Communication Satellites – an overview

S C Pascall BSc PhD CEng MIEE, in Telecommunications Engineer’s Reference Book, 1993

51.4.6 Multiple access methods

Communication satellites are designed to relay several, or more usually many, signals simultaneously. In some cases there may be a separate transponder for each carrier; this is typical of broadcasting satellites and of satellites used for distributing television signals to terrestrial broadcasting stations. More usually, each transponder will relay, not one carrier, but several or many. This is called ‘multiple access’. There are three basic techniques for achieving multiple access without unacceptable interference between the various signals involved.

In frequency division multiple access (FDMA) the carriers that will be relayed by a transponder are assigned carrier frequencies within the transmission band of the transponder, the frequency separation between assigned frequencies being sufficient to avoid overlap of emission spectra. The travelling wave tube (TWT) and solid state power amplifiers which are used in transponders have relatively constant gain characteristics within a certain range of drive levels, but they become non-linear, then saturate, as an upper limit is approached. Therefore the output of the transponder will contain the input carriers, amplified, plus distortion products, such as harmonics of the carriers and the products of intermodulation between them, the level of which will be high if the input carrier aggregate is powerful enough to drive the amplifier close to saturation (Westcott, 1972; Chitre and Fuenzalida, 1972).

For a transponder operating in the FDMA mode, the power level of each up-link carrier reaching the satellite must be set with two objectives. The first is to obtain at the output of the amplifier the optimum ratio between useful carrier power and noise due to the distortion products in the vicinity of the carriers. This involves backing-off the aggregate input level from the point where the amplifier would be driven to maximum total output, in order to obtain a larger reduction in distortion products. The output backoff necessary for TWTs is typically in the range 6dB to 10dB, although the available useful power output can be increased above that level by optimising the assignment of frequencies to carriers and by the use of TWT linearising networks. The second objective is to divide the available output carrier power between the carriers in accordance with their down-link transmission needs.

FDMA may be used for groups of carriers which have been modulated in any way, analogue or digital. Some of the carriers assigned frequencies in an FDMA system may themselves be multiple access systems, using time division multiple access (TDMA). Furthermore, if the C/N ratio in the output of the transponder is not too high, it may be feasible to overlay the FDMA signals with spread spectrum signals, forming, in effect, a code division multiple access (CDMA) system.

A time division multiple access (TDMA) system, operating alone in a transponder, allows the full power to the transponder to be used, that is, no backoff is required. This is because only one carrier is present in the transponder at any instant in time. Each earth station in the system transmits its signals in turn, in bursts, in assigned time slots, typically using PSK modulation, a brief guard time being assigned between each pair of burst slots to ensure that the bursts do not overlap even if small timing errors arise. Figure 51.11 illustrates the frame structure of a high capacity TDMA system.

Figure 51.11. Frame and burst format of the INTELSAT TDMA system. RB1 and RB2 are the reference bursts from reference stations 1 and 2 respectively. The drawing is not to scale

Signals which are to be transmitted over a TDMA system must be digital. Bits within a frame are stored at the transmitting earth station, then assembled into a burst with the necessary preamble bits and transmitted at high speed at the appropriate time. At the receiver the reverse process puts the signal bits into store, then reads them out at the appropriate lower speed, frame by frame. The characteristics of TDMA systems vary over a wide range because the principle can be applied in many different circumstances, ranging from the transmission of low information rate monitoring or control signals with an aggregate bit rate of a few kbit/s, probably transmitted on a frequency assigned within a FDMA system, to the high capacity international telecommunications network TDMA systems operating at 120Mbit/s in the INTELSAT and EUTELSAT systems (INTELSAT, 1972; Eutelsat, 1981; Hills and Evans, 1973).

On board switched TDMA has become feasible in multi-beam satellites like INTELSAT VI, using switch matrices which can operate within the TDMA frame to route one burst to down-link beam A and the next burst to another down-link beam, B.

