system Requirements THERMAL Only essential operations, if any (battery heating,thermal control, tbd).

COMMS LGA RX on. Other functions switched off.

DATA HANDLING Functional CDMU capable of executing telecommands and elaborations

PDHT All functions - OFF

AOCS All functions - OFF

PROPULSION All functions - OFF

POWER Battery provides necessary power for S/S.

PYROTECHNICS All functions - OFF

1 LM

PAYLOAD All functions - OFF

THERMAL Maintain operational temperature of services and instruments.

COMMS TT&C via LGA.

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

PDHT All functions - OFF

AOCS All functions - ON

PROPULSION

POWER Switch from batteries to SA.

PYROTECHNICS Release of deployed mechanisms: SA, antenna and instruments(tbc).

2 IM

PAYLOAD STAND-BY MODE

THERMAL Maintain operational temperature of services and instruments.

COMMS TT&C via LGA.

DATA HANDLING All Functions- ON

PDHT Handling of science data, including storage and transmission

AOCS All functions - ON

PROPULSION Operational power to service module and instruments, not including SEP.

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

PYROTECHNICS

3 NOM

PAYLOAD All functions - ON Max Duty Cicle:10%

THERMAL Maintain operational (& non) temperature of services and instruments.

COMMS TT&C via LGA.

DATA HANDLING Handling of science data, including storage and transmission

PDHT All functions - ON

AOCS SUN SENSORS OFF.

PROPULSION Operational power to service module and instruments, not including SEP.

POWER Battery provide power for Payload and S/S.

PYROTECHNICS

4 EM

PAYLOAD All functions - ON Max Duty Cicle:10%

THERMAL Maintain operational (& non) temperature of services and instruments.

COMMS TT&C via LGA - Only essential functions

DATA HANDLING Error detection and re-configuration (autonomous, tbd) - Only essential functions

PDHT Instruments kept on standby or off.

AOCS Depending on Pointing Requirements (Trade-Off)

PROPULSION Supply to essential functions.

POWER Battery provide power for S/S.

PYROTECHNICS None

5 SM

PAYLOAD STAND-BY MODE

Table 2.8: Modes of operation at subsystem level

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2.6 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

-Antenna - PDHT -Instruments (SAR)

Fuel tanks full (Fuel Decreasing)

Table 2.9: Mechanical states of S/C

2.7 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.

2.8 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.

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2.9 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.

2.9.1 Mass evolution

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

926,0

962,0970,0

1008,0

1040,0

1030,7 1033,2

1050,0 1050,0 1050,0 1050,0 1050,0 1050,0 1050,0

1 2 3 4 5 6 7

Total Launch Mass

Target Launch Mass

Figure 2.2.2: Mass evolution of S/C during CDF sessions

2.9.2 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%.

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S/C Subsystems S/S Mass

(kg)

S/S Mass Margin

(%)

S/S Mass Margin

(Kg) S/S MASS with Margin

Structure 168,56 6,4 10,83 179,39 kg

Thermal Control 29,82 20,0 5,96 35,78 kg

Mechanisms 10,00 0,0 0,00 10,00 kg

Pyrotechnics 0,00 0,0 0,00 0,00 kg

Communications 6,98 5,0 0,35 7,33 kg

Data Handling 26,38 10,0 2,64 29,02 kg

PDHT 47,79 5,2 2,47 50,26 kg

AOCS 37,32 5,0 1,87 39,19 kg

Propulsion 14,64 6,5 0,96 15,59 kg

Power 105,20 8,2 8,63 113,83 kg

Harness 45,00 0,0 0,00 45,00 kg

Payload 297,00 12,0 35,70 332,70 kg

Supports & Miscellanea 20,00 0,0 0,00 20,00 kg

Balancing Mass 10,00 0,0 0,00 10,00 kg

Total S/C Dry Mass 818,68 69,40 888,08 kg

Propellant (main) 85,11 kg

Total S/C Wet Mass 973,19 kg

System Margin 10% --

S/C Wet Mass with Margin 1070,51 kg

Adapter Mass 60,00 kg

Total S/C Launch Mass 1130,51 kg

Table 2.10: S/C mass budget

Total Spacecraft Dry Mass Breakdown

AOCS

4%Power

13%

Harness

5%

Payload

38%

Structure

20%

Propulsion

2%

PDHT

6%

Communications

1%Data Handling

3%

Thermal Control

4%

Supports &

Miscellanea

2%

Balancing Mass

1%

Figure 2.4: S/C dry mass breakdown

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2.9.3 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.

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

Power (W)

Mode 1 Mode 2 Mode 3 Mode 4 Mode 5

LM IM CM NOM SM

THERMAL 0,00 0,00 0,00 0,00 173,04

COMMS 26,00 26,00 26,00 26,00 26,00

DATA HANDLING 35,00 35,00 35,00 35,00 35,00

PDHT 0,00 0,00 166,70 166,70 0,00

AOCS 0,00 135,80 135,80 135,80 53,80

PROPULSION 0,00 40,00 40,00 40,00 0,00

POWER 80,00 80,00 80,00 80,00 80,00

PYROTECHNICS 0,00 0,00 0,00 0,00 0,00

HARNESS 6,00 16,00 20,00 20,00 10,00

PAYLOAD 0,00 104,00 2548,00 2548,00 104,00

TOTAL POWER 147,00 436,80 3051,50 3051,50 481,84

Table 2.11: 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 --

COMMS 5,20

DATA HANDLING 35,00

PDHT 33,34

AOCS 135,80

PROPULSION 4,00

POWER 80,00

PYROTECHNICS 0,00

HARNESS 20,00

PAYLOAD 255,70

Total S/C Power 569,04

Table 2.12: Mean/Orbit S/C power

18

2.9.4 Spacecraft functional diagram

Figure 2.5: S/C Functional diagram

2.9.5 Spacecraft equipments

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

FUNCTIONAL SUBSYSTEM nr Mass per each unit

(Kg)

Total Mass (kg)

Margin (%) Margin (kg) Mass with

Margin

Structure

168,56 6,42 10,83 179,38

Closure Panel (Top or Main) 4 7,34 29,38 5,00 1,47 30,845

Closure Panel (Bottom) 4 4,90 19,58 5,00 0,98 20,563

Shear Panel (Top) 4 12,04 48,16 5,00 2,41 50,568

Shear Panel (Bottom) 4 8,03 32,11 5,00 1,61 33,713

Platform (top) 1 7,78 7,78 5,00 0,39 8,165

Platform (main) 1 7,78 7,78 5,00 0,39 8,165

Platform (bottom) 1 7,78 7,78 5,00 0,39 8,165

Joints 1 16,00 16,00 20,00 3,20 19,200

Thermal Control 30,75 17,64 5,43 36,18

MLI 1 26,01 26,01 20,00 5,20 31,21

Radiator 2 1,33 2,67 - 0,00 2,67

Louver 0 0,00 0,00 20,00 0,00 0,00

Heater 1 0,00 0,00125 20,00 0,00 0,00

Doubler 1 0,16 0,15746 20,00 0,03 0,19

Heat Pipe 17 0,11 1,91165 10,00 0,19 2,10

TCS

SMU DH

AOCS GYR

RW

FSS

MTR STR

ES

MGT GPS

PDHT

TX DSHA

TT&C

XPN

XPNSAW PCDU

BAT EPS P/L

DE

Antenna

RF

Data Lines TM Lines PROP PDHT Lines

Power Lines

19

Communications 6,98 5,00 0,35 7,33

Alcatel TRC 2 3,00 6,00 5,00 0,30 6,30

model 1 1 0,50 0,50 5,00 0,03 0,53

SAAB Helix Quadrifilar Antenna

2 0,24 0,48 5,00 0,02 0,50

Data Handling 26,38 10,00 2,64 29,02

SMU 1 10,00 10,00 10,00 1,00 11,00

Cable 1 16,38 16,38 10,00 1,64 18,02

AOCS 37,32 5,00 1,87 39,19

VF MR 4.0(RW) 4 2,6 10,40 5,00 0,52 10,92

MT140-2 3 5,3 15,90 5,00 0,80 16,70

T F M 100G2 2 0,1 0,20 5,00 0,01 0,21

CT-631 2 2,27 4,54 5,00 0,23 4,76

MMS 13 410 2 0,8 1,60 5,00 0,08 1,68

JO-FSS (2) 2 0,62 1,24 5,00 0,06 1,30

GG440 GNAT 6 0,141 0,85 5,00 0,04 0,89

VF GPS 1 2 1,3 2,60 5,00 0,13 2,73

PDHT 47,79 5,16 2,46 50,25

SMJ320C6701 1 1,50 1,50 10,00 0,15 1,65

SSR 1 40,00 40,00 5,00 2,00 42,00

Encryptor 1 3,30 3,30 5,00 0,17 3,47

R Compens 1 0,25 0,25 5,00 0,01 0,26

8PSK 2 1,37 2,74 5,00 0,14 2,88

Propulsion 14,64 6,54 0,96 15,59

Pressurant 1 0,12 0,12 5,00 0,01 0,13

Propellant tank (ATK 80304-1)

2 3,85 7,70 5,00 0,39 8,09

Fill Drain Valves (Vacco LFRS-HP)

4 0,11 0,45 5,00 0,02 0,47

Latching Valves (Vacco TML-LP)

5 0,32 1,60 5,00 0,08 1,68

Filters (Vacco StB-LP) 2 0,18 0,36 5,00 0,02 0,38

Lines and fittings 1 1,50 1,50 20,00 0,30 1,80

Temp. Transducers 16 0,05 0,80 5,00 0,04 0,84

Pres. Transducers (GP50 7201)

4 0,14 0,56 5,00 0,03 0,59

Thrusters (EADS CHT-5) 6 0,22 1,32 5,00 0,07 1,39

Fill Drain Valves (Vacco LFRS-LP)

