Archive for the ‘Satellites’ Category

Lunar Orbit of the Earth

This demonstrates the orbit of the Moon around the Earth and is a good companion to the earlier post Tilt in Earths axis. NASA defines the period of the Moon’s orbit to be one complete orbit in 27.3215 days that translates to 27 days, 7 hours, and 43 minutes. The cycles or phases of the Moon are due to it’s orbit around the Earth and the angle that the light from the Sun illuminates the Moon. As demonstrated in this animation as the Moon progresses through it’s orbit it will cycle from full illumination (the Full Moon) to total darkness (called new Moon) and back to full illumination as viewed on the earth. The tilt in the Moons axis is 1.54 degrees and the orbit has a 5.14 degree inclination to the Sun line, when added together the resulting tilt in axis is a total of 6.68 degrees.

This animation shows the spin of the Earth and movement of the Moon relative to the Moons orbit period. You can also see how the the phase of the Moon is based on it’s location in the orbit.

Solar Array Deployment

At launch the satellite solar panels are folded against the body of the satellite (stowed configuration) to minimize space and allow the satellite to fit into the faring of the launch vehicle. Once the satellite has been placed into Geosynchronous orbit the ACS system is activated to point the proper axis of the satellite at the earth. This process is called ACS initialization and Earth capture activation. During this period the satellite is being powered by the batteries. Once the Earth sensors are activated and the ACS system is locked onto the Earth, the solar panel release mechanization can be actuated to allow the arrays to unfold and lock into a deployed position. With the arrays deployed the solar array drive system is activated and commanded to track the Sun. The following animation shows the release of the North panel, then the South panel, followed by the positioning of the panels to point directly at the sun.

This would be the worst case condition where the arrays are pointed 180 degrees away from the Sun, the arrays are typically deployed at a time of day that allows them to be pointed at the Sun so they only require minor pointing changes to peak their power output. With a 180 degree pointing offset as shown the arrays would be commanded to move in opposite directions (one clockwise and the other counter clockwise) at the same time. From the ACS control standpoint the torques on the body of the satellite would relatively cancel and minimize the error correction required by the ACS system. After the the arrays are peaked on the Sun the Solar array drive system is commanded to normal tracking mode to maintain this pointing through a full 360 degrees rotation over the course of the day. As soon as there is adequate power being generated from the arrays, the power load will transition from battery to array power and battery charging can be started.

Solar Array Drive

Solar Array Drive (SAD) refers to the system that positions the solar panels on the satellite to efficiently convert s the sun light into electrical energy. This system is comprised of the Solar Array Drive Assembly (SADA) and the Solar Array Drive Electronics (SADE).

The SADA includes the stepper motor, drive gear assembly, Barring And Power Transfer Assembly (BAPTA) and the mounting hardware required to mount the solar panel to the body of the satellite. The solar panel is mounted to the SADA and allowed to move by means of the barrings in the BAPTA. It also includes a slip ring assembly used to transfer the power generated by the solar panel to the EPS bus. The drive gear assembly is installed to smooth the step movement of the motor and provide a steady constant travel through 360 degrees of rotation. Stepper motors are known for their ability to provide high torque and controlled movement, this makes them ideal for this application. Rather than mounting two motors in the assembly, to increase the reliability these stepper motors are designed with redundant coils, this effectively reduces weight and provides improved redundancy.

Each SADA is controlled by independent Solar Array Drive Electronics (SADE) unit that provides the interface between it and the EPS subsystem or flight computer. It is comprised of all of the circuitry required to generate and shape the pulses needed to drive the stepper motors. Telemetry monitoring sensors collect data such as current, temperature, position, speed, rotational direction of travel, step counts, and status indicating movement or stopped for the SADA.

In operation the rotation of the solar panel is started when the SADE is enabled and step pulses are received, movement stops between steps or when it has been disabled. Control of the direction of travel, can be commanded to rotate in either forward or reverse direction to allow for pointing position optimization. The SADE is normally configured to output a consistent train of pulses at the required intervals to maintain movement through the 360 degree rotation called Normal Mode. With the panel rotating less than 15 degrees each hour it is easily commanded at less than 5 times a second allowing for rest intervals between movement. Speed and torque can be adjusted by increasing or decreasing the rate of pulses per second (pps) sent to the motor. If the motor is normally driven at 5 pps then driving it at 25 pps will increase the speed by a factor of 5. This system is typically design for the solar panel to complete a full 360 degree rotation in less than 15 minutes at the highest pps command rate. When commanding the movement of the SADA, commands are processed by the SADE, stored and then pulses are sent at a 5 pps or 25 pps rate until the complete commanded step count is reached. At completion the SADE must be commanded back to the Normal Mode or transitions back automatically.

