P20101: 3u Cubesat Flight Control
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Systems Design

Table of Contents

Team Vision for System-Level Design Phase

What did the team plan to do during this phase?

What did the team actually accomplish during this phase?

Functional Decomposition

A functional decomposition chart was created to help us better understand the customer requirements our CubeSat needs to fulfill. Our sub-functions branch off of two major functions - performing the diffractive experiment and performing mission operations. Our two main customer requirements are satisfied with these top-level functions. Dr. Barbosu and the SPEX customer would like a CubeSat from RIT to orbit the Earth (perform satellite operations) and Dr. Swartzlander would like our platform to support a diffractive solar sailing experiment.

Mission Sub-Functions:

Experimental Sub-Functions:

Functional Decomposition from CubeSat Requirements

Functional Decomposition from CubeSat Requirements

The second photo is a transform diagram showing the inputs and outputs of our system and how they relate to our sub-functions.

Transform Diagram

Transform Diagram

Benchmarking

Existing solar sailing missions were explored, as well as a few CubeSats developed by nearby universities. Our product is most like the LightSail 2 launched by the planetary society. No one has done the diffractive element experiment before, so there is not a direct comparison of that experiment, only the solar sailing deployment and control as a whole, or CubeSat as a whole.
CubeSat Benchmarking

CubeSat Benchmarking

Rough figures of sensor accuracy and price points were captured in the following table to benchmark and recognize the feasibility of these devices in accomplishing the CubeSat functions.

Sensor Benchmarking

Sensor Benchmarking

Concept Development

System Functions

Using the Transform Diagram, system functions for which concepts can be brainstormed were defined. Each one was related to a specific customer requirement and its attached engineer requirements.
 System Function Definitions

System Function Definitions

Design of Experiment

Our current design of the experiment is based on optimizing the amount of perpendicular exposure to the sun's light the sail receives during its orbit. The vector parallel to the sun's rays and running through the center of our sail/CubeSat longitudinal axis is called the heliotropic axis. Heliotropism is a property of certain flower species to move throughout the day to stay pointing at the sun. We have taken the concept of heliotropism to our CubeSat mission design, not only to maximize the effect of the solar radiation on the diffractive experiment but also to simplify the role of the Attitude Determination and Control System (ADACS) which will only have to make corrections to keep the sail pointing at the sun. This rules out the necessity for more complicated in-orbit maneuvers and orientations. The heliotropic axis will be where the diffractive torque is felt and captured with on-board Inertial Measurement Unit. An array of magnetometers is the most feasible method to capture this small experimental torque. Over multiple orbits, the magnetometers will be measuring oscillations in CubeSat's orientation in the magnetic field of the Earth. A small torque will become more substantial over time and contribute to an increase in the frequency of magnetic field oscillations caused by an increase in rotational velocity about the heliotropic axis.
Diffractive Experiment Concept & Magnetometer Based Velocity Measurement

Diffractive Experiment Concept & Magnetometer Based Velocity Measurement

Feasibility: Prototyping, Analysis, Simulation

Brushless DC Motor

Lab bench testing of a quad-copter brushless DC-motor was performed to determine if it was feasible to use for testing of spinning masses for the reaction wheel. Tests determined the motor was too powerful for the needs of ADACS prototyping and too difficult to control to the level of detail needed. We quickly programmed an FPGA to output an adjustable PWM signal to the motor controller. Resolution of PWM is down to 1% of a duty cycle, however, further resolution is easily implementable for finer motor control. Once we begin to set up our microcontroller environment, PWM control for testing will migrate to the target of our primary architecture under development.

Brushless DC Motor Testing

Brushless DC Motor Testing

Our FPGA implementation of the PWM source allows for the Duty Cycle to be input digitally using the development board's switches. As you can see from the simulation results below, the frequency required by the motor controller is maintained while the percentage of the duty cycle where the signal is high changes with respect to the switch's value.

PWM Simulated for FPGA Implementation

PWM Simulated for FPGA Implementation

Inexpensive Sun Sensor Prototype

Inexpensive Sun Sensor Prototype

Sun Sensor

Coarse sun tracking can be performed by measuring the angle of the sun to the spacecraft in the X and Y-axis. We can accomplish this with simple light-dependent resistors (LDRs), modeling a window in CAD we have 3D printed a functional prototype. Tests demonstrated that measurements of incident light angle can be made in both axes by the method of analog differential measurement of two sensors. This is a cheap way to determine the rough direction of the sun relative to spacecraft.
 $12,000 NSS Fine Sun Sensor

$12,000 NSS Fine Sun Sensor

A fine sun sensor with higher precision is more expensive but will provide precision for small attitude correction. Since our mission plan requires the satellite to spin along the axis parallel to the sun's rays, our sun sensor must be located along the axis of rotation in the front of our sail. This level of precision of the spacecraft's spin (and precession) will help guide the attitude actuators to a normal spin axis pointing the sail orthogonally to the sun.

