User’s Manual & Exercises
GEN 5.X with COSMOS
Author: Genah M. Burditt
Significant contributions of funding and course development for the first EyasSat were provided by the National Security Space Institute, Colorado Springs, Colorado, USA and the following people:
David J. Barnhart, U. S. Air Force Academy
James J. White, Timothy L. White, and Gregory D. White, Colorado Satellite Services
John B. Clark, Consultant
Jerry J. Sellers, Teaching Science and Technology Incorporated
Jesse R. Gossner, David W. Deist, and David M. Syndergaard, National Security Space Institute
Gerry Murphy, Owner, EyasSat INC.
The right to manufacture and distribute EyasSats is now held exclusively by EyasSat, Inc. Current information and technical support for the EyasSat program, which now includes additional models and accessories, can be found at EyasSat.com
Please send any suggested corrections or improvements to this manual to Genah M. Burditt.
The views expressed in this document are those of the author and do not reflect the official policy or position of the United States Air Force, Department of Defense, or the U.S. Government. Cleared for public release, distribution unlimited, USAF Academy, Department of Astronautics.
1.1. What You Need to Know.. 9
1.2. Training and Idea Exchange. 9
1.8. Recommended Accessories. 11
1.8.1. Purchase Off-the-Shelf 11
1.8.2. Specialty Items Coming Soon from EyasSat INC. 11
2.2. SS Overview Objectives. 12
2.3. SS Lab Materials and Equipment. 12
2.6.1. Determine System Center of Mass by Hanging. 16
2.6.2. Determine System Center of Mass Using Scales. 17
2.6.3. System Mass Budget Determination. 18
2.6.4. System Moment of Inertia. 20
2.6.5. System Minimum Natural Frequency. 21
4.2. EPS Overview Objectives. 25
4.4. EPS Overview Lab Materials and Equipment. 26
4.6. Battery Pack Acceptance Tests. 28
4.6.1. Battery Pack Inspection. 28
4.6.2. Battery Pack Functional Test 29
4.6.3. Battery Pack Characterization and Charging Procedure. 30
4.7.1. PDB (Top) Inspection. 30
4.7.3. Establishing Umbilical Communication and PDB Firmware Functional Test 33
4.8. Solar Array Acceptance Tests. 36
4.8.1. Solar Array Inspection. 36
4.8.2. Solar Cell Characterization. 37
4.8.1. Solar Array Characterization. 38
4.9.1. Solar Array Integration Test 42
4.9.2. Thermal Panel Functional Test 44
4.9.3. Thermal Panel Partial Integration. 45
5.2. C&DH Overview Objectives. 46
5.3. C&DH Overview Lab Materials and Equipment. 47
5.4. DH Board Acceptance Tests. 48
5.4.1. DH Board Inspection. 48
5.4.2. DH Board Functional Test 48
5.4.3. DH Board Firmware Functional Test 49
5.5. C&DH Subsystem Integration and Test. 51
5.5.1. DH Board, EPS, and Thermal Integration. 51
5.5.2. C&DH & EPS Integrated Test 52
6.1. Comm Subsystem Overview.. 54
6.2. Comm Overview Objectives. 54
6.3. Comm Overview Lab Materials and Equipment. 54
6.4. Comm Subsystem Acceptance Tests. 54
6.4.1. Comm Subsystem Inspection. 55
6.4.2. Comm Subsystem Functional Test-in Loopback Mode. 55
6.5. Comm Subsystem Integration and Test. 57
6.5.1. Comm, C&DH, and EPS Integration. 57
6.5.2. Comm, C&DH, and EPS Integrated Functional Test Using Battery Power. 57
7.2. ADCS Overview Objectives. 59
7.4. ADCS Overview Lab Materials and Equipment. 60
7.5. ADCS Hardware Acceptance Tests. 60
7.5.1. Magnetorquer (X) and (Y) Inspection. 60
7.5.2. Magnetorquer (X) and (Y) Characterization. 61
7.5.3. Top (T) and Yaw Sun Sensors Inspection. 62
7.5.4. Top (T) and Yaw Sun Sensors Functional Test 62
7.5.5. ADCS Board and Reaction Wheel Inspection. 63
7.5.6. ADCS Hardware Functional Test 64
7.5.7. ADCS FirmwareFunctional Test 64
7.6. ADCS Integration and Test. 66
7.6.1. ADCS, Comm, C&DH, and EPS Integration. 66
7.6.2. ADCS, Comm, C&DH, and EPS Integrated Functional Test 67
8.2. Structural Integration Objectives. 69
8.3. Structural Integration Lab Materials and Equipment. 69
8.4. Structural and Electrical Integration. 69
8.5. Full Functional Integrated Test. 71
8.6. ADCS Tests in a “Flight” Environment. 71
8.6.3. Calculate System Moment of Inertia (MOI) 73
8.6.4. Mode 2: Sun Track Control 74
8.6.5. Magnetorquer Tests in a “Flight” Environment 74
List of Figures
Figure 3.1 Armed for Flight (left), Mass (right) 13
Figure 3.2 RBF Panel and Lift Fixture. 14
Figure 3.3 Thermal Panel Outside & Inside. 14
Figure 3.4 Solar Panel Outside & Inside. 14
Figure 3.5 Electrical “Stack”. 15
Figure 3.6 Support Rods, Acrylic Panels, Miscellaneous Hardware. 15
Figure 3.7 (a) Lift Fixture affixed to Top Panel (Z+) 16
Figure 3.7 (b) Determining Center of Mass by Hanging. 17
Figure 3.4 Determining Center of Mass Using Weight Distribution. 18
Figure 3.9 Partially Disassembled EyasSat 19
Figure 3.10 Mass Budget Table. 20
Figure 3.1 Determining Com Port 22
Figure 3.2 Initializing COSMOS. 23
Figure 3.3 Connection Success Message. 23
Figure 3.4 Changing IP Address in COSMOS. 24
Figure 4.1 Electrical Power Subsystem (EPS) Block Diagram.. 26
Figure 4.2 (a) LED for GSE Power (red) 28
Figure 4.2 (b) LEDs for Battery Power (Green), 5V and 3.3V (Yellow) 28
Figure 4.4 Battery Test Points. 29
Figure 4.6 Power Distribution Board (PDB) 31
Figure 4.7 PDB Test Points. 32
Figure 4.8 UMB Pins on PDB and Data Interface. 33
Figure 4.10 Power Startup Message. 34
Figure 4.12 (a) Enabling Power Telemetry. 35
Figure 4.12 (b) Power Telemetry Example. 35
Figure 4.13 Turning on Power Switch 1, 3V.. 36
Figure 4.14 Solar Panel, Array and Cables. 37
Figure 4.15 Breakout Box with Corresponding Pin Diagram for Solar Array Testing. 38
Figure 4.16 Example IV Curve for 2s3p. 40
Figure 4.17 Example IV Curve for 3s2p. 40
Figure 4.18 Solar Array Characterization Schematic. 41
Figure 4.19 Integrating PDB and Solar Array. 42
Figure 4.20 (a) Telemetry Viewer, Show Screen. 43
Figure 4.20 (b) Telemetry Viewer Screen. 43
Figure 4.21 PDB & Thermal Panel Integration. 44
Figure 4.22 Thermal Panel Thermistors Pin Diagram.. 45
Figure 5.1 Data Handling (DH) Block Diagram.. 47
Figure 5.3 Data Handling Start Up Message. 49
Figure 5.4 Integrating C&DH with PDB.. 51
Figure 5.5 Attaching Thermal Panel to DH Board. 52
Figure 5.6 Telemetry Grapher 52
Figure 6.1 Comm Board with Spacecraft Radio. 54
Figure 6.3 X-CTU PC Settings. 56
Figure 6.4 X-CTU Terminal Screen Showing Loopback Mode. 56
Figure 6.5 Integrating EPS, C&DH and Comm Subsystems. 57
Figure 6.6 PDB with Arming Plug & Enable Jumper 58
Figure 7.1 Attitude Determination and Control Subsystem (ADCS) Block Diagram.. 60
Figure 7.2 Magnetorquer Board/Baseplate Assembly with Bottom Sun Sensor 61
Figure 7.3 Magnetoquers and Bottom Sun Sensor Pin Diagram.. 62
Figure 7.4 Top (T) Sun/Yaw SensorModule Outside and Inside Views. 62
Figure 7.5 Top and Yaw Sensors Pin Diagram.. 63
Figure 7.8 ADCS Startup Message. 64
Figure 7.9 Integrating “Stack” with Magnetorquer Board and Base Assembly. 67
Figure 7.10 Adding ADCS to “Stack”. 67
Figure 8.1 Securing Clear Panels. 70
Figure 8.2 Securing Solar Panel 70
Figure 8.3 Securing Top Panel 71
Figure 8.4 Suspension Stand. 72
1. Getting Started (GS)
1.1. What You Need to Know
GEN 5 is a fully functional, nano satellite. It is designed to provide users with an opportunity for understanding and exploring concepts in electrical, mechanical, and software design; and, systems engineering by using the subsystems and functionality of a Spacecraft. GEN 5 is built at its functional level much like many small Spacecraft are built today. The system is durably designed for hands-on testing and operation by multiple users. EyasSat INC. has additional products for more advanced skills training in board level hardware and software design.