2 0,11 0,23 5,00 0,01 0,24

Power 104,88 8,28 8,69 113,57

Solar Array 2 18,00 36,00 5,00 1,80 37,80

Battery 1 36,68 36,68 10,00 3,67 40,35

PDU 1 10,00 10,00 10,00 1,00 11,00

PCU 1 15,00 15,00 10,00 1,50 16,50

SADA 2 3,60 7,20 10,00 0,72 7,92

Harness 45,00 0,00 0,00 45,00

Power Harness 45,00 0,00 0,00 45,00

Payload 297,00 12,02 35,70 332,70

Instrument 1 MASTER-SAR 2006

RF Electronics 1 29,00 29,00 20,00 5,80 34,80

Digital Electronics 1 31,00 31,00 20,00 6,20 37,20

Antenna SAR 1 237,00 237,00 10,00 23,70 260,70

Table 2.13: Equipment list

20

2.9.6 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 1,62x1,55x3,67 m and the structure is composed by CRFP and aluminium. The payload dimensions are 5,6x1,0 m, is composed by 14 tiles which include 48 T/R modules. The SAR antenna guarantees 3m of ground resolution for dual polarization. The spacecraft is 3 -axis stabilized using star trackers for attitude determination and reaction wheels for attitude control, with magnetic torquers to desaturate the wheels (magnetometers measures the earth’s magnetic field). A GPS system is used for orbit determination (at this high altitude, with SAR antenna presenting its side to the velocity direction, there is no significant drag, the key attitude perturbation comes from the gravity gradient and 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 precisely over the lifetime of the mission and for the de-orbit phase. This is achieved using six 5-N thrusters. The spacecraft is powered by a 7.1 m2 two-wing, one-axis gimballed GaAs triple junction solar array 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 solid state recorded (SRR) of 220 Gbit capacity. The data handling architecture is integrated for telemetry data, command data and AOCS sensors data. Telemetry and commands are handled by an S-band telecom system with redundant transponder, while science data downlink requires the use of an X-band system, with redundant transmitter. The Thermal Subsystem uses a mixed philosophy. The passive elements of the thermal control system are multilayer insulation (MLI), OSR, heat pipes, thermal doubler, and temperature sensors. Electric heaters, both commanded and thermostatically controlled will be required for S/C eclipse and SAFE modes of operation.

Payload Summary

General characteristic

Frequency 9,65 GHz

Imaging Mode STRIPMAP

Ground Resolution 3,00 m

Minimum Ground Swath 40,00 Km Minimum Number of images per

day 50,00 A.U.

Max Observation Angle 50,00 deg

Polarization Dual Polarized

Technical Chracteristics

# Tiles 14,00 A.U.

# T/R modules 672,00 A.U.

Duty Cycle 0,10

PRF >3 <3,6 KHz

DC Peak Power 2550,00 W

Radiated Peak Power 4600,00 W

Table 2.14: Payload summary

21

Mission

Operational Orbit Description

Proposed Launch Date 15-ago-2015

Launch Vehicle VEGA

Launch Option (Type)

Design Life Time 5,00 yrs

Total Mission Time >5 yrs

Orbit Injection type Direct

Nominal Mission

Orbit Type SSO

Ground Track Repeat Cycle 16,00 day

Nominal Mean Orbit Altitude 619,42 Km

Orbit Inclination 97,87 deg

Revolution/Day 14,81 A.U.

Eccentricity 0,00 A.U.

LTAN 6,00 am

Ground Station for Operations Svalbard

Spacecraft Configuration

Heritage yes

Spacecraft Dimension 1 1,62 m

Spacecraft Dimension 2 1,55 m

Spacecraft Dimension 3 3,67 m

Max Dim. With deployed SAW 9,985 m

Central Structure type Cruciform assy

Structure material CRFP&aluminium

Stabilisation Stabilisation Type 3-axis

Pointing Accuracy 0,028 deg

Actuation system Reaction Wheels

Wheels desaturation Magnetic torquers

Actuation system SAFE Magnetic Torquers

Power Distribution type Unregulated

Bus Voltage 23-38 V

Solar Array Type AsGa MJ

Peak Power 3050 W

SAW (EOL) average power 1755 W

Battery type Li-Ion

Battery capacity 105 Ah

Total SAW Area 7,10 m^2

Propulsion

Propulsion type Chemical

Propellant Hydrazine

Number of thrusters 6

Position Thrusters 2-X face ,4-Y face

Data Handling Architecture Integrated

Autonomy 24 hours

Data Storage 220 Gbit

DH CPU throughput 14 MIPS

PDHT CPU throughput 1024 MOPS

Comms PDHT frequency 8,1 GHz

22

PDHT data rate 175,0 Mbps

PDHT antenna type Cross dipole

PDHT BER 10^-9

Modulation type 8-PSK

TT&C frequency uplink 2235,0 GHz

Telemetry data rate 16,0 Kbps

Telemetry BER 10^-6

Ratio fup/fdown 240/221

Modulation type PCM(NRZ-L)/BPSK/PM

Thermal Thermal Control type Mixed

SAR thermal control type OSR

Propulsion control type MLI

Battery thermal control type Passive

Table 2.15: S/S summary

23

3. SAR payload design

3.1 Requirements and constraints FINGERSat payload is an X-Band SAR for a LEO Hi-Res radar remote sensing

mission. In this section requirements and constraints on payload subsystem are presented. Specific mission requirements are:

• SAR imaging mode: stripmap;

• 9.65 GHz central radar frequency;

• Dual polarized configuration with four polarimetric operating modes: HH, VV, HV, VH;

• Spacecraft bus voltage is unregulated in the 23-38 V interval;

• Minimum required ground swath: 40 Km;

• Range of incidence angles: 20°-50°;

• Minimum ground resolution allowable in the 3-15 meters range;

• Minimum number of images per day: 50;

• SAR duty-cicle per orbit: ~10%;

• Sun-synchronous orbit, with height fixed by Mission CDF subsystem: 619.42 km. This choice is dictated by revisit time requirements.

An important driver for the overall instrument design is the maintenance of total payload mass around the 300 kg final target. Peak DC power required from power generation subsystem should be around 2.5 kW, and this influences the maximum radiated RF power and the SAR antenna configuration. According to the CDF design setup requirements and constraints are inputs from System and Mission subsystems.

3.2 Baseline design

Key component of FINGERSat SAR is an active phased array antenna composed by T/R modules, grouped into panels. Single modules are slotted waveguides assembled in a fully polarimetric dual-polarized configuration. Panel Control Electronics Unit for T/R modules configuration and Power Conditioning Unit are integrated on antenna panels.

The antenna is physically mounted on the upper platform of the satellite structure: its orientation is fixed at a nominal off-nadir pointing and electronic beam steering is used in order to change the observation angle. Two other units, each one equipped with its own power conditioning and distribution unit, are included inside the satellite structure and they are here briefly described:

• RF Electronics Subsystem (RFES), responsible for frequency synthesis, generation of the linear-FM (chirp) signal, up-conversion of the transmitted signal to the carrier frequency and successive down-conversion of the received echoes;

• Digital Electronics Subsystem (DES), responsible for sampling of the offset video signals, the I and Q analog components, generation of timing and

24

command signals for the RF and antenna control electronics, handling and processing of radar telemetries.

3.3 Assumptions and trade-offs

In this section the basic assumptions on payload design are reported. High resolution stripmap observations are obtained by means of single-look image acquisitions. The range of incidence angles is based on the inputs provided by mission subsystem, with 1 degree margins for the upper and lower values. The minimum required ground swath is satisfied with the addition of a 10% margin: it’s not possible to increase the swath at will, because it has a direct impact on the antenna elevation dimension (W) and on the possible PRF values, in particular on the superior constraint, that considerably decreases with increasing swath widths. The antenna width must be designed also taking in account the antenna length, constrained by the required azimuthal resolution. A margin approximately equal to 10% with respect to the theoretical dimension has been assumed, in order to realistically obtain the 3 meters ground resolution. The antenna length must be superior of a certain value because the total antenna area must observe an inferior constraint provided by timing and PRF design aspects. A 5.6 meters long SAR antenna seems to be a reasonable compromise with respect to the 6 meters maximum theoretical length, given by the SAR focused elaboration.

A 10% tolerance on ground resolution leads to 5 different chirp bands valid for the same number of incidence angles ranges. Lower tolerances lead to a higher number of different chirp bands with a consequent increased memory for the digital chirp generator. The overlap between adjacent swathes is set equal to 7 km: this value is about 15% of the swath width and leads to the need of designing 14 different observation geometries, in order to cover the entire access area.

10% margins are assumed for the inferior and superior constraints on the PRF values, with respect to the theoretical limits. The same margin can be considered as an over-sampling factor and it is applied for the computation of the maximum data-rate provided as input to the PDHT subsystem. The number of bits used to quantize the digitized analog components, I and Q, is set equal to 6: it seems to be a good compromise between quantization precision and data volumes sent to the SMU through the payload data dedicated bus.

The active antenna is composed by a series of T/R modules organized in panels: radiators and modules are designed in a dual-polarized configuration. Energy stored during the intervals between pulses is transmitted to ground by means of chirp pulses whose length is set equal to 12 µs. The chirp length has direct consequences on the peak power required to satisfy the SNR requirements and on the PRF selection: long pulse durations permit to transmit low peak powers, even if the PRF design becomes more and more complicated as echo reception can be compromised due to the reduction of RX windows, while short pulse durations simplify the timing aspects but unfortunately lead to high peak powers.

As regards SAR sensitivity requirements a -23 dB NEσ0 (Noise Equivalent Backscattering Coefficient) level at the highest incidence angle has been assumed: on the basis of existing SAR instruments performances it can be considered a reasonable value. Sensitivity, observation geometry, frequency, receiver performances and additional

25

path losses influence the computation of required peak power: a -2.5 dB propagation path loss is considered, as it corresponds to a double propagation path characterized by a rainfall intensity of about 25 mm/hr, at the maximum incidence angle [3]. Attenuation due to rain losses on principle can’t be completely neglected at X-band. Sensitivity obtained at other incidence angles is always better than -17 dB: ground swath width should always be maintained about 44 km, so a lower number of active T/R modules is involved at lower incidence angles in order to create larger angular beamwidths, and in this way it’s not possible to transmit the same peak power. It was decided not to consider a higher sensitivity value in order to avoid demanding an excessive DC power to the power subsystem.

SAR hardware design is necessarily based on analogy with existing technologies and consequent scaling and adaptation. From this point of view modularity of DESA panels is an optimal starting point for a SAR preliminary design [1]. A rough mass estimation is given in the mentioned report and there’s also a strong degree of confidence in future improvements, so 40 kg per unit area can be considered as a realistic estimation for the active SAR antenna mass. Obviously margins for mass estimations have to be considered: while a 10% margin is applied to the SAR antenna mass as it is just a scaling of an existing technology, 20% margins are assumed for the RF and Digital Electronics, as they are units to be developed, according to the ESA CDF design criteria [2].