The sun sensor's price will determine the level of precision as flight-tested sun sensing equipment can cost over $10,000. Implementing our own sensors is much cheaper but will require much more work to validate and calibrate under an appropriate testing environment. Ideally, our team would like to take the approach of finding cheap, reliable methods employed in published nano-satellite orbits. The International Journal of Aerospace Engineering has published methods of sensing solar angle which can be utilized for course (+/-5 degree) measurements.

Prototype Sun Sensor CAD Model

Prototype Sun Sensor CAD Model

Magnetometer

We prototyped two magnetometers angled at 180 degrees from each other; the Earth's magnetic field was measured using a differential amplifier in the analog domain. This cancels out any slight variation in the magnetic field, assuming the difference between the sensors is negligible (2"), to obtain an accurate relative velocity measurement over a long period of time at very low frequencies.

We plan on using an array of tri-axis magnetometers across our Intertial Measurement Unit (IMU) to sense our rotation and help guide torque movements during detumbling before sail deployment. Relatively accurate tri-axis magnetometers are available for less than $10 / piece. The most concerning hurdle to deal with is the calibration routine which must be developed and controlled by the microcontroller of the flight controller to interpret magnetic orientation. This hurdle must be overcome if we are to use any magnetic solutions, also including magnetorquers. Again, The International Journal of Aerospace Engineering has published methods of in-orbit triaxis magnetometer calibration.

Magnetometer Protoyping Circuitry

Magnetometer Protoyping Circuitry

Accelerometer & Gyroscope

Accelerometers on the earth's surface, have the ability to measure orientation with respect to the force of gravity. Satellites in orbit do not have this luxury since they are experiencing free fall. Therefore, accelerometers will be utilized as feedback from the torque mechanisms such as reaction wheels. Accelerometers will also require self-calibration routines as the ADAC System comes online. Also like magnetometers, they are relatively cheap for the accuracy desired. Accelerometers are not accurate enough to measure the experimental torque provided by the diffractive elements on the sail, so they will not be employed in measuring experimental data, only as feedback to the ADACS of torque effects.

Power System

For testing of the power management system on the CubeSat, a solar cell simulator was designed and constructed with settings allowing for all characteristics of a solar cell to be changed and tested. The simulator is capable of simulating different lighting conditions, efficiency, number of solar cells in parallel and series in the array, and temperature effects.
Photovoltaic Simulator

Photovoltaic Simulator

Solar Simulator Maximum Power Point Curve

Solar Simulator Maximum Power Point Curve

Orbit Predictions

Based on research into the orbits of other CubeSats and available launch vehicles, we have selected values for a possible orbit. There are several launch vehicles that can provide a Sun- Synchronous Orbit (SSO) at 600 km including several provided by Spaceflight. This orbit will have an inclination of 97.8 degrees. We are assuming an eccentricity of 0.00001 based on the orbit of several other CubeSats at that altitude. These values will give us an orbital period of 96.7 minutes. This orbit will always have a view of the sun and out of areas of high atmospheric drag while ensuring that communications will be possible over Rochester. The frequency of communication passes is yet to be determined. Higher altitudes would be preferred, launch permitting.
Predicted Orbit

Predicted Orbit

Diffractive Torque

Preliminary torque calculations from the diffractive sails in the z-axis are shown below. It is assumed that all sail is diffractive, and the force location is at the end of the sail. These are preliminary calculations before further analysis is done.

Preliminary Torque Calculations

Preliminary Torque Calculations

Morphological Chart and Concept Selection

Next, we took the sub-functions from our transform diagram and brainstormed ideas for each one. Shown below is our concept matrix detailing these ideas.
Concept Matrix

Concept Matrix

Concept Selection Criteria Definitions

Concept Selection Criteria Definitions

Concept Selection

 Full System Concepts Selected by Criteria

Full System Concepts Selected by Criteria

 Final Pugh Chart Concept Screening Matrix

Final Pugh Chart Concept Screening Matrix

Systems Architecture

Power System Block Diagram

Power System Block Diagram

Avionics Block Diagram

Avionics Block Diagram

Designs and Flowcharts

Power System Flowchart

Power System Flowchart

 CubeSat Electrical Systems Flowchart

CubeSat Electrical Systems Flowchart

Risk Assessment

The risk management list has been updated to reflect new concepts generated and system-specific items. Each team member has been assigned a risk to manage and thought up careful considerations to exercise in order to minimize their risk.
Risk Management

Risk Management

Design Review Materials

Plans for next phase

By the Preliminary Detailed Design Review

Amber

James

Sarah

Nick

Charlie

Jarrett


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