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Space and Astronautic Engineering | Tatiana Quercia

Supervisor: Prof. F. Santoni

February 2014

SATELLITE

DESIGN

PRELIMINARY DESIGN OF A BUS FOR

THE GEOSTATIONARY ORBIT

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Table of Contents

1. Market Research........................................................................................ 1

1.1. Funds and investments ............................................................................ 1

1.2. Satellite Manufacturers ............................................................................ 3

1.3. Satellite Buses ........................................................................................ 4

1.3.1. European Buses for GEO .................................................................... 4

1.3.1.1.Alphabus .......................................................................................... 4

1.3.1.2.Spacebus 4000 .............................................................................. 5

1.3.1.3.SmallGEO (OHB, Germany) ................................................................ 5

1.3.1.4.Eurostar ........................................................................................... 6

1.3.2. Russian Bus for GEO: Ekspress ........................................................... 6

1.3.3. US Buses for GEO.............................................................................. 7

1.3.3.1.Boeing 601 ....................................................................................... 7

1.3.3.2.Boeing 702 ....................................................................................... 7

1.3.3.3.LS-1300 ........................................................................................... 9

1.3.3.4.GEOStar ........................................................................................... 9

1.3.3.5.A2100 .............................................................................................. 9

1.3.4. Chineese Buses for GEO ................................................................... 10

1.3.4.1.CAST3000 ...................................................................................... 10

1.3.4.2.DFH ............................................................................................... 11

1.3.5. Bus for GEO: DS2000 ...................................................................... 11

1.3.6. Indian Buses for GEO ....................................................................... 12

1.3.6.1.INSAT ............................................................................................ 12

1.3.6.2.GSAT ............................................................................................. 14

1.4. Alphabus ............................................................................................. 15

1.4.1. The Alphabus platform ..................................................................... 15

1.4.2. The Alphabus structure .................................................................... 17

1.4.3. Thermal control............................................................................... 17

1.4.4. Power system ................................................................................. 17

1.4.5. Electrical propulsion......................................................................... 17

1.4.6. Chemical propulsion ........................................................................ 18

1.4.7. Electrical power architecture ............................................................. 19

1.4.8. Data handling ................................................................................. 19

1.4.9. Attitude Determination and Control System ........................................ 19

1.4.10. Antenna Tracking System ................................................................. 19

1.4.11. Key features ................................................................................... 20

2. Project Management ............................................................................... 22

2.1. Work breakdown Structure (WBS) .......................................................... 23

3. Subsystems Preliminary Design .............................................................. 24

3.1. Structure ............................................................................................. 24

3.2. Power system ....................................................................................... 28

3.2.1. Solar Panels ................................................................................... 28

3.2.2. Batteries ........................................................................................ 30

3.3. Propulsion System ................................................................................ 32

3.3.1. Chemical propulsion ........................................................................ 32

3.3.1.1.Propellant and propellant tanks ......................................................... 32

3.3.1.2.Pressurant gas and its tanks ............................................................. 34

3.3.2. Electric Propulsion ........................................................................... 35

3.4. Telecommunications System .................................................................. 36

3.4.1. Link Budget .................................................................................... 36

4. Bibliography ............................................................................................ 37

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1. Market Research

1.1. Funds and investments

Euroconsult, the leading international consulting and research and firm specializing in the

space and satellite sectors published in 2011 the report "Government Space Markets, World

Prospects to 2020." According to this report, government spending on space hit a number of

major milestones in recent years, including a historic peak in combined government spending of $71.5 billion in 2010.

However, after 10 years of spending increases across the globe, this trend is about to come to

a halt. According to Euroconsult, public space program financing will slow dramatically in the

next five years due to several factors.

Overall growth in civil program expenditures will be sustained thanks to a growing

commitment from a wider range of countries and agencies. Leading programs such as those in

the United States, Europe and Japan are expected to see continuing budget pressure. In

Russia and countries reaching 'space maturity' (e.g., China and India), space expenditures will

continue to grow, though more modestly than in the past.

Defense space programs are expected to be influenced by military agency procurement cycles

as well as the completion of most programs currently under development, especially in the United States.

Developing initiatives in other countries (such as Australia, Canada and emerging countries)

are expected to open the door to other commercial opportunities for the commercial space

industry.

A total of 692 satellites will be launched by governments in the coming decade, up 43% from

the previous decade. This is a direct reflection of the increasing number of new space-capable

countries across the globe. Civil agencies will launch roughly 75% of these satellites, a

significant increase compared to the last decade during which they accounted for 67% of all

government satellites launched.

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Applications that drove overall spending increases in 2010 included manned spaceflight,

SatCom, and Earth observation. Government spending on space security, satellite navigation

(SatNav), science and exploration, and access to space declined. According to Euroconsult, in

the coming five to ten years, a number of important trends will emerge or continue that will have a major impact on overall government space spending.

SatCom grew by 49% in 2010 to reach $8.4 billion fuelled, primarily by defense

spending. Spending should, however, return to historical levels in the short term.

Manned spaceflight spending totaled $11.6 billion in 2010, but the current transition of

the American program should lead to decreasing investment in the future.

Earth observation reached $8 billion and spending will continue to be driven by defense,

climate change and the growing participation of emerging space nations in Earth

observation, with spending likely to exceed the $9.5 billion mark by 2015.

Science and exploration budgets totaled $5.6 billion in 2010. After a period of

decreasing investment, budget growth should resume in the coming years, especially in

the United States.

Access to space (launch capability) investments reached $4.6 billion in 2010, and

should be sustained in the coming years as more governments see independent access

to space as a top priority of their space programs.

SatNav spending totaled $2.9 billion, a 22% drop due to the end of the European Space

Agency's financing of Galileo. Growth should resume in the short term with deployment

of several new domestic systems.

Space security budgets fell to $1.7 billion, a 47% decrease which is linked to program

challenges and cancellations in the United States. Governments are expected to support

funding in this area to adapt to a wide variety of threats to protect their space assets and capabilities.

With governments across the world adopting strict spending policies, the space sector will

experience a new era of cooperation as governments strive to make their programs as efficient

as possible in order to compensate for budget limitations. According to Euroconsult, these

constraints will lead to enhanced cooperation with the private sector and more multilateral

government projects. Areas of cooperation will likely favor innovation in programs, financing and management.

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1.2. Satellite Manufacturers

There are six major companies that build large, commercial, geosynchronous satellite

platforms. They are listed in the following table.

Company Location N. of satellites

launched Comments

OHB Germany

Thales Alenia Space Europe ( France/ Italy)

formerly Alcatel

Alenia Space

JSC Information

Satellite Systems

Russia 1160 formerly NPO PM

Boeing United States

Astrium Satellites

Europe ( France/ Germany/

Spain/ United Kingdom)

a business unit

of Astrium

Lockheed Martin United States

Space Systems/Loral United States 240

Many other companies have successfully built and launched satellite platforms, mostly from

USA and Russia.

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1.3. Satellite Buses

A satellite bus or spacecraft bus is the general model on which multiple-production

satellite spacecraft are often based. The bus is the infrastructure of a spacecraft, usually

providing locations for the payload.

They are commonly used for geosynchronous satellites, particularly communications satellites,

but are also used in spacecraft which occupy lower orbits, occasionally including low earth

orbit missions.

A bus-derived satellite would be used as opposed to a one-off, or specially produced satellite.

Bus-derived satellites are usually customised to customer requirements, for example with

specialised sensors or transponders, in order to achieve a specific mission.

A bus typically consists of the following subsystems:

Command and Data Handling (C&DH) System

Communications system and antennas

Electrical Power System (EPS)

Propulsion

Thermal control

Attitude Control System (ACS)

Guidance, Navigation and Control (GNC) System

Structures and trusses

Life support (for crewed missions).

Next paragraphs explore the state-of-the-art of buses for the Geostationary Orbit (GEO).

