STR-Verification 0010-0300

Each analysis subheading on this page is completed in response to a verification activity as listed on Valispace. As verification activities are completed their evidence will be summarized here. Occasionally to reduce the duplication of documentation links to other documents will be listed in these sections as well.

The structural subsystem will survive the minimum three months by achieving an orbit that will not degrade in less than three months, this is demonstrated in V-MEC-0040.

Secondly the structural subsystem will survive the minimum three months if all components are not consumed or degrade in the space environment. All materials in the Bill of Materials are capable of surviving in space for longer than 3 months.

Aluminum materials are class 2 anodized which will extend the life of aluminum components exposed to atomic oxygen. This is confirmed through V-MEC-0580.

All design scheduled for this project is done by students, and no work scheduled exceeds the minimum capabilities of any engineering undergraduate or graduate student.

As built structural mass cannot be evaluated until the structure is built after all TVAC and Vibration tests have been completed on the final launch vehicle

Mass Logging

As the modules are constructed for either Engineering model or Flight model (whichever comes first), each completed module will be measured and tallied here and compared with the mass estimates made in V-MEC-0030. If module masses are outside the range steps will be taken to determine whether the mass will affect the mass and center of mass calculations. The CAD model will have its mass updated during this verification activity so that the CAD contains the most accurate information possible for future analysis.

Mass Logging

The entire structure will be assembled and prepared for the TVAC and Vibration test. The structure will be weighed and the mass will be recorded to demonstrate Nanoracks compliance.

Mass Logging

The mission objectives for the MB-Sat are for the Payload samples to be exposed to space for a 3 month period for the mission to be a success. For that to happen the satellite must therefore stay in space for no less than 60 days. Also there is mission level requirement for the satellite to not remain in space for longer than 25 years. To determine whether or not the bounds on the flight duration will be met a preliminary analysis of the expected flight duration for the MB-Sat has been carried out using STK (Satellite Tool Kit, by AGI).

Since the MB-Sat will be launched from the ISS, flight characteristics of the ISS were used to simulate the expected flight duration. Specifically:

Apogee : 421 km

Perigee: 411 km

Inclination: 51.54 degrees

Argument of Perigee: 82.88

RAAN: 223.1235

Mean Anomaly: 277.3164

Space craft mass: 4.8 kg (upper limit for MB-Sat)

Three different scenarios of the satellites orientation were used to estimate life span of the MB-Sat mission:

All subcomponents have been given a maximum working volume as specified in the Structure ICD page. No subsystem is allowed to work outside of their volumes and their volumes were defined based on the 3U form Factor.

The exceptions to the above stated rule are the deployable arrays which work outside of the internal volume and those have been confirmed in CAD to conform to the Nanoracks deployer mechanism. Machine drawings for the entire assembled satellite will be listed on the Structure ICD page as they are completed.

This cannot be verified until the final satellite is built and confirmed. Currently the design as-planned will conform to this requirement. A Fit-Check model has been created and overall dimensions will be added to the Structure ICD page for verification that the structure is meeting the requirements LINK

During the build phase, The modules were measured with calipers to ensure dimensional tolerance compliance across the rail contact points. see testing details at the following link: Fit Check Testing

Once the entire satellite is built and assembled into a pre-launch state, it will be both measured and tested in a mock-deployer to confirm that the satellite will fit in the final nanoracks deployer.

Once the satellite has passed all vibration and thermal tests, the satellite will be re-tested to the mock-deployer model to confirm that no warping, or deformation has occurred during the testing.

The satellite will be inspected post vibration test to confirm that the satellite deployables are stable and undeployed. The tension in the burn wires should be the same as prior to the vibe testing.

The satellite will be inspected post thermal hold, and prior to shipping to confirm that the satellite still fits to the overall tolerances, the launcher, and that all deployables remain fixed in their position.

Our satellite uses the Nanoracks specification when referring to the 3 axis in which The RBF pin face of the satellite is considered to be +Y and +Z is towards the top of the satellite. the +X axis is then perpendicular to the two prior vectors using the right-hand rule.

3 Axis Orientation

This requirement has been met and is demonstrated in the structure ICD under the Full Satellite Assembly heading LINK

The finished Structural design does not contain any interfacing with the deployer. Structural ICD indicates connections are only present for U-loop and D-SUB connections in the PWR and COMS modules respectively. These connections are for pre-flight testing and are not used to link with the Nanoracks deployer. Link to ICD section.

The overall tolerances that the structure is aiming to meet are shown on the Structure ICD page under the heading "MB-Sat to Nanoracks Launch Interface". A tolerance analysis was completed in the X and Y directions which allow for a maximum machining tolerance of 0.033 mm on all components which has been verified with 3D Hubs for manufactuability.