The Gen 5 EyasSat is a versatile tool. It is designed such that it can be used in courses that provide high level overviews of Spacecraft subsystems and their interrelationships. It is also used to teach basic principles of systems engineering and detailed design engineering. The student experience level ranges from AP level high school to people who have been in the industry for years but need “refresher” courses and hands-on experience with inexpensive hardware before handling real, space-critical items.
There are many exercises in this manual, some are more advanced than others, and each illustrates operational elements of EyasSat GEN 5. The manual provides an overview of each subsystem and subsystem integration, and is designed for an instructor or user to become familiar with the components and versatility of the tool.
|This manual can be adapted for use in a classroom or training environment. It is highly recommended that instructors go through it thoroughly before developing a customized curriculum using this manual and/or EyasSat as tools to supplement and demonstrate elements of that curriculum.|
1.2. Training and Idea Exchange
EyasSat INC. can arrange for idea exchange between EyasSat users via EyasSat.com, facebook.com/eyassat, or by contacting EyasSat INC. directly.
EyasSat INC. also offers 1 & 2 day overview courses, which take participants through the exercises in this manual.
TSTI (Teaching Science & Technology, Inc.) offers a multi-day Verification & Validation course which provides participants with the processes, industry standards, information, and tools necessary to implement (or evaluate) a credible verification and validation program. Emphasis is on practice over theory using EyasSat as the system of interest.
GEN 5 Overview
GEN 5 is a fully functional satellite system. It is also a learning tool designed specifically for teaching Spacecraft systems engineering, electrical controls, and general engineering principals in the classroom and laboratory. The EyasSat can be used in teaching high school through post graduate school. It provides the perfect platform for thesis work, workforce/professional development and training.
GEN 5 demonstrates six traditional satellite subsystems: Structural, Electrical Power (EPS), Command and Data Handling (C&DH), Communications (Comm), Attitude Determination and Control (ADCS), and Thermal Subsystems.
Additional systems can be integrated, such as a student-designed experimental payload or subsystem. Contact EyasSat INC. for further details.
The name “EyasSat” has its roots in falconry. “Eyas” means “baby falcon” or “fledgling bird.” The falcon is the mascot of the U.S. Air Force Academy, where the concept of EyasSat was developed as a precursor the Falcon Sat Program. The first EyasSat was co-developed between the U.S. Air Force Academy and Colorado Satellite Services under a Cooperative Research and Development Agreement, (USAF CRADA NUMBER 04-AFA-239-1, 25 August 2004). The EyasSat concept is now owned by Gerry Murphy, of EyasSat INC., and the CRADA has been renewed. As a result, EyasSat INC. is continually evolving, improving and providing more and varied products and capability to its users. This manual applies specifically to “GEN 5”.
1.5. Remaining Current
Information on the EyasSat program can be found and shared at www.EyasSat.com and facebook.com/eyassat. EyasSat INC. welcomes feedback and encourages question and idea sharing between users, instructors, and students, via website and Facebook.
1.6. User Manual
If the exercises presented herein are done in order the user will become familiar with the “acceptance” process by examining each subsystem in turn, integrating, and testing them. By the end users will have performed several acceptance* tests; and, have a solid foundation in Spacecraft design. Ideally, users will work in teams of two to four to complete these exercises during which the EyasSat will be disassembled and reassembled. The exercises walk users through functional testing of each subsystems’ hardware and software components. The subsystems are then integrated together with the structure and then system-level integration tests are performed.
*Each EyasSat is thoroughly checked and tested before delivery; however, in industry Spacecraft go through acceptance testing routinely.
If EyasSat is intended for use in coursework or labs this manual can be adapted for that purpose. NOTE: It is recommended that it is modified and supplemented in order to best meet the goals of the course and students’ needs.
1.7. Standard GEN 5 Kit
Basic Kit Includes:
- Custom Indestructable Case
- GEN 5 Structure and Electronic Stack
- Spacecraft Radio (affixed to Comm Board)
- GSE Radio
- Serial data-USB dongle (for communicating with individual boards)
- Arming Plug
- Enable Jumper (for enabling battery power when SEP Switch is open, and/or testing comm in loopback mode)
- GSE Power Box
- Battery Charger set to 7.4V
- 9V Power Supply
- 5V Power Supply
- Breakout Box (for testing)
- 6, 8, and 14-pin extension cables (for testing)
- Reference Thermistor (affixed to Data Handling Board)
- Anti-Static Kit and Gloves
- Rare Earth Magnet
- Copper Suspension Stand
- Dedicated GSE Computer, Windows 7 or higher, Lennox.
- Multimeters (2) with test leads and clips
- Halogen Lamp
- Multi-outlet Electrical Strip
- Load Resistors (assortment)
- Kg Scale(s)
- Cold Spray (Dust off)
- Transducer board for determing center of mass THIS IS READY!!! Price = $600USD
- Single axis air bearing for single axis control experiments
- 3DoF Air Bearing for 3 axes control experimens (for use with 3 axes EyasSat3)
1.8. Recommended Accessories
1.8.1. Purchase Off-the-Shelf
1.8.2. Specialty Items Coming Soon from EyasSat INC.
2. Structural Subsystem (SS)
2.1. SS Overview
The purpose of this lab is to review the components and disassemble the EyasSat in order that each subsystem can be examined in turn and then re-integrated.
2.2. SS Overview Objectives
Objective 1: Vocabulary and structural components
Objective 2: Determine if EyasSat meets Center of Mass Requirements
Objective 3: Review the mass budget
Objective 4: Analyze the mass budget
Objective 5: Disassemble the EyasSat for further review of individual subsystems
2.3. SS Lab Materials and Equipment
From the Kit:
Fully Assembled EyasSat GEN 5
Antistatic Kit and Gloves*
Tools for testing:
|*As this is not an actual Spacecraft it is not actually necessary to operate in a static and/or FOD-free environment; however, it will help protect against blown circuitry and it is good form to practice handling sensitive equipment with gloves and to keep away from food or liquids in order to prevent unnecessary wear and tear on the unit. The anti-static kit contains a static mat and wrist strap for electrostatic discharge protection (ESD) protection.|
2.4. SS Vocabulary
ADCS=Attitude Determination & Control Subsystem, includes ADCS Board, Reaction Wheel, Top, Bottom, and Yaw Sensors, and Magnetorquer Board at the bottom of the “Stack.”
C&DH Subsystem=The Command and Data Handling Subsystem is comprised of the Data Handling (DH) Board with firmware and Reference Thermistor.
Comm Subsystem=The Communication Subsystem or Board is a mechanism for receiving data and transmitting commands between Ground and C&DH “Master” Subsystem. The Spacecraft radio is attached to the Comm Board. There is no firmware associated with the Comm Board.
EPS=Electrical Power Subsystem, includes Power Distribution Board (PDB), rechargeable battery pack (with charger) and Solar Array Panel. An extra solar array can be purchased.
GSE=Ground Support Equipment, including GSE Power Box with 9V Power Supply, GSE Computer, GSE Radio, and GSE Breakout Box.
“Stack”=Integrated, electrical subsystems, spacers, and 4 stack rods.
RBF=Remove Before Flight Panels (2).
Thermal Subsystem includes a Thermal Array Panel with two heaters beneath 2 plates, a copper rod, and heat pipe. A thermal tape package can be purchased for $75 further thermal studies.
2.5. SS Images
Figures 3.1 through 3.6 show GEN 5 EyasSat Structures and Subsystems. NOTE that some materials have changed from when these images were taken.Weights may vary.
|Heat Pipe and Copper RodPanels A & B|
|Heater Cable to PDBThermistor Cable to DH Board|
|Configuration CableCable to PDB|
|DH Board with Reference Thermistor|
|Arming Plug/ GSE Power Port|
|Reaction Wheel Module Affixed to ADCS Board|
|Comm Board with Radio & Antenna Port|
2.6. SS Mass Review
The structure of a Spacecraft has requirements and constraints in its design. Requirements may be driven by the need to support certain subsystems or payloads, handle the loads dictated by the launch vehicle, and provide for ease of assembly. Constraints include things like mass limitations, center of mass specifications, and lowest natural frequency limits.
The purpose of the following exercises is to introduce the concepts of center of mass verification and Spacecraft mass budgeting.
2.6.1. Determine System Center of Mass by Hanging
|Once CM and supplemental mass/ballast budget (section 4) are known, a simple method is employed for real Spacecraft for moving the center of mass to meet the Launch Vehicle (LV) requirements. The Spacecraft is affixed to/within a hemis/spherical Air Bearing. Ballast is then placed and refined. This is an example of one variable, among many, that has to be managed in order to meet all requirements.|
An important factor in Spacecraft design is getting the center of mass in the right place. The Launch Vehicle (LV) typically has stringent requirements. The Center of Mass (CM) can be determined along 3 axes via the methods described next.
- Measure the overall dimensions of the EyasSat and record on scale drawings that depict “looking along” the x, y, & z-axes. NOTE: Axes are labeled/printed on the panels.
- Determine the CM line along 2 planes by hanging the EyasSat from a suspension stand via the Lifting Fixture. A vertical line drawn from the hanging point will pass through the center of mass of an object.
- Hanging it from the Z axis will give the CM line in the X-Y Plane.
- Hanging from the Y axis will give the CM line in the X-Z Plane.
- Hanging from the X axis will give the CM Line in the Y-Z Plane.
