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Pioneering Communications Satellites


This page describes the evolution of key satellite technologies pioneered by the early satellite communications systems.


( The contributions and background of key players are described the History page. Follow the individual (history) links below to see more.)



Telstar Satellite


Telstar Ground Station

Telstar Ground Station Horn Antenna

Telstar 1, launched by a Delta #11 rocket in 1962, carried the first ever transatlantic television images and the first live satellite phone conversation. It was roughly spherical with a diameter of 34.5 inches (87.6 cms) weighing 170 pounds (77 kg) and was launched into an elliptical orbit with a period of 2 hours and 37 minutes, inclined at an angle of approximately 45 degrees to the equator. With a perigee of about 592 miles (952 kilometres) above the Earth and an apogee about 3,687 miles (5,933 km), its orbit straddled the border between the LEO and MEO orbits.


Due to its non-geosynchronous orbit, Telstar's availability for transatlantic signals was limited to the 20 minutes in each 2.5 hour orbit when the satellite passed over the Atlantic Ocean. Ground antennas had to track the satellite with a pointing error of less than 0.06 degrees as it moved across the sky at up to 1.5 degrees per second.

Station Keeping relied on solely on passive spin-stabilisation to resist the influence of external perturbations and keep the satellite on track with the desired orientation in space. The satellite derived its initial spin from the third stage of its Delta launch vehicle typically spinning at 200 r.p.m. before they separated on reaching the desired orbit. A viscous ring mechanical damper prevented the onset of precession or coning of the satellite itself from causing further instability.


Power Supply was by means of 3,600 solar cells covering most of the outer surface which produced a tiny 14 Watts of power which was not enough to support the continuous operation of the communications system. However, the communications electronics were only used during 20 minutes of each orbit when the satellite was visible to both the sender and the receiver. During the rest of the time the solar panels charged a 19 cell Nickel Cadmium battery which provided supplementary power when needed.


The Transponder had a single channel with a bandwidth of 50 Hz which could carry 600 one-way voice circuits or one TV channel and 60 two-way voice circuits. The uplink frequency was 6390 MHz and the downlink was 4170 MHz. The electronics were all solid state using 1064 transistors and 1464 diodes except for the transmitter amplifier which used a travelling wave tube (TWT) delivering 3 Watts, in its output stage since there were no solid state devices at the time capable of high power output and high bandwidth performance at the 4 MHz. transmitter frequency. The receiver noise figure was 12.5 dB.


The Satellite's Antennas were in the form of an array of slots or cavities spaced evenly around the waist of the satellite body, with 48 for the transmitter and 72 for the receiver. They produced a toroidal shaped pattern which was omni directional in the equatorial plane of the satellite but with a beam width of about 30 degrees in the axial direction.


The Ground Station Antennas in the USA and France were specially constructed, steerable, horizontal, conical horns, each with a parabolic reflector at its mouth which re-directed the beam to a small feed antenna at its focus providing a gain of 58 dB in the azimuth direction and 61dB in elevation. They were 94 feet (28.7 m) high and 177 feet (54 m) long with an aperture of 3,600 square feet (330 m2) and weighed 380 tons (340,000 Kg). The transmitter TWT power output was 2 kiloWatts continuous wave (CW)..

The UK antenna at Goonhilly Downs was an 85 foot (26 M) steerable parabolic dish with a transmitter power of 5 kiloWatts.


Telemetry and Control were accessed through a four-element helical antenna located on the top of the satellite. The telemetry functions were mainly status reporting on voltages, temperatures and signal levels and the command functions controlled the power management.

Telemetry was provided via a 200 mW transmitter operating at 136.05 MHz, and control from the ground station was via a 123 MHz radio link

Telstar had no capability for manoeuvring or attitude control.


See Telstar History


Syncom 1, 2 and 3

Harold Rosen with Syncom Prototype

Hughes Aircraft Company engineers Thomas Hudspeth (left) and Harold Rosen, the project leader hold a Syncom Satellite Prototype atop the Eiffel Tower during the 1961 Paris Air Show.


Syncom 2 Satellite

Syncom Satellite Components

The three Syncom satellites were experimental spacecraft built by the Hughes Aircraft Company to demonstrate the feasibility of using Geostationary Earth Orbit (GEO) Satellites to provide a world-wide communications system.


They were the first to provide solutions the problems of placing a satellite into a GEO orbit and the necessary station keeping and attitude control necessary to keep it on station in a stable orientation with its antenna beams pointed towards the ground and a minimum number of solar cells pointed towards the Sun as the satellite rotated.


Common Features


The Syncom satellite had a short cylindrical body that was spun about its axis to provide spin-stabilisation in orbit. Hydraulic dampers were used to suppress nutation of the spinning satellite. The antennas were mounted beyond one end of the body and were collinear with the satellite axis. All the satellite equipment was contained within the body. This design formed the basis for several later geosynchronous satellites. The communication subsystem had two receivers and two transmitters for redundancy; either receiver could be operated with either transmitter. It provided for two 500 kHz channels for Narrow Band two-way communications and one 5 MHz channel for one-way Wide Band transmissions. (These capabilities could not be used simultaneously.)