1.3.1. European Buses for GEO

1.3.1.1. Alphabus

It is a family of heavy geostationary communications satellites developed by a joint venture

between Thales Alenia Space and EADS Astrium Satellites in France, with support of the Centre

national d'études spatiales (CNES), the French space agency and the European Space

Agency (ESA).

The Alphabus platform is designed for communications satellites with payload power in the

range 12-18 kW. Satellites based on Alphabus have a launch mass in the range 6 to 8 tonnes,

40% more than the most powerful Spacebus 4000 (see later).

In order to cover the mission range in an optimized way, the platform product line includes

several options such as electric propulsion and features scalable resources (solar array,

radiators for thermal dissipation, etc.). The platform is able to accommodate up to 190 high

power transponders and large antenna farms and has a significant growth potential (22 kW

payload power and 9 tonnes launch mass for the extended range).

All the technical information about this bus is given in chapter 1.4.

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1.3.1.2. Spacebus 4000

The Spacebus 4000 is a medium-class telecommunication satellite (launch mass from 3000 kg

for the B2 version to 5900 kg for the C4 version) with a strong successful flight heritage and

proposed with a realistic and safe manufacturing schedule. It can easily accommodate a large

range of payloads in every band (Ku, C, Ka, X, S, L) to satisfy customer needs. Thales

Alenia Space offers a high reliability to the customer and an attractive price.

Key Features

Telecommunication Satellite; launch mass from 3000 to 5900 kg; Solar Array

power up to 15,8 kW;

Payload power up to 11,6 kW, typically 80 to 100 active channels with medium RF

power (105/110W in Ku band); standard equipment and system designs available

in Ku/C and Ka frequency bands; other frequency bands (X, S, L) can be

proposed;

Realistic and safe manufacturing schedule (typically 27 à 33 months on ground

delivery); attractive price;

Accommodation antenna (from 2.4 m to 3.2 x 2.4 m); Flight proven units

(payload and platform);

Design, manufacturer and test with an experienced teams for system and

equipment management;

Large successful experience in Orbit (6 more models under manufacturing)

Outstanding reputation for the quality of the Thales Alenia Space products (ISO

9001 successful audit in September 2003);

Launch capability demonstrated with all available launchers Avionics 4000: 100V

bus.

Key Benefits

Payload flexibility for customer requirements: Platform can accommodate different

payloads without significant cost and schedule impacts;

Risk management and reliability: Experience in orbit; Satellite functions are fully tested

before launch; Secured engineering process;

Relationship with customer: Full visibility can be offered during development as

required by customers; Continuous improvement to correct anomalies in orbit.

1.3.1.3. SmallGEO (OHB, Germany)

The platform offers an efficient modular satellite solution covering a wide range of different

geostationary applications. Highlights of the design:

use of modern, flight proven components;

alternative concepts of propulsion adapted by the mission requirements (only chemical,

chemical/electric and full electric propulsion);

fast realisation cycles from contract to launch due to e.g. paralell preintegration and

tests of payload and satellite platform;

Product line for small to medium payloads (320 to 600 kg payload mass and 2 to 7 kW

payload power);

start mass between 2 and 3,5 t;

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high reliability of the satellite platform due to adequate redundancies of all important

functionalities;

15 years lifetime in geostationary Orbit

optional adaptation to high attitude accuracy requirements, e.g. for Earth observation

or scientific missions;

compatibility to various launchers.

1.3.1.4. Eurostar

Eurostar is the brand name for a satellite bus made by Astrium which has been used for a

series of spacecraft providing telecommunications services in geosynchronous orbit (GEO).

More than 60 Eurostar satellites have been ordered to date, of which 45 have been

successfully launched since October 1990 and have proven highly reliable in operational

service. The Eurostar spacecraft series is designed for a variety of telecommunications needs

including fixed services and broadcast, mobile services, broadband and secured

communications.

Eurostar was designed in the mid 1980s for a market which at the time had a design envelope

of 1.8-2.5 tons on the proposed launch vehicles (STS PAM D2 and Ariane 4). Satellite payload

power was from 1300 to 2600 W. This was the first commercial satellite to have a digital

avionics system modular in concept. With this system architecture, all key satellite parameters

are in software, which permits mission specific requirements to be implemented without

hardware changes. The initial satellite structure and configuration designed to early

requirements had significant growth potential, which subsequently allowed the payload

capability (mass and power) to be more quadrupled between 1987 and 1992, with a minimum

of requalification. Astrium has since developed further the product line in a staged process

which mainly increases the satellite power and propulsion capability and real estate for

accommodation of equipment and antennas. The overall configuration of Eurostar satellites has

essentially not changed in 20 years through the successive generations Eurostar E1000,

E2000, E2000+ and E3000. They have just become larger, more powerful, with

implementation costs reduced through longer orbit manoeuvring lifetime, and more efficient

and powerful payloads. Nowadays the Eurostar E3000 series has been considerably enhanced

and updated with the latest technologies, still maintaining the basic proven configuration. It

offers power payload ranging from 4 to 14 kW. Satellites launch mass usually ranges between

4,000 and 6,000 kg.

1.3.2. Russian Bus for GEO: Ekspress

Ekspress is a series of geostationary communications satellites owned by Russian State

Company for Satellite Communications. The first satellite of this kind was launched on October

13, 1994. The satellites are produced by the company JSC Information Satellite Systems -

Reshetnev Company.

The first satellite of the series, Ekspress 1, had a mass of 2.5 tons, 17 channels and an

operational lifetime of at 5–7 years. Starting in the mid-1990s, NPO PM started to make

significant effort to close the technology gap between Russian and Western communication

satellites. Cooperation with the French company Alcatel (now Thales Alenia Space) was begun

in 1995. An improvement was the Ekspress AM version, first launched in 2003: it has an

operational lifetime of 12–15 years and is able to carry 38 channels, including digital TV, radio,

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broadband and internet. The launch of the Ekspress AM-3 spacecraft in June, 2005 completed

the modernization of Russia's communications satellite network.

Other versions of include the Ekspress 2000, which has a mass of 3.2 tonnes. It has up to 60

transponders, power of 25 kW and a lifetime of 15 years. Satellites using this platform are

called the Ekspress AT series and the Ekspress AM30 and AM40 series. Ekspress 1000 is

smaller than the 2000 version; 700 kg to 1,400 kg, 10 to 12 transponders, 2 kW of power and

a lifetime of 15 years. Satellites based on Ekspress 1000 are called Ekspress AK or in its

navigational version GLONASS K.

1.3.3. US Buses for GEO

1.3.3.1. Boeing 601

The body-stabilized Boeing 601 was introduced in 1987 to meet anticipated requirements for

high-power, multiple-payload satellites for such applications as direct television broadcasting

to small receiving antennas, very small aperture terminals for private business networks,

and mobile communications. The basic configuration features as many as 48 transponders

and offers up to 4,800 watts.

A more powerful version, the Boeing 601HP, made its debut in 1995. The HP versions can

carry payloads twice as powerful as the classic Boeing 601 models, through such innovations

as gallium arsenide solar cells, advanced battery technology, and an optional xenon ion

propulsion system (XIPS). The 601HP features as many as 60 transponders and provides up

to 10,000 watts.

All Boeing 601 spacecraft use the same basic bus design, enabling Boeing to realize

efficiencies gained by production volume, tooling investments and quantity buys.

The Boeing 601 body is composed of two modules:

1. the primary structure that carries all launch vehicle loads and contains the propulsion

subsystem, bus electronics, and battery packs;

2. a structure of honeycomb shelves that hold the communications equipment, electronics,

and isothermal heat pipes.

Reflectors, antenna feeds, and solar arrays mount directly to the payload module, and

antenna configurations can be placed on three faces of the bus. This modular approach

allows work to proceed in parallel, thereby shortening the manufacturing schedule and test

time.

1.3.3.2. Boeing 702

Evolved from the popular, proven 601 and 601HP (high-power) spacecraft, the body-

stabilized Boeing 702 is the world leader in capacity, performance and cost-efficiency.

