Structural
Introduction
The Mechanical/Structural subsystem ensures the internal components of the satellite are safely mounted and stowed during launch and for the duration of the mission. It will ensure that all internal components are aligned for safe operation and that they will be adequately protected from environmental hazards.
The satellite structure will conform to the Nanoracks design standard for a 3U CubeSat and will house the CDH, COMS, ADCS, Payload and Power subsystems, including the deployable antenna and two solar arrays.
Design Philosophy
One of the goals of the mechanical subsystem is to improve on the previous issues with CubeSat construction. To accomplish this task, we are attempting to 'modular-ize' the assembly process. The intent is to reduce the complexity of the mechanical model while also offering superior protection and allowing for a parallel workflow during assembly of the final flight model. The intent behind the modular design is to qualify for launch so that we will be able order the next several missions worth of structures all at once. A higher volume of parts would effectively lower the cost of subsequent missions and accelerate the design phase at the same time.
These goals are being realized by breaking the 3U CubeSat structure into six 1/2-U shells. These shells will be built in essentially two stages of machining operations. The first stage of machining creates the bulk shape, shell-to-shell alignment features and a set of five standard mounting holes. These 1/2-U shells will be identical to each other and can be ordered in bulk all at once. These shells are being designed such that they can be placed in a +-Z orientation and connect to another 1/2-U.
A second machining operation will (if necessary) be used to trim excess material and add additional mounting holes for the specific location and function for it's location in the overall satellite structure. This secondary machining operation will add the unique features for it's final location in a satellite, either for the Iris, or for future missions.
One of the top level mission requirements is for the Iris to be led by students (R-MIS-0002). By creating a standard, modular, bulk ordered mechanical shell that future teams can re-use and adapt as needed, not only will the mission level requirement be met, but will also assist future teams' structural development.
Key structural requirements for the development of Iris include withstanding launch loads (vibration, quasi-static acceleration, and static loading), meeting the NanoRacks CubeSat Deployer interface tolerances, and ensuring all deployables have redundant holding and release systems. To verify that our spacecraft will survive launch, the Iris team will be physically testing the flight structure at the University of Manitoba/Magellan Aerospace Advanced Satellite Integration Facility.
The primary science mission of the satellite is to expose payload samples to solar radiation and to observe changes in reflectance of the samples over the mission Life (R-MIS-0025 & R-MIS-003). The structure's functions are derived from this, to ensure that the samples are secured as needed to achieve this goal (R-MEC-0845).
For each "board" that is being developed by the sub-systems we are using the term Printed Wiring Board (PWB) to refer to a bare, un-populated circuit board made of FR-4 or other similar material. Once components have been assembled onto the PWB, the board gets a new designation Circuit Card Assembly to differentiate the parts in communication and to separate the parts more easily in the CAD directory.
The rest of the main Structure functions are shown in the following functional block diagram:
Functional Block Diagram
Structure Design
Iris is a 3U CubeSat, following the Nanoracks standard, and is comprised of four electro-mechanical modules using an update to the UMSATS inter-module interface standard. The module 1/2 shells have been modified to be symmetric in an attempt to lower costs of manufacturing and allow for bulk ordering of structural components for the next generation of satellites. This interface, which breaks down the spacecraft into Power, ADCS, Payload and COMS/CDH modules allows the team to perform design, assembly, integration, and testing concurrently. Additionally, this modular system allows for easier inter-subsystem testing and simulating subsystems through a single common data/power interface.
The breakdown of the sub-system interactions are laid out in the System Block Diagram here.
Each module is being designed so that it can be assembled separately from each other. This approach will allow the four modules to be assembled simultaneously and then connected together in a final operation once each module has been tested independently. The goal is to allow for a faster and more streamlined assembly process by parallelizing the procedure.
In the diagram to the right, arrows indicate which parts are being mechanically fastened to each other. Notably, the CDH board is being fastened via standoffs to the COMS board, which is then fastened to the COMS module.
Mass Budget
A preliminary mass budget for the structures systems is shown here. The upper limit for a 3U satellite is 4.8 kg, and our current estimates with a heavy 24% margin indicate that we are well under the expected mass.
The mass of the structural elements of each module are included in the module subtotals while shared elements are included in the ‘Non-Module Structural’ subtotal. A 30% mass margin was added to new components, a 20% margin was added for components that have been initially selected. 15% was added to components which have been used or built previously by other 3U CubeSats at the University of Manitoba. These margins were averaged for the preliminary mass budget summary.
Separation Switch design
The design of the separation switch includes an electrical switch with a spring loaded pin. The assembly shown in the image to the right includes an off the shelf 4mm pin with a small ring turned out with a washer and a snap ring. Each spring has a maximum of extension of 10mm and is designed for this project to have 1mm of pre-load. An additional 7mm of deflection will be applied when the switch is fully compressed, for a total deflection of 8mm. The combined force output of the four corners is 9N.
To provide the pre-load both the alignment and the pre-load on the separation-switch system a custom part has been design which will bolt to the power module shell. This part will be made of Delrin.
Four of these separation switch designs will be used on the satellite, one in each corner of the power module. Three mechanisms include electrical switches but the fourth spring mechanism will not have an electrical switch as it is only used to balance the forces for deployment.
Separation switch mechanism
Remove Before Flight (RBF) Mechanism
The RBF mechanical system will be housed on the Y face of the satellite in the lowermost portion of the power module. The system consists of a single pin with an alignment part and electrical switch inside the power module.