Satellite Solar Array positioning

As a satellite travels along the orbital path the Solar Arrays must remain pointed at the Sun to produce power. This is accomplished in a number of methods based on the satellite design. Satellites with fixed arrays must point their surfaces at the Sun or spin around the axis that allows them to rotate into the Sun light. With movable solar arrays they either actively track the Sun or are driven at a constant rate to maintain their pointing.

Spin stabilized or 2 axis satellites typically have fixed solar panels and positioning is achieved by the orientation of the spin axis of the satellite. Attitude or Spin Precession Maneuvers are periodically preformed to maintain the spin axis alignment.

Solar arrays on a 3 axis stabilized satellite track the Sun by the implementation of positioning mechanisms that either step the arrays or actively position them. To simplify the tracking system a geared drive system is driven by stepper motors that are selected to provide the required torque and move the array smoothly. The number of steps are calculated using the angular step size of the stepper motor, the gear ratio, and the angular change needed. This is used to generate the required number of pulses per second. Pulses are generated in the solar array drive electronics then applied to the motor to allow them to complete one complete 360 degree revolution a day. To increase accuracy the sidereal day measurement of approximately 23:56:04 hours is used. These systems can provide accurate positioning with minimal operational requirements. Over time periodic adjustments are made to minimize any tracking errors and optimize positioning and output power.

Active positioning systems also employ the use of stepper motors and the output of the solar array is sampled and converted to pulses that are applied to the stepper motor to maintain the peak output power. One important characteristic of this system is that it can not be used during eclipse periods, during these periods the system normally reverts to a stepped mode and the active control is resumed when the satellite returns to full sunlight. When these systems are fully optimized it significantly reduces operational intervention to maintain peak power. Failures in active tracking systems can lead to a rapid loss of power and must have antiquate procedures or automated sequences to prevent the arrays from being driven off the Sun.

Lithium-ion batteries

Lithium-ion batteries (Li-ion) are increasingly being used on satellites as a replaced for Ni-Cad and Ni-Hyd batteries. They continue to build a heritage based on high reliability. When properly operated and maintained they provide reliable reserve power. Their design life meets the 15 years life expectancy required in on orbit operations of GEO satellites.

The most significant differences in these batteries are related to their operation. The deterioration of the state of charge capacity is attributed to the chemical breakdown of the Lithium component in the battery over time. To slow this process, during periods when the batteries are not being discharged and charged (storage) between eclipse seasons, the batteries are maintained at a lower temperature and at a reduced state of charge. Typically at 50 % state of charge and the temperature is reduced by 10 degrees C, this varies based on the manufactures recommendations. Automated command sequences are stored in the flight computer and triggered by battery telemetry monitors. The flight computer maintains the battery state of charge at the recommended level and during emergencies will turn off units in a predefined sequence to reduce the power load and extend the time on batteries.

Eclipse operations have also been automated requiring additional preparations. Initially the batteries are warmed up by changing the heater set-points to bringing them up to the normal operation temperature. The battery discharge for each eclipse is calculated and then the batteries are charged to a higher state of charge to account for the expected discharge. At the start of eclipse season the periods are only a few minutes, at the center or longest eclipse is 70 minutes then gradually decrease back to a few minutes in duration. Plotting the battery charge and the battery discharge over the eclipse season shows curves that resemble a football and is loosely referred to as the football curve. These charging values are entered into a table in the flight computer and the charging profile is enabled allowing charging to be completed autonomously. Implementation of a day counter allows the flight computer to progress through the charging schedule.