The capability of the Delta rocket launch vehicle limited the size, weight and shape of the satellites which were all cylindrical in shape, with a diameter of 28 inches (71 centimetres) and a height of 15 inches (39 centimetres) and a launch weight of 150 lb (68 kilograms) fully fuelled. About half of the launch weight consisted of the apogee motor with its fuel and pulse-jet control systems necessary to achieve synchronous orbit.


Most of Syncom's interior consisted of the tanks and combustion chamber for the apogee motor, around this were arranged two hydrogen peroxide and two nitrogen tanks and the electronics.

All electrical subsystems had to be extremely small and light, yet most subsystems were made redundant to enhance reliability


Electrical Power for the satellite electronics was provided by 3840 silicon solar cells mounted around Syncom's cylindrical surface delivering 29 Watts during the 99 percent of the time that the spacecraft was in sunlight. Nickel-Cadmium rechargeable batteries provided power when the spacecraft was in the Earth's shadow.


Motive Power

The Propulsion Subsystem for Orbital Insertion was manufactured by Thiokol and consisted of a single solid-propellant rocket "Apogee Motor" extending from the bottom of the cylinder and delivering 1000 lb-thrust designed to impart an incremental velocity increase of precisely 4696 feet/sec (1431 meters/sec) when it kicked in.


Telemetry, Tracking and Command



Syncom incorporated sensors for monitoring the status of the satellite's systems and transmitting this information back to ground control. The data included voltages, temperatures, pressures and signal levels as well as timing and frequency references. The satellite's orbital position could be determined directly by ground control, but not its attitude and one of Syncom's key innovations was the method of determining the satellite's attitude coupled with a method of manoeuvring the satellite and controlling its attitude.


Attitude Sensing

In order to optimise the signal levels received at the antennas of the satellite and its ground station, the main beams of the satellite's antennas (the plane of the toroidal beams) should point towards the Earth. At the same time, to capture the maximum solar energy the maximum possible area of the solar cells (the curved surface of the satellite body) should point towards the Sun. The condition for maximum signal reception is critical and this was met when the satellite's spin axis was parallel to the Earth's axis of rotation and maintaining this required precise control of he satellite's attitude in space.

Fixing the satellite's spin axis in this orientation is however less than optimum for solar energy capture since the angle of incidence of solar radiation will vary between the winter and summer solstices. Though energy capture is important, it is slightly less critical than achieving optimum signal transmission. See Incident Solar Energy on the Satellite Technology page.


Sensors on board the satellite determined its actual attitude so that the attitude control system could detect any deviation from the desired attitude and send an error-correcting control signal to the satellite to bring it into line. Two methods were used to provide an unambiguously determination of the satellite's attitude.

  • Solar Pulses
  • The first was Syncom's ingenious "V-beam" sensor system. Its cylindrical body contained two slits, a vertical slit and a slanted slit and behind each slit was a solar cell which generated its maximum output when the slit was in line with the Sun. (See Attitude Control diagram below) The effective beamwidth of the sensors was 0.7 degrees perpendicular to the slit and ± 80 degrees along the plane of the slit so that the beam was in the shape of a thin fan.

    As the satellite rotates, the solar cell behind the vertical slit will generate an electrical pulse as it passes the Sun line and as it continues its rotation the slanted slit will generate a second pulse as it passes through the Sun line.

    The frequency of the solar pulses from the vertical slit gives an indication of the spin rate and the timing between the pulses can be used to determine the phase angle of the satellite's rotation. The interval between the pulses from the vertical and slanted slits can also be used to determine the instantaneous phase angle of the satellite about its axis, but its prime purpose is to determine the tilt angle between the satellite's spin axis and the Earth's axis of rotation.

    Because one of the slits is slanted with respect to the other, their two fan shaped beams converge towards the base of the satellite in a "V"shape. When the spin axis of the satellite is in a fixed angle with respect to the Sun, the period between the pulses from the vertical and canted slits will be constant, but if the satellite tilts away from the Sun, the elapsed time between the pulses will reduce when the Sun is below the satellite's equator. Similarly the time between the pulses will increase when the Sun is above the satellite's equator. Since the angle between the slits is known, the angle of tilt between the satellite's axis and the Sun line (the Sun angle) can be calculated.

  • Antenna Polarisation
  • The second method of determining the spin-axis attitude was by measuring the polarisation of the radio signals received from the satellite. The plane of oscillation of the electric (or magnetic) wave radiating from a simple antenna, known as its polarisation, is fixed and is determined by the position and structure of the radiating element or elements of the antenna. See more about radio wave polarisation. Tilting the radiating element of a transmitting antenna will thus tilt the orientation of the transmitted electromagnetic field. The signal output from a receiving antenna placed in the radiated electromagnetic wave will be a maximum when its plane of polarisation is in line with the polarisation plane of the transmitting antenna, reducing to zero when the polarisation planes are perpendicular to eachother.This can be used to determine the orientation of the antenna and hence the structure to which it is attached.



