Payload platform concept
The modular structural concept of BIRD allows parallel development activities in the several spacecraft segments. In the case of the payload the segment is designed as a complete subunit oriented on the functional needs of all optical instruments. That means that also two star sensors are integrated in the payload platform architecture.
Consequently it is possible to calibrate the whole system in a complex process without any uncertainties due to instrument mounting procedures. Once integrated and calibrated the payload segment according to figures 10 & 11 was integrated onto the spacecraft bus. Due to the simplified interface (plane/ plane- type) this process is very efficient and provides flexibility of refurbishment of components if needed.
Payload platform technology
Fig. 13: Design principle of the payload platform (cross section)
Early considerations showed the necessity of a thermal control of 1K for the instrument structure if a conventional iso-static aluminium plate would be used. Despite the large mass of such a solution that was not feasible within the micro-satellites resources.The resulting technical solution of the instrument platform is to combine two CF-honeycomb panels on top and below of a 3 mm-thick layer of Carbon Fibre Carbon (CFC). This layer provides heat conduction to the heat pipe interfaces on the front ends of the platform with a coefficient of 155 W m-1 K-1 - comparable with aluminium.
The instrument mounting interfaces of this multi-sandwich structure are inserts with an enlarged plate founded in the CFC-layer (see figure 12). All connections with thermal relevance are glued with silver-filled epoxy. The heat conduction reaches over 50% of that from aluminium. Its CTE is a factor of 15 less.
Another effect is the de-coupling of structural loads coming from the spacecraft bus because of separated mounting planes for spacecraft and scientific instruments.
According to the qualification procedures the pre-defined duty cycles of the complete spacecraft were simulated by the BIRD Structure- Thermal- Model. During the thermal/ vacuum program in-situ measurements of the co-alignment of all scientific instruments were implemented.
Because there is no access to the payload situated in the vacuum chamber an optical principle with the help of alignment prisms were installed which allows observation of all instruments from one point without movement of the payload segment or the measuring equipment. The misalignment caused by instrument and platform deformation in cold and warm operational phases was detected by auto-collimation (see figures 15/ 16).
As a result it could be stated that the typical alignment deviations due to instrument operations are 1 arcmin and 3 arcmin due to the change of the platform temperature of 30 K representing the shift of orbit conditions over the mission.
With the help of the measuring principle described above it was possible to check the optical orientation during all integration and test steps. This Co-alignment control is applied as an up-date of the calibration file. So effects of mounting and settlement caused by dynamical loads can be covered.