The Canadian Space Agency (CSA) has posted a new space technology development program opportunity for solar array substrate technology as an identified priority technology.
The opportunity titled STDP-26 Solar Array, was posted to the CanadaBuys website on February 14, 2025.
The closing date for submissions is March 20, 2025 and proposed start date is March 25, 2025. The contract period is estimated at 12 months. The maximum funding is $125,000, taxes excluded.
Background
The CSA provided the following background information on their needs.
The integration of electronics within solar panel substrates is a technology which the CSA wishes to develop. The concept is to embed electronic assemblies within the honeycomb core of a solar array substrate. An overview is shown in Figure 1.
The panel consists of a ¼” perforated aluminum honeycomb core with upper and lower skins. The upper skin has an insulating layer (such as Kapton or FR4) for insulating the solar cells. Cutouts are placed in the honeycomb core to accommodate electronic modules and their associated wiring which will be provided by the CSA. Wires are routed within the honeycomb core and exit on the upper skin for solar cell connections and at the side of the panel for the output wiring. Hard points for panel mounting are used to mount a fixed panel to the spacecraft structure. Hinge mounting points with tapped mounting holes are used to attach hinges between the deployable panels and the spacecraft structure. For deployable panels, an additional panel tie down hard point is included to firmly hold the panel to the spacecraft structure prior to deployment. An additional hardpoint is used to provide strain relief for the panel output wires.
An internal view of the panel is illustrated in Figure 2.
The internal view shows the electronic modules embedded within the panel. Each electronic module has a wire connection to a solar cell. These wires are mounted into narrow slots machined into the honeycomb. The modules are connected in series through inter module wires and terminate with a panel output wire which exits the side of the panel.
The electronic modules are bonded to one of the skins and cannot be relied upon for structural and thermal panel properties. The electronic modules and wiring will be provided by the CSA. The electronic modules measure 28mm x 28mm x 5.5mm high (TBC). The estimated mass for each electronic module is 7 grams (TBC). Exact measures will be provided prior to the detailed design stage.
A CSA provided deployment hinge used to mount the panel to the spacecraft (not shown) will provide the interface between the panel and the spacecraft. An additional CSA provided hinge will be used to connect a second panel for a double deployable system. The detailed requirements for the hinge will be finalized when the spacecraft bus design has been defined.
This novel packaging technique has several challenges in its implementation. The machining process leaves a honeycomb core with cavities and shallow cuts which reduce the strength of the solar array substrate. The challenge is to design the solar array substrate in such a way as to minimize the structural and thermal impact of the honeycomb cutouts. A further challenge is to develop a laminating process for the skins and honeycomb core which maintains a flat mounting surface for solar cells to be installed despite the cutouts in the honeycomb core.
The CSA intends to demonstrate this technology on an 8U CubeSat mission with two 8U double deployable arrays, two 4U fixed arrays and two 2U fixed arrays.
The CSA is also interested in exploring other solar panel substrate concepts which can house electronic assemblies and their associated wiring.
