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American Institute of Aeronautics and Astronautics

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100-300KG CLASS SATELLITE STANDARD BUS SYSTEM DESIGN AND APPLICATIONS

Keita Fukuzawa, Naohiko Abe, and Shinobu Nakamura

Mitsubishi Heavy Industries, Ltd.

1200, Higashi-Tanaka, Komaki, Aichi 485-8561, Japan

Ryutaro Suzuki and Shinichi Kimura

Communications Research Laboratory

4-2-1, Nukui-Kitamachi, Koganei, Tokyo 184-8795, Japan

Maki Akioka

Communications Research Laboratory

3601, Isosaki-Cho, Hitachinaka, Ibaraki 311-1202, Japan

Abstract

This paper describes the design and applications of

the standard bus system for 100-300 kg class satellites that is studied in MHI. It aims to meet the recent demand for fast space demonstration opportunities at low cost.

One of the design concepts of the present bus system is to separate the satellite bus module and mission payload module thermally and mechanically as much as possible. Another is to apply standard electric interface. The architecture of the bus system is kept as simple as possible, and off-the-shelf and flight-proven bus components are selected. These design approaches will make it possible to minimize the design change of

the bus system across missions. As a result, user’s procurement cost and lead-time would be reduced without losing the adaptability.

Demonstrations of On -Orbit Servicing, Corona Observation from CRL, and Optical inter-satellite communication from NeLS are given as the examples of the applications of the bus system. The bus system can be applied to practical missions by plural satellites or formation flying system in the future as well.

1. Introduction

In comparison with the ones that are installed to

current satellites, highly sensitive and fine resolution

sensors, and small and high performance electronics have already been developed on the ground. If these state-of-the-art technologies could be utilized, space science such as X-ray or infrared astronomy would dramatically progress. And smart satellite for the practical service such as communication, remote sensing, or navigation would be realized. In this way, the establishment of infrastructures that enable researchers to perform space demonstration of their new technologies in a “faster” and “inexpensive” way is desired gradually in the field of the recent space development.

A small satellite is proposed as a solution of these market’s requests. It has the advantages of short -term development, lower procurement cost, and more launch

opportunities as a piggyback satellite. Moreover, the performance of small satellites has been drastically improved with recent smaller electronic equipments with higher performance. Therefore, a single small satellite or plural small satellites will accomplish the practical mission that would be performed currently by large satellite. Furthermore, large launch vehicles such as H-IIA and ARIANE-V plan or have begun to provide the surplus launching ability to small satellites. Figure 1 shows the market trend of 100-300kg class satellite1. In U.S. and Europe, the demand for this class has increased recently for its easier development. In Japanese market, it is also expected that the demand will increase in the next decade.

21st International Communications Satellite Systems Conference and Exhibit AIAA 2003-2311

Copyright © 2003 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.


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In this paper, the design and the applications of the standard bus system for 100-300kg class satellites, which is being studied at Mitsubishi Heavy Industries, Ltd. (MHI), are described. The present bus system aims to provide space demonstration opportunities to wide range of users rapidly at lower cost.

2. Bus System Design

2.1 Design Concept

The purpose of this study is to provide users with the standard bus system, which features easy accommodation of wide range of missions. To achieve this purpose, the following design concepts are adopted. (1) To separate the satellite bus module and mission

payload module thermally and mechanically as much as possible.

(2) To apply a standard electric interface between the bus on-board computer and mission processor, which controls mission components and handles

mission data from them.

These concepts contribute to the minimization of design changes of the satellite bus module across missions. As a result, non-recurring cost can be reduced and bus system test can be simplified. Moreover the whole satellite development period can be reduced, since mission payload module and bus module are assembled simultaneously and two modules are

assembled in the end. To reduce the cost and period of the present bus system development and to provide more reliable bus system to users, the following measures are taken. (3) To reduce the number of bus components by

integrating multiple functions into one box. (On-Board Computer and Power Control Unit as

discussed later)

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Figure 1 Market trend of 100-300kg class satellite1

Table 1 Target Specifications

φ1m x H 0.6mOctagonal Cylinder

Bus Size

6 deployment panels GaAs triple-junction cells

Solar Array Panel

GTO, SSO, LEO Orbit

28 +4, –6 VNominal Voltage

Maximum 24kg (Hydrazine monopropellant)