The functioning of TDMA systems which make efficient use of the time dimension demands precise timing and complex control of access. Such systems may be costly. Where the traffic flowing through the system is light, much simpler systems which use principles first explored within the ALOHA system may provide adequate availability. In these, the transmission path is normally open and an earth station with information to send verifies that no down-link burst from another earth station is in progress; it then transmits its burst. However, several hundreds of milliseconds elapse before the start of a signal from an earth station, sent via a geostationary satellite, can be received at another earth station. Two earth stations may therefore inadvertently transmit overlapping bursts, causing both messages to be mutilated. If this happens, they are both retransmitted automatically.

CDMA systems do not structure their use of transponders either in frequency or time. Earth stations transmit spread spectrum signals which can be identified, after re-transmission by the satellite, by the coding which the signal elements carry.

These various multiple access systems differ in the effectiveness with which they use the facilities provided by a transponder. Figure 51.12 provides a measure of the capacity of a transponder having a bandwidth of 36MHz, using various multiple access and modulation techniques, as a function of the C/N ratio at the earth stations. Methods for calculating transponder performance are given in Hills and Evans, 1973, and in Bargellini, 1972.

Figure 51.12. Telephone channel capacity in 36 MHz channel

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Mark Holker CEng FIEE MBIM, in Telecommunications Engineer’s Reference Book, 1993

27.9 Propagation for satellite communications

Most communications satellites are placed in geo-stationary orbits 36000km above the equator, therefore transmitting and receiving earth stations can fix their antenna positions with only minor adjustments being required for small shifts in satellite position or changes in atmospheric propagation conditions. Such orbits also have the advantage of providing potential coverage of almost one third of the earth’s surface, but the disadvantage of high free space loss compared with lower non stationary orbits. The systems planner will need to calculate the link budget taking into account such factors as the satellite EIRP (equivalent isotropically radiated power) and receiver noise performance. The free space loss can be calculated using Equation 27.7 and distance will have to take into account both the difference in latitude and longitude of the position of the satellite on the earth’s surface to that of the transmitting or receiving station. If the distance and great circle bearing is calculated using Equations 27.10 and 27.11, the elevation and distance of the satellite can also be calculated. In addition to the free space loss, the loss due to atmospheric attenuation must be taken into account, and this will depend upon precipitation conditions in the earth station area. Typical values at 11GHz would be 1.0dB for an “average year” increasing to about 1.5dB for the worst month. The actual figures to be used should be calculated from local meteorological data and the attenuation curves given in Figure 27.12.

Calculation of satellite paths is not often required as operators usually publish “footprint” maps showing the received power contours in dBW, taking into account the path loss and the radiation pattern of the satellite transmitting antenna.

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Howard D. Curtis, in Orbital Mechanics for Engineering Students (Third Edition), 2014

Section 10.9


A communications satellite is in a geostationary equatorial orbit with a period of 24 h. The spin rate ωs about its axis of symmetry is 1 rpm, and the moment of inertia about the spin axis is 550 kg·m2. The moment of inertia about transverse axes through the mass center G is 225 kg·m2. If the spin axis is initially pointed toward the earth, calculate the magnitude and direction of the applied torque MG required to keep the spin axis pointed always toward the earth.

{Ans.: 0.00420 N·m, about the negative x-axis}


The moments of inertia of a satellite about its principal body axes xyz are A = 1000 kg·m2, B = 600 kg·m2, and C = 500 kg·m2, respectively. The moments of inertia of a momentum wheel at the center of mass of the satellite and aligned with the x-axis are Ix = 20 kg·m2 and Iy = Iz = 6 kg·m2. The absolute angular velocity of the satellite with the momentum wheel locked is ω0=0.1i^+0.05j^rad/s. Calculate the angular velocity ωf of the momentum wheel (relative to the satellite) required to reduce the x component of the absolute angular velocity of the satellite to 0.003 rad/s.

