DINO: ADCS Test Plan DINO-ADCS-PLN-TSTPLN

Deployment and Intelligent Nanosat Operations

Attitude Determination and Control System Test Plan

DINO-ADCS-PLN-TSTPLN, Rev A, 11-19-03

Prepared By: Katie Dunn and Mike Urban

Prepared For: Nanosat III DINO: ADCS Test Plan DINO-ADCS-PLN-TSTPLN

Revision Log

Revision Description Date A Initial set of test plans 11/19/03

Approvals

Originator Date Safety Date

REA Date Systems Engineer Date ______I&T Manager Date ______Configuration Manager Date

ii DINO: ADCS Test Plan DINO-ADCS-PLN-TSTPLN

Table of Contents

1 Scope...... 1 1.1 Identification...... 1 1.2 Document Maintenance...... 1 1.3 ADCS Overview...... 1 1.4 Document Overview...... 1 1.5 Definitions, Acronyms, and Abbreviations...... 1 1.6 Referenced Documents...... 2 2 Hardware...... 2 2.1 Rate Gyros...... 2 2.1.1 Overview...... 2 2.1.2 Specifications...... 2 2.1.3 Verification Test Plan...... 2 2.2 Magnetometer...... 2 2.2.1 Overview...... 2 2.2.2 Specifications...... 3 2.2.3 Verification Test Plan...... 3 2.3 Torque Rods...... 3 2.3.1 Overview...... 3 2.3.2 Specifications...... 3 2.3.3 Verification Test Plan...... 3 3 Software...... 3 3.1 Overview...... 3 3.2 Verification Test Plan...... 3

iii DINO: ADCS Test Plan DINO-ADCS-PLN-TSTPLN

1 Scope

1.1 Identification This document applies to the Deployment and Intelligent Nanosat Operations, DINO. DINO is a satellite project undertaken by the Colorado Space Grant Consortium at the University of Colorado at Boulder. DINO is a part of the Nanosat III program. The Nanosat III program is sponsored by the Air Force Research Laboratory, AFRL, the Air Force Office of Scientific Research, AFOSR, the Association of Aeronautics and Astronautics, AIAA, and Goddard Space Flight Center, GSFC.

1.2 Document Maintenance This document falls under the DINO document control requirements as specified under the DINO Configuration Management plan. All subsystems are being developed at the University of Colorado and the team members there are to be part of any changes or updates. All changes and updates must be made in accordance with the DINO CM plan.

1.3 ADCS Overview This document serves as the primary test plan for the attitude determination and control system (ADCS). Plans are developed for each of the system components to determine if the components meet their functional requirements. Plans are then developed for the system as a whole to test the integration process of the individual components. The torque rods are tested for both correct input current and magnetic dipole current. The sensors are tested for correct analog output. Concerns for testing the software system are also developed within the document.

1.4 Document Overview This document outlines preliminary test plans for ADCS. The test plans include:

 Test plans for the rate gyros  Test plans for the magnetometers  Test plans for the torque rods  Test plans for the software

1.5 Definitions, Acronyms, and Abbreviations

Acronym Name FITS Foldable Integrated Thin-film Solar Arrays DINO Deployable and Intelligent Nanosatellite Operations ADCS Attitude Determination and Control System CM Configuration Management AFRL Air Force Research Laboratory C&DH Command and Data Handling AFOSR Air Force Office of Scientific Research AIAA Association of Aeronautics and Astornautics GSFC Goddard Space Flight Center

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Acronym Name

1.6 Referenced Documents For more information about the documents referenced below, contact Colorado Space Grant at the University of Colorado at Boulder.

2 Hardware

2.1 Rate Gyros

2.1.1 Overview The angular rate sensors are used to sense the rate of the spacecraft in each axis. This measurement will be used to understand the attitude of the spacecraft and determine the amount the attitude needs to be corrected by the spacecraft through the control algorithm. The gyros being used on DINO are Analog Devices ADXRS150. This sensor outputs a voltage proportional to the angular rate about the yaw axis of the chip. With three of the sensors placed on each axis of the spacecraft, a complete measurement of the angular rate of the spacecraft can be determined.

