Payload
Phase C Deliverables
Introduction
Iris contains three payloads: The Geological Sample Exposure Mission, the Gnomon Educational Outreach Mission, and the ADCS Technology Demonstration Mission. However, as the ADCS Technology Demonstration Mission consists of an experimental ADCS component, it is documented under the ADC subsystem.
Design Philosophy
Geological Sample Exposure Mission
The primary payload, the Geological Sample Exposure Mission, will expose a set of geological samples, including lunar and asteroid samples, to low-Earth orbit (LEO) and observe changes in their optical reflectance properties over time. Several reflectance calibration standards will be included as controls as well. Measuring this change will allow us to learn how space weathering-factors such as vacuum desiccation, solar wind ion implantation, radiation damage (solar/galactic cosmic rays), micrometeorite bombardment, sputtering, amorphization, and formation of agglutinates change the optical properties and physical properties of asteroids, the lunar surface and other celestial bodies over time.
Presently, characterizing space weathering is a major challenge that restricts our ability to relate specific meteorite samples to asteroids that are still in space and impedes our ability to map the surface geology of these bodies. This is because space weathering changes the optical properties of airless bodies, and geological materials can be modified by entry into Earth’s atmosphere.
The best method of determining the surface composition of asteroids and other celestial objects is spectroscopic analysis, which compares the optical properties of astromaterials, and simulants measured on Earth to bodies observed in space. However, space weathering introduces a large amount of uncertainty into this analysis and hinders our ability to determine the surface composition of asteroids and other space objects. Studying how geological samples are affected by a space environment will allow us to more accurately relate the composition of asteroids, as seen by spacecraft and telescopes, to in-lab samples and will enhance other exploration missions, including the OSIRIS-Rex asteroid sample return mission.
Mission Concept
To measure the change in sample optical reflectance properties over time, Iris includes a sample plate containing up to 24 vacuum sintered geological samples, including material from asteroids, lunar, and Earth-based minerals. The remaining sample wells will house reflectance standards as controls. This sample plate will be directly exposed to the sun, as seen in the figure (right), and its reflectance will be measured daily using a pair of redundant 2 MegaPixel RGB cameras. The average red, blue, and green intensities of each sample will be extracted from each image for analysis on the ground. Additionally, full sample plate images will be transmitted once per week to verify the performance of our onboard image processing and for outreach purposes.
Presently, this mission is at a technology readiness level of 3, with benchtop tests being performed to collect baseline reflectance data for each proposed sample.
Sample Plate Reflectance Measurement
Gnomon Outreach Mission
Our second payload, the Gnomon Outreach Mission, is an educational outreach mission in which middle and high school students from the Interlake School Division are developing a gnomon to act as a pointing indicator for our Geological Sample Exposure Mission. A gnomon, as shown on the right, is an object that indicates time or direction by the position and length of its shadow. Typically, gnomons are used in sundials but they have also been included in space missions such as the MarsDial on the Spirit and Opportunity Mars rovers to determine the sun vector.
Gnomon
For Iris, the gnomon will be used to determine the geological sample illumination conditions by measuring the gnomon shadow’s length and direction. In addition to its technical function, our partnership with the Interlake School Division will enhance Iris’s education and public outreach impact.
The Iris Gnomon has a TRL of 3, and Interlake School Division students have developed our initial engineering model. This prototype has a target sun vector measurement range of 0 to 50 degrees with a sun vector resolution of 5 degrees and a rotational resolution of 15 degrees.
ADCS Technology Demonstration Mission
Our final payload, the ADCS Technology Demonstration Mission, consists of a set of experimental sun sensors and electromagnetic torque rods developed by York University. The sun sensor is a novel low-cost cubesat-class sensor with a 0.1 degree accuracy. In addition to its low cost, this sensor will showcase an innovative micro-fabrication approach in developing satellite attitude sensors by utilizing two orthogonal arrays of photodiodes with a slit pattern (see figure to the right) for greater sensing and calibration precision.
This payload has a TRL of 6 and will be raised to 7 soon after it is launched on the DESCENT CubeSat mission in Q2 2019.
For more information and documentation, please see the ADCS documentation.