{Ans.: 4.95 rad/s}


A solid circular cylindrical satellite of radius 1 m, length 4 m, and mass 250 kg is in a circular earth orbit with a period of 90 min. The cylinder is spinning at 0.001 rad/s (no precession) around its axis, which is aligned with the y-axis of the Clohessy–Wiltshire frame. Calculate the magnitude of the external torque required to maintain this attitude.

{Ans.: −0.00014544i^(N-m)}

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Howard D. Curtis, in Orbital Mechanics for Engineering Students (Second Edition), 2010

Section 10.9


A communications satellite is in a GEO (geostationary equatorial orbit) with a period of 24 hours. The spin rate ωs about its axis of symmetry is 1 revolution per minute, and the moment of inertia about the spin axis is 550 kg · m2. The moment of inertia about transverse axes through the mass center G is 225 kg · m2. If the spin axis is initially pointed towards the earth, calculate the magnitude and direction of the applied torque MG required to keep the spin axis pointed always towards the earth.

{Ans.: 0.00420 N · m, about the negative x-axis}


The moments of inertia of a satellite about its principal body axes xyz are A = 1000 kg · m2, B = 600 kg · m2, and C = 500 kg · m2, respectively. The moments of inertia of a momentum wheel at the center of mass of the satellite and aligned with the x axis are Ix = 20 kg and Iy = Iz = 6 kg · m2. The absolute angular velocity of the satellite with the momentum wheel locked is ω0=0. 1iˆ+0.05jˆ (rad/s). Calculate the angular velocity ωf of the momentum wheel (relative to the satellite) required to reduce the x-component of the absolute angular velocity of the satellite to 0.003 rad/s.

{Ans.: 4.95 rad/s}


A solid circular cylindrical satellite of radius 1 m, length 4 m and mass 250 kg is in a circular earth orbit with period 90 minutes. The cylinder is spinning at 0.001 radians per second (no precession) around its axis, which is aligned with the y axis of the Clohessy-Wiltshire frame. Calculate the magnitude of the external torque required to maintain this attitude.

{Ans.: -0. 00014544iˆ(N-m)}

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Dean Hope, in Space Safety Regulations and Standards, 2010

10.2 PRE-Launch Coordination Activities

Before any communications satellite is launched, the radio frequencies to be used for transmitting and receiving signals must be agreed and coordinated through the ITU. The nature of the traffic to be passed through the satellite must also be fully characterized and assessed so it is compatible with the frequency spectrum made available and the onboard radio frequency (RF) payload designed accordingly.

The process of RF coordination for a satellite-based communications system can take several years due to the detailed negotiations required between the many interested parties and so requires the assistance of a dedicated Frequency Coordination group within the applicant communications organization or company. Ultimately, the aim of such coordination is to ensure that the signals passing to and from the satellite do not cause damaging interference to other communications systems, which could affect the safety of those operations.

Launching a satellite into GEO can also take many years to prepare. Apart from the predominant cost factor, the size and mass of the satellite largely define the range of vehicles from which to choose for launch. A detailed mission analysis is required to determine the most fuel efficient orbit sequence to launch on specific days of the year in order to reach the desired target geostationary orbit. There are many constraints taken into consideration in the mission analysis, ranging from satellite apogee or perigee motor firing attitude limitations due to Sun and Earth sensor fields of view, signal strength limits at the vast apogee distances, visibility of the satellite in transfer orbits as seen from the global tracking network stations, and many other factors. The end result is what is known as a launch window for each possible launch day. The launch window represents that period of time, or times, during which the launched satellite is able to achieve its final target orbit with an acceptable on-station lifetime after executing a series of intermediate transfer orbits during the LEOP phase.

Having developed both a mature mission analysis and a mission events timeline, it is then time to generate an RF interference prediction for those periods when the satellite being launched will pass close to other geostationary satellites whose operating frequencies overlap the telemetry and tele-command (TTC) spectra allocated to the new satellite. The Frequency Coordination group of the new satellite’s operator company contacts all potentially affected operators as a common courtesy—an unwritten convention—in the weeks approaching the launch date and warns each operator of the times between which the satellite will be within ±1° of their satellite’s on-station longitude.

