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Aviation History
1962
1962 - 1453.PDF
FLIGHT International, 16 August 1962 245 All types of orbit—circular, elliptical, equatorial, polar and inclined—were considered. It had been suggested that elliptical orbits were preferable on the grounds of increased payload at irreater heights for much of the orbital period: this appeared valid if the launch vehicle could not place a reasonable payload into a sufficiently high circular orbit, but did not apply for the ELDO vehicle. On high elliptical orbits, also, perturbations of the orbit caused by the Sun and Moon would be important, as would the effect of atmospheric drag on satellites having low perigees. These effects must be taken into account if it were considered essential for satellites to keep station relative to one another. Because of the difficulty involved in satellite station-keeping, the RAE scientists considered also a random arrangement of satellites. Using the ELDO vehicle such a system was rejected because of the large number of satellites needed for near-continuous communica tions coverage. "Very much larger vehicles giving multiple satellite launchings or vehicles with a smaller payload capability have re ceived only cursory examination," the report states. "With the Blue Streak launch vehicle, launching one satellite per month would appear reasonable and economic. Thus a random system of 50 satellites does not seem sensible; in fact it appears essential to keep the numbers of satellites required in the system to less than 20 and preferably about ten." Station-keeping for satellites in elliptical orbits was considered a more severe problem than for circular orbits. It might involve fairly frequent orbit adjustments unless the perigee were increased well above 300 n.m., and, if this were done, a small decrease in payload would occur. Station-keeping in polar orbits, or in any orbit inclined at a large angle to the equator, was more complex than in equatorial orbits. Essentially, for equatorial orbits, only the orbital period need be adjusted. In the case of polar orbits it was necessary to adjust both the inclination of individual orbits and the angles between the dif ferent orbital planes, as well as the orbital period. In addition, communications ground stations for satellites in equatorial orbits could be simpler, since every satellite track to be followed would be nearly identical to the previous track. A detailed comparison based on these general considerations led to a preference for a system using circular equatorial orbits, which would be rela tively simple to develop and to operate. The general objective is a worldwide communications system capable of continuous operation. Ideally, communications between two terminals should be possible without the need for intermediate ground stations. In practice, for communications halfway around the Earth, e.g. from the United Kingdom to Australia, this would be impossible even with a stationary 24-hour orbit satellite. In addition, for telephony, the signal delay between terminals should not be greater than 0.25 sec. The traffic requirement is for about 1,000 telephone circuits, together with two television channels, for each satellite repeater in the communications system. About half this capacity would be acceptable if technical limitations precluded meeting the full re quirement initially. With the launch vehicle as envisaged (with an HTP/kerosine third stage), RAE conclude that the full communica tions requirement cannot be met. This vehicle could provide a system with several hundred telephone circuits and one television channel, however, and might be capable of development, e.g. by changing to a liquid hydrogen/liquid oxygen third stage, to meet the full requirement at a later date. A specific requirement of the system is direct communication between Western Europe and the east coast of North America. It is thought that the communications traffic on this route will be the controlling influence on financial returns for many years. Reliability is of paramount importance in developing any worldwide communications system, and the reliability of component parts of the system will be a major factor in the estimation of costs. In particular, failure of a satellite will mean its replacement. For the purposes of the RAE assessment, a satellite lifetime of five years is assumed to be a realistic target. This means that, in any period of five years, half the number of satellites may be expected to fail. The requirement for long life suggests that a minimum of electronic and mechanically moving parts should be used in the design of the satellite. One of the two main sections of the RAE report, supported by detailed appendices on specific aspects, covers the establishment of "ystem design principles. Three major technical factors are em phasized :— (1) Station-keeping satellite systems are essential to avoid the large number of satellites required with random orbits. (2) Circular, equatorial orbits are preferred on the grounds of feasibility of satellite station-keeping and the small numbers of satellites required. Station-keeping in other orbits is not considered technically sound at the present time. (3) Passive or semi-passive methods of stabilization, using natur ally occurring control torques, offer the only means of long life, although active control is probably necessary to recover from injection errors and to carry out orbit adjustment. Eighteen detailed design requirements followed from these three main principles. The requirements were:— (1) Moment of inertia about the pitch axis must be greater than that about the roll axis, which in turn must be greater than that about the yaw axis (2) Effective centre of pressure of solar and other radiation must be near to the centre of mass of the satellite in all its attitudes (3) Centre of mass of the satellite must not move appreciably, for example, as fuel is used up or due to thermal expansion (4) As large an area as possible, facing the Sun for as long as possible, must be provided for solar-cell arrays (5) Layout of the satellite must be such that a minimum of shadowing of the solar cells by other parts of the satellite occurs (6) Arrangement of solar-cell arrays should be such that only one battery-charging cycle occurs during each orbit, to avoid reduction of battery life by excessive cycling of charge and discharge (7) A clear path into space must be provided for the exhausts from all the gas jets (8) An unobstructed view towards the Earth must be provided for the communications (or other) aerial (9) An unobstructed view towards the Earth must be provided for infra-red horizon scanners (10) When installed in rocket, the satellite folded if necessary) must provide a stiff structure capable of withstanding the launching con ditions and designed to impose minimum bending loads on the solar arrays (11) Satellite is required to separate from the rocket and unfold cleanly (12) Arrangements are required for cooling the equipment compart ments before and during all phases of the launch (13) Attitude control jets should be situated as far as possible from the centre of mass to economize on fuel (14) Attitude control jets should be arranged in opposed pairs about each axis to avoid modifying the orbit during active control (15) Orbit correction jets should be situated near to the centre of mass to minimize torques due to thrust misalignment (16) // must be possible to communicate to and from the satellite throughout the launching phase as well as during orbital flight (17) For the particular design considered the all-up weight must not exceed about 4001b and must fit inside a cylinder 4ft 3in internal diameter and not more than 10ft in length (18) Liquid dampers in large diameter rings should be fitted if possible, the large diameter avoiding an excessive weight of liquid. These design principles form the starting point from which the second main section of the report, "A possible system and satellite design with time-scale for development and costs," is developed. In this detailed study the system proposed is based on a main communication repeater chain with the repeaters located in Corn wall, Karachi or Bombay, Sydney, Honolulu or Fanning Island, Vancouver and Halifax, Nova Scotia. Other links between the ground and satellites would of course also be made. The system is outlined in the following extracts from the RAE report. Overall design of the satellite. The design is based on presently known techniques, so that construction and testing is a problem of development and not of long-term research. The satellite is basically a cruciform structure with the equipment compartments located at the ends of the four solar-cell arrays. This general shape gives high moments of inertia for a given mass together with the possibility of large moment arms for gas jet control. The fuel tank for the gas jet system is mounted at the centre of the cross, which is also the centre of mass of the whole satellite. The silicon solar cells are mounted on both front and back surfaces of the solar-cell array structures. The overall span of the satellite is 18ft 3in and the width 3ft 0.5in. The equipment box at the bottom of the satellite contains the bulk of the communications equipment, including the aerial, the
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