Prepared by: Team Member A. M. Espinal Mena. Submitted: Reviewed: Revised: Approved: Team Member E.M. Portilla Matías. Team Member F. O.

Transcription

1 HASP Program Preliminary Design Review Document for the Attitude Determination System (ADS) Experiment by Team: Experiments with Quality United In Science (EQUIS) Prepared by: Team Member A. M. Espinal Mena Date Team Member E.M. Portilla Matías Date Team Member F. O. Rivera Vélez Date Team Member J. I. Espinosa Acevedo Date Submitted: Reviewed: Revised: Approved: Team Member Dr. H. B. Vo/E.G.Delgado Institution Signoff (replace with name) Date Date Institution Signoff (replace with name) Date HASP Signoff Date Team EQUIS i

2 Change Information Page Title: PDR Document for ADS Experiment Date: 04/05/2010 List of Affected Pages Page Number Issue Date Team EQUIS ii

3 TBD Number Status of TBDs Section Description Date Created Date Resolved Team EQUIS iii

4 TABLE OF CONTENTS Table of Contents Experiment by... i Title: PDR Document for ADS Experiment... ii Page Number... ii 1.0 Document Purpose Document Scope Change Control and Update Procedures Reference Documents Goals, Objectives, Requirements Mission Goal Objectives Science Background and Requirements Technical Background and Requirements Payload Design Principle of Operation System Design Electrical Design Mechanical Design Payload Development Plan Payload Construction Plan Hardware Fabrication and Testing Integration Plan Software Implementation and Verification Flight Certification Testing Mission Operations Pre-Launch Requirements and Operations Flight Requirements, Operations and Recovery Data Acquisition and Analysis Plan Project Management Organization and Responsibilities Configuration Management Plan Interface Control Team EQUIS iv

5 9.0 Master Schedule Work Breakdown Structure (WBS) Staffing Plan Timeline and Milestones Master Budget Expenditure Plan Risk Management and Contingency Glossary Appendix Team EQUIS v

6 LIST OF FIGURES Figure 1 Earth's Magnectic Field[Ref.4]... 6 Figure 2 Body fixed definition and North-East-Up (NEU) coordinate system and definition of Roll, Pitch and Yaw with rotation angles... 7 Figure 14 Position of the Photodiodes... 9 Figure 3ADS System Design Figure 4 Traceability Matrix Figure 5 Electrical Design Subsystems Figure 6 Electrical Subsystem Figure 7 Accelerometer Sensor Figure 8 SCA3000 Three axis accelerometer pins Figure 9 Gyroscope Sensor Figure 10 ITG-3200 Three Axis Gyroscope pins Figure 12 Three Axis Magnetometer Figure 11 1N457 Silicon diode Figure 13 Si S1336 Photodiode Figure 17 Control Electronics Figure 18 Heaters Subsystem Figure 19 Power System Diagram Figure 20 Flight software control flow chart Figure 23 Heater Dimensions Figure 24 Isometric view of the external structure Figure 25 Sun sensor positioning and L-shape couplers Figure 26 External Structure Drawing Figure 27 Top view payload structure Figure 28 View of punctures on structures Figure 29 Internal Structure Drawing Figure 30 WBS- Work Breakdown Schedule Figure 31 Risk Management Cycle Figure 21 Heat Loss Analysis Figure 22 External Convention Team EQUIS vi

7 LIST OF TABLES Table 1 Pins for the Arduino Atmega Table 2 Power Requirements (Amperes) Table 6 Heater Parameters Table 7 Alloy Selection Table 8 Mass Budget Table 9 Materials Acquirement & Costs Table 10 Risk Management & Contingency Team EQUIS vii