The inputs necessary for satellite tracking during the launch stage are mostly provided by the ground station. A beacon signal transmitted from the spacecraft is used by the ground stations to assist antenna tracking and as a reference signal in making range rate measurements.

  • Satellite Range - Determined by timing circuits which measure transit time delay between signals transmitted from the ground station to the satellite and received back at the ground station after retransmission by the satellite.
  • Range Rate or Velocity - Determined by measuring the two-way Doppler shift resulting from a transmission to the moving satellite and subsequent retransmission back to the ground station. This only works before the satellite enters its geosynchronous orbit.
  • Azimuth and Elevation - Derived from the azimuth and elevation of the tracking antenna which followed the movement of the satellite by keeping the signal received from a beacon on the satellite maximised within the antenna's narrow beam (locking on to the beacon). The accuracy of azimuth and elevation information was approximately one degree. Elevation measurements of less than 20 degrees were not accurate enough for automatic tracking.




On Board Control Systems

Syncom Attitude and Position Controls

Manoeuvring and Station Keeping was by means of a Spin-Synchronised Reaction Control System consisting of two Gas Jet Thrusters which together provided three dimensional control of the satellite's position as well as its attitude. See diagram opposite.

Since the satellite was spin-stabilised, control depended on knowing the precise angular position of the satellite on its spin axis as it was spinning as well as the angle (tilt) of the spin axis with respect to the Earth's spin axis.

The pulsing of the gas jet thrusters was managed from the ground by means of a synchronous controller which took its timing reference from the satellite's V-beam Sensor system. See diagram opposite.

There was no control over the satellite's rate of spin which was imparted by the launch vehicle before separation.

Because the satellite operates in a vacuum above the atmosphere it is not subject to aerodynamic drag. At the same time the Earth's gravitational pull is balanced by the centrifugal force due to the satellite's orbital motion, so that the forces necessary to move it are tiny and once it is moving there are no external forces causing it to slow down.


  • Control of Velocity Along the Orbital Path and Lateral Position Control were provided by means of a single gas jet thruster mounted with its jet opening on the outer surface of the cylindrical body of the satellite, pointing in a radial direction through the satellite's centre of gravity, Control was implemented by pulsing the thruster with a short jet of gas at the appropriate moment during the satellite's spin rotation.
  • To reduce the satellite's velocity, the radial jet was pulsed when it was pointing forwards in the direction of travel as the satellite body rotated. To increase the satellite's orbital velocity the jet was pulsed when it was pointing backwards in the direction from whence it came.

    Similarly, to increase the satellite's altitude (its orbital radius, perpendicular to the direction of travel) the radial jet was pulsed as the satellite's rotation brought the jet into the direction pointing towards the Earth and to decrease the altitude the jet was pulsed when it was pointing away from the Earth.


  • Attitude and Velocity Control were similarly implemented by means of a second thruster pointing in a direction parallel to the satellite's spin axis, mounted on the outer edge of the circular flat plate which held the apogee motor nozzle. Pulsing this off-centre axial thruster at the appropriate moment tilted the spin axis of the satellite in the desired direction.
  • By providing a constant jet thrust during a full spin revolution the satellite could also be moved in the direction of its spin axis.


    Before reorientation, operation of the axial jet in a continuous mode creates a thrust along the spin axis of the spacecraft to control its orbital speed.


    The spacecraft control system could provide velocity increments of 0.05 ft/sec (0.015 m/sec) and could control the orbit several orders of magnitude better than the orbit could be measured.


    The on-board control system was duplicated to ensure reliability. Syncom I used hydrogen peroxide propellant in one set of thrusters for coarse control and high pressure nitrogen gas for fine control in the second, (duplicate), control system. Subsequent designs dispensed with the nitrogen gas jets using only hydrogen peroxide jets in both systems and nitrogen gas was only used to pressurise the hydrogen peroxide tanks.


Ground Control Systems

There are two aspects to ground control systems, launch control and operating control.

  • Launch Control functions include tracking the spacecraft and the satellite, monitoring their status and position and controlling the timing, magnitude and direction of the propulsion components used to place the satellite in orbit. It consists of decoding equipment for the telemetry inputs from the satellite and the ground based tracking systems. This included the synchronous controller which operated the satellite's jet thrusters, command encoding equipment and auxiliary equipment necessary to place Syncom in geosynchronous orbit above its desired position of longitude.
  • The trajectory of the launch followed a pre-programmed plan with the status monitored by the ground station. Error signals generated by the satellite tracking system were used to keep the rocket launch vehicle on track whereas telemetry from the satellite itself was used to generate the signals to control Syncom's position and attitude.

