Archive for the ‘GEO Subsystems’ Category

Reaction Wheel Assemblies

Reaction Wheel Assemblies are used on LEO and GEO satellites to control errors in Pitch, Roll and Yaw axises. The assembly consists of the electronics to monitor and control the wheel, a motor and the wheel its self. A Reaction wheel works by applying current to a motor that spins a wheel with a given mass and is based on the common principal that every action has an equal and opposite reaction. By applying current to increase or decrease the wheel speed control torques can be generated in the given axis. As the wheel is spun faster it gains stored momentum in the axis and therefore in systems that employ reaction wheels, periodic momentum management is required to prevent the wheel from exceeding its recommended speed range. Momentum unload procedures typically use opposing thruster firings to drive the wheel speed down or up and reduce the stored momentum.

In a zero momentum biased control system when the unload procedure is activated, it drives each axises momentum to zero momentum. In advanced control systems this function is enabled during the standard station keeping maneuvers and can conserve fuel.

When a fourth wheel is included for redundancy, the wheel’s orientation is set such that each wheel has an influence on 2 axis and the ACS or ADACS system uses the alignment of each wheel and axis to calculate the required torque and distribute them over the wheels to control errors in an axis. In this configuration a minimum of three wheels are required to operate the satellite and the fourth wheel can be turned on when needed to replace a failed or failing wheel. The system can also distribute the torques over all four wheels if desired, providing it has been designed to do so.

TCR Antenna pattern basics

Satellite control refers to Tracking Telemetry and Command (TT&C) as the operations interfaces to control the satellite throughout its mission life cycle. These links are maintained through the communications antennas system. For launch and contingency operations additional antennas are added to allow control links when the satellites Earth deck is not pointed directly at the Earth.

During launch and contingency operations it is essential to maintain command and telemetry links to the satellite. Until the satellite is placed into Geosynchronous orbit and the sensors are locked onto the earth it is spun at a target RPM to keep it stable through the orbit and the communications antennas are stowed and can not used.  Or in cases where Earth lock is lost the satellites communications antennas are not pointed at the Earth. For these conditions additional transmit and receive antennas are selected and positioned on the satellite to provide link coverage as close to 360 degrees around the satellite.

In this case horn type antennas with a 30 degree beam width are placed on the normally Earth facing deck and the opposing Aft deck. The  Earth facing horn antennas have a beam width that allows the operations team to stabilize the satellite and reestablish pointing control in the vast majority of cases before the satellite points away from the earth. For more saver contingencies Omni type antennas are selected for their toroidal radiation pattern and larger beam width +/- 35 degrees of their centerline to transmit and receive and are positioned for use as the offset increases to either side. The Aft antenna is selected for used when the satellite rotates  to an orientation with its back to the Earth. All of these antennas are measured at the 3dB or 6dB roll off point and will provide a diminishing signal level beyond the stated beam width. The RF engineering of these designs carefully take into account the link margins required to ensure complete 360 degree coverage.

Earth Acqusition basics

In preparation for initial Earth Acquisition or as part of contingency operations to recover Earth pointing the approach is the same.  The earth sensor is configured for wide scan operations increasing its field of view.

The operations team monitors the sensor output, once earth presents is detected the sensor will provide a measurement of the pointing error that can be used to determine the rate of spin and adjust it to an acceptable range if needed.  The error will decrease as the earth approaches the center of the scan range. The ACS system is configured for Acquisition mode and then enable earth capture as the error becomes zero.  Based on the spin rate the there may be an oscillation around zero pointing as the subsystem controls the errors based on its momentum and the torque authority of the control system employed (seen in this demonstration).  The transition back to the normal operational range on the earth sensor is made either automatically or by command when the errors have settled and it is safe to do so.

Earth Sensor Error basics

The Attitude Control System ACS uses sensors to determine errors in the pointing of the satellite. The primary sensor is an Earth sensor, it has the ability to measure errors in both roll and pitch. This shows the use of 2 detectors offset to scan north and south of the equator.