Where any two CM lines intersect is the CM of the system. The third measurement is a redundancy and should interect at the same point. See Figure 3.7 (a) and (b).
- On the scale diagram indicate the center of mass as determined in the previous step and measure that distance from the geometric center. If the LV requires the CM to be within 1cm of the geometric center does the EyasSat, as built, meet the CM requirements? If not, ballast may be needed to move the CM.
The ballast budget, how much is needed, and where to place it will be covered in 8-9.SS.
2.6.2. Determine System Center of Mass Using Scales
Another method for determining System CM employs scales. This method can be applied to real Spacecraft using level load cells. Figure 3.4.
- Distribute the weight of the EyasSat on its 4 corners by placing them equidistant on four, level scales. Note the weights. Using these numbers calculate the CM as described and shown in Figure 4.
- Add the weights of two corners.
- Divide the smaller weight by the total. This number will give the distance from the heavier corner. Do this for all 4 corners.
- Draw a line between points on opposite planes. Where the two lines intersect is the CM.
|EyasSat INC. also offers a CG (center of gravity) Transducer Board that measures voltage ratios across pressure points (replaces scales) for determining mass distribution and CM. Contact EyasSat INC. for more information.|
2.6.3. System Mass Budget Determination
Assume the mass budget has been determined for GEN 5 based on maximum mass of 4kg.
- Disassemble the EyasSat structural components by removing thumbscrews* and unplugging cables between:
- Top Yaw Sensor and ADCS-1, 8-pin cable
- Solar Array and PDB-1, 8-pin cable
- Thermal Panel and PDB-1, 6-pin cable
- Thermal Panel and C&DH-1, 6-pin cable
Figure 3.9 Shows Top Panel Removed, Solar and Thermal Panels folded down. All that is left to do is remove the cables. Then, remove ththe thumb screws securing the clear panels.
*NOTE: Do NOT disassemble components that are secured with standard screws. Remove thumb screws only. Be sure to save all hardware in a secure location.
- Remove stack rods from front/X- side of the electronic “Stack.”
- Gently, so as not to bend the bus pins, pull each board from the “Stack” in turn. Do NOT remove th
|Use extreme care while separating the “Stack” as bus pins could bend and possibly break. DO NOT remove magnetorquer board from base plate.|
- e magnetorquer board from the base plate assembly.
- Weigh the various components of the EyasSat, record and compare to the individual component/subsystem budget-Figure 3.10. Ignore ballast for now.
|Component/Subsystem||Budgeted Mass (kg)||Measured Mass (kg)||% over/under|
|Top Panel Assembly||.5|
|Solar Array Panel||.5|
|Side Panels & Screws||.5|
|Magnetorquer Board & Base Plate Assembly (incl. Stack Rods)||1|
|PDB & Cables||.25|
|C&DH Board & Cable||.05|
|Comm Board & Radio||.05|
|ADCS Assembly & Cable||.25|
|RBF Panels||Non flight|
|Lifting Fixture||Non flight|
|Total (excl. nonflight items)||4|
Consider the following:
- Are some over? Some under?
- How would project management deal with overages?
- If the launch vehicle has given a 4Kg budget to EyasSat as its “launch mass” should there be any additional reserve on that 4Kg? If so, how could it be used?
- Determine if and where to add ballast to move the CM to meet LV requirements.
- Recall that the LV requires the center of mass to be within 1cm of the geometric center of the EyasSat.
- Recall that the table in Figure 3.10 has a specific row called “ballast”. This is the weight that can be added to counter balance the other masses and get the CM within specification.
|Weight estimates for actual Spacecraft are established at various stages in the design process. Initially, at the Requirements Review, a budget is established with margins and reserves. At Preliminary Design Review (PDR) design is finalized. At Critical Design Review (CDR) and then finally at the Pre-integration Review (checking each module) these numbers are updated. It is not uncommon for Spacecraft to encounter mass problems during PDR and CDR, that is why the “reserve” is larger at the beginning of the program and typically gets used up as the system matures. What matters in the end is not to exceed mass allocated by the launch vehicle.|
- Based on the measured values how much reserve mass is available for adding ballast and still not exceed the 3.7kg budget? Where should it be placed?
2.6.4. System Moment of Inertia
GEN 5 EyasSat is a single axis “spinner” Spacecraft. It is stable in space because it has a certain angular momentum about its z-axis. When designing a Spacecraft’s attitude determination system it must be sized to work well with the moments of inertia of the Spacecraft in order to provide the proper control. Typically, in addition to overall mass and the center of mass, designers must also consider the moments (and products) of inertia of the completed system. This is managed with mass distribution. Modern day Spacecraft are designed with comprehensive tool chains providing methods of calculation of all of these parameters in real time. The structures and integration engineer manages the moment of inertia.
The system moment of inertia for GEN 5 will be calculated through a series of exercises in the last chapter of this manual, section 9.6.3.
2.6.5. System Minimum Natural Frequency
It is important to note that there is one more very stringent and important aspect of the Spacecraft structure design which is beyond the scope of discussion/demonstration with GEN 5: A Spacecraft must have a minimum “natural frequency.” This is the lowest fundamental frequency that the structure will “ring” at. This requirement is in place to assure that the structure does not interact with the control system of the launch vehicle.
3. COMOS USER INTERFACE
3.1. Install COSMOS
The COSMOS user interface, or Guided User Interface (GUI), was developed by Ball Aerospace.
- Dedicate a Computer to EyasSat Operations. This computer should have Windows 7 or higher and will be referred to throughout the manual as the GSE Computer.
- Install COSMOS with installer file from Google Drive.
- Using Device Manager, check to see what COM PORT the GSE Computer assigns to each to radio and/or Serial to USB. See Figure 3.1. NOTE: the drivers for these will automatically update if the GSE Computer is connected to the internet. If not connected the driver can be found at http://www.ftdichip.com/Drivers/VCP.htm.
COSMOS only has appointed Com Ports for 1-9. If, your computer assigns a number higher than 9 to your radio or serial-USB com port, follow these instructions to edit COM PORT:
- Go to Device Manager.
- Right click on the COM PORT number.
- Click Properties
- Click Advanced.
- Change the number of the COM PORT to something between 1and 9.
- When it prompts that this COM PORT is already in use, click ignore.
- Click okay, okay, and close.
- Click the Command and Telemetry Server button. Figure 3.2
- Select the correct comm connection from drop down menu.
If “talking” with the entire system via radio, select cmd_tlm_server_com?_radio.txt
If using if using USB-Serial, select cmd_tlm_server_adcs_pwr_com?_serial.txt.
If using USB-Serial to “talk” with individual boards, ADCS or Power, select cmd_tlm_server_adcs_pwr_com?_serial.txt
- Hit Ok to use the cmd_tlm_server_comother_radio.txt config file.
- COSMOS should now connect to the EyasSat. There will be a Connection Successful message at the bottom of the Comand and Telemetry Window. Figure 3.3.
3.2. Multiple Users
The purpose of this section is to create multiple ground stations for the same EyasSat.
- On the main computer:
- Make sure the firewall is disabled or opened on port 7779
- Start COSMOS CmdTlmServer and connect to the EyasSat
- On the chained computer:
- Modify the file seen in Figure 3.4 by replacing IPADDRESSHERE with the ip address or hostname of the main computer connected to the EyasSat
- Put the modified cmd_tlm_server_chain.txt into C:\COSMOS_EyasSat_1.0.2\EyasSat\config\tools\cmd_tlm_server\cmd_tlm_server_chain.txt
- In a command prompt:
- cd C:\COSMOS_EyasSat_1.0.2\EyasSat\tools
- ruby CmdTlmServer –config cmd_tlm_server_chain.txt
4. Electrical Power Subsystem (EPS)
4.1. EPS Overview
The Electrical Power Subsystem (EPS) on the GEN 5 is representative of a typical EPS including a solar Array, battery (rechargeable), and Power Distribution Board (PDB) or module. Acceptance tests will be performed on each component to check physical compliance and basic functionality. Verification tests will ensure that the components perform as specified. The solar Array will then be characterized, as each unit will have its own “personality.” Once the component-level testing is complete the Solar Array and PDB will be integrated and tested. The Thermal Panel will then be integrated and tested.
4.2. EPS Overview Objectives
Objective 1: Perform acceptance tests on the Battery Pack (including battery charging)
Objective 2: Perform acceptance tests on the Solar Array
Objective 3: Characterize the performance of the Solar Array by creating an I-V curve
Objective 4: Integrate PDB and Solar Array and Thermal Panel
Objective 5: Verify integrated EPS operational status
4.3. EPS Block Diagram
4.4. EPS Overview Lab Materials and Equipment
From the Kit:
Solar Array Panel
GSE Power Box with 9V Power Supply
Battery Charger—set to 7.4V
6, 8, & 14 Pin Extension Cables
Antistatic Kit and Gloves
Tools for testing:
Multimeters (2) and Jumper Cables
Ceramic load resistors (variety in range of 18-150W)
4.5. EPS Background
This subsystem has been designed to be very similar in architecture to basic direct energy transfer power systems on today’s operational Spacecraft. A Solar Array should be configured such that voltage is “matched,” via choosing the right type of cells and number of cells in series, as performed in last set of exercises, to the battery voltage so that it will charge nominally near the peak power point. All power subsystems, including that on GEN 5, consist of the following components:
- Power generation (Solar Array)
- Power storage (in the form of a secondary or rechargeable battery)
- Battery charge manager
- Power switching and distribution
- Power monitoring (current and voltage at various points in the system)
Additional considerations include how a satellite is operated on the ground, how it is launched, and how it operates in orbit. The GEN 5 has two specific “plugs” that connect to the only externally exposed connector: the Arming Plug/GSE Power Socket. Just as in real satellites when “on the ground” the Spacecraft uses GSE Power. In the case of EyasSat this is GSE Power Box with 9V Power Supply and Battery Charger (set to 7.4V). In this instance the Battery is not connected to the bus and the Solar Array is not connected to the battery. The battery charger is meant to keep the battery fully “topped off” until launch. Examine the GSE Power box and note that there are two distinctly different plugs for the Battery Charger and the 9V supply. Real satellites always use different style plugs to prevent a technician from mistakenly plugging in the wrong component into a socket. This is the typical operational scenario “on the ground.” Although there are usually special provisions for operating the satellite on the battery just to test it “end to end,” most of the time the battery usually has its charge “topped off”—even on the launch pad in some cases—and a ground source supply is used to operate the satellite during all testing.
