Which of the following is not an application of communications satellite technology?

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.

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.

I.D Configuration and Function of Satellites

A communications satellite is an independent system floating in space. It provides its own electric power supply, maintains its attitude, withstands the harsh environment of space, and sees that mission devices operate normally within the required mission life. The design of a satellite consists of conceptual design, preliminary design, and critical design based on a satellite communications systems plan, plus performance-requirements design and the construction of various manufacturing models (BBM, breadboard model; EM, engineering model; PFM, proto flight model; and FM, flight model). The satellite must then pass a thermal-vacuum test that simulates a space environment in a space chamber as well as vibration and other tests before being loaded onto a rocket and launched.

A communications satellite consists of basic equipment supporting a mission (called satellite bus as a whole): the attitude control subsystem, a power supply subsystem, TT&C subsystem, propulsion subsystem, the thermal control subsystem, and structure subsystem. The mission equipment is called a communication subsystem, including an antenna and transponder. Table IV shows the functions of subsystems.

TABLE IV. Function and Typical Equipment of Subsystems of Communication Satellite

Among subsystems, it is the attitude control subsystem that has the biggest influence on both configuration and performance of the communications satellite. It is necessary to maintain satellite attitude in order to always point an antenna of the satellite at the earth to communicate with the earth stations, and to point a solar array in the direction of the sun. A satellite in space overcomes radiation pressure by the sun and the influence of disturbances generated by the interaction torque of geomagnetism with the residual magnetism of the satellite in order to make the attitude stable.

Attitude control systems for maintaining appropriate satellite orientation in space are broadly divided into the spin-stabilization type and the three-axis-stabilization type. An example of the former is the INTELSAT-VI shown in Fig. 5, and an example of the latter is the INTELSAT-VII shown in Fig. 6.

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.

Section 10.9

10.19

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) }

Section 10.9

10.19

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 )}

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).

THE INTELSAT SYSTEM

All INTELSAT satellites have operated in synchronous equatorial orbits'so-called “geostationary” orbits—35,700 km above the equator. This position allows for continuous coverage, with tracking by earth stations necessary only to maintain their radio beams pointed at the satellite through very small angles of drift. The satellites receive on frequencies in the 6-GHz band and transmit on frequencies in the 4-GHz band. All of the satellites so far launched are spin stabilized; that is, they maintain their orientation in space through rotation of the body of the satellite along an axis parallel to the earth's axis. All satellites are powered by arrays of silicon solar cells mounted on the spinning body.

The transponders employ TWTAs generating a few watts of RF output power.

The effectiveness of a communications satellite is dependent upon its capacity, which in turn depends upon radiated power and bandwidth, and also upon its lifetime in orbit. As the technology developed during the past decade, INTELSAT satellites improved in all three areas:

Power. More powerful launch vehicles placed greater weight in orbit, thus providing a greater area for solar cells and more electrical power generated and available for radio transmission.

Bandwidth. With increased power and more sophisticated transponder design, more of the allocated bandwidth was used and eventually reused.

Lifetime. Improved components, devices, and design, and more effective quality assurance techniques, all improved satellite reliability and operating lifetime.

Early Bird, built by the Hughes Aircraft Company, was launched in April 1965. It weighed only 38 kg in orbit and had a total effective radiated power of 10 W in each of two transponders, using only 50 MHz of bandwidth. Its limited power output required that the spinning antenna be “squinted” to cover the heavy traffic route between North America and Europe. Its potential capacity of 240 two-way telephone circuits could allow linking of only two earth stations at a time; i.e., there was no multiple access available, nor was it possible to carry television and telephone simultaneously.

A second series of commercial satellites was developed to support the manned spaceflight operations of NASA. Reliable communications were urgently needed to connect a worldwide network of tracking stations for Project Apollo, some on islands and others on ships at sea. The Hughes-built INTELSAT II satellites were larger than Early Bird, having more power and bandwidth, and thus were able to provide coverage of a wider area of the earth. An important innovation in INTELSAT II was the introduction of multiple-access capability: many pairs of earth stations could be connected through the satellites, each transponder carrying several radio frequency carriers simultaneously. The first INTELSAT II satellite entered service in January 1967.

The larger INTELSAT III series, built by TRW, was introduced in late 1968. By 1969, three of these satellites made possible the realization of a true global communications system. Each INTELSAT III had a nominal capacity of 1,200 telephone circuits. The increase was achieved by using a mechanically despun antenna always pointed toward the earth, providing a so-called “global beam” which covered all of the earth visible from a given position of the synchronous orbit.