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The first Boeing 702HP satellite was launched in 1999. The satellite can carry more than 100

high-power transponders, and deliver any communications frequencies that customers

request.

The Boeing 702 design is directly responsive to what customers said they wanted in a

communications satellite, beginning with lower cost and including the high reliability for

which the company is renowned. For maximum customer value and producibility at

minimum total cost, the Boeing 702 offers a broad spectrum of modularity. A primary

example is payload/bus integration. After the payload is tailored to customer specifications,

the payload module mounts to the common bus module at only four locations and with only

six electrical connectors. This design simplicity confers major advantages. First, nonrecurring

program costs are reduced, because the bus does not need to be changed for every payload,

and payloads can be freely tailored without affecting the bus. Second, the design permits

significantly faster parallel bus and payload processing. This leads to the third advantage: a

short production schedule.

Further efficiency derives from the 702's advanced xenon ion propulsion system (XIPS),

which was pioneered by Boeing. XIPS is 10 times more efficient than conventional liquid fuel

systems. Four 25-cm thrusters provide economical stationkeeping, needing only 5 kg of fuel

per year - a fraction of what bipropellant or arcjet systems consume. Using XIPS for final

orbit insertion conserves even more mass as compared to using an on-board liquid apogee

engine. Customers can apply the weight savings to substantially increase the revenue-

generating payload at small marginal cost, to prolong service life, or to change to a less

expensive launch vehicle (when cost is based on satellite mass).

For even more versatility, the Boeing 702HP also incorporates a bipropellant propulsion

system, which can lift the satellite into final orbit after separation from the launch vehicle.

Innovation extends to the Boeing 702HP power systems as well. The Boeing 702 offers a

range of power up to 18 kW. Dual and triple-junction gallium arsenide solar cells enable such

high power levels.

The Boeing 702HP separates the bus and payload thermal environments and substantially

enlarged the heat radiators to achieve a cooler, more stable thermal environment for both

bus and payload. This increases unit reliability over service life. Deployable radiators use

flexible heat pipes, which increase packageable radiator area. Further thermal control occurs

through passive primary rejection via heat pipes.

The baseline Boeing 702 is compatible with several launch vehicles. These include the Delta

IV, Atlas V, Ariane 5, Proton, and Sea Launch.

In 1997, Boeing, received a nearly $1 billion contract for a system consisting of two 702HP

satellites that will serve the provide mobile telephone service to the Middle East, North and

Central Africa, Europe, Central Asia and the Indian subcontinent. It is the largest satellite

communications project in the region and serves nearly 1.8 billion people. The first Thuraya

satellite was launched in October 2000. Thuraya-2 was launched in June 2003.

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In 2006, Boeing received a second major contract from SkyTerra LP to provide two Boeing

702HP satellites, with an option for a third. The satellites will be used to create the world’s

first commercial wireless communications service, using both space and terrestrial elements.

The Boeing 702HP geomobile satellite system features a 12.25-meter deployable antenna,

and onboard digital signal processing and beamforming. It is a mobile-communications-

supporting satellite system that integrates a Boeing geosynchronous-orbit satellite with a

ground segment and a user terminal segment.

1.3.3.3. LS-1300

The LS-1300, previously the FS-1300, is a satellite bus produced by Space Systems/Loral.

Total broadcast power ranges from 5 to 25 kW and the platform can accommodate from 12 to

150 transponders. The LS-1300 is a modular platform and Loral no longer reports designators

for sub-versions, such as: 1300E, 1300HL, 1300S, 1300X.

First available in the late 1980s, the FS-1300 platform underwent revision multiple times over

its design life, all the while remaining a popular communications platform. The earliest models

provided 5,000 RF watts of transmitter power, weighed 5,500 kg, and required a 4-meter

diameter launch fairing. Newer models provide double that, approximately 10,000 RF watts of

transmitter power, weigh 6,700 kg, and require a 5-meter diameter launch fairing.

1.3.3.4. GEOStar

Orbital is the world's leading supplier of 1.5 - 5.5 kilowatt commercial geosynchronous (GEO)

communications satellites used to provide direct-to-home TV broadcasting, cable program

distribution, business data network capacity, regional mobile communications and similar

services.

Orbital's highly successful GEOStar-2 and -3 communications satellites provide up to 7.5 kW of

payload power and can accommodate virtually all types of commercial communications

payloads.

For many applications, the lighter weight, more affordable GEOStar-2 and -3 designs are

attractive alternatives to the larger, more costly GEO satellites offered by other suppliers.

Since the first GEOStar satellite was launched in 1997, Orbital has built an impressive heritage

delivering 25 satellites with the industry’s best reliability record. As of early 2012, the

company has built or has on order 34 GEOStar commercial communications satellites,

solidifying its position as the provider of choice for 2 – 7.5 kW satellites.

Orbital GEOStar satellites are compatible with all major launch vehicles, providing customers

the widest range of launch options. GEOStar satellites have been launched on Ariane 5, Proton,

Land Launch and Soyuz rockets and are compatible with Sea Launch, Falcon 9, Atlas and H-IIA

vehicles as well. Dual or dedicated launch options are available.

1.3.3.5. A2100

Lockheed Martin Commercial Space Systems produces A2100, one of the most powerful flight-

proven commercial spacecraft currently available. This modular geosynchronous satellite has a

design life of 15 years and a flexible payload capacity ideally suited to meet the demand for

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commercial space systems well into the 21st century—a demand driven by growth in mobile

telephony, business services, direct broadcast, internet, multimedia and broadband services.

The A2100 design is highly modular at the subsystem and component levels, and the broad

inventory of fully qualified standard components allows a “configure to order” approach that

eliminates costly reengineering. The A2100 design also features a major reduction in parts—

simplifying construction, increasing on-orbit reliability and reducing weight and cost.

Lightweight all-composite materials increase strength, minimize thermal distortions

and reduce launch costs. Manufacturing A2100 spacecraft occurs in the LockheedMartin

Commercial Satellite Center. With co-located assembly and test facilities, the center was

specifically designed to dramatically cut A2100 production cycles.

A2100 satellites provide advanced services such as video on demand, video teleconferencing,

telemedicine, highspeed-internet access and multimedia services.

The A2100 is compatible with Atlas V, Proton, Delta, Sea Launch, Ariane 5 and LongMarch.

This versatility supports a wide range of mission options and increases time-to-orbit reliability.

Technical features of A2100 AX – High Power

Power: 7.5 – 12 kW

Maximum wet mass: 6000 kg

Developed: 1996 – 2003

Power subsystem enhancements

171-AHr batteries

Enhanced power regulation unit

Multi-junction solar cells

Stretched AXL structure

Stretched propellant/oxygen tanks

32-bit onboard computer upgrades

330-sec liquid apogee engine

1.3.4. Chineese Buses for GEO

1.3.4.1. CAST3000

The CAST3000 is a top class small satellite bus characterized by its high precision, high agility,

high automatization, prolonged life-span and high reliability, which can support both high

resolution imaging and agile imaging missions. The bus can perform swift and wide range of

attitude maneuver and have free space pointing capacity. It can realize multi-mode imaging

(including area target imaging, continuous strap imaging, composite straps imaging and single

orbit stereo imaging, etc). Also, it supports S/X data transmission and onboard highly effective

power supply; has the capability of automatic onboard mission management and scheduling;

adopts a highly integrated onboard synthesized electronic system; and achieves a 50-60%

carrying capacity and a more-than-5-year designed lifespan.