3.4 Performances and budgets

Table 3.1 illustrates for each swath the geometrical characteristics, the technical choices and the sensitivity evaluation, according to the previously quoted criteria. A 7 km overlap between adjacent swaths has been assumed, obtaining a global access area of 507 km. It can be remarked the use of five different chirp bands, in order to guarantee the required resolution at all incidence angles with respect to the tolerance, and the choice of suitable PRF values to respect the limits imposed by ambiguity and ALE (Altitude Line Echo) [4]. PRF values have been chosen using validity windows represented on the (PRF, θ) plane. It’s worth mentioning that all assumptions are based on theoretical support and on

analogies with similar Earth observation missions equipped with SAR systems. Mass budget is one of the most important drivers for a coherent spacecraft design. In

fact a Phase-A analysis usually stops when after a certain number of design iterations the spacecraft total mass stabilizes on a target value. Power budget is another fundamental driver for this preliminary analysis, because results concerning this aspect affect the design of other subsystems like EPS (Electrical Power System) and TCS (Thermal Control System). Table 3.2 reports masses, dimensions and power consumptions for each payload element, also considering a right margin according to the technological choices. Total payload mass budget is reported in the last row: it is obtained using a global mass margin, computed as a weighted mean of all single element mass margins.

2

6

Ta

ble

3.1

: D

iscr

ete

set

of

ob

serv

ati

on

geo

met

ries

wit

h r

ela

tiv

e p

erfo

rma

nce

s.

Sw

ath

In

dex

θ

off

-nadir

(d

eg)

Min

imu

m

dis

tan

ce f

rom

Na

dir

(km

)

Ma

xim

um

dis

tan

ce f

rom

Na

dir

(km

)

Mea

n d

ista

nce

fro

m S

/C (

km

)

Ma

x r

ad

iate

d

po

wer

(W

)

Ch

irp

Ba

nd

(MH

z)

PR

F (

Hz)

N

Eσ

0 (

dB

)

1

19,00

172

216

652

2162

139,62

3110

-17,90

2

22,40

209

253

665

2235

139,62

3008

-17,40

3

25,62

246

290

680

2328

114,23

3450

-17,90

4

28,69

283

327

697

2442

114,23

2960

-17,80

5

31,61

320

364

716

2578

93,46

3050

-18,60

6

34,39

357

401

736

2735

93,46

3430

-18,70

7

37,02

394

438

757

2915

76,47

3135

-19,80

8

39,53

431

475

780

3118

76,47

3637

-20,00

9

41,89

468

512

804

3344

76,47

3180

-20,30

10

44,13

505

549

830

3594

76,47

3637

-20,70

11

46,24

542

586

856

3869

76,47

3190

-21,10

12

48,22

579

623

883

4168

62,57

3090

-22,40

13

50,09

616

660

911

4491

62,57

3375

-22,80

14

51,00

635

679

925

4666

62,57

3170

-23,10

27

Table 3.2: Payload mass and power budgets.

3.5 Hardware architecture and equipment overview

The SAR payload architecture is essentially composed by three main parts:

• a Digital Electronics subsystem (DES);

• a RF Electronics subsystem (RFES);

• an active phased array antenna. The Digital Electronics subsystem must control data acquisitions, in order to

guarantee synchronization and to send correct data to the Payload Data Handling subsystem (PDHT) for the necessary data processing, storage and transmission to the ground stations. This unit has a dedicated CPU whose principal task is to control and manage the instrument functions and data flow. DES core is the digital chirp generator that produces chirp signals to be transmitted: a NCO-based direct digital synthesizer generates multiple chirp waveforms, commanded by an FPGA with a dedicated PROM that contains the different waveforms samples. The payload data handling hardware consists of a high-speed 60 MHz A/D converter and a data memory buffer that interfaces with PDHT. For radar control, timing, and telemetry DES includes a central processor unit (CPU), a spacecraft interface module, a radar control and timing unit (CTU). Based upon the command words generated by the CPU, the CTU generates the timing signals necessary to control the radar, including the pulse repetition frequency (PRF) in accordance with the observation geometries, gain control and phase shift settings for the antenna T/R modules. This approach ensures a simple interface to the spacecraft and aids for the ground testing of the instrument. Dedicated buses to transfer data among several payload devices and to interface with other spacecraft subsystems have to be considered in the subsystem design. .

Item

Mass

estimation

(kg)

Mass

margin (%)

Peak DC

Power

Consumption

(W)

Stand-by DC

Power

Consumption

(W)

Length

(mm)

Width

(mm)

Thickness

(mm)

SAR Antenna 237 10 2348 40 5600 1056 150

RF

Electronics 29 20 120 14 180 260 80

Digital

Electronics 31 20 80 50 220 260 150

Total Mass

(kg)

Global

Mass

Margin (%)

Total Peak

DC Power

(W)

Total Stand-by

DC Power (W)

297 12.02 2548 104

Mass Budget with global mass margin: 332.6 kg

28

The RFES contains an USO (Ultra Stable Oscillator) to generate the payload timing reference. It performs up-conversion, filtering and amplification for the transmission of the chirp waveforms generated on board. The subsystem also provides amplification, down-conversion, and filtering of the received echoes.

The radar electronics will be housed in two separated VME chassis, one for the RF electronics and one for the digital electronics. Each subsystem has its own dedicated power distribution unit to convert the raw spacecraft voltages to the required DC voltages and to condition and to distribute them to the various equipments. Full block redundancy of the radar electronics is implemented to achieve the required five-year mission life-time.

The antenna configuration is shown in Figure 3.1, while in Figure 3.2 a schematic diagram of SAR architecture is shown. For the assumed active phased-array architecture the transmission power is generated using 672 T/R modules distributed in 14 panels. In this configuration we assume roughly 7.2 W radiated power per module [5], in order to achieve the 4838 kW maximum transmission power. T/R modules are connected to the radiators, designed as dual-polarized slotted waveguides with low mass and high thermal stability, and to the RF distribution network. T/R modules are organized in columns and every column constitutes an antenna panel, 40 cm wide. A single panel is highlighted in yellow in Figure 3.1.

Use of many distributed T/R X-Band modules on the antenna provides inherent redundancy since random failures of the T/R X-Band modules result in a graceful degradation of radar performance. Every panel has its own Panel Control Electronics, commanded by timing signals coming from the digital electronics, two redundant Panel Power Conditioners and two RF distribution networks mounted on the cooler plate. The front-end electronics control the signal routing of the primary, redundant, and test/calibration signals. Gain control provides high dynamic range. The receive mode of RFD down-converts the echoes received from the antenna. These signals are then routed to the data handling system for digitalization and storage. The antenna performs the beam steering and transmission function as well as high-power amplification on transmit mode and low-noise amplification on receive mode. 22% transmission efficiency and 80% reception efficiency for every T/R module have been adopted, because they seem to be reasonable values on the basis of existing literature about the topic, while total noise factor for the receiver section is assumed equal to 4 dB, [1,5,6].

1.06 m

5.6 m

Antenna Panel

Figure 3.3.1: SAR Antenna configuration.

2

9

F

igu

re 3

.2:

SA

R a

rch

itec

ture

sch

ema

tic

dia

gra

m.

30

The antenna width (1,056 m) and length (5,6 m) are such that the complete spacecraft structure in stowed configuration can be accommodated in several existing launch vehicles that we have thought to use, for example the future small launcher VEGA. When stowed, the antenna is folded into 3 parts: the central one, measuring 1.2 m x 1.056 m, is directly fixed on the upper platform of the structure, while the other two, each measuring 2.2 m x 1.056 m, are tightly anchored to the lateral panels of the satellite.

3.6 Conclusions and remarks

A preliminary design for a stripmap X-band synthetic aperture radar has been presented. Mass budget, power budget, power losses in different operating modes, timing aspects and observation geometries have been retrieved. A conceptual schematic of the payload has also been shown in order to better explain the conceived architecture. From the satellite system point of view the design outputs are coherent and can give the green light to next developments. It’s worth mentioning that adaptation and scaling of existing technologies naturally leads to a rough design, to be accurately verified and subsequently improved. SAR system accommodation in FINGERSAT is shown in Figure 3.3.

Figure 3.2: SAR system accommodation in FINGERSAT.

31

3.7 References

[1] H. Ebner, X-band SAR technology at Astrium GmbH, technical report. [2] GESS (Global Earthquake Satellite System) Report, JPL, march 2003. [3] F. T. Ulaby, R. K. Moore, A. K. Fung, Microwave Remote Sensing, Active and

Passive, Volume I: Microwave Remote Sensing Fundamentals and Radiometry, 1981, Addison-Wesley Publishing Company

[4] G. Picardi, Elaborazione del segnale radar, metodologie ed applicazioni, 2000, FrancoAngeli

[5] C. Heer, J. Link, A potential German Contribution to the LightSAR Program, Geoscience and Remote Sensing Symposium Proceedings, 1998, IGARSS '98, IEEE International, pp. 279-281

[6] M. Gibbons, A. Stewart, M. Notter, Active Antenna Sub-System Development for Space-borne SAR, 12th International Conference on Antennas and Propagation, ICAP 2003, pp. 233-236

32

4. Power Subsystem

4.1 Requirements and design drivers

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

• Main bus must be unregulated with voltage between 23 V and 38 V

• Life Time : 5 years

• Maximum duration of eclipse is 20 minutes with 1267 eclipses each year

• Sun-synchronous orbit with inclination of 97.87° and semi axis of 6997,56 Km

• For configuration reasons, dimensions of each sub-panel of Solar Array must satisfy this constraints: Length 2.7 m, Width 1.2 m

• Minimization of subsystem mass

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

• Peak power

• Standby power

• 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 until it is possible; if the total request of power exceeds Solar Array possibilities than surplus of power is supplied by 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 PDHT Subsystems for the entire instruments duty cycle

4.2 Baseline design

4.2.1 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 4.1 illustrates Electronic Power System (EPS) architecture. Each Solar Array Wing (SA/W) is driven by Solar Array Drive Assembly (SADA). SADA rotate SA/W making normal of solar cells surface to coincide with solar light beam, in order to obtain energy as more as possible.