By measuring the length of time the detectors see the earth and comparing the results, the difference is converted into the roll error. To determine the pitch error a measurement is taken from when the detector scenes the starts of earth presents to a center of sensor reference and compares it to that measured from the center of sensor reference to the end of earth presents, the difference is converted into the pitch error. This animation shows how the scans change as the satellite moves in roll, pitch and yaw.

With the sensor pointed at the center of the earth the resulting north and south scans will be the same.  As the sensor is moved down from center the south scan will decrease and the north scan will increase. Conversely as the sensor is moved up from center the north scan will decrease and the south scan will increase.   As the sensor moves in pitch you can see how the measurements change from the starts of earth presents to a center of sensor reference and from the center of sensor reference to the end of earth presents.

In this animation I show the Earth to make it easier to depict the interaction between movement and scan changes.  The satellite movement is exaggerated due to the sensitivity of the sensor, pointing requirements are on the order of +/- 0.05 degrees and that would be difficult to detect.

There is no significant change with yaw movement. Yaw measurements require the use of data collected from Sun sensors.  As the satellite moves along the orbit, yaw will gradually translate into roll over a 6 hour period and back on the next 6 hour period. As the yaw translates to roll the ACS system will measure and manage these errors.

Orbit Eccentricity basics

Eccentricity in an orbit causes the satellite to appear to drift east and west over the course of the day. As eccentricity increases the orbit changes from circular to elliptical path. When eccentricity is zero the orbit is circular without the appearance of any drift.  The gravitational  affects of the Earth and moon on the satellite are the primary influences that result in this gradual increase in eccentricity and drift.

To control the drift, maneuvers are carefully planed and executed to fire thrusters and reduce the eccentricity returning the orbit to its circular path.  One standard approach is to plan these maneuvers in two parts separated by 12 hours where one is an East correction and the other is a West correction.  These maneuvers are referred to as Delta-V (where V is a velocity change), or East/West depending on the preferred terminology.  They are designed to maintain the satellite in a specific orbital location, plus or minus an acceptable or defined margin called the orbital box.  A typical box is +/- 0.25 to +/- 0.5 degrees this restriction can be tighter based on the the owners requirements.  This is not to be confused with attitude pointing requirements that are much tighter and on the order of +/- 0.05 degrees or less.  To conserve fuel single maneuvers can be planed to allow the satellite to drift to the edge of the box, then execute the maneuver, reversing the drift at a rate that will slow and naturally reverse again before reaching the opposing side of the box.

In addition Start and Stop Drift maneuvers utilize the same principals, typically they are longer in duration, and are preformed to move a satellite from one orbital slot to a new one.  Drift maneuvers are normally used after launch to position the satellite in it’s orbital slot or at the end of the life as part of decommissioning.

NEC Inferred Earth Sensor Basics

One of the more popular Earth sensors in use today is produced by NEC. These sensors are mounted on the earth facing deck of the satellite and are used to measure Roll and Pitch errors. This earth sensor uses inferred detectors to sense the Earth and measure the duration of its presence in the field of view. Two inferred detectors are mounted in a fixed location in the earth sensor and an oscillating mirror is used to reflect light in the inferred spectrum into the detector.

Detector locations are offset to produce north and south scans that are compared to calculate pointing errors in the roll axis , used for control of north/south pointing.  East/west error in the pitch axis is calculated by comparing the measure start of the scan to a center reference and center reference to the end of scan to produce east/west pointing errors.

While operating in the Normal mode on this sensor the oscillating mirror travels plus and minus approximately 15 degrees of its center point. For course measurements the mirror travels can be commanded to Wide scan mode widening the scan range to plus and minus 30 degrees, used during contingency operations and during initial acquisition of the earth.

Earth sensor Roll and Pitch error signals are acted on by the ACS subsystem to maintain pointing accuracy.

Orbit Inclination Basics

Inclination is the angular difference between the orbit and the equatorial planes.  Inclination Maneuvers adjust the orbital plane by aligning it with the equatorial plane.  As satellites orbit the Earth, the Moon and Sun have noticeable affects on its orbit that cause the inclination angle to increase over time.

