Satellite Subsystems
Content
Sub–Systems of a Satellite
1.
2.
3.
4.
Payload
Power (EPS)
Structures, Thermal and Mechanisms (STM)
Attitude Determination and Control System (ADCS) (or) Guidance, Navigation and Control
(GNC)
5. Telemetry, Tracking and Command (TTC)
6. Command and Data Handling (CDH) (or) On–Board Data Handling (OBDH)
7. Systems Engineering
1. Payload
The mission objective includes the reason for and purpose of the space mission, the goal and the
subject of the mission, as well as its benefits. The mission objective can be, for example,
operational weather monitoring for weather service organizations or investigation of phenomena
on the Earth’s surface or in deep space for scientific establishments or other institutions. In these
cases the mission subject would be the Earth’s surface or deep space.
The payload is directly correlated with the mission objective, because it is only possible to reach
the mission objective by using the payload successfully. In most cases it is an instrument or a
sensor which generates essential data for the satellite mission. The very strict constraints regarding
mass, volume and power drastically limit the choice of available instruments. There is a large
scope and demand for the miniaturization of instruments. If the mission is a pure technology
demonstration mission, then the component to be tested is the payload, which could be a
component of a subsystem, a subsystem or even the satellite itself. In the last case, there is no
simple distinction between satellite bus and payload.
2. Power
The Electrical Power Subsystem (EPS) provides, stores, distributes and controls spacecraft
electrical power. The most important sizing requirements are the demands for average and peak
electrical power and the orbital profile (inclination and attitude). The electrical power loads for
mission operations at beginning–of–life (BOL) and end–of–life (EOL) must be identified by the
EPS.
Typical top-level power subsystem functions:
Supply a continuous source of electrical power to spacecraft loads during the mission life.
Control and distribute electrical power to the spacecraft.
Support power requirements for average and peak electrical load.
Provide converters for ac and regulated dc power buses, if required.
Provide command and telemetry capability for EPS health and status, as well as control by
ground station or an autonomous system.
Protect the spacecraft payload against failures within the EPS.
Suppress transient bus voltages and protect against bus faults.
Provide ability to fire ordnance, if required.
3. Structures, Thermal and Mechanisms
Structures and Mechanisms Subsystem (SMS)
The SMS mechanically supports all other spacecraft subsystems, attaches the spacecraft to the
launch vehicle, and provides for ordnance–activated separation. The design must satisfy all
strength and stiffness requirements and of its interface to the booster. Primary structure carries the
spacecraft’s major loads; secondary structure supports wire bundles, propellant lines, nonstructural
doors and brackets for components typically under 5 Kgs.
Thermal Control Subsystem (TCS)
The role of TCS is to maintain all spacecraft and payload components and subsystem within their
required temperature limits for each mission phase. Temperature limits include a cold temperature
which the component must not go and a hot temperature that it must not exceed. Two limits are
frequently defined: operational limits that the component must remain within while operating and
survival limits that the component must retain within at all times, even when not powered.
Exceeding survival temperature limits can result in permanent equipment damage as opposed to
out–of–tolerance performance when operational limits are exceeded. Thermal control is also used
to ensure that the temperature gradient requirements are met. A larger gradient could cause
structural deformation such that pointing is adversely impacted, possibly permanently.
4. Attitude Determination and Control System (ADCS) (or) Guidance, Navigation and
Control (GNC)
A satellite's ADCS system is used to stabilize and orient the vehicle as required by the Concept of
Operations (CONOPs) in the presence of external disturbance torques acting on the satellite. This
requires that the satellite determine its attitude, using sensors, and control it, using actuators. The
ADCS system uses external references to determine the satellite's angular orientation with respect
to a fixed inertial reference frame, usually an Earth centered, equatorial frame for Earth orbiting
satellites. External attitude references include the Sun, Earth's horizon, the local magnetic field,
and the stars. The satellite may also use inertial sensors like angular rate gyroscopes to measure
angular rate and estimate the satellite's angular orientation between fixed inertial reference
measurements, or while fixed inertial reference measurements are unavailable. The ADCS often
is tightly couples to other subsystems on board, especially the propulsion and navigation functions.
5. Telemetry, Tracking and Command (TTC)
The Telemetry, Tracking and Command (or) Communications subsystem provides interface
between the spacecraft and ground systems. Payload mission data and spacecraft housekeeping
data pass from the spacecraft through this subsystem to operators and users at the operations centre.
Operator commands also pass to the spacecraft through this subsystem to control the spacecraft
and to operate the payload. Hardware for TTC is designed in such a way that the data transmission
is reliable for all spacecraft operating modes.
