Archive for the ‘Satellites’ Category

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.

Satellite Orbit Basics

Satellites in an equatorial orbit follow an orbital path that is synchronized with the Earth’s rotation. This is accomplished by adjusting the velocity of the satellite to complete one orbit in the same time it takes for a complete rotation of the Earth. This results in the satellite  appearing to be a stationary location in the sky.

During the spring and fall seasons in witch the equinoxes occur the satellite will pass behind the Earth during a portion of it’s orbit resulting in an eclipse as seen by the satellite.  The duration of the eclipse periods range from a few minutes to as long as 70 minutes on the actual day of equinox.

For more information please see my post titled Earth’s Orbit of the Sun or Eclipse.

Eclipse

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.

TCR Command

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

TCR commanding is accomplished by transmitting the modulated commands on a RF carrier that is tuned to the command frequency. The command carrier is received at the antenna and is applied to the input of the command receiver. By design the command receiver is a dual function unit, it will lock onto a receive carrier and demodulate valid command signals to produce a digital output. Once the command receiver is locked onto the command carrier, it verifies and demodulates commands and outputs them to a Command and Data Handling subsystem or Flight Computer for execution.

As additional units and antennas are added to this subsystem to provide redundancy and flexibility hybrid devices are installed to split the received signals. The use of switches in this receive path could result in a potential single point failure and are therefore avoided. The use of hybrids allows the signal to be divided and applied to each receiver and since hybrids are passive devices the chances of failures are minimized.

TCR Ranging

During the launch and early operations phase a path through the TCR subsystem is provided to collect phase angle range data for orbit determination. This path can also be used during the mission life if desired.

TCR ranging is accomplished by transmitting the ranging signal on a RF carrier that is tuned to the command frequency and measuring the phase angle difference of the receive signal on a RF carrier tuned to the telemetry frequency. For this reason it is normally recommended that no commands are transmitted while ranging is being preformed. By design the command receiver is a dual function unit, it will lock onto a receive carrier and demodulate the signal to produce a digital or analog output signal. Once the command receiver is locked onto the ranging carrier, it demodulated the range signal and outputs it to the input of the telemetry transmitter. The telemetry transmitter (another dual function unit) then modulates the ranging signal onto a transmit carrier and amplifies it to the transmit level and is then sent to the transmit antenna. Ranging carriers are received on the ground at the assigned telemetry frequency.

To accurately process the range data all variables must be accounted for. During the integration and test phase of the satellites assembly each range path through the TCR subsystem is calibrated. To calibrate each path the subsystem is configured for ranging, the path configuration is noted and ranging is preformed. The resulting range measurement is the delay through that path and is referred to as the satellite delay.

When processing range measurements, to improve accuracy you must account for delays through the satellite and ground station. Once these delays have been accounted for then the range measurement is the distance to and from the satellite. By dividing the measurement by 2 you get the range from the ground antenna to the satellite.

Ranging can also be preformed by transmitting and receiving the range signals through a given payload channel.

Flight Software systems

Just like any other computer the satellite has it’s own version of software. Each satellite has at least one computer on board used for its control along with RAM (random access memory) and PROM (programed read only memory) . The flight software is resident in PROM and is copied to RAM to run at system start-up.  Some satellites also have the ability to modify the software while on orbit.  This is accomplished by use of EPROM (erasable programed read only memory) or EEPROM. When EPROM’s are used there is still a PROM for start-up then it copies the software from EPROM to RAM to run.  Once the flight software (FSW) process have been started it will run in RAM until it is restarted. The flight software collects data or issue commands to and from each unit on a scheduled basis. Once collected the data is either processes it for transmission as telemetry or passes it to the C&DH system to complete that process.  Commands from the ground or commands passed from the C&DH system are processed and executed in FSW. The ACS software also run as part of FSW to collect sensor data, produce error control signals and apply the signals to the control actuators to maintain the Attitude and pointing of the satellite. The FSW in complex satellite architectures will also interface with independent processors for each subsystem to collect telemetry and process commands. Flight recorders have been employed to store data on the satellite for retrieval at a later time, or store Attitude parameters used at start-up to minimize transients. Without the use of EPROM ground commands may be issued to change the software running in RAM on a contingency basis to correct problems, if this is done then the same commands will be required after every re-start.