Propellant Capacity

1 tank, 12 thrustersPropulsion

S-band, X-band (Option)Downlink Band

47000 NsecTotal Impulse

Better than 0.1degPointing Accuracy

Zero Momentum3 axis Stabilized

Attitude Control System


Total Size

70WBus Power

Average 120W (GTO)Payload Power

S-bandUplink Band

More than 6 month (GTO)Mission Life

Around 150 kg Total Mass

50 - 70 kgPayload Mass

Less than 100 kg (WET)Bus Mass


Bus Size

6 deployment panels GaAs triple-junction cells

Solar Array Panel

GTO, SSO, LEO Orbit

28 +4, –6 VNominal Voltage

Maximum 24kg (Hydrazine monopropellant)

Propellant Capacity

1 tank, 12 thrustersPropulsion

S-band, X-band (Option)Downlink Band

47000 NsecTotal Impulse

Better than 0.1degPointing Accuracy

Zero Momentum3 axis Stabilized

Attitude Control System


Total Size

70WBus Power

Average 120W (GTO)Payload Power

S-bandUplink Band

More than 6 month (GTO)Mission Life

Around 150 kg Total Mass

50 - 70 kgPayload Mass

Less than 100 kg (WET)Bus Mass


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(4) To select off-the-shelf and flight-proven components. (such as attitude sensors, attitude actuators, secondary batteries, S-band transponder, and components for propulsion subsystem)

2.2 Overview

The bus system for 100-300kg class satellites is currently being studied in MHI. It is designed on the assumption of the joint mission from Communication Research Laboratory (CRL) and Next -generation LEO System Research Center (NeLS) as discussed later. This will be the first mission of the present bus system and the target launch year is 2006. In the design process,

however, the future standardization is considered to retain the flexibility to accommodate various missions.

Table 1 shows the target specification of the system. The bus module is shaped roughly as an octagonal cylinder with the diameter of 1m and the height of 0.6m. The mass including mission payload module is around 150kg. This comes from the piggyback launch

limits of H-IIA or ARIANE-V. Moreover the mass of around 150kg is judged to be a proper balance between technical performances and demands for technology demonstration satellite from wide range of users. Since the main application of the present bus system is demonstration of new technologies and components in space, it is designed for more-than-6-month mission life in Geosynchronous Transfer Orbit (GTO).

Exploded view and Functional block diagram of the present bus system are shown in Figure 2 and Figure 3, respectively. The On -Board Computer (OBC) handles attitude control, data processing, and telemetry command processing. The Power Control Unit (PCU) manages the power, thruster and latch valve drive, and heater control according to the design policy to reduce

the number of equipments. As for other bus components , existing off-the-shelf and flight-proven ones are selected to minimize non-recurring costs, development period, and risks .

To reduce launch costs as well, the present bus system is designed as a piggyback satellite of H-IIA instead of dedicated launchers. One or two satellites

Star Tracker

Sun Sensor Assembly

Propellant Tank

Reaction Wheel Drivers

Inertial Measurement Unit

Reaction Wheels

S-band Transponder

S-band Diplexer

S-band Antenna

S-band Hybrid

Power Control Unit

Solar Array Paddle

Valve Module

Battery

On-Board Computer

Solar Array Drive Electronics

Solar Array Drive Assembly

Star Tracker

Sun Sensor Assembly

Propellant Tank

Reaction Wheel Drivers

Inertial Measurement Unit

Reaction Wheels

S-band Transponder

S-band Diplexer

S-band Antenna

S-band Hybrid

Power Control Unit

Solar Array Paddle

Valve Module

Battery

On-Board Computer

Solar Array Drive Electronics

Solar Array Drive Assembly

Figure 2 Exploded view


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ride inside payload attach fitting (PAF) of primary satellite. 2.3 Attitude Control and Data Handling Subsystem

The target value of pointing accuracy for the present bus system is less than 0.1 deg. This is because the pointing budgets of recent technology demonstration and observation satellites are from approximately 0.025 to 0.5deg. The requests from CRL and NeLS joint mission are also considered in setting targets of attitude control.

To meet the target specification, the attitude and orbit control subsystem (AOCS) mode is 3-axis stabilization. High precision attitude control is achieved by the zero momentum system of three reaction wheels. In accordance with mission requests, attitude sensors can be chosen from various combinations of Sun Sensor Assembly (SSA), Earth Sensor Assembly (ESA), Star Tracker (STT), and Inertial Measurement Unit (IMU).

All of the attitude control functions are performed by OBC. OBC is constructed in a layer configuration as shown in Figure 4. Various sensors and actuators that have different electrical interfaces can be easily

connected to OBC by means of exchangeable interface boards. As for inertial pointing, for example, the nominal attitude determination is performed by STT and IMU to achieve high pointing budgets. The main actuator of the

present bus system is Reaction Wheel (RW). The present bus system has thrusters in order to unload RWs and to serve rendezvous or fly-around maneuvers . In acquisition and safe hold mode, SSA will provide spacecraft attitude with respect to the sun with IMU providing rate information.