York University Sun Sensor Slit Pattern Design
Functional Block Diagram
The following functional block diagram summarizes the functions of the payload: holding and protecting both the gnomon and samples during launch loads, exposing them to direct solar radiation, cosmic rays, and micro-meteorites and orbital debris. The payload must also take and store images the gnomon and samples, as well as temperature data from the payload control board and sample plate and voltage-good data from the payload control board. Additionally, the payload must communicate with CDH: sending data and executing commands.
Geological Payload Design
The following details the design of the Iris geological payload, including the payload controller, sample plate, and cameras. The payload controller is responsible for controlling the payload cameras, sending images and temperature sensor data from the payload to the spacecraft’s Command and Data Handling Unit (CDH) via a CAN bus transceiver, as well as receiving and executing commands from CDH over CAN. The sample plate holds the various payload samples, protects the samples from launch loads, and ensures that the samples are visible to the payload cameras. Additionally, the sample plate is the mounting location for a gnomon developed by middle and high school students as an educational outreach mission. Finally, the cameras take payload images as commanded by the payload controller.
Payload Apparatus
As previously mentioned, the payload apparatus consists of a payload sample plate, which holds the mineral samples, a pair of imaging cameras, a controller, and the surrounding structure.
Sample Plate
The payload sample plate holds our geological samples and will contain 24 vacuum sintered pellets and reflectance standards as well as a gnomon designed by middle and high school students as a part of an educational outreach mission. The plate is mounted at a 30o angle with respect to the +Z plane.
The sample plate assembly includes the sample plate itself and a sample plate lid. Both of these parts are made from 6061 alloy and are held together using machine screws. Locking helicoils are used as a secondary locking feature. There is also a thermistor epoxied to the sample plate to measure the plate’s temperature.
Sun vector angle with respect to the sample plate
Geological Samples
The samples themselves are circular disks with nominal diameters of 10 mm and nominal thicknesses of 2 mm. However, due to inconsistencies in vacuum sintering, the actual diameters and thicknesses are variable. The mass of each sample will vary with mineral composition but will be between 0.31 and 1.25 grams. The other sample mechanical properties are not known as these samples were custom sintered for the Iris mission. . Click here for the sample ICD page.
To ensure that the samples survive the expected launch loads, we must ensure that the samples do not move relative to the sample plate. However, we cannot use adhesives, which could affect the sample material properties. Instead, we are using disc springs to prevent relative movement between the samples and sample plate. Since the samples are variable in composition and thicknesses we are using aluminum support disks underneath the samples to concentrate the compressive load along the edge of each sample, preventing shear stresses in the samples. Using this method, the samples are only held in compression between the support disk and sample plate lid along their edges.
We have sized the disc spring to have a working load of 1.8 N to counteract the 0.086 N acceleration-induced force on the samples during 7g acceleration loads. This preloading provides a factor of safety of approximately 20.9. No smaller disc spring could be acquired.
As the frictional coefficient of the samples was determined experimentally and we found that the compressive preload is sufficient to prevent lateral sample movement during the expected launch loads. For a 7g acceleration load, we can expect a maximum 0.086 N lateral force. Our preload provides a normal force of 1.8 N but, if the sample is simultaneously accelerated in the opposite direction to our normal force, this value drops to 1.71 N. To prevent the sample from moving relative to the sample plate, the static coefficient of friction must be greater than 0.05 (slightly greater than the friction between two teflon plates). From physical testing, the actual static coefficients of friction between the samples and aluminum is between 0.28 and 0.81, depending on the sample composition. These coefficients of friction are sufficient to prevent any lateral sample movement.
To ensure that the samples do not fail or create debris, we have conducted three rounds of random vibration testing on candidate samples using an engineering model of the sample plate. For more information on these tests, please click here.
Gnomon
The gnomon is a secondary educational payload being developed by middle and high school students from the Interlake School Division. The gnomon functions in a similar manner to a sun dial and allows an observer to determine the local sun vector by observing the shadow cast by the gnomon onto a marked plate.
The gnomon consists of a 6061 aluminum measurement plate and a press-fit stainless steel gnomon pin. The gnomon is mounted directly to the sample plate using machine screws, with locking helicoils for secondary locking. The gnomon pattern is laser-etched onto the gnomon base.
