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Aviation History
1964
1964 - 1761.PDF
976 FUGHT International, 11 June 1964 EXPLODED VIEW OF THE PARTIALLY INSTRUMENTED STV 1 PAM commutator 2 Clock 3 PCM encoder 4 Recorder programmer 5 Recorder 6 Main programmer 7 Command decoder 8 Command receiver 9 I36MC/S telemetry transmitters 10 136Me/s tracking transmitters 11 4,000Mc/s tracking transmitter 12 Antenna multiplexer 13 Batteries under voltage detector and switch 14 Voltage regulator 15 INT-EXT power supply relays l£ Derandomizer for recorder play- back 17 Vibration sensors 18 4,0O0Mc/s tracking antenna ring 19 Chemical batteries 20 Inertial package (rate gyros and accelerometers) 21 I36MC/S turnstile antennas for tele' metry, tracking and command EUROPA 1 ... STV design objectives are: to establish environmental conditions for a typical satellite from launch to injection into orbit; to investi- gate separation techniques and to measure dynamic residue after injection, having in mind future users' requirements; to determine the injection, velocity vector, final guidance and cut-off accuracy; to acquire space-technology techniques; to be a means of training and familiarization with satellite launching techniques. The system has been designed to meet all possible requirements of the initial ELDO programme. This flexibility is very important in STV system design. Measurements of environment and separation parameters will be taken during the pre-orbital flight to ensure better reliability. Redundancy will be introduced in the design of the instrumenta- tion. During the orbital flight, experiments compatible with space- craft instrumentation will be performed in order to acquire infor- mation on space technology. The requirements of ESRO and future users will be considered. An experimental plan has not yet been completely denned; however, experiments will be performed to check the satellite thermal design and behaviour of components. Flexibility of the telemetry system affords great possibilities in orbital experiments. Measurements will be transmitted in real time during the trajectory boosted phase and in orbit. During such events as the jettisoning of fairings and satellite separation, data will be recorded, and played back on command during orbit. The satellite has a polyhedric configuration on an octagonal basis. The faces are trapezoidal or rectangular and allow the installation of solar cells and panels. Total weight will be between 1,100 and 1,3231b. The structure is particularly suitable to the aims foreseen, and it provides a low structure weight and relatively high usable volume. The structure consists of a central pipe with vertical reti- cular panels and horizontal plates. The central pipe stiffly connects elements and allows the axial installation of instrumentation. In- strumentation and batteries will also be installed horizontally and vertically. Separation philosophy has been established in the b'ght of require- ments of future users, especially the attitude control requirements of observatory or communication satellites. Correlating the separa- tion theoretical model, ground simulation and flight tests will per- mit special requirements determined in future missions. The satel- lite is attached by means of eight peripheral bolts plus a central one. The time sequence of bolt release and the position of the device which gives the separation push has been designed to minimize angular residue after separation. Fairings protect the satellite both aerodynamically and thermally during the boosted phase, and are made of electromagnetically transparent material. The structure consists of two cone-cylindrical shields with a spherical cap. The shields will be of reinforced glass- fibre plastic sheets, connected by a honeycomb core. Longitudinal joints will be of glass-fibre, and the spherical nose of stainless steel. A glass-fibre ring, with holes to balance inner and outer pressures, separates fairing and third stage. The shields are connected by two locks; 12 double-squib explosive bolts disconnect from the third stage; then the locks open and the shields separate by four torsion springs. Studies indicate a non-significant influence of aerodynamic forces and heat transfer; only the heating of the conical portion may be critical. CRA facilities cover a Mach range of 1,5 to 12 by means of three h-p wind tunnels and of 15 to 20 by means of an arc-driven "Hot Shot" tunnel. These facilities provide full similarity in Rey- nolds numbers and very high stagnation temperatures. Aim of the thermal control system is to keep the interior tem- perature within 5 ± 40°C. The first scheme studied is passive con- trol of the outer coating and adequate insulation of the instrumen- tation. Another analysis is an active control system with variable insulation. Suitable materials are also being studied. Power supply consists of batteries, solar cells and voltage regu- lators. A series-parallel group of silver-cadmium batteries will pro- vide power during the launch phase. During the orbital phase, silicon solar cells, with "buffer" batteries, will be used. Communications and data-handling has been designed to pro- vide for telemetry, command and tracking. The telemetry is a PCM system with 8-bit words. Data from temperature, pressure, dynamic sensors (accelerometers and rate gyros) etc, at low or high level, may be sub- or super-commutated. After satellite injection it is possible to utilize channels formerly allotted to pre-orbital mea- surements. Transmitters are duplicated and automatically com- muted in case of failure. The frequency (136Mc/s) is compatible with the Minitrack network. Data related to satellite separation and the jettisoning of fairings are recorded on magnetic tape and played back in orbit on 148Mc/s command. To compensate speed variations of the recorder motor a derandomiser circuit is used to restore the bit rate. There are two tracking beacons: one on the 136Mc/s band, and a second on the 4,0O0Mc/s band compatible with communication-satellite trackers. Ground tests are the only means to improved performance of the whole satellite system. The most important are pertinent to dyna- mic and thermo-vacuum environmental simulations. The CRA has facilities to test satellites on a system level. 5: DOWN-RANGE GUIDANCE By A. C. PATERNOTTE DE LA VAILLEE, Counseiller d'Ambassade, Directeur du Service Scientifique, Ministere des Affaires Entrangeres et du Commerce Exterieur, Belgium THE ORBIT CHOSEN for the first European satellite will be polar; in- jection with minimum consumption of fuel demands a strictly pre- calculated trajectory, with external correction to ensure the highest accuracy. This is the function of the guidance station. To achieve this, the station determines the instantaneous position and velocity of the vehicle, and compares these with the pre-calcu- lated values. It computes a new trajectory that will lead the vehicle to the chosen orbit, and issues the guidance commands. When the point of injection is reached, it orders the satellite separation from the third stage.
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