Over charging of Lithium-ion batteries will lead them to catastrophic failure. The battery is comprised of modules vs cells on other types of batteries. Independent charge/discharge circuits are included in each module and to protect the battery, bypass relays are installed to isolate week or failed modules from the circuit. If a module is bypassed the module is completely discharged forcing it to permanently fail. Once a module has been bypassed it can no longer be used. Each module operates at 4 to 4.5 volts one design for a 36 volt battery has 9 modules. With all modules functioning, each module is charged to 4.0 volts and with one module in bypass the remaining 8 modules are charged to 4.5 volts to compensate.

Prolonged short discharge and recharge cycles do not have a significant affect on the capacity of these batteries (noted at this time).

Manual battery recondition has been replaced by automated sequences with ground commanded table values.

Due to technological advancements in battery chemistry and design, Nickel hydrogen batteries batteries are starting to be replaced by more efficient Lithium-ion batteries as they prove their reliability.

Additional information about Lithium-ion batteries is available on Wikipedia

Nickel Hydrogen battery

Nickel hydrogen battery (NiH2 or Ni-H2) are used extensively on satellites. These batteries have replaced the use of NiCad batteries in almost all cases and continue to build a heritage based on high reliability. When properly operated and maintained they have provided reliable reserve power. Their design life exceeds the 15 years life expectancy required in on orbit operations of GEO satellites.

One of the most significant differences in these batteries is the power density to weight ratio allowing them to store more energy while reducing weight, These batteries are pressure vessels and the state of charge can be derived from the pressure and temperature of each cell. Another unique feature is that modules are available that contain 2 cells in one vessel. The design, testing and assembly of cells into a completed battery utilize similar process standards with minor changes. In addition to voltage, current and temperature, pressure monitoring sensors are included on the cells.

When monitoring cell voltages during eclipse or discharging, the voltage drop profile follows that similar to NiCad battery cells. The pressures will drop on a liner slope. Nearing the depletion of capacity the voltage will drop by as much as 0.1 to 0.2 volts over a few minutes resembling the initial rate at the start of discharge and below 1.0 volts dramatically drop off. These battery cells can be discharged safely below the 1.0 volt limitation placed on NiCad battery cells.

During charging, cell pressure and temperatures have to be closely monitored to ensure maximum charge and to prevent over pressure conditions that could lead to venting or bursting the pressure vessels. After discharging the batteries pressure and temperature are used in the determination of when the full state of charge is reached. Initially the battery will become endothermic and cool as it charges resulting in the battery heaters cycling on and off to maintain temperature. Once the battery state of charge nears 80% the battery will become exothermic and the temperature will start to rise. At this point the temperature rise rate should be monitored along with voltage and pressure. Monitoring the battery and cell voltages, the voltage will increase until they reach full charge and then slightly decrease before charging is complete. The pressure will rise and as it approaches full charge the ability to store energy at the high charge rate diminishes and the excess energy starts to be converted to heat account for the temperature increase and an increase in the rate pressure increases.

Prolonged short discharge and recharge cycles do not have a significant affect on the capacity of Nickel hydrogen batteries.

Battery recondition is not required and in most cases is preformed to obtain measurement and verification of battery aging. The open circuit stand is beneficial in minimizing the difference between the highest and lowest cell voltages, known as cell spreading and allows cells to reach a chemical balance and independent cell voltages to equalize.

Self discharge will occur in Nickel hydrogen batteries during storage periods due to the internal resistance of the battery. To overcome this affect a supplemental charge current at low level is applied to the battery known as a trickle charge.

Due to technological advancements in battery chemistry and design, driven by power storage density Nickel hydrogen batteries batteries are starting to be replaced by more efficient Lithium-ion batteries on GEO satellites.

Additional information about Nickel hydrogen batteries is available on Wikipedia

NiCad Batteries

Nickel-Cadmium batteries (NiCd or NiCad) have been used extensively on satellites. These batteries have a heritage based on high reliability that has been proven over time. When properly operated and maintained they have provided reliable reserve power well in excess of 15 years of on orbit operations on GEO satellites.

When monitoring cell voltages during eclipse or discharging operations, the voltage will drop by as much as 0.1 to 0.2 volts over the first few minutes then stabilize and decrease only 0.004 to 0.1 volts until it reaches less than 20 % of it’s capacity. Nearing the depletion of capacity the voltage will drop by as much as 0.1 to 0.2 volts over a few minutes resembling the initial rate at the start of discharge. It is critical to stop discharging when the voltage on any cell drops to 2/3 of the initial voltage. If the cell is rated at 1.5 volts the discharge termination voltage would be at 1.0 volt. To continue discharging below 1.0 volts could result cell reversal, cell failure or complete battery failure. During discharging of the battery the temperature will increase and stabilize.

