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Aviation History
1962
1962 - 0613.PDF
FLIGHT International, 19 April 1962 613 The Delta vehicle's low-drag shroud accommo dates the satellite and its folded booms, antennae and solar paddle arms in this configuration during the launch phase. Dia meter of the satellite body is 23in, and the electron-density and electron-temperature booms are 4ft long when extended. The "stretch yo-yo" de-spin system consists of two steel springs, wound one-half a turn around the base of the cylindrical satellite body and carrying relatively heavy weights PAYLOAD ENVELOPE DE-SPIN TERMINAL BOARD (2) SOLAR PADDLE No. 3 ANTENNA SPRING ANTENNA STAND-OFF PAYLOAD SURFACE TRUING SOLAR PADDLE ARM SENSOR BOOM ELECTRON TEMPERATURE DE-SPIN MECHANISM LYMAN-ALPHA SOLAR PADDLE MASS SPECTROMETER PROBE COSMIC RAY ANALYZER SOLAR PADDLE HINGE DE-SPIN TERMINAL BOARD (2) SENSOR BOOM, RELEASE ELECTRON DENSITY MECHANISMS The satellite uses thermal coatings of evaporated gold with about 25 per cent of the total surface area covered with a combination of black and white paints to achieve the desired absorptivity/emissivity ratio. This is a compromise between the experimenters* requirement for a conducting surface with a preference for gold or rhodium over other metals, and a thermal requirement for the maximum painted areas. Telemetry The radiating system on the satellite is a slightly modified crossed-dipole or turnstile array mounted on the upper part of the spacecraft. The telemetry transmitter operates at a frequency of-136.410Mc/s and will be used for both data transmis sion and as a signal source for tracking. The output power to the antenna system is 250mW with an overall transmitter efficiency of 35 per cent. The transmitter is designed for use with a phase-lock type receiving system in the ground stations. The command receiver is a single-channel type operating on the standard NASA command frequency, which is amplitude modu lated by an assigned sub-carrier tone. The standby power con sumption is approximately 50mW and the sensitivity is lOOdbm. Upon interrogation of the command receiver from a ground station, information which has been stored in the flight tape recorder is transmitted back to the ground station. Some 66 separate para meters will be telemetered from the experiments. Power System The main elements of the power system are: (a) four solar paddles, (b) shunt regulator and battery charge current limiter, (c) battery switching network, (d) batteries, (e) under-voltage system (under-voltage sensing circuit, and timer for shutdown of satellite for 18hr for battery recharging). The solar paddle outputs are combined and connected to a voltage regulator which is in parallel with the solar cells and the remaining power system. The solar cells charge the batteries and supply all operating power to the payload while the spacecraft is in sunlight. Test Programme Qualification testing was conducted at ' joddard Space Flight Center on the prototype spacecraft complete with a simulated experiment, on the individual prototype experiments, and on prototype sub-assemblies. Qualification tests included environmental tests as well as tests that were entirely functional. Environments to which prototype items were subjected included: sinusoidal and random vibration, acceleration, shock, static loading, humidity, temperature, and thermal vacuum. The hermetically- sealed tape recorder was given a leak detection test. The environ ments simulated by this series of tests were those expected to be imposed by storage, handling, transportation, pre-launch, boost, separation, and orbital flight. In mechanical tests, the test levels were 1.5 times the maximum levels expected during transportation, handling and launch. In temperature tests at atmospheric pressure and in thermal vacuum tests, temperatures above the maximum and below the minimum orbital operating temperatures were imposed on operating proto type items. In the thermal vacuum test, these temperature condi tions were combined with space simulations at a vacuum of at least 1 x 10"5mm of Mercury. Prototype items, while non-operative, were exposed during the atmospheric temperature test to the maximum and minimum temperatures to which they could be subjected during storage and transportation. Humidity tests were performed as applicable. Equipment was designed to be operable between - 5°C and + 50°C in orbital environment. "Acceptance testing, conducted on flight items, was limited to the environments of vibration, temperatures, and thermal vacuum. The levels of test environment were no greater than the environ ments expected during launch and orbit. THE BRITISH EXPERIMENTS British scientists have a traditional interest in electrical and other properties of the upper atmosphere, and the university groups in the S-51 project have for mariy years been concerned with labora tory studies in this field. Radio investigations conducted from the ground in past decades have given rise to the need for apparatus
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