The amount of Earth orbiting debris objects has grown steadily since the early 1960s and so the possibility of a satellite colliding with debris at the time of its launch has also increased. Many of these objects are regularly tracked by the US Space Surveillance Network (SSN) and are cataloged. Their orbit elements can be checked against the orbit elements expected for the satellite after launch and proximity assessments are made. If the estimated object versus satellite minimum separation distance is found to be below a given warning threshold, it is then possible to adjust the launch window opening or closing times or apply an intermediate window cut-out to avoid that particular conjunction.

Recent Inmarsat policy has been to contact USSTRATCOM using their Form-1 process to request a collision avoidance (COLA) analysis for the separated satellite to be performed in the days leading up to launch. Because of the dynamic nature of the forces affecting an orbiting body, a more meaningful COLA analysis can be achieved using the best orbital position and velocity data predictions available for both bodies.

Hence, the closer in time that the observational data for the debris population are to the predicted satellite transfer orbit parameters at launch, the higher the accuracy of the conjunction assessment. Typically, the first COLA request is submitted 48 hours before launch and another is submitted some 6 hours before launch, although the timing of the latter is somewhat dependent on the number of potential collision candidates appearing in the first set of results. Launch providers generally request a COLA analysis from USSTRATCOM at the same time to assess the collision risk for the launch vehicle itself, since the final rocket stage enters a similar transfer orbit to the separated satellite payload with an apogee approaching GEO altitude. Both types of COLA analyses can result in a last minute change to the nominal lift-off time.

In the months prior to launch, it is customary to submit specific satellite and mission details to a launch-licensing authority. This is usually a national quasi-government body charged with the legal responsibility for ensuring that the company whose satellite is being launched is compliant with all the regulations and policies that the Launching State has committed itself to through international treaties and obligations. In the case of Inmarsat, for example, the launch-licensing authority in the UK is the British National Space Centre (BNSC), which was established in 1985 [5]. In the USA, this role is performed by the Federal Communications Commission (FCC).

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Walter E. MorrowJr., in Space and Energy, 1977


To date, most communication satellite service has been between large, fixed terminals. In the future, a growing desire for new forms of service can be anticipated. A list of some of these possible future services is shown in Table 3.

TABLE 3. Possible Future Communication Satellite Services


Single-Channel Circuits to Ships


Single-Channel Circuits to Aircraft


TV Broadcast Service to Small Installations


Single-Channel Voice Circuits to Land-Mobile-Installations


Single-Channel Voice Circuits to Land-Fixed-Installations

Some experience with these classes of service has been obtained as a result of experimental satellites launched by NASA and by the Department of Defense. Generally, these experimental systems have used frequencies between 130 Mc and 1000 Mc because the mobile terminals can use inexpensive omni-directional antennas. Some of these early systems are shown in the following figures.

Lincoln Laboratory Experimental Satellite No. 6 (Fig. 8) has provided service to mobile Department of Defense platforms over the past 6 years at frequencies of about 300 Mc.

Fig. 8. Lincoln Experimental Satellite No. 6 (LES-6)

TACSAT (Fig. 9), another Department of Defense satellite, launched after LES-6, provided a somewhat greater capacity to mobile terminals by means of a more directive 300 Mc antenna.

Fig. 9. TACSAT

MARISAT (Fig. 10), which is planned for launch in the Summer of 1975, will provide services to civil and Navy ships.

Fig. 10. MARISAT

The general characteristics of these satellites are more varied than the INTELSAT series. Table 4 indicates some of the general features. It should be noted that the capacities to the mobile terminals are quite limited compared with that between large, fixed INTELSAT terminals. Of course, considerable sharing between mobile users is possible because of the intermittent nature of the traffic. Even so, the effect of the small antennas used by the mobile terminals is clear.