8 1.0 Document Purpose This document describes the preliminary design for the Attitude Determination System (A.D.S.) experiment by Team EQUIS for the HASP Program. It fulfills part of the HASP Project requirements for the Preliminary Design Review (PDR) to be held June 4, Document Scope This PDR document specifies the scientific purpose and requirements for the A.D.S. experiment and provides a guideline for the development, operation and cost of this payload under the HASP Project. The document includes details of the payload design, fabrication, integration, testing, flight operation, and data analysis. In addition, project management, timelines, work breakdown, expenditures and risk management is discussed. Finally, the designs and plans presented here are preliminary and will be finalized at the time of the Critical Design Review (CDR). 1.2 Change Control and Update Procedures Changes to this PDR document shall only be made after approval by designated representatives from Team EQUIS and the HASP Institution Representative. Document modification requests should be sent to Team members and the HASP Institution Representative and the HASP Project. 2.0 Reference Documents The following websites are references for relevant scientific information as well as sources of electronic components, their specifications, part numbers, products availability and prices. Bibliography: [1] (USRA), R. N. (2002, November 25). Astronomy Picture of the Day. Retrieved May 24, 2010, from Astronomy Picture of the Day Web site: [2] Analog Devices Corporate Headquarters. (2008, September). Analog Devices. Retrieved May 24, 2010, from Analog Devices Web site: [3] Arduino Mega. (2010). Retrieved March 28, 2010, from Arduino:Blog: March 28, 2010 [4] Biology Blog. (n.d.). Retrieved May 24, 2010, from Biology Blog Web Site: [5] Field Code. (2008). Retrieved February 3, 2010, from Whatis?com Web site: Team EQUIS 1

9 [6] kowoma.de. (2009, April 19). Retrieved May 24, 2010, from kowoma.de Web site: [7] Metrolab Instruments. (2006, October). Metrolab Instruments. Retrieved May 24, 2010, from MetroLab Manufactures of MRI systems: [8] Mystery Class. (2008). Retrieved May 24, 2010, from Mystery Class Journy North Web site: [9] Ritter, M. E. (2009). The Physical Environment an Introduction to Physical Geography. Retrieved May 24, 2010, from sons.html [10] Sparkfun. (2003). Retrieved May 24, 2010, from Sparkfun website: [11] Wikipedia (DC-to-DC Converters). (2010). Retrieved April 4, 2010, from Wikipedia Web site: Team EQUIS 2

10 3.0 Goals, Objectives, Requirements 3.1 Mission Goal The goal of this mission is to create an attitude determination system (A.D.S.) for a balloonborne experiment. The devices of this system will provide cross-reference information required in determining the platform s position in the atmosphere. 3.2 Objectives Science Objectives The ADS experiment will accomplish the following objectives: Obtain the platform characteristics following: o Direction of movement o Rate of rotation o Position and altitude Compare/contrast the intensity of the Earth s magnetic field as the altitude increases o To determine the platform s position in regard to the Earth s magnetic field. Monitor the external temperature o Learn and corroborate the atmosphere s temperature range o Plot the temperature s variations in combination with altitude versus time Contrast the platform s position versus time To determine the platform s orientation in regard to the sun Technical Objectives Payload must weigh 3 kg or less Include a GPS to record the platform s actual position Collect data from: o three axis accelerometer o three axis gyroscope o three axis magnetometer o internal temperature sensor It will be used to maintain a suitable operating temperature range o external temperature sensor o sun sensor Complete CDR document Complete FRR document Analyze flight data & Prepare a Performance/ Science Report 3.3 Science Background and Requirements Proper background information needs to be collected to understand the science behind the objectives defined by this experiment. This is important to understand the essential mechanical and electronics systems to accomplish this mission. Team EQUIS 3