If the architecture includes a C&DH system the actual Flight Computer hardware is included in that system.

Satellite Payload systems

Satellites provide a platform for a wide variation of applications.  The equipment placed on the satellite to accomplish the intended mission is what pays for the satellite and is designated as the Payload. Although communications is one of the most common uses, payloads can include imaging of the Earth or objects in space, global positioning systems (GPS), ranging and altimeters used in mapping, spectral analysis devices to determine the composition of the atmosphere.  In the past the payloads have been selective in three primary categories, Communications, Imaging and Scientific. I would consider GPS satellites as a forth category unto itself. Some satellites are now being built with shared payloads to provide a more cost effective means to collect scientific data for long term studies.

Communications satellite payloads range in complexity from a few channels in one frequency band to hybrid systems that have multiple channels using a number of frequency bands including frequency conversion and switching systems to translate and rough the traffic to selective areas. These satellites can receive signals from the ground and retransmit them to the ground or another satellite. Applications are also in place to communicate with ships, trains, trucks and airplanes. The basic types of communications are voice, video and data.

Imaging satellite payloads are governed by the resolution of the ground image they can provide. The images can be produced in the visible light spectrum, normal color or black and white (for higher resolution), inferred spectrum for tracking temperature differentials or other ranges based on the design of the camera used. Space imaging also uses Ultraviolet and Gama-ray spectrum among other types. These satellite may also include mapping instruments that produce topographical mapping information utilizing for example, radar or laser mapping techniques.

Scientific payloads are just that, and include a diverse range of equipment. These satellites are designed to collect very specific information. Some of the applications have been to collect data and measure the irradiation from the Sun, Sun spot activity, solar flair and wind measurements, measure the ozone layer, map carbon dioxide concentrations, search the stars and the list goes on. Scientists are developing new missions every day to gain more knowledge.

Propulsion systems

On a satellite the propulsion system is designed to provide a means to control the satellite and maintain it’s orbit.  The two most common applications are mono-propellant and bi-propellant. The mono-prop system utilizes one type of fuel, where the bi-prop system uses a fuel and an oxidizer combination to produce thrust. In it’s simplest form this system is comprised of thrusters,  a fuel tank, tubing for fuel lines, heaters, valves and filters. More complex designs also include a dedicated processor, oxidizer tank, pressurization tanks and lines, additional power supplies, and safety relays and fuses. The propulsion system can be pressurized prior to launch and operate in a blow down mode. When pressurization tanks are added to the system it can operate in a combination of pressure regulated and blow down modes. Electric Ion thrusters have been used, they have a low specific impulse and require high current and long maneuver durations so are therefore normally reserved for use in planetary missions. The mission life is determined primarily by the efficient use of the fuel loaded into this system.

In some applications due to short mission life, on the order of 6 months to a year the propulsion system is or can be eliminated. These are normally proof of concept, flight qualification and special scientific missions that do not require long durations.

Satellites in LEO orbits use thrusters to adjust the satellites velocity, orbiting altitude and inclination.

On GEO satellites the designs normally include the use of thrusters for Attitude control as a back-up system. Under 2-Axis control the fuel tanks can be oriented to utilize the centrifugal effect of the spinning satellite to force the fuel out to the thrusters.  While in 3-Axis control the thrusters can be used to reduce the momentum stored in the reaction wheels. During the launch phase the use of solid fuel motors to achieve a circularized orbit at GEO has been replaced in some cases with bi-propellant thrusters to conserve weight and potentially extend mission life.

Command and Data Handling subsystems

Command and Data Handling subsystems are the portion of the satellite that acts on commands from the ground or internal commands generated by the Flight Computer (FC) and collects telemetry from the other subsystems and prepares it for transmission to the ground. This subsystem can be as simple as one Flight Computer processor that controls the entire satellite. In distributed architecture applications there can be a Flight Computer, a processor for each subsystem, a data bus controller that provides data and command transfer between subsystems and a C&DH processor to interface with the TCR subsystem. In the more complex systems there are backup systems to minimize impact due to unit failures. Flight recorders that can also be utilized to store data and memory areas to store command tables for scheduled execution.

This subsystem contains the processor or processors, interface bus, the means to receive, decode, process and execute commands and the ability to collect, process, store and transmit telemetry from the satellite through the TCR system.

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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.