Structural Analysis & Testing

Introduction

Various tests and analysis will be performed on the structure and its constituent parts throughout the projects lifetime. A summary of the tests, their status, and their results once completed are found on this page. This page is still under active development as part of finishing the Phase D schedule and deliverables.

Phase C Tests and Analysis

Creep Test

Purpose

The Solar-Panel release mechanism is based on a melt-wire mechanism. A tab is expected to support a certain pretension in order to constrain the solar-panels over the pre-launch and launch period where the satellite will be inaccessible by the Iris team. This test serves the purpose to:

• Determine the amount of initial pre-tension needed on the melt-wire system to keep the Solar-Panels constrained.

• Determine the degree of creep that is expected (how much the pretension will relax) in the melt-wire.

Materials

Test Specimen:

• Delrin

• Teflon

• Fishing Line

Additional Material:

• Dummy Load Tray

• Mount Hooks

• Washers

• Scale/Force Guage

• Calipers

• Location to leave alone for 2 week period

Test Setup

The test setup will consist of a fixed bar, a sample melt-tab (or wire), and a loading tray. The sample melt-tabs will be connected to the bar using hooks, if the sample material is a wire it will simply be tied to the bar. Hanging below the bar the sample materials will have a loading tray fixed to their lower ends. In the loading tray a prescribed mass will be loaded to simulate a fixed tension load. A sketch of the test setup is shown in the following figure:

Test Procedure

For each sample melt-tab or wire:

1. Melt-tab (wire) will be cut to a prescribed length defined by the loading conditions

2a. If sample is a wire the melt-wire will be tied off at both ends with a loop

2b. If sample is a tab a small hole will be drilled in each end of the melt-tab to attach a hook

3. A dummy load tray will be connected to the free end of the melt-tab (wire) via the hook or a washer respectively

4. As-built measurements of the length will be taken to account for variability and be measured from hook-to-hook (knot-to-knot)

5. The dummy load tray will be removed and have a prescribed dummy mass added to it (nuts, bolts, scrap metal, etc.)

6. Record both the dummy load actual mass and the as-built melt-tab (wire) actual hook-to-hook (knot-to-knot) length in the Melt-Wire Creep Test Measurement spreadsheet

7. Record the time and place the dummy load on the melt-tab (wire)

8. Immediately measure the hook-to-hook (knot-to-knot) length after starting the time

9. Leave the system until the time variable is met

10. At the pre-set time intervals record both the actual time measured and the hook-to-hook or (knot-to-knot) length

11. After the time limit is over, measure the hook-to-hook (knot-to-knot) length and remove the load

12. Then take a final measurement of the unloaded hook-to-hook (knot-to-knot) length [Note: not used in this analysis but may be of use for future applications]

This test procedure will be completed three times per material and load combination. I.e. material A load X will be done three times, then material A load 2X will be completed three times as well, to determine the variance in each experiment. Multiple experiments can be run simultaneously, space permitting.

For each material, the minimum load determined to hold the solar panels closed will be used (0.3019 [N]). The maximum load being 2x the minimum load. A third load case for each material will be used which is 90% of the maximum rated load. This upper limit will be used to determine the worst-case loading and creep values for each material and will be used to place an upper limit on the pretension. If 90% of the rated load is less than the maximum calculated load, then only two loads will be required for that specific material

An accompanying measurement log has been premade with expected calculations populated, however additional sheets, and additional time may be required depending on the ongoing results of the experiment. The spreadsheet has been populated assuming that the creep will level off after two weeks, however, this timeline may need to be extended if no steady state is found.

Results

In all 0.3 Newton load cases the creep measured was below 2mm. The Teflon and Delrin tabs experienced almost no creep in all loading cases after 2 weeks. The Berkley fishing line performed the worst, experiencing the most creep out of the test samples, reaching between 8 to 17 mm of creep on the heaviest load case. the Spiderwire fishing line experienced about 2mm of creep.

From this test one can select the Spiderwire fishing or one of the two tab materials as optimal choices dependent on the results of the burn test.

Melt Wire Test

Purpose

The purpose of this test is to determine which materials selected from the creep test are the best choices to use as a melt wire. The materials will be subject to heating by a resistor and the time it takes to melt through the material will be logged. The material that breaks in the shortest time will be the optimal choice.

Test System

The test system will include a Resistor rated to heat up past the melting point of the materials and the melt material laying in tension flat against the resistor.

Test Procedure

For each combination of resistor & melt wires:

1. Cut the melt wire to any length.

2. Place the resistor perpendicular to the melt wire and ensure taut.

3. Heat up the resistor using the power supply at the prescribed current of 2.0 Amps.

4. Record how long it takes the resistor to melt through the melt wire.

The combination with the best combination of these two parameters should be chosen for the final design.

Results

The results of the burn test were as follows:

Spiderwire melted within 2 seconds

Teflon Tab melted within 15 seconds

Delrin Tab Melted within 10 seconds

From these results the Spiderwire shall be chosen as the melt wire material in the final design.