TABLE 4. Communication Satellites Providing Mobile Service

Launch Date 1968 1969 1975
Weight (lb) 350 1500 720
Frequency (MHz) 300 300 300 1600
Antenna Gain (dB) 10 14 10 13
Transmitter Power (W) 100 240 100 50
Approximate Capacity to Small Terminals in Numbers of Simultaneous Half-Duplex FM Voice Channels 1 5 1 5

If satellites of the TACSAT class have lifetimes and costs similar to INTELSAT IV, the yearly costs per half-duplex voice circuit amount to about $1 million or nearly 1000 times the cost of satellite circuits between large, fixed terminals. The difference is due, of course, to the very small size of the surface terminal antennas. Some communication users, such as the Department of Defense and owners of large ships and transoceanic aircraft, can profitably employ such costly circuits, particularly if they share the circuit among a number of terminals. More general application of satellite communications to inexpensive mobile and fixed terminals awaits a substantial reduction in cost through technical innovation.

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Frank R. Vigneron, R.A. Millar, in Space and Energy, 1977


An upper bound on spacecraft power for the flight-proven versions of current communications satellites, which are typified by Anik or Intelsat IV, is dictated by the choice of the “dual spin” configuration (in which the maximum spacecraft power is governed by the number of cells which can be mounted on the rotating drum of the structure) and the choice of launch vehicle. The bounds appear to be about 500 watts and 1.2 kw for dual spin satellites launched by the Thor Delta and Atlas Centaur vehicles, respectively. 3-axis stabilized spacecraft (i.e., spacecraft which employ a configuration in which a central body is stabilized with respect to the earth by active autonomous means and has deployed from it a lightweight sun-tracking solar array in the arrangement shown in Fig. 1) enable a power level of 1.2 kw to be achieved within the limitations imposed by a Thor Delta launch. 3-axis stabilized spacecraft hold the promise of being more cost effective for many communications missions than dual-spinners, but at present do not have the flight qualification status (in communications missions) associated with the latter type. Of particular concern with regard to 3-axis stabilized spacecraft are factors associated with the attitude control, and related structural flexibility implications of the large lightweight solar arrays.

Fig. 1. CTS Configuration, On Station, Array fully deployed.

Active 3-axis stabilization has been successfully utilized for a number of missions in orbits other than synchronous, for example, Nimbus, OGO, OAO, Ranger, Mariner, and Apollo. The problem of adapting 3-axis stabilization to commercial communications missions lies mainly with the requirement for operation in synchronous orbit with long life (in excess of five years), high reliability, low weight, and high boresight pointing accuracy, to the same degree as required for dual-spin stabilized spacecraft. However, the current consensus is that the development of 3-axis control for communications satellite systems does not necessitate a major technological breakthrough, but rather demonstration and accumulation of a flight record with one or more of many promising stabilization systems appropriate for this type of application.

Phenomena relating to structural flexibility of spacecraft contribute a major operational uncertainty to mission planning. The available design tools for large flexible vehicles are based on analytical modelling, related computer simulation and test data on components. Ground-based confirmation of the design at the systems level is not possible because the configurations are not generally structurally self-supporting in the earth’s one-g environment (system tests based on air bearing which are of value in design of small relatively rigid spacecraft, will play almost no role in design of large flexible spacecraft). Of a number of flexible deployable spacecraft arrays of the type appropriate for communications missions, the FRUSA is the only one upon which flight data has been reported (1); thus for the ‘Deployable Solar Array’ class of appendage, a data base essential to confident spacecraft design is not yet available. Qualification of higher power 3-axis stabilized spacecraft with regard to spacecraft flexibility must necessarily involve future flight demonstration and acquisition of flight-derived structural dynamics data.