33

Figure 4.1: Power Subsystem block diagram

4.2.2 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 1680 W

• The technology selected for Solar Array is GaAs triple junction cells with an efficiency of 25.1 % at 28 °C, which is the minimum efficiency required to support the mission requirements;

• The resulting solar array area is 7.1 m2, with 121 cells in parallel and 17 in series

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

• Solar array is characterized by an incident power per unit area of 291 W/m2 at BOL and 247 W/m2 at EOL

• Solar array design includes also a mechanism for deployment and maneuvers. Each wing is driven by a SADA which mass is 3.6 kg

4.2.3 Battery

The main battery design drivers are the following:

• Battery has been dimensioned to generate the power required in eclipse mode and to meet peak power demand

• The battery design have been based on Li-ion technology, with a DOD of 26% and an efficiency of 96%

• The required capacity is 150 Ah, with 3 string in parallel and 11 cells for each string.

• Battery mass is 36.68 kg

• Operational temperature is 20 °C

L

O

A

D

SA/W

SA/W

SADA

SADA

PCU

BATTERY

PDU

34

4.3 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 23V Distribution concept Centralised PDU Power subsystem characteristic and units Power s/s Power s/s mass 105,2 Kg (8,2% mass margin) Harness mass 32 Kg (20% margin) Solar Array (SA) SA type Flat SA Technology GaAs (TJ) Area 7,10 m2

Number of wings 2 Number of sub-panel per wing 4 Sub-panel dimensions 0,78 x 1,14 m BOL power 2066 W EOL power 1755 W SA mass 35,68 Kg SADA mass 3,6 Kg (10 % margin) SADA dimensions 0,14 x 0,14 x 0,23 m Power Control Unit (PCU) SA regulator modules PPT (Peak Power Tracking) Battery charge BCR (Battery Charge Regulator) PCU mass 15 Kg (10 % margin) PCU dimensions 0,60 x 0,32 x 0,36 m Batteries (BTRs) BTR technology Li-Ion Number of BTRs 1 Nominal cell capacity 38,6 Ah BTR capacity 105 Ah Number of series cells 11 Number of parallel cells 3 DOD 26 % BTR dimensions 0,25 x 0,32 x 0,32 m BTR mass 36,68 Kg (10 % margin)

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

• SAR (Synthetic Aperture Radar) request a peak power of 2548 W with a duty cycle of 10%. The keep-alive requested power is 104 W

• 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 135.8 W

• Thermal Control needs 173 W of power supply (only in safe mode )

• Propulsion subsystem requests 40 W with a very small duty cycle

• DH subsystem requests 35 W and reach a maximum of 201.7 W when PDHT is active (with a duty cycle of about 20% of orbit time)

4.4 Other options

In this section two alternative solutions for the EPS are described. The first one assumes an equal subdivision of SAR duty cycle in eclipse and

illumination orbit periods. Moreover PDHT and TT&C duty cycles are scheduled during illumination time. This choice leads to a worst solution because the Solar Array area becomes larger while Battery dimensions remain unchanged (with the same technological choices even if Battery performances are not completely exploited).

The second solution assumes that none of previous subsystems (SAR, PDHT and TT&C) demands peak power during eclipse time. This choice implies the minimization of Solar Array and Battery masses but doesn’t allow the same possibility of exploitation.

Table 4.1 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] 1680 1725 1416

Solar Array Area [m^2] 7.1 7.27 6

Solar Array Mass [Kg] 35.68 36.18 32.54

Battery Dimensions [m] 0.32 x 0.32 x 0.25 0.32 x 0.32 x 0.25 0.16 x 0.21x 0.25

Battery Mass [Kg] 37 37 12

Table 4.1: Comparison between different solutions for EPS design. Data are compared for the same technological choices (GaAs TJ cells and Li-Ion battery).

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

constraints. The use of 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 4.2.

Figure 4.1: Power subsystem accommodation in FINGERSAT.

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5. AOCS Subsystem

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

5.1 Requirements and design drivers

Design drivers for the FINGERSAT mission are listed below:

• Sunsynchronous orbit (inclination = 97,86 deg )

• Five years life

• Pointing accuracy = 0,028 deg

• 3-axis stabilized S/C

• Fixed nominal off-nadir pointing

• Launcher : Vega

• Altitude ≈ 619 km

5.2 Control modes specifications

Initialization: Vega leaves the S/C with a tip-off rate; de-spin manoeuvre can be

neglected in the sizing process due to the use of magnetic torquers for momentum dumping. For this reason the initialization mode is not considered in this preliminary analysis.

Safe Mode: During safe mode, the AOCS task is that of maintaining the spacecraft in a Sun-pointing attitude using a minimum of the on-board resources while ensuring power generation and ground communication. During this mode a coarse attitude control is accomplished using sun sensors, earth sensors and magnetometers, and the control law for the 3-axis stabilization is actuated by magnetic torquers. Reaction wheels are in standby and star trackers are switched off.

Nominal: We consider the SAR acquisition mode as nominal mode with the required pointing accuracy.

Eclipse mode: Same specifications as those of nominal mode. De-Orbiting: Same specifications as those of nominal mode.

5.3 S/C features

S/C properties are provided below. S/C mass properties:

• S/C mass = 970,737 kg

• Ixx = 1187 kg*m2

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• Iyy = 2297 kg*m2

• Izz = 1373 kg*m2 Center of mass coordinates:

• cmx = 0,011 m

• cmy = 0,001 m

• cmz = 0,372m Cross section effective area AD= 3,6 m2 Sun exposed surface As= 11,05 m2 S/C reflectance = 0,6 Solar array reflectance = 0,6 Worst moment arm = 0,6 m Z-axis off-nadir deviation = 39 deg

5.4 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, in a sunsynchronous orbit, 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.

Resulting perturbation torques values are presented in:

Gravity Gradient Torque (Nm) 0,00031750

Solar Radiation Torque (Nm) 0,00002999

Magnetic Field Torque (Nm) 0,00046467

Aerodynamic Torque (Nm) 0,00000156

Total Torque (Nm) 0,00081373

Secular Angular Momentum (Nms) 2,03345442

Cyclic Angular Momentum (Nms) 0,95689365

Table 5.1: External Perturbations

To estimate the magnetic torque, a 10 Am2 S/C residual dipole is considered. Due to the 35 deg off-nadir pointing requirement, the gravity gradient torque is of the same

39

order of magnitude as the magnetic torques. The others torques are of one or two orders of magnitude smaller.

Reaction wheels torque authority is estimated as three times the perturbation total torque. To evaluate the required momentum capability 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 wheels desaturation is frequent, smaller wheels will be required but a larger amount of fuel will be consumed. So if the momentum dumping is actuated by thrusters, 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.

5.5 Baseline design

5.5.1 Reaction wheels sizing

Reaction wheels are required to counteract external perturbation torques times the

margin coefficient. So the minimum required torque authority is 0,00244 Nm. To estimate the required momentum capability the sum of cyclic and secular

momentum, stored over one orbit period, is considered. The cyclic angular momentum is calculated averaging the magnetic torque over ½ orbital period. By applying a 20% margin the resulting required momentum capability is 3,5884 Nms .

Four reaction wheels are used. Three of them are oriented with their axis parallel with the S/C principal axis; the forth one is a redundant one.

The reaction wheels configuration is shown in Figure 5.0.1.

Figure 5.0.1: Reaction Wheels Configuration

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5.5.2 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 and at zero magnetic latitude. The required torque authority is estimated to be the same as that of the wheels (0,00244 Nm).

5.5.3 AOCS sensors

The required pointing accuracy is obtained using star trackers to calculate the S/C attitude with an error smaller than 0,004 deg. In safe mode sun sensors and earth sensors are used for a coarse attitude control. Magnetometers provide an accurate estimation of magnetic field. Six gyros realize a redundant system which evaluate the S/C angular rate. The GPS system provides the navigation parameters.

5.6 AOCS equipments

In Table 5.2 all the components of the AOCS subsystem are presented. For each unit

the number of elements, mass and required power are provided.

Class Model Number Mass (kg) Power (W) Company

Reaction Wheels VF MR 4.0 4 2,6 19 Valley Forge

Magnetic Torquers MT140-2 3 5,3 1,9 Zarm Technik

Magnetometer TFM 100G2 2 0,1 0,85 Billingsley

Star Tracker CT-631 2 2,26 9 Ball A&T

Earth Sensor MMS 13 410 2 0,8 3 Goodrich

Sun Sensor FSS 2 0,62 0,2 Jena−Optronik

Gyro GG440 6 0,141 2 Honeywell

GPS VF GPS 1 2 1,3 8 Valley Forge

Table 5.2: AOCS Equipment

Each element is managed by the Spacecraft Management Unit (SMU). Selected star trackers and reaction wheels are connected by a standard MIL 1553B BUS, whereas the other elements are linked to the SMU by a serial connection.

The AOCS hardware configuration is described in Figure 5.2.

41

Figure 5.2: AOCS Configuration

5.6.1 CT-631

The CT-631 star tracker was delivered for use on the Near Earth Astreroid Rendezvous (NEAR) program launched February 17, 1996. The CT-631 is a compact package including:

• Optics

• 512x512 pixel CCD detector

• Thermo-electric cooler

• Drive and readout electronics

• Digital processor

• Command and data interface

• Spacecraft power interface

• Mechanical interface to precision spacecraft mount

• Optical alignment surfaces The overall sensor length is 142,2 mm and the optics diameter is 134,6 mm. The

update frequency is 5 Hz while the acquisition time is 0,8 sec.

5.6.2 MT 140-2

This relatively simple device interacts with the Earth’s magnetic field and create control torque, which can be adjusted to a specified value. Combined with one or more

42

wheels, it provides all the control needed to maintain the spacecraft’s attitude, from low-Earth orbit up to geostationary orbit. Unlike thrusters, torquers do not use valuable consumables, are low power, and very reliable. Microcosm has teamed with German production technology institute ZARM to provide this high-performance device.