Burn Resistor Test

Purpose

The purpose of this test is to determine resistor characteristics for the deployment mechanism. Resistors of different resistance and power rating were tested against both delrin tabs and spiderwire.

Test System

The test system will include a resistor rated to heat up past the melting point of the materials and the melt material laying in tension flat against the resistor. The resistors with the best time will be chosen.

Test Procedure

For each combination of resistor & melt wires:

1. Cut the melt wire to any length.

2. Place the resistor perpendicular to the melt wire and ensure taut.

3. Heat up the resistor using the power supply at the prescribed voltage of 6.4 Volts.

4. Record how long it takes the resistor to melt through the melt wire or tab.

The resistor with the best time to melt should be chosen.

Results

From these results, the CMF558R0000FKEK and RN60C10R0FB14 are both considered valid resistor choices.

FEA Analysis

Static and Vibrational FEA analysis was conducted on the main body of the satellite to ensure that the structure would capable of handling launch conditions. the force conditions are defined by the Nanoracks NR-NRCSD-S0003 Standard. The FEA was done through NX software and the detailed report can be found in the Payload analysis page linked Here.

The results of the FEA analysis showed a maximum vibrational Von-Mises stress of about 2MPa and a maximum Static Von-Mises stress of 0.400 MPa. The maximum Nodal deflection experienced by the satellite mesh was 2 mm. The maximum stresses experienced by the satellite body is well under the yield strength of aluminum of about 275 MPa.

Stress and deflection propagation is given below in the following animated pictures.

Radiation Analysis

The planned wall thickness of the MB-Sat uses 2 mm wall thickness on average, and in some areas will be thicker. The results from the SPENVIS analysis indicate that we should expect a radiation dose of approximately 2*10^3 to 4*10^3 (rad) over a 3 year flight duration.

Phase D Tests and Analysis

Antenna Deployment Test

Purpose

The purpose of this test is to ensure that the Final Antenna Deployment design will function as intended per V-MEC-950 when built to final launch specifications. Release and hold testing are both required to fully confirm the design.

Tools and Materials Required

• The Satellite

• Force gauge

• Extra Spiderwire Fishing line

• Power Supply

• Extra Antenna Coils

Test 1 Release:

When controlled by a suitable stand-in for the power module the antenna system will deploy perpendicular to the side of the satellite. The test will be considered a pass if the antenna is deployed to +- 10 degrees to the satellite side surface in the XY plane about the Z-axis. Electrical contact to the antenna must be maintained. This test will be performed five times and if the antenna performs ten successive tests without repairs to any component (aside from resetting the consumable release mechanism) the design will be considered acceptable for flight.

Test Setup

1. The COMS Module must be partially assembled. The CDH board may be excluded for the purposes of this test.

2. The Burn resistor must be connected to the external power supply.

Test Procedure

For each of the five trials:

1. Tie and Knot two Spiderwire fishing lines to the antenna door.

2. Coil the antenna.

3. Carefully Close the antenna door while threading the fishing line into the COMS module.

4. Run each fishing line in contact with the burn resistor knob.

5. Tie and Knot the two fishing lines onto the fishing line knob.

6. Turn on the power supply

7. Turn off the power supply once the antenna deploys.

8. Repeat the process for the next trial.

Test 2 Holding:

The antenna will be coiled, stowed and the burn-wire mechanism will be manually reset in the launch state. Once in the stowed position, the antenna 'door' will have a force gauge attached to the furthest point from the axis of rotation. The force gauge will then be used to pull against the door at a slow steady rate until the door opens. If the force required to break the holding mechanism and force the door open is more than the maximum force expected during launch (2N) then the test will be considered a pass. 5 separate tests are required with new, untested consumable burn wires for each test. If less than 5 tests pass, the mechanism/melt-wire will be reevaluated and the count will be reset to zero.

Test Setup

1. The COMS Module must be partially assembled. The CDH board may be excluded for the purposes of this test.

Test Procedure

For each of the five trials:

1. Tie and Knot two Spiderwire fishing lines to the antenna door.

2. Coil the antenna.

3. Carefully Close the antenna door while threading the fishing line into the COMS module.

4. Run each fishing line in contact with the burn resistor knob.

5. Tie and Knot the two fishing lines onto the fishing line knob.

6. Pull the Antenna door open from the tip of the door farthest from the hinge.

7. log the force required to pry open the door

8. Repeat the process for the next trial.

Solar Panel Deployment Test

Purpose

The purpose of this test is to ensure that the Final Solar Deployment design will function as intended per V-MEC-960 when built to final launch specifications. Release and hold testing are both required to fully confirm the design.