  • Operating Control includes all the systems needed to keep Syncom in its desired position once it is orbit. It also includes the systems used to monitor the status of its propulsion, power, electronics and communications systems and the facility to switch between the operating and standby control systems as required or to select between on board communications channels. Station keeping uses the same position and attitude control systems as launch control.

Syncom Transponders

  • The Communications Channels
  • The prime objective of the Syncom satellites was to demonstrate the feasibility of geostationary satellite systems so the specification of the transponders was rudimentary.

    • Syncom 2 was only capable of carrying a single two-way (duplex) telephone conversation, or 16 one-way Teletype connections (slow speed digitally coded ASCII text streams) just to show that the system worked.
    • The communication subsystem had two receivers and two transmitters for redundancy. Either receiver could be operated with either transmitter. Two 500 kHz channels were available for narrow band two-way communications and one 5 MHz channel for one-way wide band transmissions but these capabilities could not be used simultaneously. Selection of the active receiver and transmitter was made by ground command.

    • For Syncom 3 the 500 kHz channel was eliminated and replaced by a 10 MHz bandwidth channel for television transmission tests with a 50kHz option for small station testing.

Communications Signal Transmission

    • The Uplink (satellite receiver) frequency was 7363 MHz and the receiver noise figure was 10 dB.

    The receiver in the transponder amplified the signal and translated its frequency to 1815 MHz in a frequency down-converter for onwards transmission, to avoid the higher power transmitted signal from interfering and swamping the very weak received signal.

    The signal was further amplified by a travelling wave tube (TWT) with an output level of 2 Watts in the transponder's transmitter. A second TWT was kept on standby.

    • The Downlink (satellite transmitter) frequency was 1815 Mhz.

The Antennas

  • Communications Antennas
  • The radiation patterns of fixed transmitter and receiver antennas used in spin-stabilised satellites, or "spinners' as they are often called, must be symmetrical about the satellite spin axis so as to be pointing at the Earth at all times even though the satellite is spinning. They were therefore mounted in a single, in -line, structure on top of the satellite, co-linear with the spin axis. The transmitting antenna was a coaxial 3 element slotted array with a gain of 6 dB which produced an omnidirectional toroidal shaped pattern with its plane perpendicular to the satellite's spin axis, centred around the antenna with a beamwidth of 23 degrees along its axis. The receiving antenna was a slotted dipole, protruding from the top of the transmitter's antenna, with a gain of only 2 dB providing a similar omnidirectional pattern but with a wider vertical beamwidth. With this arrangement, maximum signal levels were received when the satellite's spin axis was parallel to the Earth's axis of rotation so that the plane of the toroidal antenna pattern pointed directly towards the Earth and the signal remained constant as the satellite rotated.

  • Telemetry and Command Antennas
  • Four simple monopole, whip antennas oriented normal to the spin axis were provided for 136 MHz telemetry and 148 MHz command transmissions. The tracking beacon transmitted at 1820 MHz which was close to the frequency of the satellite's communications down link and benefited from the large directional ground station antenna.


Ground Stations

Various ground stations around the world were used to communicate with Syncom by means of parabolic antennas with diameters ranging from 60 feet (18 metres) down to 10 feet (3 metres). A typical ground station used a 30 foot (9metre) diameter antenna with a transmitter power of 20 kW and a receiver noise figure of between 3 and 5 dB.


Orbital Injection - Placing Syncom 3 in Geosynchronous Orbit

To place the Syncom into an equatorial synchronous orbit from a launch site off the equator two orbital plane changes were required. See diagram of Orbital Manoeuvres.

The satellites were launched into the transfer orbit by the three-stage Delta rockets.The Syncom on board orbit injection propulsion subsystem provides the boost necessary to inject the spacecraft into a nominally synchronous, circular orbit, after the vehicle has reached the apogee of the transfer orbit at the required altitude. The launch stages were as follows:


  1. The first and second stages of the Delta rocket place the third stage into a parking orbit at an altitude of about 215 miles (346 kms) inclined at 28.7 degrees to the Earth's equatorial plane (which is the latitude of the Delta rocket's launch pad at Cape Kennedy) with the rocket's axis aligned with its orbital plane and perpendicular to the Earth's axis (that is perpendicular to the Earth's spin axis).
  2. The first orbital plane change was made at.the perigee of the parking orbit. During the coasting period of the parking orbit after the stage 2 burnout, the vehicle was yawed to the left just before the third stage burn which lifted it into a highly elliptical transfer orbit with a perigee of 700 miles (1127 kms) and an apogee equal to the geostationary altitude while simultaneously reducing its equatorial inclination from 28.7 degrees, to 16.5 degrees. At the same time secondary jets, in the Delta rocket's third stage spin, the spacecraft and its payload up to 165 rpm with its spin axis aligned with the direction of travel to impart the desired spin stabilisation to the satellite just before it separates from the spacecraft.
  3. The second plane change was made by aligning the satellite towards the equator and firing Syncom's on board "apogee motor" with a very precise burn when the satellite was at the apogee of the transfer orbit thus increasing its velocity to its synchronous velocity and pushing it into a circular orbit while at the same time reducing its inclination to near zero.
  4. In order to reduce the orbital inclination to zero the satellite spin axis was tilted using the axial jet in pulsed mode to enable the satellite's orbital plane to be tilted into the equatorial plane by means of subsequent velocity adjustments.
  5. Further velocity adjustments were made by the Syncom's axial jet operating in continuous mode to reduce the eccentricity of the orbit and to bring it to its geosynchronous position over the equator at 180 degrees longitude in order to optimise it's footprint over the desired hemisphere of the Earth.
  6. (The stationary orbital position of 180 degrees, the International Date Line, was chosen for Syncom 3 in order to transmit television images from the 1964 Tokyo Summer Olympics in Japan to the United States.)

  7. Re-Orientation - Once Syncom was in position, the axial jet system was used once more, this time in pulsed mode to tilt its spin axis by 90 degrees to bring it parallel to the Earth's spin axis and normal to the equatorial orbital plane to optimise the communications signal transmission and the solar energy capture. See Satellite Attitude
  8. Final adjustment of the satellite's longitudinal position was accomplished by using the radial (lateral) jet in pulsed mode.


Syncom Milestones

Syncom 1

Launched in February 1963, after successful separation from the Delta B #16 launch rocket, one of the nitrogen tanks in Syncom’s control system ruptured during the crucial apogee burn to boost the craft from its transfer orbit into geosynchronous orbit and communications were lost.


Syncom 2

Launched in July 1963 by a Delta B #20 launch rocket, Syncom 2 the world's first geosynchronous satellite was successfully placed into orbit at an altitude of 22,230 miles (35,786 kms), demonstrating conclusively the practicality of achieving a geosynchronous orbit and the possibility of global telephone, television and data transmissions via satellite. Unlike Telstar, launched the previous year, it provided 24 hour continuous communications transmission. Because of the limited propulsive energy of the Delta B rocket it was not able to achieve a geostationary orbit and so Syncom 2's designated orbit was therefore inclined at 33 degrees to the equator so that it moved in an elongated figure eight pattern 33 degrees north and south of the equator centred around a longitude of 55 degrees which brought it over the Atlantic ocean and Brazil.


Syncom 3

Launched in August 1964 by the more powerful Delta D #25 launch vehicle equipped with three strap on booster rockets, Syncom 3 was the first geostationary communication satellite. Although Syncom 3 is sometimes credited with the transmitting the first television program over the Pacific Ocean, however the Relay 1 satellite with an orbital period of 185.09 minutes was the first to transmit television from the United States to Japan in November 1963 but like Telstar it only provided intermittent reception and required a tracking antenna to receive the signals.

Syncom 3 was a great success and set the standard for a generation of spin stabilised geostationary satellites.


See Syncom History


Intelsat 1 - The Early Bird

Intelsat 1 - The Early Bird

Intelsat 3

Intelsat III

Intelsat I - The Early Bird


Launched on 6 April 1965 by a Delta D rocket, Intelsat I was a larger, commercial version of Hughes experimental Syncom satellites, basically an up-rated model of Syncom 3 with a pre launch weight of 328.5 lbs (149.0 kgs) accommodating higher capacity transponders designed to carry commercial traffic.

Its two 6 Watt transponders operating at C-Band (6GHz uplink-4GHz downlink) each with 50 Mhz bandwidth could carry either 240 voice circuits or one TV channel but not simultaneously.

Primary power was provided by a larger array of solar cells delivering 45 Watts, increased from Syncom's 29 Watts..


See Intelsat 1 History


Intelsat III

Launched 19 September 1968 by a more powerful Delta M rocket, Intelsat III was designed by TRW but bore a striking resemblance to Hughes Intelsat I. It was double the weight of Intelsat 1 weighing 646 lbs (293 kgs) including propellant and incorporated several performance improvements, the most important of which was the technology enabling the use of high gain antennas.

Previous spin-stabilised satellites used antennas with radiation patterns which were omnidirectional about the satellite's spin axis in order to maintain communications with the Earth as the satellite body rotated, but this was very wasteful since most of the satellite's RF power was radiated into space.


The De-Spun Antenna: To enable the use of a high gain directional antenna, the antenna must be kept in a fixed direction pointing towards the Earth at all times otherwise communications would be intermittent if the narrow beamed antenna rotated with the satellite body.