5.6.3 FSS

FSS is an internally redundant analogous sun sensor for the medium/high accuracy range, combining low cost with short lead−time. This sensor is suitable for small and communication satellites´ applications. It can be used for all types of missions requiring sun sensors of medium/high accuracy. Sensor features are:

• No. of measuring axes 2 (a,b)

• Detector Photo diode array

• Axis range: ± 64°

• Accuracy a−channel: ± 0.3° (3 sigma) ± 0.06° (1 sigma)

• Accuracy b−channel: ± 0.3° (3 sigma) ± 0.06° (1 s)

• Power consumption : 400 mW

• Analog supply voltage: ± (13.2 ± 0.5) V DC

• Output voltage range : ±5V

• Mass: 620 g • Dimensions: 160 x 145 x 56 mm

5.7 AOCS data rate

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

software handle data from instruments and produces internal variables. Processing for gyros, Sun and 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)

Rate Gyros 4 12 32 1536

Sun Sensors 4 8 16 512

Earth Sensor 4 32 16 2048

Magnetometer 4 3 16 192

Star Tracker 1 4 32 128

Table 5.3: Sensors bit rate

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Table 5.3 presents the estimated bit rate for each sensor. Table 5.4 provides an estimation of the required functions and their operational frequency. In kinematic integration current attitude is estimated by integrating sensed body rates using gyros. Using error determination, it is possible to find how far the spacecraft’s orientation is from that desired.

Determination & Control Frequency (Hz)

Kinematic Integration 4

Error Determination 4

Magnetic Control 4

Reaction Wheel Control 4

Ephemeris Propagation 4

Table 5.4: AOCS Functions

5.8 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 Table 5.5 preliminary results are presented. Mass and power budget are evaluated for both nominal and safe modes. A 5% margin factor is applied.

Mass Budget (kg) 37,322

Power Budget-Nominal (W) 135,8

Power Budget-Safe (W) 53,8

Mass Margin (%) 5

Power Margin (%) 5

Table 5.5: AOCS Budgets

5.9 Conclusions The architecture of the AOCS for the FINGERSAT mission is presented. It is 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. AOCS subsystem accommodation is shown in Figure 5.3.

44

Figure 5.3: AOCS subsystem accommodation in FINGERSAT.

45

6. Propulsion Subsystem

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

the FINGERSAT mission feasibility study performed in the CDF. The propulsion subsystem is designed to meet the mission requirements of VEGA launcher scenarios.

The purpose of the FINGERSAT propulsion subsystem is to provide adequate forces during the mission lifetime (5 years) and to complete the following manoeuvres:

• To correct launcher dispersions

• Orbit maintenance manoeuvres

• De-orbit manoeuvre Requirement 1: The propulsion system shall be as simple as possible and consume as little

propellant as possible. 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.

Requirement 2: The propulsion system used shall be able to re-orbit and correct the inclination of the spacecraft.

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.

Requirement 3: The propulsion system used shall be able to de-orbit the spacecraft. Cold gas is ruled out. Ion thrusters require a longer time to de-orbit. A

mono/bipropellant seems to be feasible.

6.2 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.

Specific impulse: 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.

Thrust level: The required thrust level to provide the adequate forces are between 15 mN and 20 N (defined by mission). The maximum thrust of the selected ion engine (QuinetiQ T5) is approximately 20 mN while a typical monopropellant system can provide the adequate forces in the defined thrust range.

46

System complexity: 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.

Mass: 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 (See Table 6.2). Therefore, the marking of the ion propulsion system is lower compared to monopropellant system.

Power demand: 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 6. 1

Monopropellant Ion engine (T5)

Dry mass 15.62 Kg 67.06 Kg

Propellant 85.11 Kg 10,89 Kg

Total Mass 100.73 Kg 77.95 Kg

Table 6. 2

6.3 Baseline design

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

The propellant is stored in two tanks, of 58.5 litres each. The propellant is supplied to two thruster branches, main and redundant, constituted by three 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 thrusters are qualified for a maximum single burst duration of 4500 s. The maximum number of pulses allowed is 44000. The system is designed to satisfy the following safety factors:

- Proof pressure = 1.5 * MEOP

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- Burst pressure = 2.0 * MEOP Once the tank is integrated into the system, the allowable proof pressure value,

using gas to pressurise, is 1.1 * MEOP = 26.4 bar. The subsystem shall be loaded up to valves 1, 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 two tanks are 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. 85 Kg of propellant will be stored in the two tanks. The propellant can be evacuated from the tanks 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.

The system has two identical branches, main and redundant, isolated by a bistable Latch Valves. Each branch has a latch valve (Valve1) 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 overpressurization of the downstream lines due to temperature increase.

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 6.1

The pressure inside the tank is monitored by one Pressure Transducer (P), while two more pressure transducers monitor the pressure in the thruster branches. A

48

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 each 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 two branches enables the subsystem to perform all the manoeuvres required by the mission requirement. The thrusters are the CHT 5, manufactured by EADS space, four mounted on the –Y closure panel, firing towards the -Y direction, and two mounted on the –X side, firing towards the -X direction .

Fault tolerance requirements dictate redundant branches, each to accommodate an engine failure. If the thrust chamber are placed in series, their dominant failure mode (leakage or failure open) can be tolerated. Thrust chamber valve failure closed can be tolerated because the engine are redundant. Because the two tanks are cross-strapped, a failure of either propellant system can be accommodated. The following table (Table 6.3) summarize the logic used to establish the system schematic in a failure modes and effects analysis (FMEA). Valves 1A and 1B are opened at the start of in space operations; therefore, they should be considered open as you read in Table 6.3. Failure types are considered only once; symmetrically failures are not tabulated.

FAILURE CORRECTIVE ACTION MISSION EFFECT

Thruster A failure-off Operate with thruster B None

Thruster A failure-on Close TCV 1A and TCV 2A None

TCV 1A fail-open Operate with TCV 2A None

TCV 1A fail-closed Operate with thruster B None

Leakage Branch A below valve 3A Close valve 3A None

Leakage Branch A above valve 3A Close valve 2 Reduced duration

Diaphragm rupture Close valve 2 and 3A Reduced duration

Valve 1A fail-closed Use branch B Reduced duration

Table 6. 3

FUNCTION REDUNDANCY

Thruster branches 2 fully redundant branches comprising 2 LV and 3 thrusters, each

Catalyst Bed Heaters Internally redundant

Temperature transducer � One on each thruster chamber (both A and B branches) � One on each valve (both A and B branches) � Two on each Propellant Tank � Total of 6 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 2 units in the down-stream

Table 6. 4

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

49

liquid fills the volume behind frozen hydrazine and is trapped. The solution is to provide heather and thermostats on the lines, tanks and thrust chamber valves. Catalyst bed are also heated to increase performance and bed life.

In Table 6.4 are described the equipments and their redundancies: The propulsion system provides the following telemetry that can be used for failures

identification:

• temperature transducer

• downstream Pressure Transducer

• Latch Valve switching status

6.4 Equipments

In the following there is a description of the single equipment selected for the FINGERSAT propulsion subsystem.

6.4.1 Propellant Tanks

The Propellant Tanks proposed are ATK PMD tanks. 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.

6.4.2 Thruster

The thrusters used for FINGERSAT spacecraft are 5 N thruster units manufactured by EADS Space; they were designed and tested for blow-down applications for a nominal inlet pressure range of 22 bar down to 5.5 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 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.

6.4.3 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

50

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).

6.4.4 Liquid Filter (LF)

Due to the high flow rates needed during the de-orbit manoeuvres (4 x 5N thrusters ON), low pressure drop liquid filters are used. 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.

6.4.5 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.

6.4.6 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).

6.4.7 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.

6.5 Budgets

This section outlines the various budgets for the FINGERSAT propulsion subsystems. The budgets presented are the ∆V budget, propellant budget, dry mass budget.

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6.5.1 ∆V budget

The following ∆V budget has been considered for the FINGERSAT. Table 6.5 shows a summary of the manoeuvres required by the propulsion subsystems.

∆V Budget

Launcher dispersion 7.60 m/s

Re-orbit 4.84 m/s

Inclination 24.7 m/s

De-orbit 151.29 m/s

Total 180.83 m/s

Table 6. 5

6.5.2 Propellant budget

The propellant budget is presented in this section.. The major part of hydrazine gas required comes from the de-orbit manoeuvres (75.2%), while the remaining represents the necessary propellant for the orbit maintenance. In total, including a 6% propellant reserve, 85.1 kg of hydrazine are required to complete the mission.

Propellant budget

Launcher dispersion 4.17 Kg

Re-orbit & Inclination 13.27 Kg

De-orbit 63.98 Kg

Reserve 0.81 Kg

Trapped & uncertainty 2.88 Kg

Total 85.11 Kg

Table 6. 6

6.5.3 Dry mass budget

The propulsion system dry mass budget is shown in Table 6.7. The total dry mass for the system is 15.6 kg. The propellant tanks dry mass adds up to 8.09 kg, which is equivalent to 52% of the total, dry mass.

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Unit Element 1 Unit Name Quantity Mass per quantity

excl. margin Margin

Total Mass incl. margin

1 Pressurant 1 0,120 5 0,13

2 Propellant tank (ATK 80304-1) 2 3,850 5 8,09

3 Fill Drain Valves (Vacco LFRS-HP) 2 0,113 5 0,24

4 Latching Valves (Vacco TML-LP) 5 0,320 5 1,68

5 Filters (Vacco StB-LP) 2 0,180 5 0,38

6 Lines and fittings 1 1,500 20 1,80

7 Temp. Transducers 21 0,050 5 1,10

8 Pres. Transducers (GP50 7201) 4 0,140 5 0,59

9 Thrusters (EADS CHT-5) 6 0,220 5 1,39

10 Fill Drain Valves (Vacco LFRS-LP) 2 0,113 5 0,24

SUBSYSTEM TOTAL 14,7 6,5 15,62

Table 6. 7

6.6 Conclusions

The design process showed in this chapter has been performed under the key constraints of design simplicity and redundancy. The reliability of the system is improved by the redundancy planned and by the selection of components internal redundant. The use of two tanks meets the layout configurations. 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. Propulsion subsystem accommodation is shown in Figure 6.2.

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Figure 6.1: Propulsion subsystem accommodation in FINGERSAT.

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7. Data handling

7.1 Requirements and constraints

PDHT subsystem has two components: AOCS Data Handling and Payload Data Handling and Transmission. The Spacecraft computer is required to be integrated for payload data and AOCS.

AOCS Data Handling has the aim to manage the automated function that are performed on the satellite. These function 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) Datarate (bps)

Rate Gyro frequency 4 1536

Sun Sensor frequency 4 512

Earth Sensor frequency 4 2048

Magnetometer frequency 4 192

Star Tracker frequency 1 128,00

Telemetry (Kbit) 1 0,053333333

Command (Kbit) 1 0,004166667

Table 7.1

Some of the previous sensors are compatible with 1553 bus standard, the others will be connected to the spacecraft computer with serial connections.