：DevelopedComponents

THERMAL

Heater､MLI etc.

Thruster

PROPULSION

TankTube

S-ANT

X-ANT

SAP

BAT

PCU

OBC

PayloadCOMMUNICATION

ATTITUDE CONTROLDATA HANDLING

Structure

STRUCTURE

Bracket etc.

MISSION

POWER

SHNT

RS-422

【LEGEND】

：Off-the-shelf Components

：OPTIONGPSR

ESA

X-TX

SAP

SSA

STT

IMU

RW

【ABBREVIATION】BAT : Secondary BatteryESA : Earth Sensor AssemblyGPSR : GPS ReceiverIMU : Inertial Measurement UnitMLI : Multi-Layer InsulationOBC : On-Board ComputerPCU : Power Control UnitRW : Reaction WheelSADA : Solar Array Drive AssemblyS-ANT : S-band AntennaSAP : Solar Array PanelSHNT : Shunt RegulatorS-HYB : S-band HybridSSA : Sun Sensor AssemblySTT : Star TrackerS-XPDR : S-band TransponderX-ANT : X-band AntennaX-TX : X-band Transmitter

Wire Harness

INSTALLATION

Latch Valve

SADA

……

X-TX

S-XPDR

：DevelopedComponents

THERMAL

Heater､MLI etc.

Thruster

PROPULSION

TankTube

S-ANT

X-ANT

SAP

BAT

PCU

OBC

PayloadCOMMUNICATION

ATTITUDE CONTROLDATA HANDLING

Structure

STRUCTURE

Bracket etc.

MISSION

POWER

SHNT

RS-422

【LEGEND】

：Off-the-shelf Components

：OPTIONGPSR

ESA

X-TX

SAP

SSA

STT

IMU

RW

【ABBREVIATION】BAT : Secondary BatteryESA : Earth Sensor AssemblyGPSR : GPS ReceiverIMU : Inertial Measurement UnitMLI : Multi-Layer InsulationOBC : On-Board ComputerPCU : Power Control UnitRW : Reaction WheelSADA : Solar Array Drive AssemblyS-ANT : S-band AntennaSAP : Solar Array PanelSHNT : Shunt RegulatorS-HYB : S-band HybridSSA : Sun Sensor AssemblySTT : Star TrackerS-XPDR : S-band TransponderX-ANT : X-band AntennaX-TX : X-band Transmitter

Wire Harness

INSTALLATION

Wire Harness

INSTALLATION

Latch Valve

SADA

……

X-TX

S-XPDR

Figure 3 Functional block diagram

Figure 4 On-Board Computer (Breadboard model)


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For nadir pointing, ESA may be installed on the satellite as an option sensor. And fine SSA and ESA combination performs the attitude determination. If the satellite orbit is inside GPS orbit shell such as LEO or SSO, GPS receiver (GPSR) is also applicable. In this case, STT and IMU combination can achieve high pointing accuracy, while the orbit position of the satellite is obtained from GPSR.

OBC also performs Data Handling (DH). The CPU is selected from radiation-hardened Digital Signal Processor. Dual CPU redundancy improves the DH and AOCS. RS-422 is used for serial data interface between OBC and mission processor to keep electrical interface simple. The memory of OBC is around 16Mbyte for House Keeping (HK) data.

2.4 Power Subsystem

The power subsystem of the present system is a 28V, unregulated power bus. The target power budget of the present bus system is 120W average for mission payload module and 70W for bus module, respectively under the condition of GTO.

The main component of this power subsystem is Power Control Unit (PCU). The main function of PCU is to perform power control including battery charge/discharge control while meeting the requirement of the voltage of electronic components . PCU interfaces with the solar array, secondary batteries and the electrical loads. And it also manages thruster and

latch valve drive, heater control. The required power is generated by 2.2m2 solar

arrays that consist of triple junction GaAs cells mounted on the 6 deployment Al honeycomb panels. The high efficiency of triple junction GaAs cells contributes to smaller solar arrays. The solar array paddles utilize a single axis solar array drive assembly (SADA) in order to keep the panels best pointing to the sun. As for batteries, Lithium-ion cells are well suit ed for a small satellite because of their high energy density. It is planned that approximately 15Amps-hour Lithium-ion battery will be adopted for the present standard bus system. 2.5 Propulsion Subsystem

The present bus system requires propulsion

subsystem for translational acceleration as well as

unloading RWs. The schematic diagram of the propulsion subsystem is shown in Figure 5. It is hydrazine monopropellant system and consists of a propellant tank, thrusters, latch valves and some miscellaneous devices. The propellant tank and the thrusters are shown in Figure 6. The propellant tank can be loaded with the maximum 24kg of propellant. The thrusters can generate the thrust from 3N (BOL) to 1.5N (EOL) per one thruster and the minimum impulse bit of 0.06Ns.

