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Electrical Power System

Purpose

All subsystems have to be integrated in the mechanical structure, which is the main task of this project.

Design

The main demands from the subsystems are:

  • Large PCB area as possible
  • Mounting of 3 momentum wheels and DC motors
  • Mounting of 3 electromagnetic coilsli
  • Mounting of 6 sun sensors at each side panel
  • Most solar arrays as possible
  • Simple integration as possible
  • Deployable antenna mechanism

Image:SatelliteDEF.jpg

The above picture shows the final design of the satellite. All subsystems are connected on 4 mounts, which is connected to the main load bearing frame. By arranging the subsystems in this way, it is possible to test the subsystems outside the frame. All wiring between subsystems is done by connectors, which insures flexible error determination later on.

Present work

Different kinds of analysis were used to estimate the strength, and to dimension the structure. On the way to orbit the load environmental is high G-forces combined with acoustic and vibration loads and temperature changes. The main load for the design is random and harmonic vibrations which have been modeled via. The Finite Element Method (FEM), and other numerical methods.

Here is the response of a central point of the OBC/PLS PCB with the harmonic load at launch from a FEM analysis:

Image:mechB.jpg

As the graph indicates if the vibrating frequency is nearing 100 Hz, the structure has massive deflection, which only is the case if the satellite gets launched with EUROCOT.

Furthermore a random vibration analysis was performed of every part and the total structure. The picture below shows the von Mises stress response from the random vibrations.

Image:mechC.jpg

The structure can withstand the loads that act on the way to LEO. A transient thermal and stress analysis were preformed to verify that the structure can withstand the thermal cycles over orbits. The analysis reveled that the structure will be at equilibrium after approximately 6 orbits after deployment, and the stress results for the 6.th. Orbit can be seen below.

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The above image show von Mises stress distribution, and as it can be seen the major thermal stresses is at the connecting points between the PCB´s and mounting. The main load carrying frame is almost unloaded. The picture below shows the stress variation over the 6 th. Orbit at the EPS PCB and the mounting.

Image:mechE.jpg

Status

  • All simulations indicates that the structure can withstand the quasi static loads that act at launch, and the endurance loads under operations.
  • The simulated internal satellite temperatures indicate that the battery will lay in the required interval of -10 and +50 Celsius.
  • The production of the components will start the 7 of February.
  • Meanwhile we will design different kinds of deployable mechanisms, which can be used on later cubesats, or this one if the mass budget allows . Below there is a picture of a preliminary design with 28 extra solar cells, when deployed, the total mass of this unit is without wiring 85g.
  • We have applied for the ESA student parabolic flight campaign together with ADCS were we will test different kinds of deployable mechanisms and the change of dynamics due to these.

Image:solar.jpg

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Documents

No documents has been released yet.

For work sheets and more information contact the system engineer for this subsystem

Active Members

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