Before EyasSat is “launched” GSE Power is removed and the Arming Plug is attached. This plug connects the Solar Array to the system and “arms” the battery—that is, the battery is now ready to supply power to the satellite. When the satellite separates from the launch vehicle, a separation switch, or SEP switch, is closed and the satellite is powered. This allows the satellite to “launch dead” and not deplete its battery while waiting on the launch pad.
Examine the EPS Block Diagram (Fig 4.1) again and note how these two “plugs” operate and how the closure of the SEP Switch activates the electronics.
To operate EyasSat in “flight” mode suspend it from a Suspension device so that its SEP Switch is closed.
The power from the battery is distributed to two voltage converters that create a 5V bus and a 3.3V bus. The 5V bus and the raw battery power are both supplied to the EyaBus at the back of the “Stack” so that power is distributed throughout when the system is active. Power that is always active and cannot be turned off as it is a part of the “essential” bus. The radio, C&DH, and EPS all use essential bus power.
Note that the 5V and 3.3V outputs from the converters are also routed through switches and then to the same EyaBus. These are sources of power which are controlled via the EyasSat Control Panel or GUI. Th
|A separation switch is used to tell the Spacecraft and operators when it has separated from the launch vehicle. There are two different types, logic and power-in-the-loop. EyasSat demonstrates power-in-the-loop, in which the power to run the satellite is connected through the SEP Switch which acts like a power-on switch when the satellite is separated from the launch vehicle. This is OK for very small satellites but larger satellites which consume a lot of power use a “logic” switch which, via some circuitry, enables power to the system.|
ey activate parts of the satellite or its payload that do not need to be on all the time.
On the PDB there is a series of 5, blue LEDs that will come on when specific switches are activated—See Section 25.EPS:.
Additional LEDs will indicate when the EPS is operating as it should.
- Red LED: When powered by GSE Power (9V supply) a red LED is illuminated. Figure 4.2 (a).
- Green LED: When operating on the Battery power a green LED (SEP switch closed or “flight” mode) is illuminated. Figure 4.2 (b).
- 2 Yellow LEDs: Whether operating on GSE Power or Battery Power there will be two yellow LEDs, indicating that the PDB is powered and that the 5V and 3.3V supplies are operating normally.
Note that there are current and voltage monitors on a number of key test points. All of these analog signals are routed either though an external multiplexor (MUX) first or directly to one of the Analog to Digital Converter (ADC) channels on the microcontroller.
Each EyasSat board uses an Atmega 128 microcontroller programmed in C to operate. All communication over the backplane connector is via an SPI bus (for more on SPI search “Standard Peripheral Interface”).
With that background thorough testing of the PDB/EPS can proceed.
4.6. Battery Pack Acceptance Tests
4.6.1. Battery Pack Inspection
The purpose of this test is to ensure that the Lithium Ion Battery Pack is affixed to PDB.
- Visually inspect the battery pack to ensure that there are no loose parts, wires, connectors or solder connections. Figure 4.3.
|The purpose of secondary (rechargeable) batteries on a Spacecraft is to provide power during eclipse times and to provide peak power when the payload and/or subsystems draw the greatest load.|
4.6.2. Battery Pack Functional Test
The purpose of this test is to ensure that the wiring to the Battery Pack is properly connected. The PDB has a number of test points on it. These allow easy access to test voltages. NOTE: Do not attempt to remove the batteries or disconnect the wiring. And, never charge the batteries with anything other than the GSE Power Box and Charger (set to 7.4V) supplied in the kit.
- Using test leads and a multimeter connect the black lead to the pin on the PDB labeled “gnd” for ground, and the red lead onto the battery test point labeled “BATT+” as shown in Figure 4.4.
|Test Points on PDB|
- Set the multimeter to the VDC ( ) mode and note the voltage.
4.6.3. Battery Pack Characterization and Charging Procedure
If the initial voltage from the Battery Pack Functional Test is greater than 7.5V the module most likely has enough charge to perform additional experiments. If the voltage is less than this the battery needs to be charged. In any case the battery will eventually need to be charged, therefore, the method for charging is covered next.
- Connect GSE Power Box to 9V Power Supply and PDB. Figure 4.5.
|Charge Port, use ONLY charger provided in kit, set to 7.4V|
- Connect the Battery charger to the port indicated. Note that the charger connector cannot be interchanged with the 9V power connector—it is good design practice to assure that technicians cannot plug in a wrong connector by mistake. DO NOT plug the charger into an outlet yet.
- Using a multimeter and test leads, as in step 3.EPS, verify the battery “pre-charge” voltage.
- Connect the charger to a wall outlet.
- Verify that the battery voltage starts to go up by observing the battery voltage at 5 minute intervals, this will give an idea of how long it takes to charge. The charger prevents over-charging.
- Stop charging the battery when the green light appears on the charger and note the final battery voltage—it should be near 8.5V.
4.7. PDB Acceptance Tests
4.7.1. PDB (Top) Inspection
The purpose of this test is to ensure that the PDB has proper configuration.
- Compare top of PDB to Figure 4.6. Ensure there are no loose solder connections or debris.
- Note the firmware and hardware versions.
|Arming Plug & GSE Power Port|
|SEP Switch Jumper Pins|
|Data Interface Pins|
|Thermal Panel Connection|
|Solar Panel Connection|
|Firmware & Hardwared Versions|
4.7.2. PDB Functional Test
The purpose of this test is to verify that the PDB can deliver all the necessary voltages. Before commencing this test in earnest it may be useful to once again review the EPS Block Diagram (Fig 4.1).
- Power up the EPS with GSE Power Box and the 9V Supply.
- Verify that the red power LED illuminates and that the two yellow (5V and 3.3V) also illuminate.
- Set the multimeter to the VDC ( ) mode.
- Using test leads check the voltage across the GND and 5Vcc voltage test points on the PDB as shown in Figure 4.
- Move the positive test lead to measure the voltage on 3Vcc then MAIN, then to BATT+ to verify battery voltage-PDB is still being powered by ground. NOTE: SA+ voltage can be measured in this manner when Solar Array is connected to PDB and running on GSE Power.
- Move negative test lead to AGND and the positive test lead to 5V to measure the analog voltage. This is used to power analog sensors.
- Remove GSE Power from the PDB.
- Replace it with the Arming Plug.
- A single green LED should illuminate indicating that the bus is “armed” and ready to go as soon as it separates from the launch vehicle.
- Now lift the PDB board so that the SEP Switch is closed as in separation; OR, bypass the SEP Switch by installing the Enable Jumper. Two yellow LEDs should illuminate indicating that 5V and 3.3V power are active. Recall Figure 4.2 (b).
4.7.3. Establishing Umbilical Communication and PDB Firmware Functional Test
The purpose of this test is to verify that the software running in the onboard microcontroller is able to correctly process commands, turn on and off the power switch hardware, report telemetry (voltage and current) of the battery and solar panel inputs, and 5 and 3.3 V regulated lines.
The following steps, powering up and starting communication between Ground and EyasSat, will be repeated throughout this guide.
- Attach the Data Interface shown in Figure 4.8 to PDB. Use cable included in Kit to connect Data Interface to GSE Computer.
|UMB pins where Data Interface attaches to PDB|
- Open COSMOS and Command and Telemetry. Select cmd_tlm_server_adcs_pwr_com?_serial.txt
- Open Data Viewer. Figure 4.9.
- Power the PDB using GSE Power and 9V supply.
- A startup message will appear in Data Viewer. Figure 4.10. This will indicate the firmware version on the board. This should match the sticker on the board.
- Open Command Sender. Figure 4.11
- To retrieve additional Power Data select that command from the drop down menu. And select appropriate value/state. Click Send to send the command. Figure 4.12.
- Verify that the 5 power lines can be switched on and off by selecting appropriate command and value/state. Figure 4.13.
The 5 power lines are:
- Pwr_3V for adding an experiment
- Pwr_ADCS (9V)
- Pwr_EXP (5V) for adding an experiment
- Pwr_HTR1 (5V) powers the heater behind Panel B (black)
- Pwr_HTR2 (5V) powers the heater behind Panel A (white)
Reference the EPS Block Diagram. Each of these switches is connected to a different pin on the bus connector and provides power to different devices. A blue light on indicates that the switch has closed and voltage is provided to that bus pin.
