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Availability & Reliability of Satellite Constellation, Satellite System and Subsystems

Over the past three years, SpaceX has deployed thousands of satellites into low-Earth orbit as part of its business to beam high-speed internet service from space. But the company’s latest deployment of 49 new satellites after a Feb. 2021 launch did not go as planned.


As a consequence of a geomagnetic storm triggered by a recent outburst of the sun, up to 40 of 49 newly launched Starlink satellites have been knocked out of commission. They are in the process of re-entering Earth’s atmosphere, where they will be incinerated.


The incident highlights the hazards faced by numerous companies planning to put tens of thousands of small satellites in orbit to provide internet service from space. And it’s possible that more solar outbursts will knock some of these newly deployed orbital transmitters out of the sky. The sun has an 11-year-long cycle in which it oscillates between hyperactive and quiescent states. Presently, it is ramping up to its peak, which has been forecast to arrive around 2025.


Challenges of Space

Those can be listed as vacuum, high temperature changes regarding nonconductive thermal feature of vacuum typically between −150 and 150°C, outgassing or material sublimation which can create contamination for payloads especially on lens of cameras, ionizing or cosmic radiation (beta, gamma, and X-rays), solar radiation, atomic oxygen oxidation or erosion due to atmospheric effect of low earth orbiting.


The first hurdle for  space systems  to overcome is the vibration imposed by the launch vehicle. Rocket launchers generate extreme noise and vibration. When a satellite separates from the rocket in space, large shocks occur in the satellite’s body structure. Satellite  must survive the extreme vibrations and acoustic levels of the launch. Pyrotechnic shock is the dynamic structural shock that occurs when an explosion occurs on a structure. Pyroshock is the response of the structure to high frequency, high magnitude stress waves that propagate throughout the structure as a result of an explosive charge, like the ones used in a satellite ejection or the separation of two stages of a multistage rocket. Pyroshock exposure can damage circuit boards, short electrical components, or cause all sorts of other issues.


Then, as it quietly circles the earth doing its job, it has to operate in very harsh conditions. It must function in an almost complete vacuum, while handling high levels of electro-radiation and fluctuation in temperatures that range from the hottest to the coldest. Outgassing is another major concern. The hard vacuum of space with its pressures below 10−4 Pa (10−6 Torr) causes some materials to outgas, which in turn affects any spacecraft component with a line-of-sight to the emitting material. Plastics, glues, and adhesives can and do outgas. V apor coming off of plastic devices can deposit material on optical devices, thereby degrading their performance.


High levels of contamination on surfaces can contribute to electrostatic discharge. Satellites are vulnerable to charging and discharging. Discharges as high as 20,000 V have been known to occur on satellites in geosynchronous orbits. If protective design measures are not taken, electrostatic discharge, a buildup of energy from the space environment, can damage the devices. A design solution used in geosynchronous Earth orbit (GEO) is to coat all the outside surfaces of the satellite with a conducting material.


The atmosphere in LEO is comprised of about 96% atomic oxygen. Atomic oxygen can react with organic materials on spacecraft exteriors and gradually damage them. Plastics are considerably sensitive to atomic oxygen and ionizing radiation. Coatings resistant to atomic oxygen are a common protection method for plastics.


Another obstacle is the very high temperature fluctuations encountered by a spacecraft. Because it is closer to the Sun, the temperature fluctuations on a satellite in GEO stationary orbit will be much greater than the temperature variations on a satellite in LEO. Thermal cycling occurs as the spacecraft moves through sunlight and shadow while in orbit that can cause cracking, crazing, delamination, and other mechanical problems, particularly in assemblies where there is mismatch in the coefficient of thermal expansion.


Radiation effects (total dose, latchup, single event upsets) are one of the main concerns for space microelectronics.


The reliability of a system is defined by the probability of correct operation of the system during a given lifetime. Reliability is defined as “the probability that a given component or system performs its functions as desired within a specific time t. Reliability theory is nothing but the mathematical attempt to predict the future of the satellite.