The first INTELSAT IV satellite was launched in January 1971. Also built by Hughes, and weighing 720 kg in orbit, these satellites are currently the mainstay of the global network. The major advance of INTELSAT IV over its predecessors is the use of “spot-beam” antennas covering only a small portion of the visible earth, in this case a beam angle of about 4.5°. The resulting concentration of radiated energy provides another increase in capacity. The INTELSAT IV satellites are therefore rated at about 4,000 circuits or greater, depending on the number of transponders connected to spot beams and the multiple-access system in use. The INTELSAT IV electrical power subsystem provides about 470 W generated by some 45,000 solar cells and includes Ni-Cd batteries used during solar eclipse.

INTELSAT IV is the first communications satellite to be bandwidth limited. The communications subsystem is channelized into 12 transponders of 36-MHz bandwidth. All transponders receive from a global-beam antenna. Both global- and spot-beam transmit antennas, providing about 180 and 2,500 W of equivalent radiated power (22.5- and 34-dBW e.i.r.p.), respectively, are available.

Extended capacity satellites, improved versions of the present INTELSAT IV models (to be known as INTELSAT IV-A), have been under construction by Hughes Aircraft Company for three years. These satellites are expected to see service in late 1975 in the Atlantic region to meet that area's higher level of traffic requirements. As already mentioned, these will be the first satellites to incorporate a frequency reuse technique, expanding the number of simultaneous transmissions per frequency bandwidth through spatial separation. The spacecraft body will be similar to that of INTELSAT IV, but will include a modified communications subsystem and advanced antenna array. The antenna patterns will allow simultaneous use of the same portion of the spectrum in two separated areas; e.g., in the Atlantic region one shaped beam will cover Europe and Africa, and the other will cover North and South America. Global beams will also be included. The satellite will have 20 transponders connected in various combinations to global, eastern, or western receive or transmit beams. With approximately the same weight and power as INTELSAT IV and using the same 6/4-GHz frequency band, the “A” version satellite will have 50 percent greater capacity (about 7,000 circuits) through frequency reuse.

Table I shows how the first five generations of INTELSAT satellites have increased in size and weight, communications capacity, and design lifetime with an attendant increase in satellite construction and launching costs. However, the overall cost effectiveness of each successive generation has improved, and during the 10-year period, the investment per circuit year in orbit has decreased from $30,000 to about $1,000 from INTELSAT I to IV (Edelson, 1975).

TABLE I. The Five Generations of Satellites in the INTELSAT System

Along with the spacecraft, the earth segment has grown enormously. The global satellite communications system has grown from 5 earth stations in 1965 to 112 earth station antennas in 60 countries.

Figure 3 and 4 show the INTELSAT system today, and its traffic growth from 1967 to 1975. Figure 5 depicts the distribution of the communications paths in the three ocean regions: Atlantic, Pacific, and Indian.

The INTELSAT system provides continuous global telephone service in addition to television, telegraph, and data transmission. Temporary service is also available on a short-term basis during special world events and times of special communications requirements. All such services meet or exceed international standards of quality (Browne, 1974).

The primary operational mode of the system is multiple access, which is preassigned in the frequency domain. A demand-assignment multiple-access system (known as SPADE) was introduced in 1973 and has been implemented in the Atlantic Ocean region with 21 terminals in operation. Direct digital transmission has been established between a number of points, and more are planned.

The system has provided restoration service for interrupted submarine cables on very short notice, and permitted the rerouting via satellite of a number of such cable circuits which would have remained suspended indefinitely while the repairs were made at sea. During the first half of 1974, temporary service for submarine cable restoration via satellite comprised 24, 926 half-circuit days. At one period in 1972, 34,100 half-circuits were used in this way.

All real-time transoceanic television is transmitted via satellite. TV service is characterized by large fluctuations, caused by singular world events such as international political happenings and sports events.

FUTURE SERVICES

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

A.

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.

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.

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

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

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.

1962—Communications Satellite Act

Also in 1962, Congress passed the Communications Satellite Act108 in which it opened commercial use of space for communications and created the Communications Satellite Corporation (Comsat). Organizationally, Comsat was a unique entity. It was a new, semiprivate corporation, half owned (50%) by the international communications carriers, including AT&T, and the other half (50%) owned by public investors.109 As a semiprivate corporation, Comsat was required to report to the Congress and the president each year.110

Congress delegated responsibility for satellite regulation to the FCC in three ways. First, Congress declared Comsat to be a common carrier and thus included satellite service in Title II, the common carrier portion of the Communications Act of 1934 regulated by the FCC. Second, the FCC's authority over satellites derived from Title III of the 1934 Act, as amended, which gave the FCC broad power to regulate use of the radio-frequency spectrum for communications purposes.111 Third, under its enabling statute, Congress delegated to the FCC responsibility for implementing all communications law, including the Communications Satellite Act. However, to keep a monopoly from developing in the satellite industry, the FCC regulated details in satellite communications such as Earth station ownership. For example, the FCC required that Comsat never own more than 50% of each satellite Earth station, while the other 50% be owned by the remaining international carriers.

INTRODUCTION

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.

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.