- Power down the PDB by removing GSE Power.
4.8. Solar Array Acceptance Tests
4.8.1. Solar Array Inspection
|Always handle the Array with gloves and keep dust and finger prints off the surface. Never lay the solar Array down on a hard surface with the cells facing downward. It can be laid on a soft cloth.|
The purpose of this exercise is to “accept” the condition of the hardware. The jumper cable allows switching from 2s3p (3 parallel sets of 2 cells in series) to 3s2p (2 parallel sets of 3 cells in series).
- Compare the Solar Array, inside and out, to Figure 14 to ensure components match. Note the Array configuration.
- Visually inspect (both sides of) the Solar Array Panel to ensure that there are no loose parts, wires, connectors or solder connections.
4.8.2. Solar Cell Characterization
The purpose of these exercises is to observe Solar Array temperature under different conditions and to characterize individual solar cells. Best results will be achieved when performed outdoors, in bright sunlight. A high intensity light source such as an LED or halogen light will also work. NOTE: A halogen lamp can be used; however, it is advised to proceed quickly, yet safely, in taking all the measurements because Array temperature will rise rapidly.
Real Spacecraft have systems in place to keep panels cool because solar cells are less efficient when hot. This is easier with deployed Arrays than body-mounted panels. The thermistor will give the temperature of the Array.
In order to conduct these tests the Breakout Box, 15-pin extension cable, and an intense light source will be needed.
|A thermistor is simply a resistor that has a temperature dependent resistance.|
- To check invidual cells connect the Solar Array Panel to the Breakout Box (Figure 4.15) using 14-pin extension cable.
- Set the multimeter to ohms (Ω) to measure the thermistor across pins 7 & 8 at room temperature.
- Set multimeter to volts ( ) to measure voltage to for each cell across the pins indicated in Figure 4.15. Take the first measurement in ambient light. NOTE: If voltage reading is negative, the polarity is reversed
- Cover the Array and observe differences in voltage.
- Now, illuminate the Array with a bright light source such as LED flashlight, halogen light, or direct sunlight.
- Switch multimeter back to ohms (Ω) to observe any temperature change that may have been induced by light source.
- Remove the 14-pin cable from the Breakout Box.
4.8.1. Solar Array Characterization
The purpose of these tests is to determine maximum power point and ideal solar cell configuration, to match charging range of battery.
In order to conduct these tests the Breakout Box, 8-pin extension cable, 2 multimeters, and resistors with a variety of values, and an intense light source will be needed.
| Characterizing a Solar ArrayThe purpose of finding an I-V (current – voltage) curve of an assembled Solar Array is to determine at which point maximum power is delivered from the solar Array and to understand how to “match” an Array to a battery.
Most EPS systems use a method called “direct energy transfer.” In this method, the engineer designs and “tunes” the system to operate most efficiently under expected flight conditions.
In a more complex system in which power must be optimized, “peak power tracking” (PPT) is used: dynamic measurements are made and the interface adjusted to assure peak power delivery in all lighting and load conditions.
Voc (open circuit voltage) and Isc (short circuit current) are the endpoints on the I-V curve on the axes.
Higher values of resistance (lower current) will produce data points to the right of the “knee” of the curve while lower values of resistance (higher current) will produce data points left of the “knee” of the curve.
Direct Energy Transfer from Solar Array to Battery
The Battery and Solar Cells in GEN 5 are designed so there is no need to perform voltage or current conversions in order to send energy back and forth. Some systems have “step up” voltage circuits because the solar cells produce inadequate voltage to power the battery. Systems like these are complex; greater complexity can lead to greater number of problems.
NOTE: in a Direct Energy Transfer System the solar array voltage telemetry will read roughly the same as the battery when operating on battery power. To determine actual voltage produced by solar array it must be disconnected from the battery or separate from the system.
For user reference representative data was collected for two Solar Array configurations, Figures 16 & 17. In both cases the data was obtained under the following conditions:
- Clear skies
- Near summer solstice
- Elevation = 5,675 ft/1,729.74 m
- Lat/Long = 7469° N, 105.2108° W
NOTE: Results will vary with each EyasSat and testing conditions.
- Start by putting the 14-pin configuration cable on the back of the Array in 2s3p position in order to characterize the entire Array in this configuration.
- Connect the Array and Breakout Box using 8-pin cable.
|Jumper cables between pins|
- Prepare the Breakout Box with Ohm (Ω) and Voltage meters and jumper cables as in Figure 4.18.
|Gnd side (cathode)|
|Plus side (anode)|
- First measure the Voc leaving the terminals open (no load). The meter designated as “I” should measure 0 mA DC.
- Measure Isc by shorting out the test leads where the load resistors will be placed. The meter designated as “V” should show 0 VDC (or nearly 0).
- Now, using a range of resisitors (18-150W) place resistors alone, in series, and/or in parallel and note current (I) and voltage (V) for each instance.
- After several data points have been collected an I-V curve can be plotted showing I vs. V and P vs. V. The curve should be similar to the one in Figure 16 though exact data points will be different.
- Repeat this exercise with 3s2p configuration by moving the 14-pin configuration cable on the back of the array.
- Compare the I-V curves to determine which configuration more closely matches the range of battery voltages for ideal Direct Energy Transfer.
4.9. EPS Integration Test
4.9.1. Solar Array Integration Test
This purpose of these tests is to observe how a Solar Array works on a Spacecraft while in orbit, going through sunlight and eclipse cycles. The EyasSat Solar Array is approximately 10% efficient which is enough to provide some “help” to the battery in powering the bus when properly configured; but, not as efficient as cells on an actual space craft. In an actual Spacecraft the design would support the Solar Array in charging the battery during times when the bus is drawing little power. RECALL: The battery should always be charged with the charger provided and set to 7.4V.
- Connect the solar Array to the PDB as shown in Figure 4.19.
|Solar Array Panel|
- Connect the Data Interface Dongle to PDB and GSE Computer.
- Open COSMOS and connect to PDB as before
|In a direct engery transfer system the Solar Array and Battery Voltages are tied together and therefore will give the same voltage telemetry. When operating on ground power there will be independent telemetry.|
- Power up PDB with GSE Power. NOTE: All Solar Array tests must be completed while using GSE Power. When PDB is powered by battery the Solar Array voltage telemetry will be the same as the battery.
- Open Telemetry Viewer as Figures 4.20 (a), (b), and (c).
- Observe V_Batt, V_SA, I_Batt, I_SA, and I_MB in ambient light. NOTE: I_SA will be 0 when on GSE power.
- Cover the Array completely. Observe the same data points.
- Illuminate the Array with a bright light source, i.e. direct sunlight, LED flashlight, or halogen lamp and observe the change in telemetry.
- Note T_SA.
|Imb * Vmb > I3.3 * I3.3 + I5.0 * V5.0 because of losses in the converter.Also, the user will notice when operating on battery power that VSA = VBatt; but now ISA > 0.|
- Remove GSE Power.
4.9.2. Thermal Panel Functional Test
The purpose of this test is to ensure the thermistors within the Thermal Panel are functional. NOTE: The thermal panel temperatures are read by the C&DH Subsystem therefore complete integration and functional testing of the thermal system will take place in the next section.
- Using 6-pin cable, connect Thermal Panel to Break out Box. See Figure 4.21.
|PDB Attachment Point|
|Thermistor or Breakout BoxCable|
|Solar Array Panel|
- Observe resistance for various thermistors at ambient temperature across pins shown in Figure 4.22.
4.9.3. Thermal Panel Partial Integration
- Attach the Thermal Panel to PDB, as in Figure 4.21. This connects the heaters from switches 4 & 5 to the thermal panel. NOTE: the resistance of these heaters is approximately 10Ω.
- Switch 4 powers the heater behind Panel B (black)
- Switch 5 powers the heater behind Panel A (white)
- Attach the serial data to UMB interface to PDB if is not already.
- Open COSMOS and connect as before.
- Now, power up EPS with GSE Power
- Turn on the 5V power to heaters 1 (Panel B, Black) and 2 (Panel A, White) using the commands Pwr_HTR1, value/state, ON and Pwr_2, value/state, ON.
- Observe changing resistances/temperatures for Panels A and B. NOTE: The heaters warm up slowly. When operating on battery power note that the heaters will draw more power from the battery fairly quickly which can lead to odd behavior in any of the subsystems. In an actual space craft this would be mitigated.
- Turn off the heaters and allow Panels to return to ambient temperature. NOTE: Observe that the two panels emit heat at different rates.
- Hold a heat source, such as a halogen light, over the Thermal Panel and observe resistance/temperatures for Panels A and B. NOTE: Observe that the panels will absorb heat at different rates.
Thermal tapes, with varying emissivities/absorptivities are available for purchase from EyasSat INC ($75USD) for more comprehensive thermal studies.
- Next note the ambient resistance for Base, Top A and B.
- Now spray base with cold spray (i.e. Dust Off) and observe changes in the three temperatures. NOTE: One rod is a hollow heat pipe, the other is a solid, copper rod. Engery transfer from base to top will be different for each.
EPS integration and functional testing is now complete.