After discharging the batteries the amount of energy removed must be calculated and 110 to 120% returned to reach the full state of charge. This is based on the internal resistance and inherent characteristics of the NiCad battery. Initially the battery will become endothermic and cool as it charges resulting in the battery heaters cycling to maintain temperature. Once the battery state of charge nears 80% the battery will become exothermic and the temperature will start to rise. Charging of NiCad batteries at higher temperatures can cause a chemical reaction that produces hydrogen or oxygen gasses in the battery. At this point the temperature rise rate should be monitored and if it exceeds a rate of 5 degrees per hour then charging should be terminated to minimize potential of gas buildup that could lead to cell failure or rupture. Monitoring the battery and cell voltages, the voltage will increase until they reach full charge and then slightly decrease before charging is complete.

Prolonged short discharge and recharge cycles can lead to a diminished capacity over time, this condition is known as memory discharge. In this condition the battery will discharge normally and then prematurely discharge rapidly to a secondary voltage level creating what appears as a step in the plotted voltage over time. By completing 2 full deep discharge and recharge cycles this condition can be minimized or eliminated. The process is called battery reconditioning it involves placing a large load on the battery and discharging it until the first cell reaches 1.0 volts (or 2/3 initial voltage) then reduce the load by one half. The voltage will slightly increase then decrease back to 1.0 volts again where the discharge is terminated. At this point the battery is left with no charge or load for a hour to allow the cells to stabilize. This open circuit stand is intended to allow cells to reach a chemical balance and independent cell voltages to equalize. This also minimizes the difference between the highest and lowest cell voltages, known as cell spreading. Charging current is applied and maintained until the battery reaches a full state of charge. This cycle is then repeated for a second time.

Self discharge will occur in NiCad batteries during storage periods due to the internal resistance of the battery. To overcome this affect a supplemental charge current at low level is applied to the battery known as a trickle charge.

Due to technological advancements in battery chemistry and design, driven by power storage density to weight ratio NiCad batteries have been replaced by more efficient Nickel hydrogen, and Lithium-ion batteries on GEO satellites.

Additional information about NiCad batteries is available on Wikipedia

Batteries used on satellites

Satellite EPS systems are designed to operate on the power produced from the Solar Arrays with a battery system to store and provide power during eclipse periods and emergencies. During initial launch and commissioning the batteries are charged and discharged in the process of supplementing the power generated from the Solar Arrays. Once in a stable GEO orbit the batteries are only used to provide power during eclipses and emergencies. In LEO orbits the batteries are used every orbit to provide power as the satellite passes behind the Earth and charging is started as soon as it re-enters Sun light.

To achieve the high reliability and exceptional performance requirements over the extended life time of the battery, strict standards are maintained in the assembly process. Individual battery cells are rigorously tested. The maximum charge capacity, peak voltage, internal resistance, discharge rate and end of discharge voltages are closely matched between all cells to insure that the cells are as close to identical as possible. Independent cell voltage sensors are installed on each cell to provide the ability to monitor their performance. Monitors are also installed to measure the total battery current providing a negative reading when the battery is discharging and a positive reading when the battery is being charged. Temperature sensors are strategically located on the completed battery to sample the temperature at significant locations. Excess heat is generated during the discharge and charging of the battery and extreme low temperatures are experienced during eclipse periods. Management of the recommended operational temperature range is achieved by mounting the battery on a thermally conductive surface with the ability to radiate excess heat into cold space. To counteract extreme cold conditions supplemental heaters are selected and installed to insure that the battery can be maintained above the minimum operational temperature requirements during these periods.

The three most commonly used battery types are Nickel-cadmium, Nickel hydrogen, and Lithium-ion batteries. The batteries are designed and manufactured with the storage capacity required to power the satellite during eclipse with additional margin to account for potential emergencies. To provide added protection, the system power requirements are normally supported by two batteries to protect against failure of a single battery.

Satellite operations staff closely monitor battery and cell voltages, temperatures and currents in addition to load current and voltage during eclipse periods and battery charging. This data is collected and used to calculate the proper charge to recharge ratio based on the battery type, charging current and time required to complete charging.