Reliability is done for following reason:

(1)To know what is the probability that the system will work after a given period.

(2)To provide redundant subsystems when the probability of failure is too great to accept.


By duplicating the less reliable and critical components, the overall reliability of the system could be improved. If any failure occurs in operational unit, then the standby unit takes over to develop a system with redundant components, its redundant elements are considered in parallel.


Figure 3 from Extracting best reliable scheme for Electrical Power Subsystem (EPS) of satellite | Semantic Scholar


Satellite Relaibility

The reliability of a complete satellite communications system depends on the reliability of its two principal constituents—the satellite and the ground stations.


As far as a satellite is concerned, it is necessary to consider its reliability, which is determined by breakdowns of on-board equipment, breaks during an eclipse if the only source of power for on-board equipment is solar power and the lifetime of the satellite.

For complex equipment such as that of a satellite, two types of breakdown occur:
—coincidental breakdown;
—breakdowns resulting from usage (examples are wear of mechanical devices such as bearings
and degradation of the cathodes of travelling wave tubes (TWT)) and exhaustion of energy
sources (such as the propellant required for station keeping and attitude control).


For space equipment, failures due to ‘infant mortality’ are eliminated before launching by means
of special procedures (burn-in). Hence, during the period of useful life, most of the electronic and
mechanical equipment has a constant failure rate Lamda. The instantaneous failure rate is thus often expressed in Fit (the number of failures in 10 exp(9) h).


For a satellite, the designed maximum satellite lifetime U can be defined as the time interval at the end of which the service is no longer provided, usually due to exhaustion of the propellants. After time U, the probability of survival is zero.


Hybrid reliability can be defined as the product of the reliability considering only wear-out
and the reliability which characterizes random failures. Equipment is generally designed in such a
way that the lifetime determined by wear-out, is long compared with the maximum satellite


Bath–tub curve

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  1. For most electronic equipment, the probability of failure is higher at the beginning of life and then
  2. The bath tub curve is a graph of probability of failure versus time.
  3. Three types of failures are identified from the bath tub curve:

1.Initial failures:

These appear shortly after t=0 and are characterised by a high failure rate. The initial high failure rate is attributed to manufacturing faults, defects in materials, etc.

2.Chance failures:.

These occur during a period in which probability of failure is constant and characterized by low failure rate. Low failure rate is attributed to random component failures. They are caused by severe and unpredictable environmental conditions.

3.Wear out failures:

It follows the region of chance of failures with an increase in the failure rate until the end of life of the component is observed. This can be attributed to component wear out.

  1. Therefore the early failures can be reduced by testing the components to determine reliability, causes of failure and expected lifetime. This testing is performed in vigorous condition experienced by the components.
  2. The random failures can be reduced by using reliability engineering techniques.
  3. The beginning of the wear out failure can be delayed by improving the type of materials used to make satellite components.

The mean time to failure (MTTF) is the mean time T of the occurrence of the first failure after
entering service. If the failure rate  is constant, T=l / failure rate (Lamda).

For equipment which is repaired after the occurrence of a failure, the Mean Time Between
Failures (MTBF) is defined in a similar manner.

Satellite Availability

The availability is the ratio of the actual period of correct operation of the system to the required
period of correct operation. Availability A is defined as A = (required time – down time)/(required time), where required time is the period of time for which the system is required to operate and downtime is the cumulative time the system is out of order within the required time.

To provide a given system availability A for a given required time L, it is necessary to determine
the number of satellites to be launched during the required time L. The number of satellites to be
launched will affect the cost of the service.


The availability of a complete satellite communications system depends not only on the reliability of the constituents of the system but also on the probability of successful launching, the replacement time and the number of operational and back-up satellites (in orbit and on the ground).


The availability of the ground stations depends not only on their reliability but also on their
maintainability. For the satellite, availability depends only on reliability since maintenance is not
envisaged with current techniques.