5. Command & Data Handling Subsystem (C&DH)
5.1. C&DH Overview
This purpose of the following exercises is to become familiar with the Command and Data Handling Subsystem (C&DH) which is representative of any nanosat C&DH. The DH Board is the “Master” in the electronic “Stack.” It distributes commands and collects telemetry and sends it back to Ground through the Comm Subsystem (or Serial-USB). It also has a thermistor reading-bank for the following temperatures:
The DH Board will be characterized individually. Then, once component level testing is complete, it will be integrated into the “Stack”. Following that, integrated Thermal Panel functional testing can take place.
5.2. C&DH Overview Objectives
Objective 1: Perform acceptance tests on the C&DH Subsystem and its components
Objective 2: Integrate C&DH with EPS and Thermal Panel
Objective 3: Verify integrated operational status
5.3. C&DH Overview Lab Materials and Equipment
Integrated EPS and Thermal Panel
DH Board with affixed Reference Thermistor (middle pins)
GSE Power Box with 9V Supply
5V Power Supply
Data Interface and Cable
Antistatic Kit and Gloves
Tools for testing:
5.4. DH Board Acceptance Tests
|5V Power Port|
|Reference Thermistor, Center Pins|
|Thermal Panel Connection|
|UMB/Data Interface Pins|
|Firmware & Hardware Versions|
5.4.1. DH Board Inspection
The purpose of this test is to ensure that the DH Board is in the proper configuration.
- Visually inspect the DH Board to ensure that there are no loose wires, connectors or solder connections or bent pins.
- Compare the DH Board to that in Figure 5.2. Ensure that the type, location, and number of components match exactly. Ensure the Reference Thermistor is affixed to middle pins.
- Note the firmware and hardware versions.
5.4.2. DH Board Functional Test
The purpose of this test is to confirm that the DH Board meets basic performance requirements.
|Always use Power Supplies provided in the Kit.|
- Connect the 5V Power Supply provided in the kit to the DH Board.
- Verify that a red power LED illuminates on the DH Board.
- Disconnect the 5V Power Supply.
5.4.3. DH Board Firmware Functional Test
The purpose of this test is to verify that the firmware running in the onboard microcontroller is able to correctly interpret and execute call sign, time set, and telemetry delay commands.
- Connect the Data Interface to the UMB pins on the DH Board (reference Figure 5.2) and attach to the GSE Computer.
- Open COSMOS and connect as before.
- Reconnect the 5V Power Supply to the correct power port on the DH Board.
- A start up message should appear in the Text Only Interface Screen, similar to that in Figure 5. This will denote the firmware version and should match that on the printed label. NOTE: There is no need to enable telemetry. Telemetry will automatically begin streaming, updating every 2 seconds-the default telemetry delay value/state.
- Verify that each telemetry line starts with an “ES” followed by the call sign digit (between 0 and 9-default is 0); followed by a time stamp in 24-hour format. Note the time will be arbitrary if the clock is not set.
- Note the 2 telemetry lines in Figure 5.3: “I,” information, and “T” for the Thermistor Bank on the DH Board. The “—-“ is just a separator line between sets of telemetry.
| The fields in the C&DH telemetry line are as follows:ES is followed by a single digit value/state, 0-9, that represents call sign
00:00:00 is the hour, minutes, and seconds. Each parameter is set with a distinct command.
TelemDelay= number of seconds between telemetry reports
Pwr: 0, not attached, 1, attached
ADCS: 0, not attached, 1, attached
EXP: 0, not attached, 1, attached
- The “I” line is telemetry from C&DH that relays the status of the at subsystem:
The “T” line gives temperatures of various thermistors. When only the DH Board is being monitored, as
| The fields in the Thermal telemetry line are as follows:DH: Data Handling Board
Exp: Experiment Board
Ref: Ambient Thermistor
The telemetry value for unconnected thermistors will be a nonsense number, i.e. -4.7.
in this case, the only two meaningful temperatures are DH Board and Ref (Reference).
- Verify the functionality of the telemetry delay using command “Pwr_Tlm.” Select a value/statue, 1-9. Click “Send.” This will configure how frequently, in seconds, telemetry is reported. Experiment with different values to determine which is optimal.
- Verify the functionality of the call sign change command using command “CALL_SIGN,” select a value/state, 0-9. Click “Send.” This command is useful if multiple EyasSats and multiple users are operating in the same vicinity. The Call Sign distinguishes each group of operators.
- Verify the functionality of the clock set commands using the following commands, “HOUR,” “MINUTE,” and “SECOND.” Select the value/state of each, 12-23H, 1-59M, 1-59S, correspondingly. “Send” after each.
- Disconnect the 5V power supply.
5.5. C&DH Subsystem Integration and Test
5.5.1. DH Board, EPS, and Thermal Integration
Now the C&DH Subsystem will be integrated with EPS and Thermal Panel. Once it is integrated, functional tests will be performed to ensure the C&DH Subsystem is communicating properly with EPS a
|GEN 5 was designed with a distributed, master-slave computer processing architecture. Most subsystems have their own microcontroller. These exercises should help the user to better understand how a master controller, like the one on the DH Board, is able to control slave/subsystem controllers, such as within EPS & ADCS.|
- Align the Bus pins on the underneath side of the DH Board with the PDB, checking in both directions, then press together firmly. Figure 5.4.
- Attach 6-pin cable from Thermal Panel to DH Board. Figure 5.5.
5.5.2. C&DH & EPS Integrated Test
- Connect Data Interface to DH Board.
- Open COSMOS and connect as before.
- Power up the system with GSE Power Box and 9V Power Supply.
- Verify DH Start up message and streaming telemetry from “I” and “T.”
- Click Show Updates. If communication is working then Rx Count field should begin updating. If not re-verify and open COM Port. Tx will update following a command.
- Enable Power Telemetry as before.
- Verfiy that C&DH subsystem is commandable by repeating 13.C&DH of the DH Board Firmware Functional Test. NOTE: Time should be preserved; however, high temperatures (above 30°C) can cause crystal in the DH clock to drift, which can lead to a drift in time.
- Verify EPS is commandable by repeating PDB Firmware Functional Tests.
- Lastly, verify that Thermal Panel is integrated and that the thermistor bank within C&DH is functioning properly by repeating the steps 63-68EPS.
- Opent Telemetry Grapher to view changing temperatures in graphical form. Figure5.6.
- Remove GSE power.
- Disconnect the Data Interface from the “Stack” and the GSE Computer.
- Leave the PDB and C&DH integrated as a “Stack”.
C&DH, EPS, and Thermal integration and functional testing is now complete.
6. Communications Subsystem (Comm)
6.1. Comm Subsystem Overview
These exercises provide a platform for understanding a small Spacecraft communications subsystem (Comm). First, acceptance tests will be performed on the Comm Subsystem. Once this component level testing is complete Comm will be integrated with the rest of the “Stack,” enabling wireless operation.
6.2. Comm Overview Objectives
Objective 1: Perform acceptance on the Comm Subsystem
Objective 2: Integrate the Comm Subsystem with current “Stack”
Objective 3: Verify integrated operational status
6.3. Comm Overview Lab Materials and Equipment
Integrated EPS, Thermal, C&DH Subsystem
Comm Board with affixed Radio and Antenna (if applicable-most radios have embedded trace antennae)
GSE Radio with Antenna (see above)
GSE Power with 9V Power Supply
5V Power Supply
Tools for Testing:
6.4. Comm Subsystem Acceptance Tests
|Loopback Jumper Pins, No Jumper Installed|
|Spacecraft Radio (must match GSE)|
|5V Power Port|
6.4.1. Comm Subsystem Inspection
The purpose of this test is to ensure that the Comm Board and Spacecraft Radio are in the proper configuration.
- Compare the Comm Subsystem to that in Figure 6.1. Ensure that the type, location, and number of components match. Check to see whether a jumper is installed on 2-pin jumper post. This jumper should be installed in next set of tests. If it is not installed then it can found in a hardware baggie within the kit. NOTE: It should not be installed during normal
- Verify that there are no loose wire clippings or other material in the EyaBus and that the connector pins are not bent.
- Note the hardware version.The Comm Board has no firmware.
6.4.2. Comm Subsystem Functional Test-in Loopback Mode
The purpose of this test is to ensure that the Comm Subsystem meets basic performance requirements.
- Install Loopback jumper across correct pins on the Comm Board (reference Figure 6.1).
- Connect the GSE Radio, shown in Figure 6.2, to GSE Computer.
- Ensure that there is a red LED on the GSE Radio indicating there is power to the unit.
- Use Device Manager on GSE Computer to determine which COM PORT has been assigned to the radio.
|This number must match Space Craft Radio|
- Close (do not minimize) COMSOS (if it is open).
- Open “X-CTU” utility. This utility will enable communication testing between radios.
- Select the COM PORT which matches that for GSE Radio. Ensure Baud is set to 19200, Data Bits to 8, and Stop Bits to 1. Figure 6.3.
- Plug 5V Power Supply into Comm Board.
- Open the Terminal Tab and type a “command” (any text will work as there is nothing attached yet; command text is blue) and a copy (text will be red) will be returned if communication between radios is working in this mode: “Loopback.” Figure 6.4. NOTE: it may take a few seconds to begin receiving mirrored text.
- Remove the Loopback Jumper.
- Sending “commands” now should not return mirrored text (there will be blue text only).
- Close X-CTU facility.
- Remove 5V Power.
6.5. Comm Subsystem Integration and Test
In this section the Comm Subsystem will be integrated with the rest of the “Stack.” Once it is integrated, functional testing, using radio communication versus umbilical, will ensure the Comm Subsystem, “Stack,” and Control Panel continue to perform as designed. This will also be the first time the system is powered using battery.