Technical Specifications

Bus mass: 300-400kg

Payload capacity: >380kg

Orbit type: low, medium and high orbits

Attitude control ability: 3-axis stabilization

Pointing mode: earth pointing or inertial space pointing

Attitude maneuver ability: maneuver range: ±45° (along roll and pitch axes)

Stereo speed: 5° within 11s, 15° within 16s, 30° within 21s, 45° within 25s, 60° within 30s,

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90° within 35s (including stabilization time, along roll and pitch axes)

Attitude measurement accuracy: < 0.001° (3-axis@3σ);

Attitude pointing accuracy: < 0.05°(3-axis@3σ);

Attitude stability: < 0.0005°/s (3-axis@3σ)

Orbit control: hydrazine propellant, 32.5~130kg optional

Solar array output power: 945~1148W (EOL)

Designed lifetime: >5 years

Applications

The satellite bus can be widely applied in different missions, such as earth observation,

technological demonstration, scientific exploration, earth environmental exploration, formation

flight and networking, meteorological research and application, communications, navigation,

and so on. And it is mainly applicable to the high resolution optical imaging and radar imaging

missions.

1.3.4.2. DFH

DFH bus is a communications satellite bus whose capability is between high and medium ones.

It adopts hexahedral structure, consisting of propulsion, service and communication modules,

communication antennas and solar arrays and adopts 3-axis stabilized attitude control.

Technical Specifications of DFH-3B

Dimensions:2200mm(X)×2000mm(Y)×3100mm(Z)

Mass: 3800kg

Payload bearing capacity: 400~450kg

Orbit type:GEO and other orbits

Antenna pointing error:

Pitch ,Roll ≤ 0.06°(3σ)

Yaw ≤ 0.2°(3σ)

Station keeping precision: ± 0.05°(3σ)

Solar array output: 5500W

Payload power consumption: 3000W~4000W

Lifetime: 12~15 years

Applications

Applicable to communications and navigation satellites and deep space probes through

adaptive modification.

1.3.5. Japanese Bus for GEO: DS2000

Heeding the call for large-capacity, high-speed communications capable of satisfying the many

increasingly diversified needs in the commercial communications satellite market, Mitsubishi

Electric developed the DS2000 standard satellite platform.

The DS2000 platform was developed based on a design originally created for the DRTS and ETS-

VIII platforms through development by JAXA. After winning an international bid competition for

the MTSAT-2, a Japanese commercial satellite launched in 2006, the company incorporated

evolutionary changes to match the requirements for standard commercial communications

satellites and introduced the DS2000.

An original program management system developed based on Mitsubishi Electric’s years of

experience in the communications satellite business allows customers to access design data and

processes and request changes during development, production and testing. This high level of

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visibility ensures that each platform is tailored to exact requirements and is completed in time to

meet the delivery schedule, thus allowing the DS2000 to meet the needs of communications

satellite operators around the world.

Distinctive Features

Highly reliable design and production based upon rich experience derived from

participation in more than 280 satellite projects worldwide.

Capable of providing an output of up to 15 kW, satisfying the power

requirements for powerful and multiple communications transponders.

Flexible design matches various applications including hybrid communications

payloads.

System

Satellite Power Up to 15 kW

Launch Mass 5,000 kg

Transponders 72 (nominal)

Life 15 yr

Launch Vehicle H2-A, ARIANE-V, DELTA-IV, ATLAS-5, Sea Launch, etc.

Subsystem

Electrical Power System 100 V regulated dual bus with high-efficiency silicon,

multijunction GaAs solar array and NiH2 or Li-Ion batteries

Onboard Control System Attitude control, data handling and satellite management

Attitude Control Controlled bias momentum or zero momentum with

satellite controller (SC) and four skew reaction wheels

Telemetry Tracking &

Command 1553B interface with SC

Structure Carbon-fiber reinforced plastic, central cylinder type

Thermal Control Heat pipe embedded payload panel, OSR, blanket and

heater

Propulsion

IIntegrated Bi-propellant with apogee and attitude control

thrusters

Ion engine is available

1.3.6. Indian Buses for GEO

1.3.6.1. INSAT

INSAT or the Indian National Satellite System is a series of multipurpose geo-stationary

satellites launched by ISRO to satisfy the telecommunications, broadcasting, meteorology,

and search and rescue operations. Commissioned in 1983, INSAT is the largest domestic

communication system in the Asia Pacific Region. It is a joint venture of the Department of

Space, Department of Telecommunications, India Meteorological Department, All India

Radio and Doordarshan. The overall coordination and management of INSAT system rests with

the Secretary-level INSAT Coordination Committee.

INSAT satellites provide transponders in various bands (C, S, Extended C and Ku) to serve the

television and communication needs of India. Some of the satellites also have the Very High

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Resolution Radiometer (VHRR), CCD cameras for metrological imaging. The satellites also

incorporate transponder(s) for receiving distress alert signals for search and rescue missions in

the South Asian and Indian Ocean Region, as ISRO is a member of the Cospas

Sarsat programme.

The Indian National Satellite (INSAT) system was commissioned with the launch of INSAT-1B

in August 1983 (INSAT-1A, the first satellite was launched in April 1982 but could not fulfill the

mission).

INSAT-4B Technical features

Mission Communication

Weight 3025 Kg (at Lift – off)

Onboard

Power 5859 W

Stabilization

It uses 3 earth sensors, 2 digital sun sensors, 8 coarse analog sun sensors,

3 solar panel sun sensors and one sensor processing electronics. The

wheels and wheel drive electronics were imported with indigenous wheel

interface module to interface the wheel drive electronics and AOCE.

Propulsion

The propulsion system is employing 16 thrusters, 4 each located on east,

west and AY sides and 2 each on north and south sides. There is one 440 N

liquid apogee motor (using Mono Methyl Hydrazine (MMH) as fuel and

oxides of Nitrogen ( MON3 as oxidizer) and three presurant tanks mounted

on the LAM deck.

Payload

12 Ku band high power transponders covering Indian main land using

140W radiatively cooled TWTAs.

12 C band high power transponders with extended coverage, covering

southeast and northwest region apart from Indian main land using 63 W

TWTAs

Launch date March 12, 2007

Launch Site French Guyana

Launch

Vehicle Ariane5

Orbit Geostationary (93.5o E Longitude)

Mission Life 12 Years

14

1.3.6.2. GSAT

GSAT-14 is the twenty third geostationary communication satellite of India built by ISRO. The

main objectives of GSAT-14 mission are: i) To augment the In-orbit capacity of Extended C and

Ku-band transponders ii) To provide a platform for new experiments.

The payloads of GSAT-14 are:

Six extended C-band transponders for Indian mainland and island coverage with 36

dBW Edge Of Coverage- Effective Isotropic Radiated Power (EOC-EIRP)

Six Ku-band transponders covering the mainland India with 51.5 dBW EOC-EIRP

Two Ka-band Beacons operating at 20.2 GHz and 30.5 GHz to carry out attenuation

studies

Some of the new technologies being tested on GSAT-14 are:

Fiber Optic Gyro

Active Pixel Sun Sensor

Ka band beacon propagation studies

Thermal control coating experiments

Mass At Lift-

Off 1982 kg

Overall Size

(m) 2.0 X 2.0 X 3.6

Power 2600 W

Attitude and

Orbit Control

System

(AOCS)

Momentum biased 3-axis stabilized mode

Propulsion

System

Bi propellant-Mono Methyl Hydrazine and Mixed Oxides of Nitrogen

(MON-3)

Antennae One 2m and one 2.2 m single shell shaped reflector Antennae

(transmit and receive)

Launch date January 05, 2014

Launch site SDSC, SHAR

Launch

vehicle GSLV-D5

Orbit 74 deg East longitude in geostationary orbit

Mission life 12 Years

15

1.4. Alphabus There are times when size really does matter, especially when it comes to ‘satcom’, or satellite

telecommunications. Larger and more complex satellites are required by a demanding market,

and Alphabus is Europe’s answer to meet this need.

The Alphabus market Telecommunication satellite systems provide telephony, internet, radio,

video and data services to millions of people around the world.

Alphabus has come a long way since 2002, when the cooperation between ESA and CNES

started to develop the large platform. Phase C/D began in September 2005, the Critical Design

Review was closed in February 2008, and the Alphabus Qualification Review in the second half

of 2010 marked the achievement of a major objective of this important programme.