11 3.3.1 Science Background The attitude of a spacecraft is its orientation in space. The motion of a rigid spacecraft is specified by its position, velocity, attitude, and attitude motion. The position and velocity quantities describe the translational motion of the center of mass of the spacecraft and are the subject of what is variously called celestial mechanics, orbit determination, or space navigation, depending on the aspect of the problem that is emphasized. The attitude and the attitude motion quantities describe the rotational motion of the body of the spacecraft about the center of mass. Generally, knowledge of the spacecraft orbit is required. In general, the orbit and attitude are independent. For example, in a low altitude Earth orbit, the attitude will affect the atmospheric drag which will affect the orbit; the orbit determines the spacecraft position which determines both the atmospheric density and the magnetic field strength which will, in turn, affect the attitude. However, the dynamical coupling is usually ignored and assumes that the time history of the spacecraft position is known and has been supplied by some process external to the attitude determination and control system. One distinction between the orbit and attitude problems is in their historical development. Predicting the orbital motion of celestial objects is one of the oldest sciences and was the initial motivation for much of Newton s work. Consequently, although the space age has brought with it vast new areas of orbit analysis, a large body of theory directly related to celestial mechanics has existed for several centuries. In contrast, although some of the techniques are old, most of the advances in attitude determination and control have occurred since the launch of Sputnik on October 4, The results of this is that relatively little information is recorded in books or other coordinated, comprehensive reference sources. The language of attitude determination and control is still evolving and many of the technical terms do not have universally accepted meanings. Attitude analysis may be divided into determination, prediction, and control. Attitude determination is the process of computing the orientation of the spacecraft relative to either an inertial reference or some object of interest, such as the Earth. This typically involves several types of sensors on each spacecraft and sophisticated data processing procedures. The accuracy limit is usually determined by a combination of processing procedures and spacecraft hardware. Attitude prediction is the process of forecasting the future orientation of the spacecraft by using dynamical models to extrapolate the attitude history. The limiting features are the knowledge of the applied and environmental torques and the accuracy of the mathematical model of spacecraft dynamics and hardware. Attitude control is the process of orienting the spacecraft in a specified, predetermined direction. It consists of two areas- attitude stabilization, which is the process of maintaining an existing orientation, and attitude maneuver control, which is the process of controlling the reorientation Team EQUIS 4

12 of the spacecraft from one attitude to another. The two areas are not totally distinct, however. For example, consider stabilizing a spacecraft with one axis toward the Earth, which implies a continuous change in its inertial orientation. The performance of the maneuver hardware and the control electronics are generally the limiting factor for attitude control, although with autonomous control systems, it may be the accuracy of orbit or attitude information. Nearly all spacecraft requires some form of attitude determination and control. For engineering or flight-related functions, attitude determination is required only to provide a reference for control. Attitude control is required to avoid solar or atmospheric damage to sensitive components, to control heat dissipation, to point directional antennas and solar panels (for power generation), and to orient rockets used for orbit maneuvers. Typically, the attitude control accuracy necessary for engineering functions is on the order of 1 deg. Attitude requirements for the spacecraft payload are more varied and often more stringent then the engineering requirements. Payload requirements, such as telescope or antenna orientations, may involve attitude determination, attitude control, or both. Attitude constrains are most severe when they are the limiting factor in experimental accuracy or when it is desired to reduce the attitude uncertainty to a level such that it is not a factor in payload operation. These requirements may demand accuracy down to a fraction of an arcsecond (1 arc-second equals 1/3600 deg). There is no single profile characteristic of all space missions. However, most missions have in common three more or less distinct phases: 1) Launch, consisting of the activities from lift-off until the end of powered flight in a preliminary Earth orbit. 2) Acquisition, consisting of orbit and attitude maneuvers and hardware checkout 3) Mission operations, consisting of carrying out the normal activities for which the flight is intended. Launch is the most distinct and well-defined phase and is normally carried out and controlled primarily by personnel concerned with the rocket launch vehicle and who will not be involved in subsequent mission operations. A limited number of launch vehicles are in use. The objective of attitude determination is to determine the orientation of the spacecraft relative to either an inertial reference frame or some specific object of interest, such as the Earth. One or more reference vectors i.e. unit vectors in known directions relative to the spacecraft must be available to accomplish the objective. A platform attitude determination system generally uses a variety of sensors and actuators. The state of the systems is measured by the sensors, while the actuators are used to adjust the state of the system. Since attitude is described by three or more variables, the difference between the desired states and measured states are slightly more complicated than that of other systems, such as a thermostat or even the position of the satellite in space. Team EQUIS 5