Satellite constellation

The demand for broadband satellite communication has driven companies to design large space-based communication constellations, termed mega-constellations.


Large space-based communication networks have been growing in numbers of satellites,
with plans to launch more than 10,000 satellites into Low Earth Orbit (LEO). LEO is defined as an orbit with an altitude of 160 km to 2,000 km. At these altitudes, the one-way communication latency is 0.5 ms to 7 ms. Although the footprint of these orbits is extremely small, companies have opted to use more satellites to take advantage of the lower communication latency and cheaper cost to launch. SpaceX has been approved by the FCC to launch 7,518 satellites into a Very Low Earth Orbit (VLEO), operating below 350 km altitude. Satellites in the VLEO regime face the additional obstacle of combating atmospheric drag, which is a significant factor for satellites below 500 km altitude.


While these constellations offer many advantages over ground-based communication systems, they pose a significant threat when they fail and generate space debris. LEO is becoming more crowded with orbital debris, which raises the concern of debris from mega constellations. The more satellites we launch into orbit, the more likely collisions are, generating yet more debris


Iridium Communications found that almost 30 percent of their satellite fleet failed in orbit, despite the fact that they used “highly reliable components” (Foust, 2019), and filed for bankruptcy in 1999. Amazon reported to the U.S. Federal Communications Commission (FCC) that if 15 percent of their 3,236-fleet Kuiper System broadband internet satellites fail, there is a 12 percent chance that one of those failed satellites would collide with a piece of space debris.


We can think of a satellite constellation as a system, where multiple components interact
with each other to produce a desired output. The system’s performance is evaluated based on the
desired system output. In an ideal system, where components do not fail, the system’s performance remains constant over time.


Researcher to design a useful model for broadband communication networks is comprised
of four key elements: (1) the satellite constellation model; (2) the communication network model;
(3) the satellite failure model; and (4) a definition of the network performance metric.The communication network model also defines the volume of data sent between groundstations. The critical part of the satellite broadband communication network model is the satellite failure model The performance metric is based on how much data the network is delivering to groundstations, with component failures


In a real system however, individual components fail. These failures occur from improper
use, component deterioration, or external forces. As components fail fewer interactions between
system components produces a lower output.


The two major LEO regular constellation types used in practice are: (1) Walker star constellations; and (2) polar constellations. Both Walker star constellations and polar constellations have circular orbital planes of the same altitude with equally spaced satellites within each plane by their mean anomaly, and equally spaced orbital planes within the plane of reference by their right ascension. The total number of satellites in the Walker star constellation network is 𝑀𝐿 × 𝑁𝐿
, where 𝑀𝐿 is the number of identical orbital planes with the same inclination, and 𝑁𝐿
is the number of satellites in each plane.


Communication links in satellite constellations can be categorized as groundstation-tosatellite communication links, satellite-to-groundstation communication links, and satellite-tosatellite communication links. Each of these communication links operate with a specific radio frequency (RF) to overcome the obstacles of each data link environment. The most common RF band for communication satellite constellations is the K-band, which is further categorized into Ku
and Ka bands.


Satellite communication networks can be categorized into two basic network topologies: (1) hub/remote networks; and (2) point-to-point networks. In a hub/remote network topology, a single hub ground station serves as the central ground station which transmits data to smaller remote ground stations. The hub/remote network topology is typically used for applications requiring asymmetric data exchange, such as remote internet access. In a hub/remote network
topology, the hub ground station is the most important node of the network. Unlike
the remote ground stations, a failure of the hub ground station would cause the entire network to


Most communication satellite constellations use a point-to-point network topology, where
communication exists between any two groundstations. No single groundstation
failure would cause the entire network to fail.


Satellite failures are independent. Most communication satellites are operationally independent, meaning that one satellite failure does not impact the reliability of other satellites. Ground stations do not fail. Ground stations can be maintained regularly, unlike satellites, which makes them significantly more reliable that satellites.