Alphabus has been jointly developed by EADS Astrium and Thales Alenia Space as co-prime

contractors, leading a European-wide industrial consortium.

1.4.1. The Alphabus platform

Typically, communication satellites consist of two main parts: the ‘platform’ (or service

module) which, in simple terms, supplies the mechanical and thermal control, together with

electrical power, attitude control and data handling to the second part, known as the ‘payload’.

After the transfer into geostationary orbit (36 000 km away from Earth), the platform keeps

the satellite in this orbit, pointed towards Earth while providing a stable thermal environment

for the payload by radiating heat generated into

space. Telecom payloads receive signals from Earth,

process and amplify them and then transmit them

back to Earth.

Alphabus has been developed to support large

payloads with power levels ranging from 12 kW up

to 18 kW, and a payload mass of up to an

impressive 1500 kg. It can accommodate a large

amount of payload equipment with a large thermal

rejection capability. As an example, it can hold 190

transponders, allowing the transmission of more

than 1000 television channels and more than 200

000 audio channels.

Alphabus can cope with large ‘antenna farm’

configurations, of up to 12 antennas with rigid

reflectors of up to 3.5 metres in diameter.

Alphabus has been designed to fit inside the 5 m

fairings on Ariane 5 and Atlas 5, but remains also

compatible with a 4 m fairing on Proton. Alphabus is

compatible with an Ariane 5 ECA dual launch

configuration at the low end of the range and up to

8800 kg on Ariane 5 single launch or Atlas 5 at the

high end.

The overall Alphabus platform qualification was

completed in 2010.

16

17

1.4.2. The Alphabus structure

The Alphabus platform is based on a scalable and modular design allowing parallel integration

and tests of the standardised Service Module (SM) as well as the mission-specific three-floor

Repeater Module (RM).

The SM is built around a large central tube (1.6 m) embedding two large propellant tanks with

a maximum capacity of 4200 kg, and provides the mechanical interface with the launch vehicle

for a launch mass of up to 8,800 kg.

The RM itself is split in two halves, allowing parallel integration of the repeater units with an

accommodation capacity doubled compared to what is available today. It is mounted on top of

the SM. The use of an ‘ultrastable’ antenna module structure for the Earth-facing side allows

the efficient mechanical alignment of the antennas and their accurate pointing towards Earth.

It can also include lateral arms to hold large antenna reflectors on the lateral sides of the

satellite, if required by the mission. The antenna module structure is under development by

RUAG Space (Switzerland). The overall physical configuration of Alphabus built on these three

modules – the SM, RM and the antenna module – allows for maximum payload unit

accommodation. The overall structure has been developed by Thales Alenia Space Cannes

(France). The central tube was developed by EADS Casa (Spain) and is based on state-of-the-

art carbon fibre placement technology offering high strength and low mass.

RUAG Space (Switzerland) developed various structural panels. The RM structure thermal

control integration and assembly were performed by Thales Alenia Space Turin (Italy).

1.4.3. Thermal control

The Alphabus platform configuration is designed to provide both physical and thermal

separation between the mission-specific RM and the SM, minimising the design and analysis

required to adapt the platform to a dedicated mission. The SM thermal design uses 3D heat

pipe network linking the East/West and the North/South panels. A 3D surface heat pipe

networks are also installed on the backside of the North/South panels and on the shelves of

the RM. A high level of payload thermal dissipation is ensured by adaptation of the fixed

radiators surface.

1.4.4. Power system

The solar generator, inherited from Eurostar E3000, is scalable from four to six panels per

wing. The 10 m2

panels are fitted with triple junction solar cells from Azur Space (Germany)

and will benefit from the continuous efficiency improvements of gallium arsenide (GaAs) cell

technology. When needed, the fifth and sixth panels are deployed laterally from the third in-

line panel. During launch and early orbit operation, one panel per wing is deployed and

provides sufficient power for the satellite electrical power balance. The solar array was

developed by EADS Astrium in Ottobrunn (Germany). The low-shock release mechanism was

designed also by EADS Astrium Ottobrunn and developed by RUAG Space. The high-power

solar array drive mechanism was developed by EADS Astrium Stevenage (UK).

1.4.5. Electrical propulsion

The Electrical Propulsion subsystem is based on four plasma thrusters used for north/south on-

orbit ‘stationkeeping’ of the satellite. The system is an evolution of an existing design and

consists of proven hardware with flight heritage. The plasma thruster, PPS 1350-G, was

developed by Snecma (France) and its control electronics (Power Processing Unit, PPU) come

from Thales Alenia Space ETCA (Belgium). Two thruster orientation mechanisms, developed by

Thales Alenia Space Cannes, hold two thrusters each and allow orientation in two

18

perpendicular directions. The system includes two off-the-shelf xenon fuel tanks of 68 litres

each; however, larger xenon tanks each of 105 litres are being developed at Thales Alenia

Space in Italy.

1.4.6. Chemical propulsion

The Alphabus Chemical Propulsion System (CPS) is a helium-pressurised bipropellant system

using monomethylhydrazine (MMH) as the fuel and mixed oxides of nitrogen (MON-3) as the

oxidiser. It has a 4200 kg total propellant mass capacity scalable down to 3500 kg, with 16

reaction control thrusters and a 400 N apogee engine. EADS Astrium Lampoldshausen

(Germany) is responsible for the CPS subsystem and provides the reaction control thrusters

and apogee engine along with most CPS components. A new high-efficiency 500 N apogee

engine is being developed at EADS Astrium Lampoldshausen. The titanium carbon-fibre over-

wrapped propellant tanks were developed by MT Aerospace. With a volume of up to 1925 litres

and a dry mass of less than 85 kg, they are among the world’s largest yet lightest satellite

tanks ever built.

The system includes three helium tanks of 90 litres each from EADS Astrium Aquitaine

(France). An upgrade to a larger tank configuration of two tanks of 150 litres each is planned.

The development of the larger helium tanks is in progress at Thales Alenia Space in Italy.

19

1.4.7. Electrical power architecture

The overall electrical configuration has been designed to allow efficient powering of payload

units. A primary 100 V regulated power bus with structure return is distributed to payload

units with aluminium bus bars protected through decentralised fuse boxes. The 100 V Power

Supply Regulator (PSR) and the lithium ion modular battery configuration allows for efficient

power regulation. The PSR was developed by EADS Astrium (France). The battery modules and

their latest generation ‘G5’ lithium ion battery cells were developed by Saft (France) for

Alphabus. The fuse box was developed by EADS Astrium Crisa (Spain).

1.4.8. Data handling

The data handling subsystem uses maximum synergy with EADS Astrium and Thales product

lines. It is composed of the onboard computer, the Satellite Management Unit (SMU), the Data

Bus Network (DBN) and the Platform Interface Unit (PFDIU), inherited from the Thales Alenia

Space Spacebus® 4000. The modular concept of the PFDIU is scaled to platform needs,

embedding interfaces for Plasma Propulsion Mechanism control, propulsion hardware control,

heater lines, pyrotechnic lines, and up to eight antenna 2-axis pointing mechanism controls

under a fine-tracking control loop. The DBN will use 1553 and OBDH/RS485 data buses.

The SMU was developed by Thales Alenia Space (France and Italy); the PFDIU was developed

by Thales Alenia Space (France) and Thales Alenia Space ETCA (Belgium). This data handling

system is complemented by a Payload Interface Unit, which is inherited from EADS Astrium’s

Eurostar E3000 and allows for Payload Unit TM/TC devices with discrete or Low Speed Serial

Bus interfaces. The on-board software was developed by Thales Alenia Space Cannes and

validated by Critical Software (Portugal).