Cameras
The payload includes dual COTS 2 MP cameras to observe changes in mineral sample reflectivity. Measurements will be taken with a single camera, with the second included for redundancy. A baffle is included on each camera to ensure that neither can image areas outside of the structure and are not capable of imaging the Earth. As such, a license from Global Affairs Canada will not be necessary.
We have selected the ArduCam MT9D111 camera because it is capable of outputting raw images, has a known Bayer pattern, and has a sufficiently high signal to noise ratio and pixel count to provide statistically meaningful measurements. It is based on a camera-on-a-chip with a standard s-mount lens housing. For other camera parameters, see the table to the right.
This camera has an unknown radiation susceptibility. The camera baffle will prevent direct solar light exposure which should prolong the orbital life of the cameras and reduce UV light from degrading their Bayer patterns.
ArduCam MT9D111 Parameters
Payload Controller
The primary function of the payload controller is to command the two Arducam MT9D111 cameras and transmit images from these cameras to CDH via a CAN bus transceiver. The controller must also measure temperature readings of a thermistor mounted to the sample plate and a second thermistor mounted near the two cameras, as well as a voltage-good signal from the board's buck converter transmit this data to CDH via CAN bus.
The payload controller shall be capable of performing the following when commanded to do so by CDH via the controller’s CAN transceiver:
Turn either camera on/off
Take a .RAW image with either camera and transmit the resulting image to CDH
Reset either camera
Turn off both cameras and enter a low-power mode
Exit a low-power mode
Send thermistor temperature data to CDH
Send power-good data to CDH
The Payload module will be powered by a single 6.4 VBAT (nominal) power line. Voltage will be regulated by the payload controller using a buck converter. Each camera requires 348 mW peak power and will be powered off when not in use. The payload controller itself, not including the cameras, should use less than 500 mW at peak consumption and no more than 50 mW in low-power mode. All components have storage temperature ranges of at least -20C to 50C and operating temperatures of at least -10C to 45C.
The Payload Controller has a footprint of 90 mm by 90 mm and will be mounted to the structure using machine screws and locking helicoils for secondary locking.
Payload Sample Sizing
The payload samples are 10mm nominal diameter discs. The outer 1mm will be used for clamping and the inner 8mm will be exposed to the LEO environment.
The analysis calculated here was performed to ensure that these samples were sufficiently large to ensure that the resulting data will be statistically valid.
Budgets
The following section summaries the various engineering budgets for the geological payload.
Financial Budget
The financial costs for the geological sample mission are summarized below. Note that costs for samples and ASIF times are in-kind values.
Power Budget
The geological sample mission has three major subassemblies that use power: the two cameras and the payload controller.
The total peak power use will be no more than 900 mW for a duration of 2.5 minutes once every 2 hours during the first 5 days of the mission and once every week for the rest of the mission. The payload will be in a low power mode when not in Science Mode, waiting for commands from CDH.
As shown in the camera characteristics table, each camera consumes a maximum of 350 mW while in use. However, only a single camera will take one image every two hours during the first five days of the mission and once every seven days for the rest of the mission lifetime. While not taking images, the cameras can be powered down to reduce power usage. The time to take and transmit an image to CDH has not yet been determined, but it should take no more than 2.5 minutes for an orbital duty cycle of ~2.5% (for orbits in which images are taken). For the first five days, we can expect a daily duty cycle of approximately 2.5% and a weekly duty cycle of 0.025%.
The payload control board has a targeted peak power usage of less than 500 mW. The controller will use peak power when capturing and saving images and when transmitting images to CDH. Physical testing of the EM payload board is not yet complete but will include measuring the actual power usage so that a more refined power budget can be determined. However, we are confident that the current estimates are conservative and the payload will use less than values stated below.
Mass Budget
The following table details the payload mass budget. Note that the payload module structural mass is included in this budget as well.
Subsystem CONOPS
After ejection from the Nanoracks CubeSat Deployer (NRCSD), Iris will enter a post-ejection hold for thirty minutes. After this period, the spacecraft will detumble and begin commissioning. As a part of this commissioning process, the payload will perform the following:
Check that each camera is functioning correctly.
Take initial images, process initial sample reflectance data and:
Send these images and data to ground.
Calibrate the cameras via software updates (from ground) as/if needed.