Functions of the TTC subsystem:
Carrier tracking (lock onto the ground station signal)
Command reception and detection (receive the uplink signal and process it)
Telemetry modulation and transmission (accept data from spacecraft systems, process them,
and transmit them)
Ranging (receive, process, and transmit ranging signals to determine the satellite’s position)
Subsystem operations (process subsystem data, maintain its own health and status, point the
antennas, detect and recover faults)
6. Command and Data Handling (CDH) (or) On–Board Data Handling (OBDH):
The CDH performs two major functions. It receives, validates, decodes, and distributes commands
to other spacecraft systems and gathers, processes, and formats spacecraft housekeeping and
mission data for downlink or use by an onboard computer. This equipment often includes
additional functions, such as spacecraft timekeeping, computer health monitoring (watchdog), and
security interfaces.
Classical OBDH architectures are based upon a central processor, typically connected via a video
or digital path to the RF communications subsystem. This central processor will communicate with
the platform subsystems such as the AOCS and the payloads using a serial bus with high data
integrity.
7. Systems Engineering
System engineering is often characterized as both an art and a science. This apt characterization is
appropriate because good system engineering requires the creativity and knowledge of systems
engineers, but it also requires systems management or the application of a systematic disciplined
approach.
“A logical process of activities that transforms a set of requirements arising from a specific mission
objective into a full description of a system which fulfils the objective in an optimum way. It
ensures that all aspects of a project have been considered and integrated into a consistent whole.”
System engineering encompasses the following functions:
Requirement engineering including requirements analysis and validation, requirement
allocation, and requirement maintenance.
Analysis in order to resolve requirements conflicts, decompose and allocate requirements
during functional analysis, and assess system effectiveness (including analyzing risk factors).
It is also used to complement testing evaluation and to provide trade studies for assessing
effectiveness, risk, cost and planning.
Design and configuration in order to define a physical architecture, and its complete system
functional, physical and software characteristics.
Verification, to demonstrate that the deliverables conform to the specified requirements,
including qualification and acceptance.
System engineering integration and control to ensure the integration of the various engineering
disciplines and participants throughout all the project phases.
References:
1. Space Mission Analysis and Design, Larson and Wertz; Kluwer; 2005
2. Handbook of Space Technology; Ley, Wittmann and Hallmann; Wiley; 2009
3. Spacecraft Systems Engineering; Fortescue, Swinerd and Stark; Wiley; 2011
1.
2.
3.
4.
Payload
Power (EPS)
Structures, Thermal and Mechanisms (STM)
Attitude Determination and Control System (ADCS) (or) Guidance, Navigation and Control
(GNC)
5. Telemetry, Tracking and Command (TTC)
6. Command and Data Handling (CDH) (or) On–Board Data Handling (OBDH)
7. Systems Engineering
1. Payload
The mission objective includes the reason for and purpose of the space mission, the goal and the
subject of the mission, as well as its benefits. The mission objective can be, for example,
operational weather monitoring for weather service organizations or investigation of phenomena
on the Earth’s surface or in deep space for scientific establishments or other institutions. In these
cases the mission subject would be the Earth’s surface or deep space.
The payload is directly correlated with the mission objective, because it is only possible to reach
the mission objective by using the payload successfully. In most cases it is an instrument or a
sensor which generates essential data for the satellite mission. The very strict constraints regarding
mass, volume and power drastically limit the choice of available instruments. There is a large
scope and demand for the miniaturization of instruments. If the mission is a pure technology
demonstration mission, then the component to be tested is the payload, which could be a
component of a subsystem, a subsystem or even the satellite itself. In the last case, there is no
simple distinction between satellite bus and payload.
2. Power
The Electrical Power Subsystem (EPS) provides, stores, distributes and controls spacecraft
electrical power. The most important sizing requirements are the demands for average and peak
electrical power and the orbital profile (inclination and attitude). The electrical power loads for
mission operations at beginning–of–life (BOL) and end–of–life (EOL) must be identified by the
EPS.
Typical top-level power subsystem functions:
Supply a continuous source of electrical power to spacecraft loads during the mission life.
Control and distribute electrical power to the spacecraft.
Support power requirements for average and peak electrical load.
Provide converters for ac and regulated dc power buses, if required.
Provide command and telemetry capability for EPS health and status, as well as control by
ground station or an autonomous system.
Protect the spacecraft payload against failures within the EPS.
Suppress transient bus voltages and protect against bus faults.
Provide ability to fire ordnance, if required.