1.4.9. Attitude Determination and Control System

The very accurate and flexible Alphabus Attitude Determination and Control System (ADCS) is

inherited from the zero-momentum, four-reaction-wheel control concept of the Spacebus®

4000, with three-axis determination using a star tracker and an accurate on-board orbit

propagator and precise on-board time. It also includes a gyroscope and a coarse sun sensor.

The reaction wheel has been developed by RCD (Germany) in different versions (angular

momenta of 18, 25 and 50 Nms); a new active pixel sensor-based star sensor was developed

by Galileo Avionica (Italy); a new Hemispherical Resonator Gyroscope was developed by

Sagem (France) with Syderal (Switzerland) and a new coarse sun sensor was developed by

TPD TNO (The Netherlands).

1.4.10. Antenna Tracking System

Alphabus is designed to accommodate an Antenna Tracking System (ATS) based on radio-

frequency sensing in order to reach 0.05° half-cone target performance. Up to eight antennas

can be controlled by the ATS. This system allows for individual pointing control of each

antenna, compared to the standard ADCS body control, hence compensating for each beam-

specific error with a higher control bandwidth. An onboard closed loop is implemented in the

onboard software.

20

1.4.11. Key features

Alphabus Service Module in geomobile configuration, compatible with Ariane 5 and

Proton 4 m-diameter fairing;

Repeater Module built in halves, for ease of payload accommodation;

Total launch mass: more than 6.5 tons;

Total electrical power: 12 kW.

Nominal lifetime 15 years

Payload power 22 kW

Satellite mass 8.8 tons

Payload mass 2 tons

Typical payload capacity 230 transponders, equivalent to:

more than 1000 television channels

more than 200 000 audio channels

Payload control data handling 1553 bus

Structure

Spin-formed carbon fiber central tube

and additional carbon and aluminum

panels

Cross section: 2800 x 2490 mm

Launcher Interface: 1666 mm

Modular concept Antenna module for:

Easier antenna accommodation

Efficient assembly and test

Attitude and orbit control

Gyros

Star and Sun sensors

Reaction wheels

Chemical propulsion

Apogee engine and 16 thrusters

2 propellant tanks (with up to 4200 kg

of bi-propellants)

Helium tanks (3 x 90 liters)

Electrical propulsion

Xenon tanks (max. 350 kg)

PPS 1350 thrusters on orientation

mechanisms

Power generation and distribution

2 gallium arsenide solar wings with 4–

6 panels

Power supply and power distribution

offering both 100 V and 50 V regulated

buses

Modular lithium-ion battery

21

Equipment Supplier

STRUCTURE

Central tube EADS CASA

Primary structure Thales Alenia Space Cannes

Secondary structure Ruag Space (Switzerland), Thales Alenia

Space Cannes (France), EADS CASA (Spain)

ELECTRICAL POWER SYSTEM

Solar array EADS Astrium Ottobrunn (Germany)

Low-Shock Release Unit for solar

array deployment

Ruag Space (Switzerland), EADS Astrium

Ottobrunn (Germany)

Power System Regulator EADS Astrium (France)

Battery modules based on 5th-

generation lithium ion technology

Saft (France)

High-power Solar Array Drive

Mechanism

EADS Astrium Stevenage (UK)

High-power equipment EADS Crisa (Spain)

CHEMICAL PROPULSION SYSTEM (CPS)

Apogee boost motor to transfer

satellite into geostationary orbit

EADS Astrium Lampoldshausen (Germany)

Reaction control thrusters for

keeping satellite on station

EADS Astrium Lampoldshausen (Germany)

Propellant tank MT Aerospace (Germany)

Helium tank (90 litre) EADS Aquitaine (France)

New helium tank (150 litre) under development

Propellant and helium filters Sofrance (France)

Pressure transducers Bradford Engineering (The Netherlands)

Various CPS items EADS Astrium Lampoldshausen (Germany)

ELECTRICAL PROPULSION SYSTEM (EPS)

Xenon tank (68 litre) Off-the-shelf

New xenon tank (105 litre) under development

Various EPS propulsion items Ampac (UK and Ireland)

Plasma thrusters, PPS 1350-G Snecma (France)

Power Processing Unit, Filter Unit

and Hot Interconnection Boxes

Thales Alenia Space ETCA (Belgium)

Thruster Orientation Mechanism

(TOM)

Thales Alenia Space Cannes (France)

ATTITUDE DETERMINATION AND CONTROL SYSTEM

High-momentum reaction wheel RCD (Germany)

Coarse Sun sensor TPD TNO (The Netherlands)

Hemispherical resonator

gyroscope

Sagem (France) with Syderal (Switzerland)

Active pixel sensor-based star

tracker

Galileo Avionica (Italy)

DATA HANDLING SYSTEM

Satellite Management Unit Thales Alenia Space (France and Italy)

Platform Data Interface Unit Thales Alenia Space (France and Belgium)

Payload Interface Unit EADS Astrium (France

22

2. Project Management

The European Cooperation for Space Standardization (ECSS) is the organization established to

develop a coherent, single set of user-friendly standards for use in all European space activities.

ECSS documents (Standards, Handbooks and Technical Memoranda) are organized in four branches:

Space engineering

Space project management

Space product assurance Space sustainability

The branches are themselves organized in several disciplines, as follows.

23

WBS

Satellite Platform Design

WP 1000

Management

Procurement

Interface Management

Customer Relations

WP 2000

Mission Analysis

Literature Review

Requirements Definition

Operative and Enviromental

Contraints

WP 3000

Subsystems Design

Power System

Structure

Telecommunications System

Propulsion

WP 4000

Assembly Integration and

Verification

Structural Tests

Acoustic Tests

Electromagnetic Tests

Thermal Tests

WP 5000

Hardware and Software

Development

WP 6000

Logistics

Transportation

Test Ground

Planning Human Resources in Situ

2.1. Work breakdown Structure (WBS)

The WBS is the principal structure used in managing a project and provides a framework for

managing cost, schedule and technical content. It divides the project into manageable work

packages, organized according to the nature of the work by breaking down the total work to be

performed into increasing levels of detail.

The WBS is derived from the product tree, selected elements of which are extended to

include support functions (i.e. management, engineering, product assurance) and associated

services (e.g. test facilities).

An example of a ECSS WBS applied to Spacecraft Design is shown in the figure below.

24

3. Subsystems Preliminary

Design

3.1. Structure

The structure is thought to be similar to the one of alphabus and adaptable to Ariane 5. It is

composed by two modules:

- a central tube (diameter: 1.666 m, height: 3 m) with four shear panels and 4

rectangular panels as its basis (height: 1 m), called ‘service module’;

- two rectangular panels (height: 4 m) with three shelves, called the ‘repeater

module’.

25

The material is supposed to be carbon fiber (graphite/epoxy HTS) for the entire bus and its

properties are supposed to be isotropic. The real material has the following properties:

- Ultimate tensile strength: 1337 MPa;

- Ultimate transversal strength: 66 MPa;

- Longitudinal Young’s Modulus: 151 GPa;

- Poisson’s Modulus: 0.3;

- Density: 1490 kg/m3.

The assumption is obviously rough, thus the results must be considered with big values of

safety factors.

A constraint of total rigidity is applied to the basic circle of the structure, where the bus is

assembled to the launcher.

Static accelerations due to gravity loads are applied to the whole structural mass according to

the data given by Ariane 5 User’s guide:

- Longitudinal load: 5 g;

- Transversal load: 0.5 g.

Nonstructural distributed masses are added:

- on the central tube, representing the full chemical propellant tanks (2 tanks, see

par. 3.3.1.1.);

- on the shear panels, representing the Helium and Xenon tanks (2 tanks each,

see par. 3.3.1.2. and 3.3.2.);

- on the rectangular panels, representing the solar arrays and the batteries (see

par. 3.2.1. and 3.2.2.);

- on the shelves, representing payload (2 tons, as maximum satellite capacity).

The FEM analysis is conducted in order to size the thickness of the structure. The mesh is

chosen linear tetrahedral of size 100 mm and absolute sagittal of 30 mm. Finer meshes are not

supported by the calculator used.