A simplified battery design example would be, if the satellite uses 3000 watts of power (a small satellite) at a voltage of 36 volts using NiCad batteries with a cell voltage of 1.2 volts. The number of cells in the battery would be (36/1.2) 30 cells. The current draw form the batteries would be approximately 83.3 amps. The longest eclipse is 70 minutes, this converted to amp hours is 1.167 hours, with the resulting current draw of (83.3 x 1.167) 97.25 amps from the batteries. Applying a margin of 25% to this shows that the battery would have to be rated for over 121.6 amp/hours of capacity or if two batteries are employed, a minimum of 60.8 amp/hours each to adequately power the satellite. To minimize cost or increase the margin this value or the next larger standard battery capacity value can be selected for use.

Solar Array basics 3-axis

The solar arrays are designed to provide power based on the load requirements of all units including the payload operating on a 3-axis satellite. To determine the load all operational configurations are assessed based on the units that will be powered on and the power they require. Solar array efficiency is maximized when the Sun’s angle of incidence is at a 90 degree angle to the panel, therefore the maximum power generated will be at equinox and the minimum power produced will be during the solstices. During eclipse periods, power is supplied by the batteries, this will require additional power for battery charging post eclipse. Eclipses occur each day over a period of approximately 45 days centered around the equinox. The Power budget is developed for each configuration and the solstice period efficiency and maximum eclipse recharging requirements are used to determine the power required from the solar arrays during the worst case. A margin is applied to account for potential damage and solar cell degradation to ensure antiquate power is available throughout the mission life.

Solar arrays are comprised of a number of panels and each panel is broken down into a number of strings. Starting with the strings, they are made up of a number of solar cells connected in series to obtain the necessary voltage. For example if you require 36 volts and each cell produces 0.5 volts, then the string would have 72 cells connected in series (in series voltage adds and current remains the same). To protect the string from a damaged cell or during eclipse diodes are installed. By connecting strings in parallel (the current adds and the voltage remains the same) the output current is increased. By use of Ohm’s Law, you can use the voltage and current to determine the panel’s output power generated. Panels are then added in parallel to achieve the total power needed.

Any excess power generated by the solar arrays is controlled by the voltage regulator and power distribution section of the EPS subsystem. Based on design a number of approaches have been taken to address this, ranging from shunting the excess current, to the use of switching transistors that add or remove strings connected to the power bus.

To maximize performance the solar arrays must maintain pointing at the Sun. Solar array drive motors are used to actively track the Sun or can be stepped to keep pace with the Sun. The drive motors are connected to the solar array drive electronics providing an interface for control by the EPS subsystem, ACS subsystem or the Flight Computer based on design.

Momentum Wheels

Momentum Wheels are used in a momentum biased control systems. Similar devices are used in zero momentum bias control systems and are called Reaction Wheels, they differ in use by centering their operational RPM range around zero and therefore do not provide significant gyroscopic stiffness in the spin axis.

Momentum Wheels function by spinning a wheel of a given mass at a targeted RPM to maintain gyroscopic stiffness in a given axis, and can induce movement in a secondary axis by varying the RPM. By orientating the location of the unit to provide stiffness in the Roll axis, variations in RPM will provide control over the Pitch axis. As a satellite progresses through its orbit, Yaw error will translate into Pitch error, therefore actively controlling Pitch errors effectively manages errors in Yaw, thereby achieving influences over all three axis through the course of the orbit.

With the orientation of the Roll axis held in a relatively fixed position by the momentum of the wheel any variations in the inclination of the orbit will appear as Roll errors. Periodic attitude adjustments are made to the momentum wheel spin axis to minimize Roll errors and momentum adjustments are made to maintain the wheel speed within the optimal RPM range.

Normal maneuvers are preformed to maintain the North/South (inclination) and East/West (eccentricity) drift of the satellite.

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INFORMATION

Shining light on satellites and how they operate. Drawing from over 30 years of knowledge and experience in all phases of the life of a satellite from concept, to operations, and through end of life. You will find short topics intended to give you an understanding of how they work, the general concepts, and principals used along with information on ground systems. There is also a section dedicated to topics that can be used as basic concept training along with links to animations and 3D models I have created.