In a simple satellite constellation where one groundstation is sending data to another groundstation, one or more satellite failures in a communication path will cause the entire path to
fail. A failed path results in complete loss of the data being routed through the path. Once a data
packet has been sent from a groundstation, the data packet will attempt to take the communication path assigned regardless of failed satellites. When a satellite along the path fails, the receiving groundstation does not receive the sent data.


The satellite constellation network can be abstractly thought of as a network made up of
groundstation nodes and satellite nodes. In a simple scenario, two groundstation nodes can be
connected through at least two possible data paths. Broadband constellations in practice are often
large enough to offer multiple paths between groundstations.


One of the most common performance metrics used to evaluate communication constellations is the availability of the system, which is the fraction of total uptime communication hours over some time period divided by the total time period length. A point-to-point network’s life-cycle availability, 𝐴PTP network, can be calculated using the ground station availability, 𝐴GS life cycle, and the satellite availability, 𝐴satellite life cycle.
𝐴GS life cycle = GS Uptime / Total Time
𝐴satellite life cycle = Satellite Uptime / Total Time
𝐴PTP network =  Sqr(𝐴GS life cycle) * (𝐴satellite life cycle)

The GS Uptime and Satellite Uptime are the median time to failure of each component, which is less than the total life cycle time for each system. This measure of network performance is a scalar value measuring the network performance over its lifetime.


Subsystem Reliability

Calculation of the reliability of a system is performed from the reliability of the system elements. As far as the satellite is concerned, except in the special case where elements in  parallel can independently fulfil a particular mission, most subsystems are essentially in series from the point of view of reliability. This indicates that correct operation of each subsystem is indispensable for correct operation of the system.


Satellite Power system reliability

Power electronics is the engineering study of converting electrical power from one form to another. At a world-wide average rate of 12 billion kilowatts every hour of every day of every year, more than 80% of the power generated is being reprocessed or recycled through some form of power electronic systems. A lot of energy is wasted during this power conversion process due to low power conversion efficiency. It is estimated that the power wasted in desktop PCs sold in one year is equivalent to seventeen 500MW power plants! It is therefore very important to improve the efficiency of these power conversion systems. It is estimated that with the widespread use of efficient and cost-effective power electronics technology, the world could see a 35% reduction in energy consumption.


Power electronics is the technology associated with the efficient conversion, control and conditioning of electric power by static means from its available input form into the desired electrical output form. Power electronic converters can be found wherever there is a need to modify the electrical energy form (i.e. modify its voltage, current or frequency.) With “classical” electronics, electrical currents and voltage are used to carry information, whereas with power electronics, they carry power. Some examples of uses for power electronic systems are DC/DC converters used in many mobile devices, such as cell phones or PDAs, and AC/DC converters in computers and televisions. Large scale power electronics are used to control hundreds of megawatt of power flow across our nation.


Satellite present in an orbit should be operated continuously during its life span. So, the satellite requires internal power in order to operate various electronic systems and communications payloads that are present in it. The Electrical Power System (EPS) is a vital subsystem whose primary role is to supply other systems in the satellite with the necessary electrical power to operate effectively.


The EPS is one of the most critical systems on any satellite because nearly every other subsystem requires power. This makes the choice of power systems the most important task facing satellite designers. The main purpose of the Satellite EPS is to provide continuous, regulated and conditioned power to all the satellite subsystems within the specified operation period and during ground testing.


The Electrical Power Supply, or EPS of the Satellite is composed by three modules which are the PCC (Power Control Circuit), the PV (photovoltaic panel) and the BAT (Battery). The role of the EPS is to generate, store and distribute the electricity produced by the solar panels.


The electrical power generated at the solar arrays (SA) must be controlled to prevent the storage battery (SB) from overcharging and creating undesired spacecraft heating. The two main power subsystem configurations are a peak -power tracker (PPT) and a direct-energy transfer (DET). A PPT is a non-dissipative subsystem and its disadvantages appear at End Of Lifetime (EOL). An SA and/or SB has sensible values of degradation, while the DET is a dissipative subsystem since its shunt regulator operates in parallel with the solar array to dissipate power if the loads do not require it. The advantages of DET are as follows: fewer parts, lower mass and higher total efficiency at EOL.

