The satellites subsystems

artist rendering of  the satellite in space
Artist’s rendering of Jason-3. Image credit: NASA/JPL-Caltech

Most of the challenges facing satellites are the same for every mission and thus we can distinguish few subsystems which we can find in almost every design of artificial satellite

Payload

The most important part of satellite. Payload realize the satellite mission: makes scientific experiments, take pictures or videos, transfer data etc… . Generally satellite consist two parts: the bus and the payload. The bus transports the payload to the desire position and ensures conditions (e.g. attitude, temperature, electric power) which are required to execute payload tasks. The bus without payload does not make sense, it exists only because payload needs it.

Mechanical structure

The framework for mounting other subsystems and an interface between the satellite and launch vehicle. Its functions are:

  • Links the satellite with launch vehicle (or canister)
  • Supports all equipment caring by the satellite
  • Being shield against radiation, dust, micrometeorites

The mechanical structure is facing few different challenges: it must be light as possible to reduce price of lift of, it must be strength enough to withstand vibrations and acceleration during the launch phase, it must survive very width range of temperature- from few hundreds of Celsius degrees (during facing sun) to several tens under zero (in the shadowed side) . Moreover it must be resistant to mechanical impacts of micrometeorites, dust or space junks. Mechanical structure has to ensure reliability of operations in space e.g separation from launcher, deployment of solar panels, operations of rotating parts etc..

NanoAvionics Small Satellite Structures
NanoAvionics Small Satellite Structures

Propulsion subsystem

Some missions requires to change velocities of the satellite after its detaching from the launcher. Propulsion subsystem gives possibility to execute maneuvers during life of the spacecraft e.g. transport the satellite from its transfer orbit to target orbit, station keeping operations, decommissioning.

Propulsion systems can be categorized by propellant and the mechanism used to produce thrust. Generally we can distinguish solid fuel, liquid fuel, electric and ion propulsion systems. In theory there are also other methods to produce the thrust like solar sail, speed up vessel with laser or even with nuclear explosions.

During missions 95% of propellant are used for east-west(longitudinal) manoeuvres and only 5% for north-south(latitudinal) station-keeping operations. Small amount of fuel is required to move the satellite from its orbit after the end of mission.

Thermal control subsystem

Thermal control must ensure that equipment of the satellite works in correct range of temperatures. The spacecraft can exchange thermal energy with space only by radiation. When the satellite is exposed to sun, then its temperature will grow, either fall when is in shadow. We can distinguish two kind of thermal systems: active and passive.

Passive thermal systems have no any moving elements and do not use electrical power. The systems are in form of multilayer insulated surfaces which absorb or reflect thermal energy from external or internal sources. The passive thermal control elements are blankets, coatings, reflectors, insulators, heat sinks, etc… This kind of thermal control requires very good layout plan and carefully choosing of materials for structure..

Aquarius instrument thermal blanketing
Aquarius satellite instrument thermal blanketing. NASA

Active thermal system consist heaters, refrigerators and remote thermal pipes. Heater and refrigerator may be controlled by on-board systems or by commands from ground. Thermal pipes are pipes fulled with a liquid which has low boiling point. Inside the pipe liquid evaporate(takes energy) in its warm end, then move to cold end where it is being condensed (release energy), and then backs to the warm end.

Satellite heat pipe
Satellite heat pipe. B. Anderson/Wikipedia

Power supply subsystem

Power supply subsystem generates and distributes electrical power to other subsystem of the satellite. There are three possible sources of electrical power on spacecraft: solar energy, chemical energy and nuclear energy.

Solar energy

The sun is a constant source of energy. The energy is cough by solar panels of satellite. Solar panels are set of connected solar cells – semiconductors which makes voltage with reaction for hits by solar photons. To get more power, cells are connected into string, and string are connected on their ends and all together are the panel. If satellite equipment require more power, then more solar cells are required, and total area of the panels are bigger.

Solar panels are heavy, what means that if spacecrafts requires more power, then it lift of is more expensive. Because of this fact 3 axis stabilized satellites are preferred more than spin stabilized. Spin stabilized satellites cannot utilize at the same moment all its cells, and thus require more panels than 3-axis stabilized satellite which can change attitude of panels to effectively utilize whole panels area.

One disadvantage of solar panels is their susceptibility to mechanical damage and damage caused by cosmic rays.

Chemical energy

Batteries are used on satellites to ensure required level of electrical power in case when the main energy source cannot work effectively. Typical usage of rechargeable battery is system of solar panels connected with battery: battery is charged when solar radiation works on panels and discharged during eclipse phase of orbit or is employed during short term power peeks. Battery also ensure power supply during launch phase, when solar panels are inactive.