Intelsat III was the first communications satellite to solve this problem which it did by de-spinning the antenna using a motor to rotate the antenna at the same speed as the satellite spin but in the opposite direction so that it appears stationary. Infra-red solar sensors which detected the Earth's horizon were used to synchronise the motor speed with the satellite's speed of rotation. It used a de-spun directional horn antenna, 34 inches (86 cms) tall, with a gain of 15.6 dB replacing the previous slotted dipoles which had a gain of only 4 dB, increasing its effective radiated power and greatly enhancing the transponder's signal to noise ratio and thus its usable bandwidth. See Signals and Noise.

Instead of the hydrogen gas jet thrusters as used in Intelsat I, Intelsat III used hydrazine propellant for station keeping.

Primary power was increased to 178 Watts peak by increasing the number of solar cells with the energy being stored in a 9 AmpHour Nickel Cadmium battery.

This extra power was needed to drive the de-spinning motor but it also allowed the use of higher power TWT amplifiers in its C-Band (6GHz uplink-4GHz downlink) transponders. The up rated power output of 12 Watts per channel improved the signal to noise ratio even further.

These potential improvements in noise performance enabled the bandwidth of each channel to be increased to 300 MHz, sufficient to carry 1,500 telephony circuits or four television channels.


See Intelsat III History



Molniya Satellite

Molniya 1 Satellite

Molniya 1 Satellite

Launched on 23 April 1965, the Molniya 1 (Russian "Lightning") was Russia's first communications satellite. It's design concept and its orbits differed in almost every aspect from the early spin stabilised and later three axis stabilised systems developed in the USA. It was also much heavier than contemporary US systems with a launch weight of 3630 lbs (1650 kgs), more than 10 times the weight of the Early Bird satellite launched the same year. Its transmitter power outputs of 40 Watts and 20 Watts were also 10 times greater than the Early Bird's.

Electronic systems were contained in a cylindrically shaped body with conical ends 14.4 feet (4.4 m) tall and 4.6 feet (1.4 m) in diameter. An active liquid cooling system kept the components at a stable temperature during the day-night cycle.


The Molniya Orbit

Molniya satellites used a highly elliptical orbit which enabled them to concentrate their signal coverage into a footprint spanning the whole of Russia and its immediate neighbours including much of the Arctic polar region.


See more about the Molniya Orbit and its benefits.



Molniya's body was designed with the flexibility to house different civil and military applications. The first examples carried out experimental TV transmissions with the uplink transmitting TV signals from Moscow to the satellite and the downlink transmitting to 20 ground stations in cities in Siberia and the Russian Far East including Norilsk, Khabarovsk, Magadan and Vladivostok.

In 1967 Orbita, the world's first national satellite television network was set up with Molniya satellites relaying the Moscow transmissions to ground stations across the country which received the signals through 40 to 50 foot (12m to 15m) parabolic antennas and converted the signals to frequencies suitable for reception by domestic TV receivers and re-broadcast the programmes to local communities through conventional TV transmitters.

The other major early application was long range military communications.

Later examples were used for multi channel telephony, mobile radio systems, monitoring weather systems, Earth observation and photography.


On Board Power

One of Molniya's design goals was to minimise the use of propellants for attitude control and the weight penalty they incurred by maximising the use of renewable solar power where possible for this purpose. It therefore incorporated six large windmill-type solar panels, fixed to the satellite body, with a span of 26.90 ft (8.20 m) which provided up to 1 kW of electrical power.

Molniya's lifetime was limited by the vulnerability of its solar cells and other electronic components to electromagnetic radiation as it passed four times per day through the inner and outer Van Allen radiation belts.


Stabilisation System

As well as the use of very large solar panels, a second design goal was the use of high gain antennas but neither of these goals could easily be accomplished with a spinning satellite. The body of the Molniya satellite was therefore deigned to be static to avoid these limitations. Nevertheless, it still used gyroscopic stabilisation, but instead of spinning the satellite body, it used an internal gyroscope aligned with the satellite's axis which achieved the same effect.


Attitude Control

The Molniya satellite used a three axis attitude control system. Attitude sensing for the satellite body was by means of a Sun sensor mounted near the centre of the solar arrays. Two pairs of small reaction thrusters, one pair on each of the two axes orthogonal to the main satellite axis, were used in a control system to adjust the orientation of the body so that its axis and thus its solar panels were pointing directly towards the Sun to optimise the capture of solar energy. The pointing accuracy of 10 degrees was low but sufficient.

Two more jet thrusters were used to control the angular position of the satellite body around the gyro spin axis and to damp any tendency for the body to spin about its axis.


The Antennas

Molniya's two antennas spaced 180 degrees apart were electrically steerable but only one was employed at any one time and because of the satellite's highly elliptical orbit, the active antenna was only used eight hours per day. The second antenna was kept on standby.

During communications sessions the antenna pointing system used two optical horizon sensors to detect the position of the Earth and an electrical motor control system to point the active antenna towards the Earth's centre. Because the angle between the Sun line and the Earth line varied as the Molniya moved around its orbit, the antenna positioning system was in constant action while the satellite was communicating with the ground stations, but this usually only happened for eight hours per day while it was passing over Russia. The rest of the time it the transmitter and the antenna control motors were switched off to conserve power.