26

The following thicknesses are set:

Service Module

- tube: 8 mm;

- shear panels: 6 mm;

- rectangular panels: 8 mm.

Repeater Module

- rectangular panels: 8 mm;

- shelves: 8 mm.

The results show a little deformation of the structure, with a maximum displacement of about

1 mm. The stresses are calculated by Von Mises Criterion, therefore they are not very

meaningful. However, the result of about 30 MPa of maximum stress reveals the structure

should be strong enough to sustain the transversal and shear loads, which are the most

destructive for composite materials (orthotropic).

The resulting total mass of the structure is 811.2 kg (about 12% of the total satellite mass).

All the analysis has been conducted in CATIA v5 r18.

27

28

3.2. Power system

3.2.1. Solar Panels

The mission is supposed to last 15 years. The power requested by the bus (always on) at the

end of its life is 12 kW. The balance of the energy per orbit at beginning of life:

The damage of solar cells is supposed to be equal to 3% per year (from Larson & Werz), of

which up to 0.5% is due to radiation, assuming a total radiation degradation of 6% on solar

cell voltage when exposed to a total equivalent fluence of 1014 cm-2 (@1 MeV, from Spectrolab

29.5% XTJ Solar Cells datasheet; all next assumptions about solar cells will be referred to this

datasheet).

By SPENVIS Information System (ESA) these requirements are verified and met by a

coverglass thickness of 3 mils (76.2 micron) made of fused silica.

Cell: Spectrolab 3J EOL

Electron/proton damage ratios:

Pmax = 870.0

Voc = 1020.0

Isc = 565.0

Coverglass

material: fused silica

density: 2.20 (g/cm3)

Summary of 1 MeV equivalent electron fluences (cm-2

)

Coverglass thickness Total Trapped electrons Trapped protons

g cm-2

mils micron Pmax Voc Isc Pmax, Voc, Isc Pmax Voc Isc

0.0000 0.0 0.00 4.741E+18 5.558E+18 6.361E+18 1.451E+14 4.740E+18 5.558E+18 6.361E+18

0.0056 1.0 25.40 1.284E+14 1.296E+14 1.310E+14 1.217E+14 6.746E+12 7.909E+12 9.303E+12

0.0168 3.0 76.20 9.810E+13 9.810E+13 9.810E+13 9.810E+13 1.000E-12 1.000E-12 1.000E-12

0.0335 6.0 152.40 7.701E+13 7.701E+13 7.701E+13 7.701E+13 1.000E-12 1.000E-12 1.000E-12

0.0671 12.0 304.80 5.326E+13 5.326E+13 5.326E+13 5.326E+13 1.000E-12 1.000E-12 1.000E-12

0.1118 20.0 508.00 3.645E+13 3.645E+13 3.645E+13 3.645E+13 1.000E-12 1.000E-12 1.000E-12

0.1676 30.0 762.00 2.478E+13 2.478E+13 2.478E+13 2.478E+13 1.000E-12 1.000E-12 1.000E-12

0.3353 60.0 1524.00 9.903E+12 9.903E+12 9.903E+12 9.903E+12 1.000E-12 1.000E-12 1.000E-12

From the first table it is clear that proton fluences can be negligible.

Since

Then .

29

Using equation (1) it is possible to find The following parameters were

assumed:

-

-

-

Since GEO with respect to the Sun’s orbit has the maximum inclination of 23.5°, the solar

panels also have such maximum inclination with respect to the solar rays, therefore

The power generated by solar panels per unit area can be calculated as

[

]

Being

-

(at minimum Sun activity)

-

Therefore, the total area of solar cells is

In order to estimate the total area of the solar panels, it is necessary to calculate the number

of solar cells. Solar cells are installed both in series and in parallel.

Being

-

-

Each cell is and produces 1 W, therefore 45 W are produced by solar cells in series.

Series are linked in parallel one another, so the number of series in parallel are

The total number of cells is then

The effect of installation decreases the power generated by the solar panel till 376 W/m2 (by a

factor of 0.95). Therefore the total area of the panel is

30

The efficiency of solar arrays decreases also because of high temperature. All the calculations

made so far have been supposed at the standard condition of 25°C. Efficiency decreases of

about 0.5%/°C over 25°C.

The equilibrium temperature at Sun’s exposure can be calculated by the heat balance:

Being:

- , the equivalent absorptance of the panels;

- , the electrical efficiency of the panels;

-

, the average solar power constant per unit area

-

, the area exposed to the Sun with respect to the emitting area, which

includes both front and back areas;

- , the factor of view, representing the percentage of area effectively seen

by the solar rays (the product is the Sun projected area); it is assumed

1 as the most frequent condition;

-

, the Stefan Boltzmann’s constant;

- , the equivalent emittance of the panels.

Neglecting, with respect to solar radiation:

- Earth radiation;

- Albedo;

- Internal power radiation;

- Space emission (Space is supposed to be at 0 K, because usually it’s at 3÷4 K).

, so a variation of almost 10° can be supposed, producing a degradation of 5%.

Therefore

Finally the total solar arrays area is about 50 m2.

The extension of the satellite with both the solar wings deployed is about 40 m (from Alphabus

datasheet), so together they are about 35 m long and 1.4 m wide. Each wing hosts 4 panels of

about (1.4X4.4)m2 area.

Supposing a specific weight of 2 kg/m2 the total weight of solar arrays is about 100 kg, that is

25 kg per solar panel.

3.2.2. Batteries

Since the mission length is 15 years, secondary power storage is required. The storage is

needed for two 45-day eclipse periods per year with no more than 72 minutes each day. Saft

Li-Ion batteries for GEO are selected. Their operating range of temperature is between 10°C

and 40°C and their cell voltage between 3.6 V and 4.1 V.

First of all, the number of batteries in series is calculated as

31

considering the minimum cell voltage.

Therefore the energy stored by one line of batteries in series is

Being the capacity of a cell (choosing the model Saft VES 180 and assuming a

Depth of Discharge of 60%).

The number of batteries in parallel is then

The total energy is then

Since the specific energy for such batteries is about 165 Wh/kg, the weight of the batteries is

about 122 kg, that is about 1.1 kg each cell ( ).

32

3.3. Propulsion System

3.3.1. Chemical propulsion

Chemical propulsion is used to transfer the spacecraft from GTO to GEO (apogee motor) and

for reaction control (16 thrusters). It is chosen a regulated-pressure liquid bipropellant system

using MMH as fuel, N204 as oxidizer and Helium as pressurant gas. The two propellant tanks

are stored into the central tube of the service module, while the two Helium tanks are attached

to its shear panels.

Sizing the system means finding the following masses:

- Propellant mass (fuel + oxidizer)

- Propellant tanks mass

- Helium mass

- Helium tanks mass

The sum of such masses makes the total weight of the chemical propulsion system.

3.3.1.1. Propellant and propellant tanks

Given the ΔV budget of the mission, by Tsiolkovsky equation it is possible to find the required

propellant mass.

The ΔV budget is composed by the contribution for GTO/GEO transferal and the one for orbital

control:

The Tsiolkovsky equation is

(

)

Since

Then

[ ]

Being

- the satellite mass at launch;

- the apogee engine specific impulse (it is assumed equal to the one of

thrusters);

- Universal gravity constant g0 assumed constant with altitude.

A value of mixture ratio O/F must be chosen in order to find the mass of fuel and oxidizer

separately. It is chosen O/F=1.64 because this value results in tanks of equal size.

But

33

Then

(

)

Given the density of fuel and oxidizer, it is possible to quantify their volume.

Being

-

(MMH);

-

(N2O4).

Given the maximum propellant operating pressure (MEOP) and chosen the shape and the

material, it is possible to size the propellant tanks.