An RBF pin will be inserted into the power module and held in alignment in two places, on either side of the U-shaped alignment part. When the pin is inserted the side of the pin will depress the electrical switch. The U-Shaped alignment part will ensure that the pin does maintain contact with the switch and that it remains depressed until it is intentionally removed. The forward U bracket hole will contain threading that matches with threading on the pin itself to ensure that the pin remains solidly in place until removal.
This U-Shaped part bolts to the PWB which will have the electrical connector for the RBF switch. The U-shaped part does not press on the switch in any way and only encases the switch to ensure that the location of the pin will always be directly above the switch.
The pin design contains a groove that will align with the switch when it's depressed and threading near the head to screw into the U-Bracket. Together these elements will ensure the pin does not move during flight or under vibration.
Remove Before Flight Assembly
RBF Location
Solar Panel Deployment Mechanism
The chosen design for the Solar Panel deployment is torsion spring hinge mechanism held down by a burn wire mechanism. The panels are held down by a fishing line which is in contact with a resistor capable of heating up to the melting temperature of the fishing line. when ordered by the flight computer, the resistor will heat up and cause the fishing line to lose integrity. the hinges torsion springs will then force the solar panel wings to deploy.
The solar wings themselves both use two torsional spring hinges capable of moving 90 degrees. Also pictured is the wire clamp on the wing itself which holds onto the fishing line.
The overall design of the solar panels is 2 wings on the minus and positive Y faces as well as 3 static solar panel arrays positioned on the positive and minus X faces. The 3 static panels share a design and have a notch cut into the panel to allow access to the RBF pin.
Static Solar panel
Solar Panel Knotting Plan
Each wing has a burn wire and they will use identical knotting strategies to secure the wing stowed and around the central burn resistor column.
Use a double-Davy knot to attach the wire to the hole in the Solar panel
Pull the wire taut from the inside and thread it towards the central burn wire column.
Secure the wire to the holding knob inside the module using a round turn & two half hitch knot while ensuring that the wire is in contact with the resistor.
Ensure the burn wire is pulled taut to secure the wing in place as the knot around the central burn resistor column are tightened.
Double-Davy Knot: https://www.animatedknots.com/double-davy-knot
Round Turn & Two Half Hitch Knot: https://www.animatedknots.com/round-turn-two-half-hitches-knot
Antenna Deployment
The Chosen design for the antenna will use a hinge mechanism held under tension with a melt wire system, similar in design to the solar panels. The antenna when undeployed is housed within a bucket shaped like an involute curve. the hinge door is held down by fishing line which is in contact with a resistor. when ordered, the resistor will heat up to the melting point of the fishing and thus release the antenna.
Resistor and Burn Wire
Note: Fishing line does not actually intersect resistor
The antenna gate contains a hole for tying the fishing in place. the gate contains side walls that align with the bucket to ensure the antenna deploys smoothly.
Antenna Door
The Antenna is housed within a container containing an involute curve to ensure a smooth exit from the containing as the hinge is deployed.
Top View of Antenna Container
The overall antenna deployment mechanism is located on the negative Y face in the communication module. The stowed and deployed states can be seen below.
Stowed configuration
Deployed configuration
Antenna Knotting Plan
The antenna burn wires will use identical knotting strategies for each of the two wires. Links are provided from a knot database website on how to tie each knot mentioned.
For each burn wire, use a double-Davy knot to attach the wire to the hole in the antenna door.
Pull both wires taut from the inside such that wires both meet the burn resistor.
Secure each wire to the holding knob inside the module using a round turn & two half hitch knot.
Ensure wires are pulled taut to secure the door in place as knots around the knob are tightened.
Double-Davy Knot: https://www.animatedknots.com/double-davy-knot
Round Turn & Two Half Hitches: https://www.animatedknots.com/round-turn-two-half-hitches-knot
Antenna Ground Plane
The Chosen design for the antenna ground plane will be a 1mm copper plate that lies below the antenna doorway. This ground plane has dimensions of 49mm by 202mm.
Antenna Ground Plane
Risks Registry
The following table contains the risks identified as being related to structure.
Key Risks
The two highest risks for the structure subsystem are the antenna and solar panel deployment systems
D21 - Antennas Fail to deploy
D22 - Solar Panels Fail to deploy
These systems were tested to ensure their hinge mechanisms and the burn wire holding mechanisms both function correctly. Material creep tests have confirmed that our selected burn wire material will not fail due to mechanical creep.
Both mechanisms will also be included in vibration tests during phase D to ensure that they do not fail under load.
Should deployment still fail on orbit, despite rigorous testing and prototyping, the team shall attempt to operate in a low power state.
Assembly
Structure's role in the AIT plan is largely to ensure proper integration of all the submodules into one working satellite. The satellite design philosophy is to utilize modular half U-shells for each subsystem. These subsystems can then be constructed and tested independently prior to the final combination into a final satellite.
Steps
Each Subsystem assembles individual modules.
Subsystems are connected electronically.
Modules are stacked while lying horizontal.
Rear rails are screwed on.
Superstructure is rotated.
Frontal rails are screwed on.
Vertical Torque rod screwed to power module.
Wing hinges screwed to structure.
Wings screwed to hinges.
Static solar panels screwed onto structure.
Antenna Ground Plane Epoxied onto structure.