Function Frequency (Hz)

Kinematic Integration frequency 4

Error Determination frequency 4

Magnetic Control frequency 4

Reaction Wheel Control frequency 4

Ephemeris Propagation frequency 4

Table 7.2

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:

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Telemetry Word Length (bit) 16

Telemetry frequency (Hz) 1/60

Number of telemetry 200

Table 7.3

The requirements that influence directly PDHT components sizing are:

• SAR duty cycle 10% an orbit

• Down link data rate 155 Mbps

• Mean number of images >50 a day

• Transmission band X-band During CDF meetings it was defined that payload data are quantized with 6 bits per

sample and during SAR acquisition their data rate is about 700 Mbps. Besides it was chosen only one earth station to collect payload data; an STK simulation gives us the visibility time for spacecraft PDHT antenna. There are no precise constraints for image quality.

A high data downlink capability is required, considering the large data volumes generated by the instruments. Nevertheless, downlink capacity is limited by the short available field-of-view time of satellite to the ground station. On the other hand the earth-to-satellite distance is enough low and, then, is possible to download data to chosen ground station without peculiar conditions of antenna pointing. Mission requirements have imposed X-Band to perform datalink but an encoding use is possible to have more affordable information data. An important constraint is the use of only one ground station. Another driver of the mission is the use of already existent ground station X-Band antenna. A mission requirement was to verify if it has been possible to develope S-Band link or improve this one using another band.

The assigned mass and power budget are summarized in the table below:

BUDGETS Mass (Kg) Power (W)

DH 29

PDHT 52

Table 7.4

7.2 Baseline design

A requirement was to have an integrated processor for AOCS and payload data. This choice was confirmed from our analysis.

Usually AOCS computer has a quite standard architecture (primarily composed by a set of general purpose processors, GPP, and memory banks), instead payload data computer for SAR mission has several DSP processors with very high throughput to get the image data compressed quickly and efficiently. Compression algorithms are strictly dependent from the whole system architecture (SAR data rate, analogue quantization bits, image quality desired), so the “compressor” processor must be redefined each new mission.

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Figure 7.1

The set of DSP processors and the memory required is generally coordinated by a “glue logic” that is well managed by a Field Programmable Gate Array (FPGA) or by some functions of a GPP. We choice the latter option thinking to keep some spare throughput on AOCS computer to satisfy this request.

As described in the image above we choose two redundant 1553 bus to connect payload interface equipment, star trackers and reaction wheels to the spacecraft computer, we are using a series of serial connectors for other sensors and actuators. Moreover there are many thin cables from other subsystems which carry telemetry data.

A huge mass memory is necessary to store SAR compressed image data. The size of this component is quite important because of its relevant mass. The parameters that affect this dimensioning are: visibility time of earth station (payload data transmission properties), downlink data rate which is fixed at 155 Mbps, SAR maximum duty cycle and effective SAR acquisition time (mean number of images a day). We designed a simulator to get the right size of mass memory in function of these parameters. An STK simulation gives us the visibility time for antennas with different half-power beam-width (in the range 90 to 120 degrees) for all the orbits in a revisit time. We supposed a probability distribution for SAR acquisition duration time in a orbit, the statistical mean is defined by the mean number of images a day. The result of the simulation is a probability distribution of the mass memory needed onboard with given parameters. The simulator designed is very useful as a validation tool at the end of subsystem sizing with the parameters of the equipment chosen.

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To transmit to the ground station, PDHT subsystem is developed following a standard architecture. The data signal is stored in the memory of subsystem until it is established a link between ground station tracking antenna and spacecraft. When there is synchrony, the satellite can transmit data. PDHT control unit transmit payload data to the main transmitter. We need to select a modulation and coding technique.

During the choice of the modulation and technique type we considered its Eb / N0 ratio to achieve a required bit error rate (BER) at the ground receiver.

Chosen this value, fixed the download data rate, we can compute the effective isotropic radiated power, or EIRP. This permits to determinate link budget and, thus, what type of transmitter and antenna are adequate for receiving data signals with chosen probability of error and the half beam width angle needed to transmit data, which is a product of simulation described above.

In fact we can estimate how high transmitter power for antenna gain product is required.

During link design it is estimated the propagation path length, atmospheric losses, as a function of several factors such as rainfall density, and all the other parameters which introduce some losses in the link. If the link margin is too much high or too much low, it is possible to try a new simulation session modifying some parameters or change architecture components to achieve correct results.

The transmission part of the architecture is represented in the following diagram.

Figure 7.2

7.3 Assumptions and trade-offs

The first aim of the analysis consisted in giving computational power and memory needs to elaborate AOCS functions. The assumption made in this part is summarized by the table taken from [Larson; Wertz; “Space Mission Analysis and Design” – 3rd

58

edition Chapt. 16] and used as a reference for typical computer requirements for sensors and functions. We assumed to use a GPP processor and estimated the operative system requirements with reference found in [Larson; Wertz], the supposed word length is 32 bit, the most common value in present processor’s architecture.

During the analysis we suppose 50% spare computational and memory requirements because this is only the first part of the design and many functions are not well defined.

During the component choice it has been considered that some spare computational capacity is necessary for DSP coordination. Moreover redundancy and radiation hardening must be assured from the component chosen.

All the components chosen both for managing the spacecraft and for payload data handling are assembled in a unique spacecraft computer commonly named SMU (Spacecraft Management Unit).

During PDHT sizing we assumed that SAR raw data are compressed with a special algorithm chosen between the following that represents different theoretical approaches:

• Fixed Block Adaptive Quantizer (FBAQ)

• Normalized Block Adaptive Quantizer (NBAQ)

• Fast Fourier Transform BAQ (FFT – BAQ) No others elaboration are needed on SAR raw data. It is considered the use of a

memory buffer to reduce computational complexity supposing that SAR acquisition is made of several 7-8 seconds burst, the time to acquire a 40 km x 40 km image.

During the process of sizing mass memory we used the simulator described in the previous paragraph assuming to use as minimum memory required the necessary memory to keep SAR data coming from an orbit in which the instrument works at its maximum that is 10% orbit time (maximum SAR duty cycle). As a result we guarantee that at least for one orbit the SAR can reach its greatest data rate even if there is no visibility with the ground station. We notice that this value of data acquisition for one orbit, if sustained, consists in having more than 2000 40 km x 40 km images a day, this is largely greater than system requirements on mean number of images a day (>50).

For transmissions components assumptions derive from the following mission requirements:

• the operative frequency is 8.100 GHz;

• the spacecraft altitude is 619 km;

• the payload datarate is 704,1 Mbps;

• the chosen ground station is Svalbard. The five-year mission duration requires the use of functional redundancy in design,

so PDHT subsystem is made by two transmitters, but antenna is a single point of failure, so in future analysis another antenna and a variation in connection may give more availability to the entire system.

In link budget we considered atmospheric losses but when the frequency is so much low, below 10 GHz, the effects of rainfall and ionosphere are very reduced and it isn’t necessary to estimate these losses exactly. In fact we can consider these losses through a rough value, as we also introduce a total link margin of 6 dB.

Note that the values Eb / N0 are theoretical, based on infinite bandwidth transmission channels and ideal receivers. In practice, we should account for band-limiting effects, deviations from ideal filter responses, phase noise and frequency drift in local oscillators, noise in carrier tracking loops, and bit synchronizing errors. Thus,

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we must add 1 or 2 dB to the theoretical Eb / N0 to allow for these losses. There is in link budget a typical estimated value of the feeder loss between the transmitter and the antenna. 3 dB of margin is necessary to consider the geometry for the case of imperfect alignment of transmitting and receiving antenna when the receiving antenna is pointing on edge of coverage of the half-power beam width of transmitting antenna beam.

For data reception we use an already existent antenna at Svalbard. The figure of merit of ground station receiving system GR / TR is 35.4 dB/K including radome losses. It’s possible to compensate polarization losses by rotation the feed system of the receiving feed system of antenna.

Considering the large data volumes, the high data rate on the downlink and the short available field-of-view time of satellite to the ground station an S-Band link would not give required performance with similar costs. A Ku-Band link would be more interesting but it is necessary a steering beam of the antenna or an antenna pointing mechanisms to overcome the problem of a narrow half power beam width.

A trade-off is the choiCe between an higher data rate with more performed devices and a higher image number or lower datarate by simpler devices and less image number. Higher rates of data transmission mean higher system costs.

7.4 Performances and budgets

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

Requirements Component Chosen

Code Memory – ROM (Kbit) 3736 4096

Data Memory – RAM (Kbit) 4377

Telemetry Memory – RAM (Kbit) 1600 8192

Throughput (MIPS) 2,01 14

Table 7.5

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. However it is necessary to insert another 4 Mbit memory bank to give TT&C subsystem a one day autonomy.

The component has a computational power of 14 MIPS much greater than requirements, we accepted this choice because it will manage also the DSP for payload data compression. We added some memory banks to reach the minimum quantity needed.

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 16 1,50 0,04

Bus 2 3,00 0,35

Actuators 22 1,50 0,04

TL sensors 200 1,50 0,04

TOTAL 16,38

Table 7.6

The second part of the analysis concerned with the choice of PDHT components. To compress payload data we selected a classical algorithm (FBAQ) that has not great computational requirements but don’t have excellent performance, however there were no precise constraints on SAR image quality so we give a good quality using a compression ratio of 6/3 (that is 2). The improvement in quality image (see table below) using a more complex algorithm (NBAQ) doesn’t justify the increase (400%) of computational power in MOPS (Million Operation per Second).

Algorithm Quality – SNR (dB) Computational Power

(MOPS)

FBAQ 14,79 1174

NBAQ 15,75 4224

Table 7.7

The choice of FBAQ is coupled with the use of a memory buffer to reduce elaboration rate. We choose a buffer of 640 Mbit and computational power was reduced to 1021 MOPS and we have a DSP processor that has a computational power that satisfies this requirement with a throughput of 1024 MOPS. This processor is redundant.

Another option was to select a series of ten 120 MOPS DSPs which satisfy the computational requirement and have an intrinsic redundancy. However it is convenient to keep the subsystem simple with only one DSP processor.