6.5.1. Comm, C&DH, and EPS Integration
- Align Bus Pins underneath the Comm Board with the DH Board, checking in both directions, then firmly press the Comm Board onto the DH Board. Figure 6.5.
6.5.2. Comm, C&DH, and EPS Integrated Functional Test Using Battery Power
- Verify communication between Spacecraft and GSE Radios by opening COSMOS and connecting as before. RECALL that the radio will have a connection address in this format: cmd_tlm_server_com?_radio.txt
- Power up the “Stack” by plugging in the Arming Plug and installing Enable Jumper on PDB (the same Jumper used in Loopback Testing) in order to bypass SEP Switch. Figure 6.6.
NOTE: There is a built-in, 10 second delay in gathering telemetry when communicating via radio, therefore, there will be no Start Up Message.
- Enable Power Telemetry as before.
- Verify commanding and telemetry.
RECALL that the system is now on battery power. The heaters can drain the battery if left on for an extended period. This could lead to odd behaviors in the system. Battery voltage below 7.5V is unreliable, especially when operating ADCS.
Comm, C&DH, EPS, and Thermal are now integrated and functional testing is complete.
7. Attitude Determination and Control Subsystem (ADCS)
7.1. ADCS Overview
The following exercises will expose the user to a small Spacecraft attitude determination and control subsystem (ADCS), which is composed of the ADCS Board, and integrated, solid-state, three-axis Magnetometer and Accelerometer, two actuators (Reaction Wheel and Magnetorquers), and multiple external sensors (Top, T, and Bottom, B, Sun Sensors and 4, Yaw Attitude Sensors). Acceptance tests will be performed on the Magnetorquer Board, ADCS Board, ReactionWheel, and Sensors. Then the system will be characterized. Once the component level testing is complete, the ADCS will be integrated to create the final “Stack.”
7.2. ADCS Overview Objectives
Objective 1: Perform acceptance tests on the ADCS Board, Sensors and Actuators
Objective 2: Integrate ADCS with current “Stack”
Objective 3: Verify integrated operational status
7.3. ADCS Block Diagram
7.4. ADCS Overview Lab Materials and Equipment
Integrated EPS, Thermal, C&DH & Comm Subsystems
Top Panel with Yaw Sensor Module
Magnetorquer Board/Base Plate Assembly
GSE Power Box
5V Power Supply
8-pin and 6-pin extension cables
Tools for testing
High intensity light source
7.5. ADCS Hardware Acceptance Tests
This section reviews acceptance on the various hardware components of the ADCS including: the ADCS Board with Reaction Wheel Module, Magnetorquer (X), Magnetorquer (Y), Sun Sensor Top (T), Sun Sensor Bottom (B), and Yaw Attitude Sensors 1-4.
7.5.1. Magnetorquer (X) and (Y) Inspection
The purpose of this test is to ensure that the Magnetorquers and Bottom Sun Sensor are in the proper co
|Magnetorquers are actuators that are used to change the orientation of the Spacecraft about its center of mass by acting against the Earth’s magnetic field. They are suitable for Low Earth Orbit (LEO) applications only.|
|Bottom (B) Sun Sensor|
- Compare Magnetorquer (X), Magnetorquer (Y), and Bottom Sun Sensor to those in Figure 7.2.
- Visually inspect Magnetorquer (X) and Magnetorquer (Y) to ensure that there are no broken wires, connectors or solder connections.
- Use a compass to check each Magnetorquer for residual magnetism.
7.5.2. Magnetorquer (X) and (Y) Characterization
The purpose of this test is to characterize the resistance of the Magnetorquer coils and Bottom Sun Sensor using a multimeter.
- Using 6-pin extension cable, connect the Magnetorquer Board to the Break Out Box.
- Set multimeter to resistance mode (Ω).
- Connect leads to corresponding pins shown in Figure 7.3 to measure resistance for magnetorquers X, Y, and Bottom (B) Sun Sensor:
- Note the resistance for the Bottom Sun Sensor in ambient light, in darkness, and in bright light.
7.5.3. Top (T) and Yaw Sun Sensors Inspection
The purpose of this test is to ensure that the Sun Sensors are in the proper configuration.
- Compare Top (T) Sun and Yaw Sensors to that in Figure 7.4.
|Yaw Sensors 1-4 (0-270°)|
|Top (T) Sun Sensor|
- Ensure the 8-pin plug is firmly attached to Top Panel.
7.5.4. Top (T) and Yaw Sun Sensors Functional Test
The purpose of this test is to ensure that the Sun Sensors are functioning properly and telemetry reflects changing from dark to light.
- To test the Sun Sensor Module connect the 8-pin extension cable from the inside of the Sun Sensor Module to the Breakout Box.
- Set multimeter to resistance mode (W).
- Measure resistance across pins 1 & 2, as shown in Figure 7.5, for Sun Sensor “T.” Note the resistance in ambient light, darkness, and in bright light (using an LED Light or direct sunlight) conditions.
- Following the pin map from Figure 7.5 measure resistance for the 4, Yaw Sensors under different conditions.
|The yaw sensor module is composed of five, small photo resistors that can precisely determine the location of a bright light source by using the power output cosine rule.|
7.5.5. ADCS Board and Reaction Wheel Inspection
The purpose of this test is to ensure that the ADCS Board and Reaction Wheel Module are in the proper configuration.
- Compare the ADCS Board and Reaction Wheel Module to that in Figure 7.6. Ensure that the type and location of components match.
|5V Power Port|
|9V Power Port|
|8-pin Connection to Top panel|
|Firmware Version, Behind Bus Pins|
- Note the firmware and hardware versions.
- Verify that there are no loose wire clippings or other material in the EyaBus connectors and that the connector pins are not bent.
7.5.6. ADCS Hardware Functional Test
The purpose of this test is to ensure that the ADCS Board meets basic performance requirements.
- Connect 9V and 5V Power Supplies to appropriate ports as indicated in Figure 7.6.
- Verify that LEDs illuminate for both power supplies.
- Spin the wheel by hand and ensure the Tach LED blinks.
- Disconnect all power.
7.5.7. ADCS FirmwareFunctional Test
The purpose of this test is to verify that the software running in the onboard microcontroller is able to properly start up, display telemetry, and properly interpret commands.
- Connect the Data Interface.
- Open COSMOS and connect as before (using serial com port again).
- Open Data Viewer
- Reconnect all Power Supplies. Confirm start up message.
- Enable ADCS telemetry using ADCS_TLM command, value/state, on.
| The fields in the ADCS telemetry line are as follows:
s_T: Top sun sensor “counts” (no units)S_B: Bottom sun sensor “counts” (no units)
s0, s90, s180, s270: Yaw Sensors 1-4 (0-270°)
ya: Solar or Yaw angle derived from s0-s270
M_X , M_Y, and M_Z: Magnetometer telemetry in units of Gauss
X and Y: Magnetorquers 0 (off), 1 (positive), 2 (negative)
rps_out: Revolutions per Second of the Wheel (range: -50-+50)
PWM_o: Pulse Width Modulation. PWM = (offset+gain) requested RPS
alg: given as 0, 1, 2, or 3 and indicates the type of close loop control (modes) algorithm 0=default, 1=wheel control is using PID loop, 2=using sun pointing algorithm
P, I, D: Floating Point Numbers
Delta T: given in seconds indicating the time to integrate before updating wheel speed
DB: (deadband) in units of degrees
g: (gain or slope) any number, slope, or gain, is part of PWM calculation (see above)
e: (extra) any number that gives an extra boost that is given to wheel when starting from 0 RPS (to overcome “stiction”.) Only applies when not in PID mode.
hyst: (rps_hysterisis) Revolutions per second leeway between commanded speed and output speed
- Send a wheel command using ADCS_WHEEL_SPD, value/state, 20
- Set value/state to -20 to reverse the wheel.
NOTE: Each wheel and motor module has its unique character; therefore, various parameters, such as gain, extra, and intercept may need to be tweaked to achieve appropriate wheel speed, within the deadband.
- Set value/state to 0 to stop the wheel.
- Verify torque rod switches/LEDs work by sending the following commands:
- ADCS_X_ROD, value/state POS = Green LED, value/state NEG = Red LED
- ADCS_Y_ROD, value/state POS = Green LED, value state = Red LED
- If accelerometer/magnetometer instrument is installed, verify acceleometers by orienting the board on all four edges and top/bottom, creating 1g acceleration along the corresponding X, Y, or Z-axis.
- Verify, magnetometers are working in orientation to the earth’s magnetic field.
- Disconnect Power Supplies and Data Interface.
7.6. ADCS Integration and Test
The ADCS will now be integrated with the rest of the “Stack”. Once it is integrated, functional tests will be performed to verify proper “Stack” operation.
7.6.1. ADCS, Comm, C&DH, and EPS Integration
- Add ADCS Module to the “Stack,” carefully aligning bus pins, as before.
- Affix X+ stack rods (2) to Magnetorquer Board/Base Plate Assembly.
- Using Stack Rods as a guide, add the “Stack” to the Magnetorquer Board/Base Plate Assembly. Figure 7.9. Be sure to align bus pins before pressing down firmly.