13 The attitude determination is either undetermined or over determined; consequently, the mathematical analysis of attitude determination is complicated. A combination of sensors and mathematical models are used in attitude determination to collect vector components in the body and inertial reference frames to increase the precision in the analysis. Several different algorithms use these components to determine the attitude, typically in the form of quaternion, Euler angles, or a rotation matrix. A sun sensor measures the components of S in the body frame,, while a mathematical model of the Sun s apparent motion relative to the spacecraft is used to determine the components in the inertial frame,. A magnetometer measures the components of m in the body frame,, while a mathematical model of the Earth s magnetic field relative to the spacecraft is used to determine the components in the inertial frames. The attitude determination analyst needs to understand how various sensors measure the bodyframe components, and how standard attitude determination algorithms are used to estimate. The magnetic sensor will measure the earth s magnetic field which will produce an inclination compass that will react to the inclination of the earth s magnetic field. This will distinguish the pole ward and the equator ward depending on the field s strength of the local earth s magnetic field. Figure 1 Earth's Magnectic Field[Ref.4] Figure 2 shows a definition of body fixed and North-East-Up (NEU) coordinate system. The coordinate system shown in Figure 2 is defined by the pointing of one axis towards the geographical North Pole, one towards East and one in the opposite direction of the gravity force vector g. This image simulates the ADS payload during the flight time. This draw helps to visualize the results of the ADS experiment, since this model was use to determine the rotations that should be measure by calibrating the gyroscope to gather information about the selected Team EQUIS 6

14 rotations. The tilting is also shown in Figure 2 by a rotation in the body fixed frame of the gondola indexed with b around the Roll-Axis equal to X b with the angle Q and the Pitch- Axis equal to Y b with the angle Φ. Figure 2 Body fixed definition and North-East-Up (NEU) coordinate system and definition of Roll, Pitch and Yaw with rotation angles Science Requirements To Record the time during the flight To analyze the measurements gathered from the sensors and convert the quantity values into physical quantities To understand the environment at which the payload is going to be exposed Determine the platform s o Rotational & translational motion during flight o Position o magnetic field intensity o direction to the Sun 3.4 Technical Background and Requirements The principle of operation for the ADS experiment consists on the instruments integration to collect measurements from several sensors that will be analyzed to comprehend the balloon s dynamic that will be affected during the flight Technical Background The ADS payload will gather the following measurements with a: o The three-axis accelerometer is an electromechanical device that will allow us to obtain acceleration, by measuring the amount of static acceleration due to the gravity; this amount will permit us to determine the angle of tilt with respect to the earth of Team EQUIS 7

15 the platform. This device will allow us to understand the payload surrounding better, thus allowing better knowledge of attitude motion. o The three-axis gyroscope is a device for measuring or maintaining orientation, based on the principles of conservation of angular momentum. The gyroscope will provide information about the rotation on each one of the axes. The orientation can be determined by combining the rotation data of the gyroscope with the acceleration data of the three axis accelerometer. o Three-axis magnetometer The Earth s magnetic field will be used as a reference to determine the position of the platform. The magnetometer will measure the magnetic field intensity; as a result, the magnetic field obtained will be used to acquire the position with respect to the Earth s magnetic field. o Global Positioning System (GPS) The GPS will provide the altitude, position, latitude & longitude and time from the large amount of data that the GPS sends on each measure. o Temperature sensor (two) Internal temperature sensor is used to monitor the internal environment at which the devices will be exposed. This will be used to obtain the data that will turn ON and OFF the heater to maintain the proper temperature for the main components. The external temperature will be gathered to compare the readings of the internal and external temperature obtained during flight and ensure that the selected alloy will isolate the components from the external temperature The software code will classify the raw data into analog and digital outputs; therefore, if it is an analog output it will be converted into digital using an ADC and then it will be stored in an SD card memory module. If it is digital then it will be stored directly in the SD card. The sun sensor uses a photodiode which has been used successfully in similar projects. This photodiode is subject to change depending in the different parameter of the payload. The photodiode is used because it is simple and lightweight compared with the star tracker that is more expensive and heavier. Five photodiodes will be used and place strategically in each side and top of the payload as shown in Figure 14. Team EQUIS 8