The table below summarize the results:

Requirements Component Chosen

Code Memory – RAM (Kbit) 11387 4x4096

Buffer Memory – RAM (Mbit) 656 3x256

Throughput (MOPS) 1021 1024

Table 7.8

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During the process of sizing mass memory we used the simulator and considered a data rate of 155 Mbps as fixed by requirements, the payload data rate is 704,1 Mbps and compression ratio is 2. During the first session of simulation we found that with an antenna with an half-power beamwidth of 120 degrees and a statistical distribution that leads to about 150 images a day we can reach good performance with 211 Gbit, the picture below shows the probability to need more memory than the available; 211 Gbit corresponds to a probability of 0,04%. The value of 211 Gbit is the minimum required to keep SAR data coming from an orbit in which the instrument works at its maximum that is 10% orbit time. As a result we guarantee that at least for one orbit the SAR can reach its greatest data rate even if there is no visibility with the ground station. However the requirement of 50 image a day is satisfied with a large margin. During the choice of real component we consider 5-10% spare to keep the value of memory space until the end of life. The radiation hardening is a feature of the component considered.

Figure 7.3

The results of simulation session gave some requirements to transmission component sizing. Initially we choose a transmitter with a 7,78 dBW RF power out, supporting QPSK modulation at 140 Mbps, lower than required data rate. This solution was compliant with the link budget analysis and gave 0,38% as probability to need more memory than available, validated by our simulator. Another option was to select a transmitter with a higher data rate, compliant with EIRP specifications and giving the possibility of more images a day. Therefore we choose a 175 Mbps 8PSK Transmitter, changing the modulation scheme with 5/6 Trellis coding. This, coupled with the high gain 120 degrees beam-width antenna (5dbW) give us about 7 dB margin on link budget. This choice allows to have about 190 images a day with a low and acceptable probability to go out of memory (<3%).

The following image shows simulation validation of these data.

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Figure 7.4

The antenna chosen provides a hemispherical radiation pattern. The used crossed-dipole antenna is a low cost, high performance unit for high-rate data transmission applications.

Security of transmitted data is guaranteed by the use of an encryptor inserted serially in the downlink chain.

Every used device is already space qualified, however two transmitter units are combined to produce a cold dual-redundant configuration.

The analysis of subsystem and the design choices gave the following budgets.

MASS Requirements (Kg) Components (Kg) Margin

DH 29 26,4 10%

PDHT 52 47,8 5,2%

POWER Requirements (W) Components (W)

DH 35

PDHT 167

Table 7.9

7.5 Hardware architecture and equipment overview

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

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Unit Unit Name # Mass (Kg) Mass

Margin (%) Power

(W) Features

1 ICS 1 10 10 35 MIPS: 14

RAM: 8 Mbit ROM: 4 Mbit

2 Cable - 16,38 10 - Data harness

3 DSP

SMJ320C6701 2 1,50 10 3,7

MOPS: 1024 RAM: 16 Mbit

4 Mass

Memory SSR 1 40 5 110 220 Gbit

5 Encryptor 1 3,3 5 18 -

6 Transmitter

8PSK 2 1,37 5 35

Data rate: 175 Mbps RF out: 7,78 dBW

7 Range

Compensated Antenna

1 0,25 5,00 - θ: 120 degrees

Gain: 5 dB

Table 7.10

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

7.6 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.

During analysis and design of the subsystem we pointed the attention on heavier component because of mass target to achieve. Therefore in a deeper analysis it should be important to study the following subjects:

• Updating reference table taken from [Larson; Wertz]

• Analysis of bus throughput and bus technology

• Studying the radiation hardening of components, making an analysis on Single Event effects on operational units.

Data handling subsystem accommodation is shown in Figure 7.5.

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Figure 7.5: DH and PDHT subsystem accommodation in FINGERSAT.

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8. TT&C Subsystem

8.1 Requirements and constraints

Basic requirement for TT&C subsystem is to provide a telecommunication link in S-band between FINGERSat 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. Moreover link margin must give the possibility to download stored telemetry data to ground within a day (24 hours autonomy is one of the system constraints). Downlink design has to guarantee a BER for telemetry rate up to 10-7. 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 23 V

Voltage Maximum

Voltage Power Vmax 38 V

Mass Budget MTT&C 7,33 kg

System Budgets

Power Budget WTT&C 26 W

Telemetry Bit rate

Bit Rate Rb 4 Kb/sec

Telemetry BER Bit Error Rate BER 1,00E-07 adim

MACRO INPUT INPUT VARIABLE NAMES

VARIABLE VALUE MEASURE UNIT

Orbit Orbit

Height H 619,42 Km

Frequency TT&C up-

link Frequency

flin 2050 MHz

Table 8.1

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8.2 Baseline design 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.

8.3 Assumptions and trade-offs 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. Supposing an average data rate around 100 bit/sec, daily stored housekeeping data are less than 5 Mbits. A possible subsystem tradeoff is to establish the downlink strategy: a daily ground contact to download all the stored telemetry data or a ground contact every time ground station is in visibility. Obviously the latter solution allows more updated housekeeping data.

Ground Station Ground operations are supported by Svalbard Station that has an ideal location for

sun-synchronous orbits having visibility up to ten passes per day, each of ten minutes in average. According to the link margins, we are able to select the antenna with the minimum allowable diameter in order to limit operative costs.

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).

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8.4 Performances and budgets

Link budgets

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

• An S-Band 5 meters antenna presents an EIRP of 58 dBW and a G/T of 17 dB/K at 10 degrees of elevation angle;

• Slant range is calculated for a circular orbit of 619,42 Km and at 10 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 240/221 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 58,00 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 1976,58 Km PLL-BDW 2Bl0 2000,00 Hz

TOTAL PROP. LOSS 166,94 dB IMPLEMENTATION LOSS 3,00 dB

POW. FLUX at S/C -51,25 dBm/m^2 REQ C/N in PLL BDW 25,00 dB

Spacecraft Receiver CARRIER MARGIN 18,85 dB

RX ANT GAIN -2,46 dBi Telecommand Recovery TOTAL CIRCUIT&CABLE LOSS 4,30 dB MODULATION LOSS 8,97 dB

S/C RX G/T -36,38 dB/K IMPLEMENTATION LOSS 3,00 dB

RX POWER -85,70 dBm BIT RATE 4,00 kbps

CAR ACQ THRSH -128,00 dBm REQ Eb/N0 14,3 dB

TC RX THRSH -118,00 dBm

TELECOMMAND MARGIN 20,99 dB

REQ RX POWER -118,00 dBm

RX POWER MARGIN 32,30 dB

RX S/N0 83,28 dBHz

Uplink

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S/C Transmitter Carrier Recovery

S/C TX POWER 5,00 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 -0,78 dBW REQ LOOP S/N 25,00 dB

Propagation Losses CARRIER MARGIN 13,96 dB

FREQUENCY 2234,50 GHz Telemetry Recovery SLANT RANGE 1976,58 Km TLM MODULATION LOSS 4,88 dB

TOTAL PROP. LOSS 167,69 dB DEMODULATOR TECH LOSS 3,00 dB

POW. FLUX at G/S -107,69 dBm/m^2 BIT RATE 4,00 kbps

Ground Station MODULATION TYPE BPSK dB

RX G/T 17,00 dB/K REQ Eb/N0 14,30 dB

RX S/N0 77,13 dBHz

Modulation Indices TELEMETRY MARGIN 18,93 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 108046,44 mHz

Table 8.2

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 5% 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 5 6,30

Radio Frequency Distribution Unit 1 0,50 0,50 5 0,53

S/S

S-Band Helix Low Gain Antenna 2 0,24 0,48 5 0,50

TOTAL MASS 7,33

TARGET MASS 7,33

DIFF. 0,00

DownLink

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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 26,00 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. 0,00

Table 8.3

8.5 Hardware architecture and equipment overview

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 8.1

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Equipment overview 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 8.2 and in Table 8.4.

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 8.4

Figure 8.2

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Transponder

The Diplexer allows to operate simultaneously the Receiver and the Transmitter with just a single RF antenna input/output port. The Diplexer also isolates Transmitter and Receiver and helps to filter the SSPA spurious frequencies.

The receiver assures the reception of signals in the range of 2025-2120 MHz. Its main function is to demodulate the RF signals with a 16kHz subcarrier from PM/BPSK to NRZ-L which is sent to Data Handling Unit (DHU) afterwards. The receiver lock threshold is -129dBm while the demodulation threshold has been restricted to –19dBm. The transmitter section generates a carrier in the 2.2-2.3 GHz band in coherent mode. It also performs the PSK modulation of the carrier with the base band modulating signal coming from the DHU.

The Power FET based SSPA amplifies this signal up to the required output power. Then, an output filter reduces the harmonics generated by the SSPA, a directional coupler with one detector allows to measure the output power, and an output isolator assures the output impedance and protects the output transistor from accidental short circuits. The modulation is directly performed at the output carrier frequency. The specified TM downlink BER is 10-7. The maximum available RF power at the diplexer output is 37 dBm in the TM frequency range of 2200 to 2290 MHz. The RF power applied to each antenna becomes 33,7 dBm due to power divider and coaxial cables losses. Since pulse shaping is not used, the transmitter output power amplifier can operate in saturation without increasing significantly the technological loss.

8.6 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 than strictly requested by mission.

TT&C subsystem accommodation is shown in Figure 8.3.

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Figure 8.3: TT&C subsystem accommodation in FINGERSAT.

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9. Thermal control subsystem

9.1 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 off 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.

9.2 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 2.1 gives typical component temperature ranges for representative spacecraft components.

Operating Temperature Non Operating Temperature Subsystem

Tmin (°C) Tmax (°C) Tmin(°C) Tmax(°C) Electronics -55 115 -55 150

Batteries 11 21 10 30 Hydrazine Tanks

and Lines 10 30 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 9. 1: Spacecraft subsystems typical temperature requirements

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

• Sunsynchronous orbit (altitude: 619 Km);

• An inclination of 97.56 deg;

• A payload (SAR) duty cycle of 10% for each orbit;

• 31 deg off-nadir pointing;

• Eclipse duration about twenty minutes for each orbit;

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• A SAR operative temperature and non operative temperature respectively between –20 °C and 60 °C, and between –55 °C and 80 °C;

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

9.3 Thermal design assumptions

The following assumptions have been used in the design process:

• The sun vector lies between –30 deg and +30 deg with respect to the normal to the orbit plane:

• The platform’s components are supposed to be at the same temperature between 0 °C and 35 °C, in the nominal phase, and between –20 °C and 50 °C, in the safe phase;

• Only the external view factors between the platform, the SAR, 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 parallelepiped (size: 1.2 × 1.2 × 3 m3).