- Now slide the ADCS Board onto the stack rods. Align the Bus pins, checking in both directions, then firmly press the ADCS Board onto the Comm Board.
|8-pin Connection for Top Panel|
- Lastly, plug 8-pin cable from Top Panel into ADCS Board. The panel can remain flat on the table, use extension 8-pin cable if necessary. Figure 7.10.
7.6.2. ADCS, Comm, C&DH, and EPS Integrated Functional Test
- Commence Radio Communication by opening and connecting COSMOS as before.
- Power up the EyasSat with GSE Power. NOTE: the ADCS switched power line (switch 2) will be slightly illuminated once the ADCS module is in the stack; however, it is NOT fully powered until PWR_ADCS command, value/state, ON is selected.
- Power telemetry and ADCS telemetry must also be enabled, with PWR_TLM and ADCS_TLM, value/state, ON.
NOTE that due to asynchronous nature of EyasSat bus vs. synchronous natue of COSMOS, and/or radio frequency interference, ADCS telemetry is sometimes lost or truncated. COSMOS is designed to ignore 0 values.
- Verify EPS, Data Handling, and Thermal sub systems.
- Place a compass near the magnetorquers and observe the needle as the magnetorquers are turned on and off and polarity is reversed.
- Verify Sun and Yaw Sensors by illuminating each and then covering while observing how telemetry changes.
- Finally, verify that the “Yaw Angle” telemetry properly reports the direction of the light source (+/- 180) normal to Yaw Sensor 1.
- Remove GSE Power to power down the “Stack.”
Full electrical integration and functional testing is nearly complete. The last section will add and test the final components.
8. Final Integration
This section will take the user through the process of integrating the “Stack” with the Structure. There is also an integrated functional test and complete ADCS testing in “flight” environment.
Before commencing exercises review the block diagrams for the following:
- Figure 4.2: EPS
- Figure 5.1: C&DH Subsystem
- Figure 7.1: ADCS
8.2. Structural Integration Objectives
Objective 1: Integrate the full “Stack” into the Structure
Objective 2: Verify integrated operational status
Objective 3: Verify ADCS functionality in “flight” environment in 3 modes
8.3. Structural Integration Lab Materials and Equipment
Integrated “Stack” with Magnetorquer Board/Base Assembly, Thermal , Solar, & Top Panels
2 Clear Panels (-X and +X)
Thumbscrews removed during disassembly
Tools for Testing:
8.4. Structural and Electrical Integration
- Insert and finger tighten Front (X-) Stack Rods. Secure front and back Stack Rods with nuts.
- Place front (X-) and back (X+) Clear Panels and secure to structure with bottom, center thumbscrew. Figure 8.1.
- Secure Solar Panel to the Clear Panels with thumb screws. Use the outer most holes on 2 sides and one thumb screw, bottom, center. The inner holes along the sides are for affixing the RBF Panels. Figure 8.2.
|RBF Attachment Points|
- Secure the Thermal Panel in the same manner.
- Secure the Top Panel on four sides with thumbscrews. Figure 8.3.
8.5. Full Functional Integrated Test
The purpose of this section is to verify that all systems are functioning properly following structural and electrical integration.
- Repeat verification of each system. This time using battery power. To close/bypass sep switch, use jumper. Connect GSE Radio to GSE Computer. Verify that red LED is illuminated.
8.6. ADCS Tests in a “Flight” Environment
The purpose of the following exercises is to experiment with the attitude determination and control using Reaction wheel, magnetorquers, and sun/yaw sensors when the EyasSat is a “separation” or “flight” environment. NOTE: These tests require that the EyasSat is suspended from a semi-frictionless bearing (included in the kit, Figure 8.4) or floated on a single axis air bearing (available soon).
- Suspend EyasSat from Stand using 20 pound fishing line.
8.6.1. Mode 0: Bang Bang
- Steady the EyasSat as it hangs; release when stable.
- Enter a wheel speed between +15 and +30 using command ADCS_WHEEL-SPD.
- Observe the movement of the EyasSat, yaw reading, RPS Commanded, and RPS Out.
- Reverse the wheel by entering a wheel speed between -15 and -30.
- Observe the Reaction Wheel slowing down and reversing direction. The EyasSat itself will now reverse direction as well. Notice how yaw readings change.
- Experiment with the ability to control EyasSat using this technique.
- Turn the wheel off (set to zero speed) and steady the unit before proceeding.
| Zero Bias and Positive BiaseThe preceding exercise demonstrated the use of a zero bias momentum wheel. This means that the wheel is at rest and moves clockwise or counterclockwise as appropriate to move the Spacecraft.
The next steps demonstrate a positive bias momentum wheel. This means that the wheel is already spinning and therefore increases or decreases speed will cause an attitude change.
- Stabilize the EyasSat while entering a speed of +15.
- Once the wheel has spun up release the unit.
- Increase the wheel speed to +30
- Observe the change in movement of the EyasSat.
- Decrease the wheel speed back to +15.
- Observe the change in movement of the EyasSat.
- Experiment with the ability to control the unit using this technique.
- Turn the wheel off (set to zero speed) and steady the EyasSat again before proceeding.
8.6.2. Mode 1: PID Control
| A PID (proportional-integral-derivative) controller calculates an error value as the difference between a measured process variable and a desired setpoint. The PID controller algorithm involves three separate constant parameters. P depends on the present error, I on the accumulation of past errors, and D is a prediction of future errors, based on current rate of change.
The proportional (P) factor acts like gain in the equation PWM = (a+g)commanded RPS.
The differential (D) factor will boost the P or gain in the short term taking into account where RPS is at and where it needs to go.
The integral (I) factor considers the longer term error and helps to narrow the offset further refining the stability of the wheel as it approaches its commanded speed.
- Change to PID control using the command ADCS_CTRL_ALG, value/state, 1.
- Experiment with different values for P, I, and D. Starting with P equal to 1 and I and D at some fraction of 1. I and D should never be higher that P.
8.6.3. Calculate System Moment of Inertia (MOI)
The purpose of this test is to determine the system moment of inertia (MOI) since the relationship between the rate of rotation of the wheel (RPS or degrees/second) and the rate of rotation of the satellite is the same as the ratios of the MOIs. Once this is known it is straight forward to calculate the RPS needed rotate the satellite a predetermined number of degrees per second. These rations can be applied when the ADCS is in control Mode 2 (next section). The MOI of the system will be determined by solving the ratio of the wheel RPS to the system RPS in the opposite direction.
- In either Mode 0: Bang Bang (with hysteresis ≤ 2) or Mode 1: PID Control (if refined for stability) command a wheel speed of 25 RPS.
- Once the system is spinning at a stable speed shine a light on the Yaw Sensors and track the YAW data over time (logfile). Determine the rotation rate of the system-by plotting the yaw angle as a function of time. To recover log files go the COSMOS file on GSE Computer’s C: drive. C:\COSMOS_EyasSat_1.0.2\EyasSat\outputs\logs.
- Calculate the MOI of the system which is now just the ratio of rotation rate of wheel to the system multiplied by the MOI of the wheel. Higher accuracy can be obtained by trying different rotation rates and/or plotting the data over a longer period of time.
8.6.4. Mode 2: Sun Track Control
- In order to experiment with Sun Track Mode: Mode 2, start in Mode 0, Bang Bang, with text command: “ac0” or by entering a 0 in Closed Loop Control Field.
|GEN 5 can be commanded to point at a specific angle to a light source, within a certain deadband (some number of degrees.) However, to achieve success in this requires careful analysis of moment of inertia of the wheel, moment of inertia of the satellite, and the correlation of RPS to degrees per second. As the minimum Delta T is one second, spinning up the wheel too fast will cause the satellite to swing right through its deadband. This also means that appropriate values for deadband and extra need to be considered, or again, the satellite will swing through its deadband before the wheel can slow down and stabilize the satellite within it.|
- Give the satellite a momentum bias by holding it steady and entering a wheel speed of 15.
- Once stable at 15, switch to Mode 2, Sun Track Mode, by entering a value/state of 2 in ADCS_CTRL_ALG.
- Shine a bright light at the Yaw sensors.
- Enter a command for sun pointing, i.e. 270° in order to point the solar array at the light source.
- Experiment with commanding different values for wheel speed, deadband, and extra to see which combination is optimal for sun pointing. RECALL that once the relative RPS for System and Wheel are known then an appropriate wheel speed can be set for spinning the system at a desired Degrees/Second.
- Turn the wheel off.
8.6.5. Magnetorquer Tests in a “Flight” Environment
The purpose of the following exercises is to test the integrated functionality of Magnetorquer (X) and Magnetorquer (Y). Because of the weak and disturbed magnetic field from the earth a strong magnet will be needed to create a stronger, localized field.
- Place the magnet underneath the suspended EyasSat; ideal distance is 1-2”/2.5-5cm from the X Magnetorquer.
- Check the EPS Screen in the Control Panel and note the value of total bus power before the next step.
- Command Magnetorquers using ADCS_X or _Y, entering value/state, POS, NEG, or OFF. This exercises Magnetorquer (X) with normal polarity.
- Observe the movement of the EyasSat. It may take a moment or two to react.
- Demonstrate the concept of “momentum dumping.” Turn on one of the Torque Rods and allow the EyasSat to become inertially stable. Try to bring the wheel speed up slowly without moving the satellite body.
- Turn off the magnetorquers and stow the magnet.
Full integration and functional testing is now complete.