The Communications Technology Satellite (CTS) will be launched into synchronous orbit with a Thor Delta 2914 vehicle in early 1976. CTS is a joint program between Canada, Department of Communications (DOC), and the United States, National Aeronautics and Space Administration (NASA). The European Space Agency (ESA) is also participating by supplying certain communications components and by supporting the development of the solar cells and flexible substrates for the deployable solar array. The agreements between DOC and NASA formally identify several communications satellite systems technologies to be developed and flight tested in the course of the program (2) among which are a deployable solar array with an initial power output greater than one kilowatt, and a 3-axis stabilization system to maintain accurate antenna boresight pointing on a spacecraft with flexible appendages. The following paper describes the plan for flight evaluation in attitude stabilization and control, and flexible spacecraft dynamics for CTS. Particular attention is given to the technical objectives, the ground evaluations and test program, various mission events planned, the baseline of anticipated results, the complement of spacecraft instrumentation, and the data handling system.

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Mamoru Ishii, in Extreme Events in Geospace, 2018

3 Action to Aviation

The International Civil Aviation Organization (ICAO) is interested in space weather information, which is important in three categories: HF communications, satellite positioning, and human exposure. Some descriptions about HF communications have already been shown in the previous section, so here the remaining two will be introduced.

The Global Navigation Satellite System (GNSS) is now important as a part of the aviation system. ICAO decided the international standards of the Satellite-Based Augmentation System (SBAS) and the Ground-Based Augmentation System (GBAS). The Japanese government launched the Multifunctional Transport SATellite (MTSAT) in 2005 and 2006 for preparing MSAS in the west Pacific area.

The significant ionospheric anomaly from the assumed shape in the model of GNSS is one of the largest factors in GNSS error. It is necessary to study the quantitative influence of ionospheric disturbances on MSAS for using the system in operational aviation.

The EPB is a type of ionospheric disturbance. EPB generates after sunset symmetrically with respect to the magnetic equator, with the size of several thousand km in North-South and several tens of km in East-West. Once generated, the EPB moves westward at about 30 km/h by the Earth’s electric field [Fukao et al., 2006]. The southern area of Japan can be affected by EPB when its activity is high.

Recently, a three-dimensional high-resolution EPB model was developed that enables us to study the generation process and key parameters for EPB growth. Yokoyama et al. (2014) show the results of EPB growth with the high-resolution bubble models that have spatial resolution as fine as 1 km. Fig. 4 is an example of the growth of EPB.

Fig. 4. An example of EPB growth calculated from the high-resolution bubble models.

From Yokoyama, T., Shinagawa, H., Jin, H., 2014. Nonlinear growth, bifurcation, and pinching of equatorial plasma bubble simulated by three-dimensional high-resolution bubble model. J. Geophys. Res. doi: 10.1002/2014JA020708.

The Electric Navigation Research Institute is responsible for research and development in the field of electronic navigation in Japan; many studies about the influence of ionospheric disturbances on MSAS are conducted. Saito et al. (2015) tested GAST-D (GBAS Approach Service Type D) at the real airport environment at New Ishigaki Airport. Their results showed that the ionospheric spatial gradient monitor (ISGM) worked satisfactory under nominal conditions, but some events of enhanced ISGM outputs, which may potentially cause an ISGM false alarm, were observed.

Saito et al. (2017) built an ionospheric delay scintillation model for GBAS specifically for the Asia-Pacific region to collect the observed data in various areas in that region.

When solar energetic particles (SEPs) are incident to the atmosphere, they can induce air showers by generating varieties of secondary particles. Such secondary particles can reach deep into the atmosphere and enhance the level of radiation doses, which can be a hazard to aircrews.

To estimate the aviation exposure from SEP, several institutes in Japan and the United States developed WASAVIES, a Warning System for AVIation Exposure to Solar energetic particles (Sato et al., 2013; Kubo et al., 2015). WASAVIES has been tested and verified by making a comparison between the measured and calculated count rates of several neutron monitors during past ground level enhancement (GLE) events. The final goal of this project is to predict the enhancement of radiation doses due to SEP exposure within 6 h from the GLE onset.

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Geoffrey Hyde, Pier L. Bargellini, in Reference Data for Engineers (Ninth Edition), 2002


Since the necessary rockets and know-how related to the complex in-orbit injection maneuvers were developed, the majority of commercial communications satellites were of the geostationary type until the 1990s. However, because the elevation angle at an earth station drops at higher latitudes, geostationary satellites cannot serve the near-polar regions. For this purpose, other orbits have been used. Polar orbits are also used for non-communications-type missions such as earth observations, weather, surveillance, etc.