Once the spacecraft finishes all commissioning activities and enters normal operations, the payload will begin early payload operations. Early payload operations, which will occur for the first seven days of operations, will consist of taking images at two hour intervals and transmitting average sample reflection intensities to ground.
After the first seven days of operations, sample images will be taken at least once per week and transmitting the average sample reflection intensities will be transmitted to ground. However, this sampling frequency may be increased, based on the observed rate of change in the sample reflectance intensities. Full sample images will be transmitted to ground once per week as well.
Component Coatings
Payload structural components must be black anodized in visible areas. Parts requiring black anodizing include:
Note that the UMS-003 Sample plate should be masked before anodizing so that the face containing the sample wells is not anodized. Anodizing this face will cause tolerancing issues.
Key Risks
The following payload-specific risks have been identified and have had mitigation strategies planned. Most of these risks are due to uncertainties in the sample material properties but these risks have decreased as testing has occurred. As these risks are better quantified, further mitigation strategies will be developed.
Model Philosophy
The Iris team has developed an engineering model (EM) of our geological payload, and in Phase D will develop the flight model, and a second copy of our flight model to act as a ground-based control.
Engineering Model
The engineering model will have several components as follows:
Engineering model sample plate;
engineering model payload controller;
engineering model structure.
The EM sample plate is used to validate our sample plate holding system, to test if this system can prevent damage to the samples and prevent FOD creation during launch loads, and to hold samples during optical sample characterization. The EM payload controller will be used to ensure that the controller design functions as intended and will be used to control the purchased camera units during functional testing. Finally, the EM structure has been used to mount the cameras and sample plate during functional testing to ensure that the cameras can image the samples as intended. The payload module is based on similar structures developed and tested by the University of Manitoba Space Applications and Technology Society (UMSATS). Due to this similarity, we believe spacecraft-level vibration and thermal testing will be sufficient.
The EM plate vibration tests were our most critical Phase B tests, as these tests helped to determine if our mounting system could protect the samples and prevent FOD creation. For these tests, the plate was be mounted to a test interface using the same method that will be used to secure it to the spacecraft structure. To detect any material that liberates itself from the samples, the plate was be encapsulated in a container which would collect any liberated material. After testing, the container was closely examined for debris and the samples were removed from the plate and examined for damage. For more details on these tests, please see the test report here.
Flight Model
The flight model payload, consisting of the sample plate, cameras, payload controller, and supporting structure, will undergo spacecraft level testing for random vibration, thermal cycling, and functional testing.
Control Model
Finally, the control model will be physically identical to the flight model but will remain at the University of Winnipeg to act as a control during the Iris mission. Sample measurements of the control model will be taken in parallel to the on-orbit mission to account for any changes in sample reflectivity not caused by the LEO environment. No physical testing will be performed on this model, but this model might be used for functional testing, training science team members, and post-mission outreach events.
Assembly Plan
The sample plate sub-assembly and the payload control board sub-assembly (including cameras) must both be assembled before final payload assembly can begin. The upper and lower portions of the payload structure must also be prepared by installing all necessary harnessing.
Next, the sample plate will be installed onto the upper payload module structure and the payload control board will be installed onto the lower payload module structure. Finally, the harnessing between the upper and lower portions of the module must be connected and then the two half-shells can be fastened together. For more details on assembly, see the Payload Flight Assembly Page here.
Pre-Close-Up Verification Activities
Before the module is closed up, two verification activities must be performed:
[V-PLD-0131] Verify that Payload measurements are taken at a 30 degrees +/- 3 angle to the samples by inspecting and measuring the mount-angles of both the flight cameras and flight sample plate within the flight structure.
[V-PLD-0200] Verify that Payload sample plate does not cause reflective interference with the samples by taking an image of the flight sample plate with the flight cameras as they are installed into the flight module-structure and confirming that no glare obscures any samples.
Controller Schematics
Payload Parts List and Bill of Materials
The following tables summarize all mechanical and electrical parts for the payload followed by the bill of materials (BOM) for these components.
Mechanical Parts List
Electrical Parts List
Electronics Part List.xlsxBill of Materials
Verification, Analysis, Testing, and Schedules
For information regarding Payload verification, analysis, testing, and schedules, click here.
References
Payload Design Document, Iris