Different orbits have different characteristic of batteries utilization. Important things are frequency of charging/discharging cycles and depth of discharge(DoD). For low Earth orbit(LEO) orbital period is roughly 100min with eclipse phase about 30-40min, but for geostationary orbit(GEO) period is 24h and eclipse phase is max. 72min. Typically DoD for LEO is 40% and for GEO is 80%. Frequency of charge/discharge cycles and DoD are decisive for lifetime of battery. The important factor is that capacity of battery depends of temperature, and thus is must be taken into consideration with choosing kind of battery and designing thermal control.

Main types of batteries

  • Nickel-cadmium (NiCd)
    Commonly used, but has memory effect – if is only partially discharged before charging, then it may forgot that it can be further discharged. Also energy per mass ratio is lower than other batteries. Advantage is its robustness versus number of discharge/charge cycles, thus they are used mostly on LEO.
  • Nickel Metal Hydride (NiMH)
    Have better energy density so its energy to mass ratio is higher than NiCd, and are not affected much with memory problem. However the batteries have problems with working on very hi or very low temperatures and are not suitable for spacecrafts.
  • Nickel-hydrogen (NiH2)
    Combination of batteries and fuel cell. The batteries have high cyclic stability (over 5000 cycles). Are resistant to deep discharge and overcharge. Very popular on aerospace applications and for GEO and LEO missions.
  • Lithium ion
    The most universe kind of battery with properties similar to NiH2, but with 30% less wight. No memory effect. Disadvantage is that special treatment is required because of lithium ignites very easily.
Nickey-hydrogen batteries
One 100Ah high-pressure type nickel-hydrogen battery unit for Engineering Test Satellite VIII(Credit: JAXA)

There are over 250 known satellite explosions, with about 10 of these due to batteries. The problem is with inactive spacecrafts which in case of explosion will pollute space with large number of debris.

Nuclear energy

For deep space missions, when The Sun is too far from spacecraft to effectively pass energy through solar panels, nuclear power is used. In radioisotope thermoelectric generator(RTG) natural decay of radioactive elements warm one end of thermocouple whereas second end emits energy to outer space, what gives even 400oC difference between two ends. Thermoelectric effect produce electric energy. RTG is very effective, stable and long term source of energy. However is very expensive and in case of problems during launch or operations it may cause nuclear pollution on The Earth. Soviet satellite KOSMOS 954 polluted 124.000 of square kilometer of Canada in 1978.

Attitude determination and control subsystem (ADCS)

Attitude of satellite is its orientation as determined by the relationship between its axes (yaw, pitch and roll) and some reference plane.

yaw, pitch and roll Gemini
Explanation of yaw, pitch and roll on Gemini spacecrafts as an example.(Wikipedia)

There are two kind of satellite stabilization: spin and three-axis stabilization

  • Spin stabilization is accomplished by setting the spacecraft spinning, using the gyroscopic action of the rotating spacecraft mass as the stabilizing mechanism. Propulsion system thrusters are fired only occasionally to make desired changes in spin rate, or in the spin-stabilized attitude.
  • Three-axis stabilization the spacecraft is held fixed in the desired orientation without any rotation.
    • use small thrusters to continually nudge the spacecraft back and forth within a deadband of allowed attitude error
    • Electrically powered reaction wheels, which are mounted on three orthogonal axes aboard the spacecraft. They provide a means to trade angular momentum back and forth between spacecraft and wheels. To rotate the vehicle on a given axis, the reaction wheel on that axis is accelerated in the opposite direction. To rotate the vehicle back, the wheel is slowed.
Wester VI spin stabilized satellite
Westar VI spin stabilized satellite (NASA)

Determination of attitude

Position of satellite is determined with respect to reference directions – commonly used are Earth, Sun or star. Sensors are used to find direction to reference objects e.g infrared detectors of emission from Earth, optical systems with map of stars to detect the star.

ADSC computes the error between desire attitude and current attitude of spacecraft and steers with torque proportionally to the sensed error. When the satellite is in orbit the subsystem maintains the antenna of the satellite pointed accurately in the desired direction. During launch phase ADCS ensures attitude which allow to correct orbital maneuvers. On LEO orbit ADCS may be used to change position of spacecraft to make higher or lower aerodynamically drag to precise position of the satellites from one constellation on orbit(differential drag).

Tracking, telemetry and control subsystem (TT&C)

The subsystem is used to monitor and control the satellite from tests on ground to the end of its operational life in orbit.

Telemetry gather information about health of various subsystems, encodes it and transmits toward the Earth control center. The data are named housekeeping – engineering parameters which are needs to be monitored to keep a check and operating status of equipment: temperatures, pressure, voltages, operational statuses etc… Sometimes “the housekeeping” may concern only the bus part of satellite, and the payloads engineering data may be threaten as a separated set of informations.

The tracking part of the system determines the spacecrafts position. It may send data for ADSC sensors.

The command element receives commands from The Earth control center and executes it to change platform configuration, position, velocity etc…

Because TT&C must be used in every stage of the satellite life it must contain communication interfaces which allow to make efficient transfer to/from control center, even when attitude of antenna cannot be set precisely toward ground station. Because of that TT&C often use omni directional antenna working on UHF band.

References

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