North/South station keeping maneuvers are designed to control inclination and are scheduled on a regular basis. Thrusters located on the North face of the satellite are used for this purpose along with a set of thrusters that will be used to control the attitude disturbances generated by the thruster firings. During preparation, thruster sets are selected, thruster firing durations are calculated and the resulting period is centered on the ascending node of the orbit to reduce the inclination angle and maintain the orbit.

SPM Orbit alignments

During the launch phase the satellite is placed into an elliptical orbit. To maintain a stable orientation the satellite is spun to add gyroscopic stiffness to the axis aligned with the orbit plain. Prior to a orbit raising maneuver or a solid fuel motor firing to circularize the orbit it is critical to align the spin axis with the targeted orbit plan. This is done by performing a Spin Precession Maneuver. Based on the satellites design, this can be accomplished by using thruster firings or momentum wheel torques.

This shows a simplified depiction of  the satellite’s movement to reinforce the concept.


One of the most active periods during the operations of a satellite in Geosynchronous Earth Orbit is during eclipse seasons. These period are centered around the vernal and autumnal equinoxes. On equinox the satellite will pass through the longest period of the eclipse season having a duration of approximately 70 minutes.

Due to refraction of the light passing through the Earth’s atmosphere the sunlight gradually fades in intensity from full sunlight to total darkness over a period of approximately 2 minuets proceeding and following the eclipse. This is call the penumbra and is depicted in gray. The area of total darkness is called the umbra and is 70 minutes in length.

The satellite operations team will prepare each satellite for eclipse before entry into the penumbra, monitor it through the umbra and  either verify or configure the charging system to recharge the batteries after the completion of the eclipse.

Before the scheduled eclipse the charging system is commanded to charge the batteries to 100% state of charge (SOC) this is shown by the increase of the green bar on the indicator. This bar turns yellow as the solar array power decreases. When the solar arrays can no longer support the power requirements of the satellite the load transitions to the batteries and this is indicated by the change to red on the status bar. During the eclipse the SOC of the batteries will decrease as the stored power is removed. By design the batteries are selected for their capacity and the ability to support the total power requirement with no less than a 25 percent margin at worst case. The SOC decreases to 25 % on exit. As the satellites exits into the penumbra the load is transitioned back to solar array power as it becomes available, the bar turns yellow again. In the absences of sunlight the solar arrays will dramatically cool to ruffly -200 degrees and will be more efficient on entry into the sun. This is shown in a slight bounce in the status indicator on exit.  When back in full sunlight the charging system is enabled and the batteries are recharged at the high charge rate.

The demo only shows the return to a 50% state of charge at the end. With an eclipse of 70 minutes, it typically requires high charge of the batteries for approximately 8 to 10 hours. This varies based on the battery type, the power load of the satellite, the initial charge state of the battery and a number of other variables.

A higher resolution AVI of this demo can be obtained through Turbosquid.

TCR Telemetry

Due to the sensitive nature of this topic, I will only address this subject from a top level general process.

Data from the satellite is collected by the Flight Computer for each subsystem and passed to the CDH subsystem where it is formatted into a telemetry stream.  This stream of data is clocked out at a defined rate to the telemetry transmitter (a dual function unit). The data is then modulated onto a transmit carrier, amplified to the transmit level and is then sent to the transmit antenna. Based on the design the telemetry transmitter can produce a single telemetry carrier, 2 telemetry carriers, a ranging carrier or a combination of both. The carriers are received on the ground at the assigned telemetry frequency.

As additional units and antennas are added to this subsystem to provide redundancy and flexibility hybrid devices or switches are installed to connect the transmitters to the antenna paths. The use of switches in the transmit path could result in a potential single point failure and therefore their use is minimized.

To accommodate higher transmit power requirements  the telemetry carriers have been passed through a payload channel. This is not a typical design, when used it is related to contingency operations configuration.

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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.