16 Figure 3 Position of the Photodiodes The reason that the photodiodes are place on the sides is to allow the system to determine the position of the sun through the measurements on each photodiode. Since the sun will emit light to the payload at certain angles, the measurement obtained by each photodiode will be different, thus allowing a more precise positioning of the sun. Knowledge of the orientation of the payload is necessary to obtain the trajectory and behavior of payload. For this reason, every payload needs a set of sensors to determine its position relative to the sun. For instance, if we desire to orientate a camera at a certain direction, this system will allow us to know our orientation and permit us to do adjustments Technical Requirements A temperature sensor to monitor the temperature at which the payload was exposed. Have an SD card with enough space to store the sensors data since the data is calculated to be stored are in the range of megas bytes but more calculation need to be considered. The idea of having a SD card with larger capacity than what is need is just in case a problem occurs. 4.0 Payload Design The ADS experiment has a sensor subsystem of eleven main sensors that will take measurements to study the balloon platform dynamics. The design arrangement for the experiment consist of rotational sensor, acceleration sensor, five sun sensors to sense the direction to the Sun, magnetic field sensor, GPS, internal and external temperature sensor to acquire data at different positions to get exact information about the ADS payload. The magnetic field sensor, the GPS sensor and the five sun sensors set will be used to determine a specific reference frame from where the others sensors will work from. The 3-axis gyroscope and the 3-axis accelerometer are used to specify the payload location with respect the reference frame previously defined. The temperature sensors will help to keep the system at proper temperature by turning ON/OFF a heater depending on the real-time environment conditions during launch and ascending. Team EQUIS 9

17 4.1 Principle of Operation The ADS experiment will take a variety of measurements, such as acceleration, precision and orientation, temperature, position, magnetic field and sense the direction to the Sun. A three axis accelerometer sensor will be use for the acceleration data. The position will be determined by a GPS. The magnetic field concentration will be determined by a three axis magnetometer. The orientation data will be obtained by a three axis gyroscope, whereas the Sun location will be acquired by five sun sensors. The internal and external temperature data is obtained by an internal and external temperature sensor respectively. Finally, a heater device will maintain the internal temperature within the temperature operating range of all instruments. 4.2 System Design Figure 4ADS System Design The system design Figure shows several elements of the ADS system. The wide clear blue connections designate the data exchange between the flight control and the sensors that will be in the payload. The black connection stands for the ON/OFF control instructions from the Atmega 1280 Arduino microcontroller to the heater. The green line represents communication of the telemetry and command interface. The dark gray line and the mechanical system box represent the ADS payload as shown in Figure 3. The two temperature sensor boxes represent the internal and the external temperature sensors. Finally the power supply box stands for the power subsystem. The power subsystem is explained in more detail in the sections power supply and power budget. Team EQUIS 10

18 4.2.1 Functional Components 1. Power Subsystem: VDC will be provided by the HASP platform during the entire flight. A power budget analysis is required to ensure that this power will be sufficient for the instruments of the ADS experiment for the entire period of flight. 2. Microcontroller: The Atmega 1280 Arduino microcontroller, which is connected to all the instruments, is in charge of collecting the measurements from all the sensors and store it in the SD card. In addition, the Atmega 1280 will send the required on and off instruction/functions to the heater, so that it can control the heater. 3. SD Card Memory: The SD card memory module will be added to the Arduino were the measurements of the sensors will be stored here as digital quantities. 4. Three Axes Accelerometer: The accelerometer will take the acceleration measurements of the payload. Each one of it axes will perform measurements related to the payload response on that axis. 5. Three Axes Gyroscope: The gyroscope is in charge to determine the orientation of the ADS payload by sensing the platform rotation during flight. 6. Three Axes Magnetometer: The magnetometer will measure the earth s magnetic field concentration. 7. External Temperature Sensor: The measurements of the temperature sensors will be compared with the sensors measurement for data analysis after recovering the payload. 8. Internal Temperature Sensor: The internal temperature will be monitored by the flight control computer using an internal temperature sensor. The internal temperature readings will be compared with a particular temperature value set in the program of the microcontroller. If the internal temperature readings are equal than the temperature value set in the microcontroller then the microcontroller will send the turn on command to the heater. 9. Sun sensor: Each one of the five sun sensors that will be located in the different sides of the ADS payload will sense the direction to the Sun. 10. GPS: Will provide the position and altitude information of the platform. 11. Heater: The heater will be controlled by the microcontroller to turn on when the internal temperature reach a particular temperature value specified in the microcontroller program and will receive the turn off instruction when the measured internal temperature is higher than the temperature value set in the microcontroller. 12. Mechanical System: The mechanical system is the enclosure of the subsystem components for the experiment. An aluminum 2014-T4 alloy will be use to construct the mechanical system to isolate the circuit components and protect them from the temperature. 13. Thermal Subsystem: The main purpose of the thermal system is to maintain the internal temperature of the payload within the operating temperature range of the sensors and the components at the same time protect them from any impact that can damage it. 14. Data Gathering and Analysis: After recovering the payload the Atmega 1280 has to execute post flight software to gather the sensors measurements stored in the SD card. Once the data is extracted from the Atmega 1280 several plots will be made to analyze the data and made new recommendations. 15. Telemetry & Command Interface: HASP has the ability to transmit data through telemetry; data will be transferred from ADS payload to the platform, then ground station thus allowing constant monitoring. Team EQUIS 11