Tools and Materials Required

• The Satellite

• Force gauge

• Extra Spiderwire Fishing line

• Power Supply

Test 1 Release:

When controlled by a suitable stand-in for the power module the solar panel system will deploy perpendicular to the side of the satellite. The test will be considered a pass if the antenna is deployed to +- 10 degrees to the satellite side surface in the XY plane about the Z-axis. Electrical and data contact to the solar panels must be tested and maintained. This test will be performed 5 times and if the model performs 5 successive tests without repairs to any component (aside from resetting the consumable release mechanism) the design will be considered acceptable for flight.

Test Setup

1. The Structure must be partially assembled. Internals may be excluded except the Solar deployment mechanism.

2. The Burn resistor must be connected to the external power supply.

Test Procedure

For each of the five trials:

1. Tie and Knot two Spiderwire fishing lines to one of the wings.

2. Carefully Close the wing while threading the fishing lines into the PLD module.

3. Run each fishing line in contact with the burn resistor

4. Run the fishing lines to the other wing.

5. Tie and knot the fishing lines onto the other closed wing.

6. Turn on the power supply

7. Turn off the power supply once the wings deploy.

8. Repeat the process for the next trial.

Test 2 Holding:

The solar panels will be stowed and the burn-wire mechanism will be manually reset in the launch state. Once in the stowed position, the solar panels will have a force gauge attached to the furthest point from the axis of rotation. The force gauge will then be used to pull against the door at a slow steady rate until the door opens. If the force required to break the holding mechanism and force the door open is more than the maximum force (10N) expected during launch then the test will be considered a pass. 5 separate tests are required with new, untested consumable burn wires for each test. If less than 5 tests pass, the mechanism/melt-wire will be reevaluated and the count will be reset to zero.

Test Setup

1. The Structure must be partially assembled. Internals may be excluded except the Solar deployment mechanism.

Test Procedure

For each of the five trials:

1. Tie and Knot two Spiderwire fishing lines to one of the wings.

2. Carefully Close the wing while threading the fishing lines into the PLD module.

3. Run each fishing line in contact with the burn resistor

4. Run the fishing lines to the other wing.

5. Tie and knot the fishing lines onto the other closed wing.

6. Pull the wing open from the tip farthest from the hinge.

7. log the force required to pry open the wing

8. Repeat the process for the next trial.

Antenna and Solar Panels are able to successfully deploy after undergoing soft stow vibration.

Structure Sharpness Test

Purpose

To ensure the structure is safe to handle, there must be no sharp edges present on the structure.

Test Procedure

Wearing a tight-fitting set of rubber/latex/nitrile gloves, run a finger along all exposed edges, corners, and serviceable areas to confirm that no sharp edges exist or can harm any user that will be interacting with the satellite outside of the Iris team.

If a sharp edge is located, sand it down to remove the sharp edge.

Fit Check

Purpose

To ensure compliance with Nanoracks requirements for fitting into the deployer. The satellite structure's maximum dimension in the x and y axis must be 100 +-0.1 mm.

Test Procedure

The satellite superstructure must be assembled. The superstructure includes the shells, rails, and solar panels. Final fit check will instead use the fully assembled satellite.

The following marked locations are measured using a caliper to ensure dimensional compliance (100 +- 0.1mm ) across the front, rear, left, and right faces.

The assembled superstructure is placed in the fit check jig for further confirmation of dimensional compliance should the caliper test pass.

Should the fit check fail, the location where the dimension is out of specifications must be identified and fixed.

Following the fix to the structure, the fit check must be repeated to ensure dimensional compliance.

Results

Fit Check 1:

Preliminary Fit check finished using calipers prior to EM vibration test to ensure no portion of satellite superstructure is past the maximum 100.1 mm dimension in any axis.

It was determined that the satellite is slightly larger than 100.1 mm on the front face of the satellite in the region of the payload hole (see PLD1 Front and PLD2 Front in results Table).

The cause is likely due to the payload upper shell having experienced deformation during secondary manufacturing when the major hole was cut out of the wall.

The payload shell will be shaved down to fix the bulge and a second fit check will be conducted.

Fit Check 2:

Fit check 2 was finished using calipers prior to deployment vibration test to ensure no portion of satellite superstructure is past the maximum 100.1 mm dimension in any axis.

It was determined that the satellite is within dimensional tolerance specifications. The results are listed in the following table in millimeters.

Mass Inspection

Purpose

the Cubesat has a requirement that it weighs less than 4.8kg. Masses of each module will be logged whenever assembled throughout phase D testing to ensure compliance. logs will be stored at the following link: Mass Logging

Additionally, the Full satellite will be weighed and logged after final assembly to further ensure compliance with weight requirements.

Materials Needed

• The Satellite or Satellite Module to be weighed

• Scale

Test Procedure

Once a submodule is fully assembled during testing throughout phase D and prior to connecting it to other assemblies during final assembly, the submodule will be weighed to ensure it is under said submodules maximum weight limit.

After the assembly of the full satellite, weigh the fully assembled CubeSat to ensure it is less than 4.8 kg.

Log the data obtained in the following excel sheet

Mass Logging Results