Some components, most usually batteries and hydrazine tanks, require special thermal control methods to reach their operative temperatures.

The hydrazine propulsion system is prevented from freezing (2 °C) by a Multilayer insulation blankets MLI (25 layers are used), whereas the batteries are maintained within their operative range of temperatures (usually between 0 °C and 20 °C) using MLI, radiators and doublers.

9.4 Baseline design

9.4.1 The platform

The FINGERSAT thermal design is based on passive thermal control techniques augmented by a controlled heater. The heater is used to avoid the platform’s cooling in the worst cold safe case.

Most of the platform dissipating units and payload’s electronics are mounted on +X, -X, -Y wall panels (anti–Sun side), on which radiators with integral heat pipes are located (total area: 1.5 m2).

The appropriate radiating area is designed for the maximum dissipation of the spacecraft.

MLI blankets (25 layers) cover all other external surfaces of the spacecraft wall panels; the external side consists of Aluminized Kapton (1 mil), whereas the internal side consists of Aclar film (1 mil).

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Homogenisation of the internal temperatures is favoured with the use of black paint (maximising radiative heat absorption) and, when necessary, doublers to direct excess heat toward radiators.

The thermal analysis has demonstrated that the platform’s temperatures can be maintained within the limits specified in Table 9.1.

9.4.2 The payload

A transient analysis is performed due to the short duty cycle of the SAR: the temperature of the instrument is checked both before the eclipse starting and at the end of the eclipse period.

Two critical scenarios are considered: the worst hot case when the SAR works and is exposed to the sunlight, and the worst cold case when the SAR does not work and the platform experiences the eclipse season. Again, to guarantee the thermal control of the SAR, passive techniques are taken in account based on the use of black coatings and radiators to eject the waste heat.

In addition a thermal link between the platform radiator and SAR has to be considered, using heat pipes. In figure 3.1 the shape of eclipse SAR temperature vs time is analysed: it can be noted that the temperature is decreasing during the eclipse season; the minimum value is obviously reached at the end of the eclipse period.

Eclipse SAR Temperature vs Time

40,3

40,4

40,5

40,6

40,7

40,8

40,9

41

0120

240

360

480

600

720

840

960

1080

1200

Time (s)

Te

mp

era

ture

(°C

)

Figure 9. 1: SAR Temperature, during the eclipse period, vs Time

9.4.3 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.

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9.5 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 FINGERSAT mission a trade-off on possible presence of louvers in the platform thermal equipment is conducted. Louvers are semi-active thermal components and allow a saving of requested heater’s power at the expense of added weight.

However, in this mission, louvers are not used because they constitute a single point of failure.

9.6 Thermal equipment

In Tables 9.2 and 9.3 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.

Table 9.2: Platform thermal equipment

Component class

Used Component

# items εεεεEOL ααααEOL εεεεBOL ααααBOL

Mass (Kg)

Mass Margin

(%)

Total Area

(m2)

Required Power

(W)

MLI (25 layers)

Aluminized Kapton (ext.

Side) ; Aclar film (int. Side)

- 0.01 0.01 0.01 0.007 26 20 - -

Radiator

Indium oxide coated

optical solar

reflector

3 0.76 0.11 0.76 0.07 0.20 20 1.5 -

Heater (safe mode)

Sylicon rubber

patch heater etched foil

- - - - - - 20 - 178.6

Heat pipe NH3 17 - - - - 0.11 20 - -

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Component

class

Used

Component # items εεεεEOL ααααEOL εεεεBOL ααααBOL

Mass

(Kg)

Mass

Margin

(%)

Total Area

(m2)

Black coating

Anodize

Black - 0.88 0.88 0.88 0.88 - - -

Radiator

Indium oxide

coated optical

solar reflector

1 0.76 0.11 0.76 0.07 1.31 20 3.2

Heat pipe NH3 17 - - - - 0.11 20 -

Table 9.3: Payload thermal equipment

9.7 Conclusions

The total preliminary and current mass budget for the FINGERSAT thermal control mass budget is provided in Table 9.4 below. The table also gives a summary of the current heater’s power budget in the nominal and safe operation modes.

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.

Total Thermal Power Budget

(W)

Total Preliminary Thermal Mass Budget (Kg)

Total Current Thermal Mass

Budget (without margin) (Kg)

Total Thermal Mass Margin

(%) Nominal and

Eclipse operative

modes

Safe operative mode

36 29.86 20 0 178.65

Table 9.4: Total thermal mass and power budgets

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10. Structures and Configuration

10.1 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.

10.2 Configuration

10.2.1 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 centered 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).

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 the power and the propulsion subsystems units are accommodated.

The second one contains the AOCS, TT&C, PDHT units. In particular in the SVM there are:

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

• 2 propellant tanks on the opposite two corners of the satellite bottom cruciform assy.

In the PLM there are:

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• reaction wheels, mounted on the central cruciform assy;

• magnetometers, accommodated on the cruciform assy;

• 2 transponders, mounted on the closure panels up –Y;

• SSR and SMJ, mounted on the closure panel top +X;

• 2 Tx and 1 encryptor (PDHT), on closure panels top –Y;

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

• the harness connecting all the instruments units.

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;

• Antenna of the PDHT subsystem is located on the 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, four panels per each wings) are hinged on the edges of the top closure panels in order to point the Sun in the deployed configuration;

• SAR Antenna;

• 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. The SAR antenna is folded up on the top platform and on the closure panels and increase the cross section of 0.5 m and the height of 0.25 m. So the final stowed satellite has cross section of 1.7 x 2.2 m and is 3.25 m high.

In the deployed configuration solar panels and SAR 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.

However, the overall layout is made symmetrical and the final position of the centre of mass of the total satellite in the stowed configuration has been determined as (-15, -1, 152) mm with respect to the standard spacecraft reference system. The offset of the COG in the X and Y directions is less than the maximum permitted by the Vega launcher that is stated to be 30 mm for a 3-axis stabilized spacecraft.

Also offset in the Z direction is less than the maximum permitted for a payload of about 1000 kg. In fact, the distance of the COG from the separation plane of the launcher has to be between 500 and 1750 mm in order to stabilize static moments which may cause bending of the spacecraft. This distance is stated to be 1652 mm.

Balancing masses can be accommodated on the bottom platform, if required. Stowed and deployed configurations are shown in Figures 10.1 and 10.2. Internal

arrangement of the elements is shown in Figures 10.3 and 10.4.

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Figure 10.1: stowed configuration

Figure 10.2: deployed configuration

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Figure 10.3: Internal arrangement of the elements

Figure 10.4: Internal view of spacecraft

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In Figure 10.5 a X-Y cross section of the spacecraft is shown.

Figure 10.5: X-Y cross section

An exploded view is shown in Figure 10.6.

Figure 10.6: Exploded view of spacecraft

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10.3 Structure

10.3.1 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. So the chosen value for X and Y sides is fixed to 1.2 m since, on the closure panels, 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 937 adapter. It is a carbon fiber 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 (f 1920 mm) by a bolted connector frame, and also provides for spacecraft separation. This 937 adapter has a 60 kg mass.

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 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. Therefore CFRP was chosen as material because of its thermal stability . Moreover CRFP has a high module-weight ratio that allows to obtain a reduction of total weight without loss of performance.

10.3.2 Baseline design

The spacecraft shape is a parallelepiped with a square cross section of 1.2 m of side and 3 m high. The load-bearing structure is made by a cruciform assy which has the function to withstand to the axial and lateral loads. These items are made with sandwich panels manufactured with skins in epoxy matrix reinforced with carbon fibers and honeycomb core in Aluminum. Dynamic, strength an stability analysis on these elements lead to assume larger skin thickness than the other elements. So the skins withstand all the axial load while the core supply enough resistance to the lateral

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loads. The shape of this items supply stiffness to the whole satellite and transfer the load toward the interface with the launcher.

On the top and bottom, panels close the spacecraft to avoid space contamination and to supply a mounting surface for the SAR. 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 at the side and other four smaller panels close the service module. This choice allows easy accessibility to the equipment if a unit has to be changed late in the program, and mounting simplicity.

Also these items are made with 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. These panels contain the equipments and instruments.

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.

10.3.3 Assumptions and trade-offs

Preliminary design has been performed considering a Safety Factor equal 2. This factor is greater than the usual values because composite materials have been used. These materials have a complex mode of failure respect the traditional ones and it isn’t easy to predict their failure.

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 materials are:

• 7075 Al-alloy

• Graphite/epoxy composite with 55% in volume of fibers. Sandwich panels with honeycomb core in Aluminum have been utilized and above-

mentioned materials have been used for the skins. A trade-off has been performed between Al-alloy and CFRP materials in order to

maximize stiffness-weight ratio. This one is nearly the same for the two classes of materials, hence the CFRP material has been chosen as baseline for the load bearing structure in order to reduce as much as possible the spacecraft thermal distortion.

Since some items don’t support a strong load, their minimum thickness is very small. This can be a physical limit to their production and for this reason a minimum suggested thickness has been adopted to manufacture these panels.

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10.3.4 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 Material Mass per

each Item (kg)

Mass with

margin (kg)

Mass Margin

(%)

Height (mm)

Length (mm)

Width (mm)

Thickness Skin or

Bulk (mm)

Closure Panel (Top or Main)

4 Al 7,344 7,711 5 1800 1200 0 0,250

Closure Panel (Bottom)

4 Al 4,896 5,141 5 1200 1200 0 0,250

Shear Panel (Top)

4 CFRP 12,04 12,643 5 1800 600 0 3,070

Shear Panel (Bottom)

4 CFRP 8,027 8,428 5 1200 600 0 3,070

Platform (top) 1 Al 7,776 8,165 5 0 1200 1200 0,250

Platform (main)

1 Al 7,776 8,165 5 0 1200 1200 0,250

Platform (bottom)

1 Al 7,776 8,165 5 0 1200 1200 0,250

Joints 16,00 19,200 20

Total Mass of Structure (kg) 168,5 179,387 6

Table 10.1: Mass budget and sizing for baseline design

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

requirements of safety is 180 kg considering a mass margin of 6%. However it’s possible to perform a trade-off on the feasible kind of configurations

both internal and external ones: configuration with a cylinder load-bearing structure or octahedral shape of the spacecraft. Moreover it’s possible to evaluate use of other kind of materials in order to maximize performance-weight ratio.