3. Structures, Thermal and Mechanisms
Structures and Mechanisms Subsystem (SMS)
The SMS mechanically supports all other spacecraft subsystems, attaches the spacecraft to the
launch vehicle, and provides for ordnance–activated separation. The design must satisfy all
strength and stiffness requirements and of its interface to the booster. Primary structure carries the
spacecraft’s major loads; secondary structure supports wire bundles, propellant lines, nonstructural
doors and brackets for components typically under 5 Kgs.
Thermal Control Subsystem (TCS)
The role of TCS is to maintain all spacecraft and payload components and subsystem within their
required temperature limits for each mission phase. Temperature limits include a cold temperature
which the component must not go and a hot temperature that it must not exceed. Two limits are
frequently defined: operational limits that the component must remain within while operating and
survival limits that the component must retain within at all times, even when not powered.
Exceeding survival temperature limits can result in permanent equipment damage as opposed to
out–of–tolerance performance when operational limits are exceeded. Thermal control is also used
to ensure that the temperature gradient requirements are met. A larger gradient could cause
structural deformation such that pointing is adversely impacted, possibly permanently.
4. Attitude Determination and Control System (ADCS) (or) Guidance, Navigation and
Control (GNC)
A satellite's ADCS system is used to stabilize and orient the vehicle as required by the Concept of
Operations (CONOPs) in the presence of external disturbance torques acting on the satellite. This
requires that the satellite determine its attitude, using sensors, and control it, using actuators. The
ADCS system uses external references to determine the satellite's angular orientation with respect
to a fixed inertial reference frame, usually an Earth centered, equatorial frame for Earth orbiting
satellites. External attitude references include the Sun, Earth's horizon, the local magnetic field,
and the stars. The satellite may also use inertial sensors like angular rate gyroscopes to measure
angular rate and estimate the satellite's angular orientation between fixed inertial reference
measurements, or while fixed inertial reference measurements are unavailable. The ADCS often
is tightly couples to other subsystems on board, especially the propulsion and navigation functions.
5. Telemetry, Tracking and Command (TTC)
The Telemetry, Tracking and Command (or) Communications subsystem provides interface
between the spacecraft and ground systems. Payload mission data and spacecraft housekeeping
data pass from the spacecraft through this subsystem to operators and users at the operations centre.
Operator commands also pass to the spacecraft through this subsystem to control the spacecraft
and to operate the payload. Hardware for TTC is designed in such a way that the data transmission
is reliable for all spacecraft operating modes.
Functions of the TTC subsystem:
Carrier tracking (lock onto the ground station signal)
Command reception and detection (receive the uplink signal and process it)
Telemetry modulation and transmission (accept data from spacecraft systems, process them,
and transmit them)
Ranging (receive, process, and transmit ranging signals to determine the satellite’s position)
Subsystem operations (process subsystem data, maintain its own health and status, point the
antennas, detect and recover faults)
6. Command and Data Handling (CDH) (or) On–Board Data Handling (OBDH):
The CDH performs two major functions. It receives, validates, decodes, and distributes commands
to other spacecraft systems and gathers, processes, and formats spacecraft housekeeping and
mission data for downlink or use by an onboard computer. This equipment often includes
additional functions, such as spacecraft timekeeping, computer health monitoring (watchdog), and
security interfaces.
Classical OBDH architectures are based upon a central processor, typically connected via a video
or digital path to the RF communications subsystem. This central processor will communicate with
the platform subsystems such as the AOCS and the payloads using a serial bus with high data
integrity.
7. Systems Engineering
System engineering is often characterized as both an art and a science. This apt characterization is
appropriate because good system engineering requires the creativity and knowledge of systems
engineers, but it also requires systems management or the application of a systematic disciplined
approach.
“A logical process of activities that transforms a set of requirements arising from a specific mission
objective into a full description of a system which fulfils the objective in an optimum way. It
ensures that all aspects of a project have been considered and integrated into a consistent whole.”
System engineering encompasses the following functions:
Requirement engineering including requirements analysis and validation, requirement
allocation, and requirement maintenance.
Analysis in order to resolve requirements conflicts, decompose and allocate requirements
during functional analysis, and assess system effectiveness (including analyzing risk factors).
It is also used to complement testing evaluation and to provide trade studies for assessing
effectiveness, risk, cost and planning.
Design and configuration in order to define a physical architecture, and its complete system
functional, physical and software characteristics.
Verification, to demonstrate that the deliverables conform to the specified requirements,
including qualification and acceptance.
System engineering integration and control to ensure the integration of the various engineering
disciplines and participants throughout all the project phases.
References:
1. Space Mission Analysis and Design, Larson and Wertz; Kluwer; 2005
2. Handbook of Space Technology; Ley, Wittmann and Hallmann; Wiley; 2009
3. Spacecraft Systems Engineering; Fortescue, Swinerd and Stark; Wiley; 2011