The propulsion subsystem has twelve thrusters. Six thrusters are mounted on the top deck of the bus module, and other six ones are mounted on the bottom deck. They are called the fore thrusters and the aft thrusters, respectively. In order to minimize the cost,

fully redundant configuration is not adopted. The fore thrusters and the aft thrusters can be isolated with two latch valves on their failure instead. If one of the thruster valves fails to close, corresponding latch valve will be closed. And then the remaining thrusters will be operated in degraded function mode. In case of the fore thrusters failure, +X acceleration, roll, pitch and yaw control can still be maintained. In case of the aft thrusters failure, -X acceleration, pitch and yaw control can still be maintained.

Aft Thrusters Fore Thrusters

P

N2H4

GFD

PFD

【Legend】

P Pressure Transducer

Fill & Drain Valve

Filter

3 N thruster

1 2 3 4 5 6 7 8 9 10 11 12

3NTHR 3NTHC 3NTHC 3NTHS

ALV FLV

ATP FTP

Latch Valve

FLT3NTHR：right angle3NTHC：45deg cant3NTHS：straight nozzle

Aft Thrusters Fore Thrusters

P

N2H4

GFD

PFD

【Legend】

P Pressure Transducer

Fill & Drain Valve

Filter

3 N thruster

1 2 3 4 5 6 7 8 9 10 11 12

3NTHR 3NTHC 3NTHC 3NTHS

ALV FLV

ATP FTP

Latch Valve

FLT3NTHR：right angle3NTHC：45deg cant3NTHS：straight nozzle

Figure 5 Schematic Diagram

of Propulsion Subsystem


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2.6 Communication Subsystem

The satellite communicates with ground stations via S-band communication subsystem. The S-band subsystem provides command uplink, mission data downlink, and housekeeping data downlink. The block diagram of the communication subsystem is shown in Figure 7.

Dual S-band transponders (S-XPDR) are installed on the present bus system to maintain communications. In addition to telecommand and telemetry functions, S-XPDR also has ranging functions for spacecraft tracking. The footprints of two omni directional antennas cover almost all around the satellite.

Data formats for uplink and downlink streams conform to the CCSDS packet. Even at the GTO apogee, telecommand and telemetry data rate to the ground stations are estimated to be more than 0.5kbps and 1kbps, respectively. If this data rate is not satisfied with mission requirements, an X-band transmitter can be added as an option. 2.7 Structure and Thermal design

According to the design policy, the present bus

module is separated from a mission payload module

mechanically and thermally in order to accommodate various missions easily. A mission payload module is allocated 50-70kg and located on the top deck of the bus module. The shape of the present bus module is roughly an octagonal cylinder and constructed with

aluminum honeycomb sandwich panel. Except for some attitude sensors and antennas, bus components are installed inside of the bus module.

Most of the bus components on the present bus system are off-the-shelf and flight-proven products. Accordingly, the bus module structure is designed under the condition of the launch environments of H-IIA and ARIANE-V so that the vibration and shock levels on each component are satisfied with its historical Acceptance Test levels.

The present bus module is mainly designed with a passive thermal control system. The thermal control hardware includes multi-layer insulation (MLI), silver-deposited Teflon films, optical coatings, thermistors, and heaters. The PCU controls heaters by

monitoring temperatures with thermistors. To separate mission payload module and bus module thermally, MLI and GFRP spacers are used between them for radiative and conductive insulation unless mission components require thermal connection. The present bus module can be adapted to various thermal environments such as LEO, SSO, and GTO. The method of thermal design is not to change the placement of bus components but to adjust the ratio of MLI to silver-deposited Teflon radiator and the heater power.