From Eq. 19, the maximum latitude that can be served is

(Eq. 22)β=cos−1[(Rcosθ)/(R+h)]−θ

Hence, for a minimum elevation angle θ of 5°, β = 76°21′. Thus, inclined orbits are required to serve those regions further north. On the basis of energy considerations, it can be shown that for a given rocket, the lower the perigee the higher the apogee. The slower satellite motion around apogee results in the dual benefit of making communications possible over greater distances on earth and for longer periods of mutual visibility; the tracking problems are also eased.

As previously mentioned, an orbit inclination of 63.5° is advantageous because of the zero rotation of the line of the apsides for this critical angle. The inclined orbit discussed above, 63.5° inclination, 12-hour period, turns out to be highly advantageous. A satellite in such an orbit spends about 8 hours near geosynchronous altitude. Thus three such satellites can provide continuous service at altitude. Further, properly adjusted, the satellites return to the same station every 24 hours. Finally, because of the inclination, at altitude the satellites can be seen in polar regions. The former USSR (and the current Russia) had extensive polar regions and used the MOLNIYA series of satellites in the above-mentioned inclined orbit to provide point-to-point telephony (fixed satellite services), data relay, mobile communications, TV distribution, and DBS services very successfully. Variants of this orbit will be used in the United States (for DARS) and in Japan (for multimedia).

Advances in technology permitting high levels of sophistication on board smaller satellites have led to a variety of proposals for constellations of satellites in low earth orbit (LEO) and in medium earth orbit (MEO) to provide mobile or data relay, remote sensing, radio location services, or combinations thereof. Some, such as data relay satellites (TDRSS), radiolocation (LOCSTAR), mobile communications (Iridium, Globalstar, and others) and messaging/paging (ORB-COMM) are in service (in all or in part) and others, such as ICO, are actively under construction at the date of publication of this book. The basic reason that these systems have emerged in the 1990s is that the cost of the technology required and the reliability of the components (largely digital), combined with their reduced weight and power requirements (of VLSICS), have made such systems practical to build, launch, and maintain in orbit.

Because of the van Allen belts with their much higher radiation levels, which significantly and deleteriously affect the satellite electronics, the choice of orbits is limited. LEOs lie below the inner van Allen belt; MEOs lie between them and GEOs above them.

The LEO systems have certain advantages. They have much shorter earth-satellite paths, which leads to lower eirp requirements and much shorter delays. Further, especially of advantage in mobile/urban communications, elevation angles to the useful part of the LEO are high, thus making for less “multipath.” However, their lower orbits require sizable constellations of satellites for continuous service (e.g., 66 for Iridium).

MEO systems are, as might be expected, intermediate in their delay and their eirp requirements, and their constellations are smaller. The trade-offs made in system design determine the altitude and inclination of the orbits and the rf frequency/elevation angle/availability considerations combine to define the constellations. (GEO systems are well known and amply described herein.)

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Garry McCracken, Peter Stott, in Fusion, 2005

The initial observations of microwave radiation, by Penzias and Wilson at Bell Telephone Laboratories, were made quite by accident. They were testing an antenna designed for communications satellites, and in order to check the zero level of their instrument they had pointed it at a region of the sky where they expected no radio sources. To their surprise they obtained a radiation signal that they could not explain. A few months later Jim Peebles at Princeton University heard of their results. He had been doing calculations based on the Big Bang theory, which predicted that the universe should be filled with a sea of radiation with a temperature less than 10 K. When they compared results, the observed radiation was in good agreement with predictions. As the measurements have improved, including more sophisticated instrumentation on satellites, it was found that not only was the intensity correctly predicted, but the measured CMBR also has precisely the profile of intensity versus frequency to be consistent with the Big Bang model.

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