For simplicity a spherical shape is chosen. The material chosen is Titanium overwrapped with

carbon fiber, which allows both strength and lightness. Its maximum tensile strength is

supposed to be 1000 MPa (being Titanium alloy’s 880 MPa) and its density is supposed to be

1490 kg/m3 (the one of carbon fiber, because its percentage is bigger than Titanium’s).

The maximum expected operating pressure (MEOP) is supposed to be 25 bar.

By the balance of external and internal forces

The thickness of the tank is

Being

- ;

- (

)

, from the expression of the volume of a sphere.

Therefore the mass of the propellant tank is

The total mass is

34

3.3.1.2. Pressurant gas and its tanks

First of all, the ratio between the pressurant gas volume and propellant volume can be

calculated by the conservation of energy from initial pressurization to final pressurization (with

the assumption of adiabatic transformation and ideal gas):

Being

- , the ratio of specific heats of Helium;

- the maximum expected initial pressure of the gas;

-

, the overpressure needed to let propellants in the combustion

chamber (assuming ).

Therefore the gas volume is

Let’s now see what changes if we assume an isotherm transformation: the temperature is

almost constant because the engine is continuously switched on throughout the mission. By

the expression pV=cost:

Last result will be considered.

Assuming the same material and shape of the propellant tanks and repeating the same

calculations, it is found

The Helium mass is

Being (@180 bar, 25°C, R=2078 J/kg K).

Finally, the total pressurant system mass is

35

3.3.2. Electric Propulsion

The electric propulsion consists in 4 plasma thrusters used for north/south on orbit station-

keeping. The system includes two off-the-shelf Xenon fuel tanks, attached to the shear panels

of the service module.

The budget is

Being the electric motors switched on for one hour a day. Then the total impulse is

Given the Isp, it is possible to calculate the power to thrust ratio and the mass of Xenon:

Being (from SNECMA PPS 1350-G/-S):

-

- , the overall system efficiency.

Therefore the given power and thrust are P=1.4 kW and F=130 mN.

The sizing of the tanks can be done as well as for chemical propulsion. Given the total Xenon

mass and dividing it per 2 tanks, the volume of one tank is obtained:

Being (liquid).

Given the same shape and material of the previous tanks, it is found

Being

- Rint = 0.219 m

- MEOP = 150 bar

Therefore

The total mass of the electric propulsion system is then

36

3.4. Telecommunications System

3.4.1. Link Budget

The sizing of the TLC system is based on the value of the link budget, which is defined as the

ratio between the received energy per bit and the noise energy per bit:

Being

- , the equivalent isotropic radiative power;

- , the required transmit power of one transponder, to be sized;

- , the propagation loss in transmission;

- (

) , the transmit antenna gain, assuming 1° as

antenna solid angle of transmission ( );

-

, assuming the diameter of the antenna equal to 0.5 m;

- the transmission path loss, a function of many factors, such as rainfall

density; this value is for transmission with rainfall at 20 GHz and it is equivalent

to a gain attenuation of 3 dB;

- typical value for the efficiency of the area of the transmit antenna;

- , the propagation loss in reception, for safety reasons less than in

transmission;

- , the average distance runnable by the signal from GEO to a point of

Earth’s surface;

-

, the Boltzmann’s constant;

- , the system noise temperature;

- , the receiver noise bandwidth; one transponder contains 12 channels

in a bandwidth of 36 MHz, so one channel has a bitrate R=3 Mbps, therefore one

receiver has R=3 Mbps; it is assumed B=R.

Assuming a bit error rate BER=10-6, as usual value for safety reasons (it’s expected one wrong

bit every million) and choosing the modern code Viterbi, it is possible to find the value of the

link budget on a diagram reported by Larson-Werz:

Therefore applying the conversion to dB to its expression, it is found the required transmit

power for :

Which must be multiplied for 190 transponders to find the total transmit power:

37

4. Bibliography

All the general information for preliminary sizing was taken from:

W.J. Larson, J.R. Werz, Space Mission Analysis and Design, 3rd edition. Space

Technology Library, 1999.

Arianespace, Ariane 5 User’s Manual, Issue 5, Revision 1. July 2011.

http://www.arianespace.com/launch-services-

ariane5/Ariane5_users_manual_Issue5_July2011.pdf

Here follows the bibliography for more specific topics.

Funds And Investments

Staff Writers, Worldwide Government Spending On Space To Flatten Over The Next Five

Years. Paris, France (SPX) Feb 18, 2011 -

http://www.spacedaily.com/reports/Worldwide_Government_Spending_On_Space_To_

Flatten_Over_The_Next_Five_Years_999.html

Satellite Manufacturers

Wikipedia, the Free Encyclopedia, Satellite Manufacturers. http://en.wikipedia.org

Buses

Wikipedia, the Free Encyclopedia, Satellite Bus.

Wikipedia, the Free Encyclopedia, Comparison of Satellite Buses.

Europe

Wikipedia, the Free Encyclopedia, Alphabus.

Thales Group, Spacebus 4000.

http://www.thalesgroup.com/Portfolio/Space/space_product_spacebus_4000/

OHB Germany, SmallGEO (Brochure). www.ohb.de

Wikipedia, the Free Encyclopedia, Eurostar (spacecraft).

Russia

Wikipedia, the Free Encyclopedia, Ekspress.

Reshetnev Company, GEO satellites. http://www.iss-reshetnev.com/?cid=53

38

USA

Boeing, Boeing 601. http://www.boeing.com/boeing/defense-

space/space/bss/factsheets/601/601fleet.page?

Boeing, Boeing 702. http://www.boeing.com/boeing/defense-

space/space/bss/factsheets/702/702fleet.page?

Wikipedia, the Free Encyclopedia, LS-1300.

Gunter’s Space Page, Space Systems Loral LS-1300.

http://space.skyrocket.de/doc_sat/ssloral-1300.htm

Orbital, GEOStar (Brochure).

http://www.orbital.com/SatellitesSpace/Communications/index.shtml

Lockheed Martin, A2100 (Brochure). www.lockheedmartin.com

Asia

China Academy of Space Technology, CAST3000.

http://www.cast.cn/CastEn/Show.asp?ArticleID=39325

China Academy of Space Technology, DFH-3B. http://www.cast.cn/CastEn/Show.asp?ArticleID=39322

Mitsubishi Electric, DS2000.

http://www.mitsubishielectric.com/bu/space/products/platform/index.html

Indian Space Research Organization, INSAT 4B.

http://www.isro.org/satellites/insat-4b.aspx

Indian Space Research Organization, GSAT 14. http://www.isro.org/satellites/gsat-

14.aspx

Alphabus

European Space Agency, Alphabus factsheet. www.esa.int

European Space Agency, Alphabus Structure (Bulletin ESA 142).

Centre National d’Etudes Spatiales, Alphabus Dossier. www.cnes.fr

Project Management

European Cooperation for Space Standardization, www.ecss.nl

Propulsion System

EADS Astrium (Airbus Group), Satellite Propulsion Systems – Alphabus.

http://cs.astrium.eads.net/sp/spacecraft-propulsion/showcase/alphabus.html

ATK, Pressurant Tanks. http://www.psi-pci.com/Pressurant_Tanks.htm

Prof. M. Di Giacinto, Dispense del Corso di Propulsione Spaziale (ITA). Sapienza

University of Rome, 2013.

SNECMA, A Complete Range of Plasma Thrusters (brochure).

http://www.snecma.com/IMG/files/gammeplasmiqueen_modulvoir_file_fr.pdf

Power System

Spectrolab, 29.5% XTJ Solar Cells and Solar Panels (datasheets).

http://www.spectrolab.com/DataSheets/cells/PV%20XTJ%20Cell%205-20-10.pdf

http://www.spectrolab.com/DataSheets/Panel/panels.pdf

SPENVIS, Space Environment Information System (ESA). www.spenvis.oma.be

Saft, Saft Li-Ion batteries (Battery Information Sheet).

http://www.saftbatteries.com/battery-search/ves-vl-batteries-satellites