Antenna gains were approximately 18 dB



A variety of electronic component technologies were used in the transponders, mostly solid state but also metal-ceramic triodes. klystrons, magnetrons and traveling wave tubes. Because of the harsh operating environment most of the electronic systems were duplicated with one system operational and redundancy provided by one system and sometimes two systems on standby.

For government and military applications the transponder uplink frequency was 1.0 GHz and the downlink frequency was 800 MHz with a transmitter power of 40 Watts

TV channel frequencies were 4.1 GHz for uplinks and 3.4 GHz for downlinks transmitting with 40 Watts power. Data and telephony were transmitted with a power of 20 Watts.

The telemetry was carried at 1.0 GHz.


See Molniya History



ATS-6 (Applications Technology Satellite-6)

ATS-6 (Applications Technology Satellite-6)

ATS 6 Satellite

Launched on May 30, 1974, the ATS-6 was the first geostationary satellite to use three-axis stabilisation for attitude control and the first to provide Direct to Home (DTH) television broadcasting, also called Direct Broadcasting Satellite (DBS).


The ATS-6 project benefitted from the use of a much larger launch vehicle, a Titan IIIC, which could carry a much greater payload. The satellite's weight at launch was 2945 lbs (1336 kgs), nearly ten times the weight of the Early Bird, and it was 28 feet (8.51 m) tall and 59 feet (16 metres) wide across the two booms holding its solar arrays. The Titan's guidance sytem enabled the satellite to be inserted directly in the geosynchronous orbit which reduced its on-board fuel requirements to less than 40 kgs.


Three Axis (Body) Stabilisation Benefits

The major advance of ATS-6 was its three axis stabilisation system, the enabling technology which made many new applications possible. Spin-stabilisation had been used in previous satellites but their spinning bodies imposed severe restrictions on the size and shape of the solar arrays and antennas which they could support. The gyro controlled three axis or body stabilisation transformed the satellite into a stable, fixed platform which no longer needed to be spinning enabling many new benefits to be realised.

  • It provided more accurate attitude control.
  • The antenna no longer needed to be de-spun saving energy and complexity.
  • Large high gain directional antennas could now replace the omnidirectional antennas previously necessary with spinning satellites, avoiding the wasteful loss of the transmitter energy into space and focusing it all onto the Earth into defined footprints. At the same time the use of higher gain antennas increased the satellite's effective radiated power.
  • Larger flat solar arrays could also be deployed with every solar cell normal to the Sun's radiation receiving the maximum possible uninterrupted solar energy from the number of cells used.
  • This was a major improvement on the spinning satellite's solar arrays which suffered from three drawbacks:

    • Their capacity is limited by the quantity of solar cells which could be mounted on its curved surface
    • Most of the cells are inclined to the direction of Sun's rays capturing less of the available energy.
    • Because of the satellite's rotation, only 50% of the cells are exposed to the Sun at any one time.

    To take full advantage of this opportunity however, the orientation of the flat solar array panels must be controllable to keep their surfaces normal to the Sun's radiation.

  • With more available solar power came the possibility of higher power transmitters, more equipment and more capabilities.
  • With more transmitter power and a high gain antenna, signals could be received by smaller antennas and less sensitive receivers on the ground.


The ATS-6 design made the best of these opportunities.


Attitude Sensing

The attitude control system consisted of a monitoring system which sensed the satellite's actual attitude and compared it to the desired attitude to provide an error signal which was used in a feedback control system to drive the error to zero. Precision attitude sensing was not only required for station keeping, but also for pointing and slewing.


Pitch and roll attitude sensing was by means of radio frequency interferometry and Polaris star tracking was used to sense the yaw attitude. (See diagram of ATS-6 Attitude Axes). The pitch and roll RF interferometers used separate C band radio uplink transmitters each transmitting continuously to three horn antennas arranged along each of two orthogonal baselines parallel to the satellite pitch and roll axes. One antenna in each trio was used as a measurement reference on each baseline, with the remaining two horns, spaced at 1.66 and 19.95 wavelengths apart to provide coarse and fine phase measurements relative to the respective reference antenna. The measured phase difference associated with each axis was digitised and transmitted back to ground control. The angular resolution was 0.017 degree in the coarse mode and 0.0014 and degree for the vernier mode.


The star tracker is an optical system which senses the satellite's yaw, that is its angular deviation from its desired attitude, by measuring the displacement of the image of a chosen navigation star from its position on a reference star map. As with the pitch and roll measurements, the result is digitised and transmitted to ground control. The yaw attitude could be determined within 0.5 degree.


The Antenna

The 30 foot diameter (9.14 metre) parabolic antenna reflector provided gains from 34 dB to 46 dB in the range UHF to C band depending on the frequency. Coupled with a 80 Watt UHF transmitter transmitting at 860 MHz, it provided the capability for direct TV broadcasting with reception by domestic receivers on the ground using small 10 foot (3 metres) antennas.