2.1.2 Specifications The sensor requires an input voltage range of 4.75V to 5.25V and draws 200μA of current. The range of the angular rate measurement is 0°/sec to 150°/sec, producing a nominal output of 2.5V with a rate of 0°/sec (see the attached data sheet for further specifications).

2.1.3 Verification Test Plan In order to verify that the rate sensor is functional for the purposes of the DINO project, a testing scheme must be developed to assure this functionality. The test will require the use of a spinning apparatus with known dimensions and a low friction in its swivel shaft.

The rate gyro will be spun up to a known rotation per minute and the power turned on. The rate gyro will output a valve proportional to the rotation rate and that can be compared to the known value of rotation, thereby verifying the functionality and accuracy of the rate gyro.

The spinning apparatus angular velocity will be measured using a stop watch and a person to count the amount of times the apparatus spins in a specified amount of time. This will give a measurement in rotations per minute, which will then be converted into a degrees per second measurement in order to compare with the output of the rate gyro. The following conversion will be used:

Omega = RPM measurement * (2*pi [rad/revolution]/ 60 [sec/min])

This will need to be done with all three rate gyros that will be used on the spacecraft to be certain the hardware is functioning correctly.

2.2 Magnetometer

2.2.1 Overview The magnetometer will be used to obtain a reading of the magnetic field of the Earth at any given instant in time. This value will be used in conjunction with the rate gyro measurement and the

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onboard software to determine the satellite’s attitude and position above the Earth. The onboard software will also use this valve as part of the control system to change the attitude and/or the position of the satellite if needed.

2.2.2 Specifications The magnetometer being used on DINO is a three axis magnetometer built by Honeywell (HMC2003). It takes a supply voltage of 6V to 15V DC and draws around 20mAmps. The resolution of the magnetometer is 40μGauss, however this may change depending on the type of analog to digital conversion used for this piece of hardware. The range of the instrument is -2Gauss to 2Gauss. The magnetometer will output the magnetic field value as a voltage between 0.5V and 4.5V with a nominal value of 2.5V.

2.2.3 Verification Test Plan The magnetometer will be tested using rated magnets. These magnets will be placed near the magnetometer and the voltage output of the magnetometer will be compared to the rated value of the magnet.

2.3 Torque Rods

2.3.1 Overview The torque rods will be used to control the spacecraft’s attitude.

2.3.2 Specifications The torque rods being used on DINO will be made of a magnesium zinc alloy. The permeability of the material is 800 W/(Am). 24 gauge copper wire will be wrapped around the alloy and the torque rods will produce 3Am2. The torque rods will operate on a 5V line and each draw 300mA.

2.3.3 Verification Test Plan The magnetometer will be used to verify to test the output of each torque rod. The torque rods will then be characterized using the magnetometer to measure the output and a circuit that can limit the current going into the torque rod.

3 Software

3.1 Overview The software used to run ADCS will be located in memory on the flight computer. The ADCS team will be writing the software for their subsystem along with the software team. The software team will be in charge of integrating this software package into the rest of the software.

The ADCS software will convert the raw data of the magnetometer and the rate gyros into useable data. It will then store this information in memory to be used later. There will be a software based orbit propagator and a magnetic field model. The software will use these models and compare them to the data collected. The comparison algorithm will determine which torque rod(s) need to be turned on and how much the current needs to be limited to.

3.2 Verification Test Plan The majority of the software testing will occur after the spacecraft has been integrated. This is primarily due to the fact the software will reside on the flight computer’s memory as a part of the Command and Data

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Handling System. The ADCS software also needs to be integrated into the rest of the satellites software before testing can being.

The first test to occur will be testing the conversion of the raw data from the magnetometer and the rate gyros. These will be tested like mentioned in section 2 but instead of using the raw values a read out from the flight computer will be used. To test the orbit propagator and the magnetic field model hardware simulators will be used in place of the magnetometer and the rate gyros. The hardware simulators will use numbers that DINO is expected to see in the orbit.