In the Z direction a looser tolerance of +-0.1mm on the corner rail lengths, however the hole dimensions for mounting the modules to the corner rails is tighter holes for each module require tolerance of +-0.01mm which can be achieved with in house manufacturing facilities.

Assuming all parts are built to spec a total clearance of 0.1mm will be met on all parts. If all parts are built to the maximum or minimum tolerances specified all parts will fit with zero clearance to the Nanoracks specification. Note that according to the Nanoracks documentation is all parts are built to the maximum size within the specification there will still remain an additional 0.5mm of clearance between the satellite and the deployer, meaning that a zero clearance between the as-built satellite and the Nanoracks Launch interface is acceptable.

Tolerances were physically confirmed using Fit Check and caliper testing. Fit Check Testing

Three foot switches have been used in the design of the PWR module. There are four locations designed, however only three will be used to actively inhibit the electrical system. The additional foot switch pin is used as an alternate location if as-built internal spacing doesn't work as planned and also to balance the forces during launch.

The four pins can be seen on the Structure ICD page in the PWR module assembly ICD page.

Switch locations

All components no less than 6.5 mm from the corner leaving an additional 0.5 mm of clearance. Drawings on the ICD will show the rail contact points as a shaded region LINK

A bill of all materials used in the Structure subsection is linked below. All materials selected are confirmed to be stress corrosion resistant.

Bill of Materials

The satellite is physically vibration tested using the soft stow vibration profile provided by Nanoracks. Results are located here: Final Vibration Test

FEA results are given below.

FEA Results

Vibe test results are given below.

Deployment VIbration Test Results

Vibe test results are given below.

Deployment VIbration Test Results

Thermistors are part of the Solar panel deployment mechanism and will be mounted to the backside of the solar arrays. During the vibration test these thermistors will be checked independently to confirm that

A) They remain attached to the solar panel

B) Solder joints survive and continue to pass the solder quality requirement (R-MEC-0990) after the vibration test has been completed

C) Electrical connections to within the structure are maintained (no damage to the thermistors or the wires has been caused by the opening mechanism, the vibration, wire insulation does not become frayed kinked or otherwise damaged as determined through visual inspection.

D1) Before beginning the vibration test the thermistors will be shown to work (they create a predictable result from a known temperature source), and then after the test the thermistors will be checked to confirm that they still produce the same signals to the same test source.

D2) If a computer interface is not available at the time of the vibration test, continuity and resistance of all wires to the thermistor will be used as a secondary means to confirm the connection. This test is a backup test if the first test (D1) cannot be performed due to lack of either a computer or software to read the test signals.

The thermostat is mounted to the battery chassis. A vibrational analysis will be performed and results will be linked here: Thermostat Vibe Report

FEA results are available below.

FEA Results

All fasteners are planned to have a secondary locking feature. Where possible and applicable all-metal lock nuts will be used. In areas where this cannot be done or lock-nuts cannot be found, either Loctite 241 or Loctite 271 will be used in place. These two types of Loctite have been previously approved by NASA. Documentation for this can be found in the Structure documentation folder here.

All components within the structure subsystem are either inert materials, or are mechanisms to work under mechanical deformation such as springs and lock nuts. There are no special or explosive materials used in this system, as shown in the Design Philosophy. Once the final system has been designed this can be further reviewed in the Bill of Materials which will be listed on this page under the heading for V-MEC-120.

The concept of operation for the structure subsystem does not produce any free material at any point. Operations Page

Design does not require any ITAR controlled products. See parts list: Part List

None of the materials or design of the structure subsystem are functions of moisture content. This can be reviewed in the BOM

Given a temperature gradient of 10 degrees celsius the largest deformation expected is 0.0067mm in aluminum parts (corner rails). In structural modules largest deformation is 0.0024 mm. All other mechanical systems deformation is an order of magnitude less than these. There are no thermal impacts to the structural system due to a thermal gradient.

The structure will be tested in the TVAC chamber and held at the Cold plateau temperature. Afterwards there must not be any failed external fasteners and deployment systems should remain undeployed.

The structure will be tested in the TVAC chamber and held at the hot plateau temperature. Afterwards there must not be any failed external fasteners and deployment systems should remain undeployed.

After the thermal cycling test is completed and the satellite is returned to normal temperature and pressure, the externally accessible structure surfaces will be checked to ensure that no cracks or other signs of visible wear are present. A macro-camera will be used to inspect the surfaces (macro-camera is an on hand tool already)

Review of the structure subsystem confirms that the structural walls have 2mm of aluminum over the satellite components except for Payload. Payload is exposed to space to facilitate taking photos and expose the samples to the space environment.

Review of Bill of Materials shows that 6061 aluminum structural components are present which can be susceptible to atomic oxygen degradation. These components were black anodized at Price to mitigate this risk. All other components are not susceptible to this form of degradation.