Reliability Specification:

The essential elements of a reliability specification are:
• A quantitative statement of the reliability requirement specified by the customer as a MINIMUM acceptable value. The Satellite Power System (EPS) under study shall provide operation within 5 years of satellite active lifetime. The reliability of the EPS, within 5 years, shall be not less than 0.97.
• Environment of the power subsystem operation: The EPS shall operate as specified under the following environments; 670 km altitude, ~98 degree inclination and AM0, temperature ~ (-80: +80 oC).

• Mission lifetime in orbit and ground operation identification
• Constituted failure definition that should be expressed in terms which will be measurable during the demonstration test.


EPS Reliability Modeling and Prediction

The conventional Probability Modeling Method is used to determine the mathematical model EPS. The EPS reliability equation can be formed as follows:
Reliability EPS (REPS) = RSA  *RSB * RPMC



The key factors affecting EPS functioning efficiency (besides random effects which may lead to EPS subsystems failures) should be considered as:

  • Illumination of orbit, illumination and temperature of solar arrays.
  • Degradation of SA characteristics from space factor effects.
  • Degradation of battery characteristics in the course of its cycling

Assessment of Quality and Reliability

  • Electrical Measurements (e.g. verification of characteristic datasheet specifications)
  • thermal measurements (e.g. by infrared thermography)
  • static thermal resistance (mW…kW)
  • heating behavior
  • transient heat distribution (thermal impedance)
  • power cycle tests up to 2000 A
  • temperature cycles

Material related evaluation of electronic assemblies

  • analysis of thermo-mechanical behavior
  • description of damage mechanisms, failure analysis and damage level evaluation in solder joints and bond contacts by metallurgical methods

Evaluation of Reliability and Lifetime Prediction

  • product lifetime prediction based on observed damage mechanisms
  • transfer of mission profile under actual operating conditions (active and passive temperature changes) to test cycles

Thermal Engineering

  • simulation of the heat balance (components, modules and assemblies)
  • modelling of static thermal resistance
  • modelling of transient heat distribution (heating behavior)
  • modelling of mechanical loads due to material incompatibility (thermo-mechanical mismatch, warpage of layer composites, solder creeping)
  • heat sink calculation, e.g. dimensioning of air- and liquid coolers
  • evaluation of cooling concept efficiency

Development of Assembly Concepts

  • optimisation of thermal resistance
  • evaluation of relevant characteristics in production quality and reliability
  • prototype manufacturing
  • power balling 



Space qualification

  1. This process of reliability testing which ensures the space worthiness of each component and sub system of a spacecraft as well as that of the entire spacecraft as a complete system is called as space qualification.
  2. Space qualification is carried out by performing stress test on three prototype models of the satellite namely,
  3. The mechanical model- containing all the structure and mechanical parts included in the spacecraft.
  4. The thermal model- which contains all the electronic packages and the other component that must be maintained at correct temperature.
  5. The electrical model -which contains all the electronic parts of the spacecraft and is tested for correct electrical performance.


The design of radiation-hardened integrated circuits ( RHlCs ) involves four primary efforts. First is the selection of a technology and process which are relatively insensitive to the projected application environment of the IC.


Second, parts representative of the selected technology must be characterized in a simulated environment that models the RHIC’s application environment in order to quantify the effects of the environment on material and device characteristics. In the third phase, the circuit design techniques which make device responses most insensitive to the radiation are selected based on the technology analyses, and implemented in an IC design. The fourth phase actually occurs throughout the design process. Computer simulations of the chip response in pertinent environments should be performed as a part of each cycle of the design, manufacture, and testing processes, write Sherra E. Kerns, Senior Member, IEEE, And B. D. Shafer, Department of Electrical Engineering, Vanderbilt University, Nashville, TN 37235, USA.




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