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In recent years, the trend in spacecraft design has been towards smaller and lighter satellites with sophisticated instrumentation. The satellites use smaller launch vehicles, reducing operational cost. Composite materials reinforced with highconductivity carbon fibres will contribute decisively to these design goals.
(Published on June-July 2005 – JEC Magazine #18)
BY JESÚS MARCOS, HEAD OF SPACE DIVISION, INASMET-TECNALIA
In recent years, composite materials have been used increasingly to reduce the structural weight and volume of satellite components. However, the structure typically represents as little as 10 to 15% of the total mass of the spacecraft, while the instrumentation and equipment units account for a considerable part. In conventional spacecraft construction, the structural, thermal and electronic functions are generally designed and manufactured into separate elements:
- load support unit consisting of shell and frame structures; - radiation shielding unit to protect against the high-energy space particles that penetrate spacecraft materials, damaging and degrading the functionality of the electronics, and shortening the mission’s lifespan; - thermal management unit to control the electronics-generated heat through heat discharge, heat compensation (additional heaters are needed to ensure minimum operating temperatures for the electronics in cold cases) and heat distribution; - electronic enclosure units consisting of aluminium black boxes, to protect the unit from electromagnetic interference (EMI); - electronic power distribution unit consisting of cable bundles / wiring harness.
This design results in significant mass penalties. For smaller satellites, a different subsystem functionality is used not only to reduce the overall satellite size and weight, but also to be able to accommodate all subsystems in the very limited volume inside the satellites. Multifunctional technology aims to integrate all these functions to reduce satellite weight and size down to a minimum.;
The solution under consideration consists of using a multifunctional structure (MFS) design concept to design the structural elements of a spacecraft, placing most of the spacecraft’s electronic components on its composite-panel walls, which also include structural, radiation-shielding and thermal-control elements. Flexible printed circuits are laminated integrally into the structural composite skins. This approach, coupled with higher-density electronics packaging technology, will substantially reduce the overall weight and volume of future nanosatellites. The challenge is to accomplish this integration of multiple functions into the spacecraft structure at an affordable cost.
The main objective of this research project is to manufacture a breadboard panel that integrates all the aforementioned functions (see figure). The panel is a structural honeycomb sandwich construction 500mm long, 280mm wide and 27mm thick. It is equipped with a flex-circuit made of multi-layer copper with Kapton laminate patches. This laminate is bonded to the inner surface of the panel, which also incorporates a passive thermal control system based on high thermal conductivity carbon fibres. The outer surface of the panel works as a radiator.
The benefit of the MFS design compared to traditional design is that it: - eliminates chassis and cabling, reducing electronic enclosures and harness; - maximizes functional elements/volume ratio for maximum integration; - reduces weight and volume (>25% increase of mass fraction, > 50% increase on S/C volume availability); - shortens thermal paths from the electronic components to the spacecraft radiators, enhancing the heat radiation capability needed for the growing dissipation densities; and - enhances robustness and reliability.
To manufacture the sandwich skins, M40J PAN carbon fibres were selected for their good structural behaviour and relatively reduced cost, and high thermal conductivity K1100 pitch carbon fibres, to improve the composite’s capacity to distribute the heat away from the dissipative components and thereby decrease the system’s resulting thermal gradients. The skins are cured in autoclave. To avoid compatibility problems, the same resin, YLA’s cyanate ester RS-3, is used with both reinforcement fibres. RS-3 is specifically developed to provide good behaviour under the space environment. This includes low outgassing, low shrinkage, low moisture absorption, low microcracking, and low modulus after radiation exposure. From both the thermal and mechanical point of view, the lay-up finally selected for the sandwich skins is as follows: [±45a / 0b / 90b/]S where a = M40J/RS-3, 0.14mm thick and b = K1100/RS-3, 0.11mm thick. Thus, each skin is close to 1mm thick.
The prototype is manufactured in three different steps: 1) and 2): the two composite skins are manufactured individually; 3): the skins are bonded to the honeycomb with an adhesive layer (Hexcel Redux 319).
The thermal control of the MFSP is based mainly on passive means. The heat generated in the components is distributed over the internal skin and transmitted through the honeycomb to the external surface, which acts as a radiator to evacuate the heat to the deep space. The remaining area of the external skin not used as radiator is covered by MLI blankets. Heaters are used to maintain the temperature of the components above their corresponding non-operational temperature limits in non-active modes. Thermal performance was analyzed with FEM software. The results obtained in the hot case (nominal operating conditions) are shown below:
Thermal testing showed good correlation between analytical and experimental results. The health tests performed ensure good thermal performance and no degradation of electrical performance.