19 4.2.2 Component Interfaces The ADS system design of Figure 3 shows the interface between the components. The Atmega 1280 is going to be connected with each one of the sensors output. The power subsystem will have connection with the sensors subsystem and with the Atmega 1280 to deliver power to both subsystems, thus permitting for each sensor to receive the power necessary to operate to specifications by the use of four DC to DC converters. From section to is explained in greater detail the power interfaces and consumption. The interface between the Magnetometer, Accelerometer, Gyroscope and GPS is a direct connection to the digital pins of the microcontroller. The Sun sensors and the external temperature sensor will be connected to a signal conditioner then to the analog pins of the Arduino. A direct connection interface with the SD card memory module is used to receive the information that should be stored. Also the information stored will be sent to the ground station through serial/usb connector cable on the platform which will be received on the ground. The serial/usb will allow any communication between the payload and HASP such as uploading commands Traceability Main Mission Goal Objectives Requirements Implementation The mission of this project is to create an Altitude Determination System (ADS) for a balloon borne experiment that would stand the hostile conditions, to obtain data through the launch time and during flight from various sensor measurement that would help for future analysis, allowing a better understanding of the dynamics experimented by the HASP platform. Determine the orientation of the payload since the launch and during flight. By combining the acceleration and movements in pitch, yaw and roll. Monitor Internal and external temperatures Determine the position of the payload with respect to magnetic field and Sun To determine the latitude and longitude coordinate. Include an accelerometer and gyroscope determines the orientation. Include a Si Diode (Temperature Sensor) for external temperature and use the accelerometer's temperature sensor for internal temperature Include a Sun sensors and Magnetometer to determine position with respect to the sun and the earth magnetic field. Include a Global Positioning System. Acquire a gyroscope and a accelerometer, develop proper circuitry and connections to allow proper measurements. Design and implement necessary circuitry to allow correct function of the external temperature sensor. Execute a proper code to retrieve the data from the accelerometer internal temperature sensor. Realize a circuitry that permits access to the proper measurements of the photodiodes through a signal conditioner. Implement a proper interface between the magnetometer and the microcontroller. Establish proper connections and employ the proper code to retrieve the GGA string. Team EQUIS 12

20 Store data for further analysis Include a SD memory card Figure 5 Traceability Matrix Develop a voltage divider circuit to allow appropriate power requirements and connection to the SPI pins on the Arduino. Figure 4 shows the table of the traceability matrix for the ADS experiment. 4.3 Electrical Design Figure 6 Electrical Design Subsystems As shown in Figure 5 the electrical design consists of five main subsystems; the power subsystem (HASP), the heater subsystem, the flight control subsystem, the sensor subsystem and the telemetry & command subsystem. The power subsystem is based on the power provided by HASP. The sensor subsystem has analog and digital devices; therefore, analog and digital signal outputs from the sensors will be received by the microcontroller as shown in Figure 5. The flight control subsystem will gather data from all instruments and will control the heater. The sensors subsystem consists of a heater device and eleven sensors; a three axis accelerometer, a three axis gyroscope, a three axis magnetometer, five sun sensors, an internal and external temperature sensor and a GPS. Finally, the Telemetry & command subsystem will allow the communication of data between HASP and the payload. Team EQUIS 13

21 4.3.1 Sensors Subsystem Legend Sensors Analog Outputs Sensors Digital Outputs ON/OFF Control Instructions Heater Subsystem GPS Sensors Subsystem 3-Axis Magnetometer 3-Axis Accelerometer Internal and External Temperature Sensors Five Sun Sensors 3-Axis Gyroscope Flight Control Subsystem Atmega 1280 Arduino SD Card Memory Analog Inputs Digital Outputs Digital Inputs Figure 7 Electrical Subsystem The electrical subsystem is shown in figure 6. The SCA3000 is a digital three axis accelerometer with an acceleration range of +2 g. The digital I/O voltage of the SCA3000 is from 1.7 V up to 3.6 V. The digital outputs range of the SCA3000 accelerometer is from 2.35 V up to 3.6 V. Team EQUIS 14