3. Applications 3.1 Demonstration of On-orbit Service Technology2

The concept of Orbital Maintenance System

(OMS) has been proposed recently. OMS supports

(a) Propellant tank

(b) Thruster

3N Thruster

(a) Propellant tank

(b) Thruster

(a) Propellant tank

(b) Thruster

3N Thruster

Figure 6 Components of Propulsion Subsystem

S-XPDR1

S-DIPS-HYB

S-ANT1

S-ANT2

S-SW

S-XPDR2

S-XPDR1S-XPDR1

S-DIPS-HYBS-HYBS-HYB

S-ANT1

S-ANT2

S-SWS-SW

S-XPDR2S-XPDR2

Figure 7 Block diagram of Communication Subsystem


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space systems by inspecting, de-orbiting useless, and repairing satellites on orbit. The key technologies to realize OMS are rendezvous maneuver to an uncontrolled satellite, remote inspection, and capture of the target satellite. As the first step, the demonstrations of rendezvous maneuver and inspection are planned by CRL. In this demonstration, two standard buses are used. One bus (servicing or chaser satellite) equipped with an optical camera flies around the other bus (target satellite) for inspection. The purpose of this demonstration is to establish the autonomous rendezvous maneuver technology using on-board camera images. This technology leads to the unmanned on-orbit servicing as “Inspector”,

“Reorbiter”, and “Repairer”. 3.2 Demonstration of Corona Observation3

It is known that the high-energy radiation such as

Solar Energetic Particle is emitted, when solar flare occurs. This phenomenon causes malfunctions of computers and sensors on satellites and jamming to radio wave communication or broadcast. In order to predict and elucidate it, the “Space Weather Forecast” project has been initiated by CRL in Japan. In this project, Coronal Mass Ejection (CME) propagation between sun-earth space is to be observed from the 5th

Lagrangian point. (L5 mission) The present bus system could carry the newly developing corona observation

system composed of a camera, a detector, and data processors . In this way, it could provide the precedence space demonstration opportunity of the L5 mission. In

this demonstration, the performance of the corona observation system can be checked and the new image-processing algorithm will be established through observation data analysis. 3.3 Demonstration of Optical Inter-Satellite Communication4

In recent years, the demands for Mobile Satellite

Service have been more personal and multimedial. To meet these demands, the realization of the Global Multimedia Mobile Satellite System (GMMSS) has been required. GMMSS could provide global personal communications service by means of a group of LEO satellites with a user data rate around 2Mbps for handy

terminals. Due to their high data rate and large capacity with more compact communication tools, an optical inter-satellite communication system could be suited for GMMSS. Moreover it has an advantage of limited risk of interference with other communication systems. The “Ubiquitous Network” concept using an optical inter-satellite communication system has been proposed by NeLS. In this project, developments of optical inter-satellite communication components and a direct radiating active phased array antenna are thought as the most important breakthroughs. Using two satellite buses, the demonstration of optical inter-satellite communication technologies could be performed. In this demonstration, one satellite transmits data. The other receives them and downlink with the

direct radiating active phased array antenna. Moreover it is required that an optical transmitter points to an optical receiver with high accuracy.

The mission components for the above-described demonstrations could be light and small enough to fit on two small satellites launched at the same time. Figure 8 shows the illustrated image of the joint mission using the present standard bus system.

4. Conclusion

The 100-300kg class satellite standard bus system, which is being studied at MHI, and the first applications have been described. The purpose of this

study is to deliver the standard bus to users rapidly at lower cost. To achieve the purpose, the mechanical, thermal, and electrical interfaces between bus module

Figure 8 Applications of the bus system


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and mission payload module are kept as simple as possible, and off-the-shelf and flight-proven bus components are selected.

The standard bus system is being studied on the assumption of the joint mission planned by CRL and NeLS. And target launch year is 2006.

Besides the three missions described in this paper, there are potentially various needs for on-orbit demonstration of new instruments and technologies. Moreover, it can be proposed to construct practical mission by plural satellites or formation flying system for communication and observation missions. The present standard bus system can be an answer to these needs under the limited budgets.

5. Acknowledgements

The authors would like to express their gratitude to members at NeLS for their cooperation in the joint mission project using the MHI standard bus system.

6. References 1. Web page at

http://www.ee.surrey.ac.uk/SSC/SSHP/mini/. 2. S. Kimura : “Orbital Maintenance System;OMS”,

Journal of the Institute of Electronics, Information and Communication Engineers, Vol.82, No.8, pp820-823, 1999. (in Japanease)

3. M. Akioka : “Operational Space Weather Activity

in CRL”, Proc. of Space Weather Workshop, Dec. 2001.

4. R. Suzuki, K. Sakurai, S. Ishikawa, I. Nishiyama and Y. Yasuda : “A Study of Next -Generation LEO System for Global Multimedia Mobile Satellite Communications”, Proc. of 18th AIAA International Communications Satellite Systems Conference, AIAA-2000-1102, 2000.

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