Tracking, Pointing and Slewing

The combination of precision sensing and three axis attitude control enabled ground control to perform accurate pointing and slewing of the satellite and with the aid of its high gain antenna ATS-6 became the first satellite capable of tracking sub synchronous S-Band satellites. This was the precursor to the NASA's Tracking and Data Relay Satellites (TDRSS) program. Using its GEO vantage point the ATS-6 could look down on LEO satellites and relay data from a LEO satellite through the GEO satellite and down to the ground. This reduced the need for NASA to maintain ground stations all over the globe to collect data from LEO satellites such as the Hubble Space Telescope (HST) and the International Space station (ISS) as they passed overhead. Similarly propagation studies demonstrated the feasibility of multiple relay links to aircraft.


The Solar Energy

The two solar arrays contained a total of 21,600 solar cells delivering an instantaneous power of 595 Watts at the beginning of life with the energy being stored in 15 AmpHour Nickel Cadmium batteries supplying a 30.5 Volt bus. The satellite did not have the capability to orient the direction of the solar array independently of the antenna so its solar panels were half cylindrical in shape with one array pointing North and the other pointing South to ensure that a sufficient number of solar cells were normal to the Sun as the Sun's apparent direction moves from 23.5 degrees North to 23.5 degrees South between summer and winter solstices. See Solar Energy Reception and ATS-6 Attitude Directions. This arrangement was also necessary to maintain the maximum possible electrical power levels when the satellite had to execute a roll manoeuvre as part of its tracking facility.


Thermal Management

One of the downsides of a static satellite is that it is subject to uneven solar heating with the fixed, Sun facing side possibly reaching very high temperatures while the opposite side, receiving no solar energy, remains very cold. In the vacuum of space this temperature difference can be very high. Spinning satellites do not suffer from this problem.

ATS-6 incorporated heat pipes and phase change materials to equalise the temperature distribution across the satellite body to alleviate this problem.


The Transponders

ATS-6 could receive in any of the VHF, L, S and C-Bands, and transmit using solid state transmitters with outputs of 80 Watts in UHF (860 MHz), 40Watts, in L-band (1650 MHz), 20 Watts in S-band (2 GHz) and, using a TWTA transmitter, 20 Watts in C-Band (4 GHz).

The transponder provided cross connections at the 150 MHz intermediate frequency (IF) so that any receiver could be connected to any transmitter.



ATS-6 carried out 23 different experiments and was the first satellite to provide DBS broadcast television to simple home receivers which it demonstrated by transmitting educational programmes to India, the USA and other countries. It was also the first GEO satellite to demonstrate electric propulsion. Tests included monitoring the space environment and it was used to carry out particle physics experiments and to measure the affect of radiation on the life of solar cells. For other experiments it carried a high resolution scanning bolometer (radiometer). Operating on two channels: infra-red (10.5 to 12.5 µm) and visible light (0.55 to 0.75 µm), it was able to scan the Earth, measuring its infra-red radiation (temperatures) and cloud patterns, techniques which were subsequently used by weather satellites.

ATS-6 was also used to also carry out air traffic control tests and to practice satellite-assisted search and rescue techniques and it played a major role in the Apollo/Soyuz docking in 1975 when it relayed signals to the Houston Control centre.


See ATS-6 History


Intelsat V

Intelsat V

Intelsat 5

Launched in December 1980 Intelsat V was the first commercial Direct Broadcast TV satellite. This was made possible by adopting three axis stabilisation using momentum wheels as pioneered by the ATS-6 satellite. Weighing 4250 lbs (1928 kgs) at launch it was stabilised to within 0.5 degrees and propulsion was by means of hydrazine thrusters.

Because it did not rely on a spinning body for stabilization, Intelsat V could be made in any convenient shape, in this case a box, onto which various appendages housing subsystems could be mounted.

An antenna farm was located on the side of the box facing the Earth with antennas optimised for global, hemispherical, zone and spot footprints with linear and circular polarisation and different frequencies to avoid interference.

Two great fields of solar panels spanning 52.1 feet (15.9 metres), delivering 1800 Watts of power, extended from the adjacent sides of the box and were kept pointing towards the Sun by electric motors as it orbited the Earth and during the Sun's apparent North - South seasonal excusions. Energy was stored in Nickel Cadmium and Nickel Hydrogen batteries.

Communications were provided by 21 C-Band (6GHz uplink-4GHz downlink) and 4-Ku-Band (14 GHz uplink 11 GHz downlink) transponders carrying 12000 voice circuits and 2 TV channels.

As in ATS-6, it used passive thermal management.

The Intelsat V configuration became the template adopted by many subsequent satellite designs.


Intelsat V was designed and manufactured by Ford Aerospace led by Robert E. Berry.


See also Satellite Technology and GPS Satellite Navigation






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