22 Figure 8 Accelerometer Sensor The sensing element of the SCA3000 three axis accelerometer consists of three acceleration sensitive masses. A capacitance change will occur due to acceleration and will be converted into a voltage change in the signal conditioning ASIC. The element s measurement coordinates are rotated 45º compared to the conventional orthogonal X,Y, Z coordinate system, due to its mechanical construction. The sensing element is interfaced via a capacitance-to-voltage (CV) converter. The SCA3000 includes an internal oscillator, reference and non-volatile memory that enable the sensor s autonomous operation within a system. The SCA3000 includes a temperature sensor; therefore, temperature stability can be reached by using the temperature information from this sensor. Since the ADS payload will include three internal temperature sensors, the accelerometer temperature sensor can be one of the three internal temperature sensors. Figure 9 SCA3000 Three axis accelerometer pins The figure 8 shows the pins for the SCA3000. The accelerometer can receive instructions from the Arduino microcontroller by connecting an output pin of the Arduino to the accelerometer Master Output Slave Input (MOSI) pin. The Arduino represents the Master and the accelerometer the slave. The SCA3000 can be set to operate at low power to save system level power consumption by using the Motion Detection (MD) mode. When the Free-Fall Detection (FFD) is enabled normal Team EQUIS 15

23 acceleration is available. In general, the accelerometer provides several modes and the complete list of modes for the accelerometer, the gyroscope and the magnetometer that will be use for the ADS experiment will be provided in the Critical Design Review (CDR) document. The full scale range of the ITG-3200 three axis gyroscope is º/s. The ITG-3200 has digital (I 2 C) outputs. The operating voltage range of the gyroscope is from 2.1 V up to 3.6 V. Figure 10 Gyroscope Sensor The EQUIS team selected the ITG-3200 three axis gyroscope for the ADS experiment. The ITG has both analog and digital outputs. The analog output of the gyroscope requires a signal condition to adjust the output signals range up to the level of the ADC input range. The pins to receive the digital instructions are shown in the following figure. Figure 11 ITG-3200 Three Axis Gyroscope pins The ITG-3200 is a three independent vibratory Micro-electro-mechanical System (MEMs) gyroscope sensitive to Coriolis forces. The ITG-3200 can detect rotational rate in three axis; the Team EQUIS 16

24 X (roll), Y (pitch), and Z (yaw). A rotation about any of the sense axes of the gyro must be induced to create a Coriolis Effect, which causes a deflection that is detected by a capacitive pickoff. The MEMs device apply the following process to the detected signal; amplification, demodulation and filtering that produces a voltage that is digitalized with an ADC. The converted voltage is proportional to the angular rate. The three axis gyroscope has a range of degrees per second (º/s). The ADC output rate is capable from 3.8 up to a maximum 8,000 sample per seconds, while allowing a wide range of cut off frequencies due to its capability of a user selectable pass filter. The micromag3 three axis magnetometer has a low power: draws < 500 µa at 3 VDC and have a field measurement range (3 VDC at Rb = 43Ω) of µt. The micromag3 has a fully digital interface: SPI protocol at 3V. Figure 12 Three Axis Magnetometer A micromag3 three axis magnetometer has been selected for the earth s magnetic field measurements of the ADS experiment. The three axis magnetometer sensor operates as an oscillator circuit composed of the internal sensors, bias, resistors, digital gates and a comparator. Only one sensor can be measured at the time. The user sends a command byte to the MicroMag3 through the SPI port specifying the sensor axis to be measured. A 1N457 silicon diode is selected to be the external temperature sensor and will be positioned outside the ADS payload to monitor the exterior temperature. Figure 13 shows a 1N457 silicon diode. The temperature measurements will allow determining the temperature that the ADS payload will be exposed during flight. Figure 13 1N457 Silicon diode Team EQUIS 17

25 The payload will be exposed to different temperatures during flight and as the temperature decrease the voltage provided by a battery can decrease, such as the batteries that HASP contains; therefore, there will be a change in voltage and in current when the temperature changes. Consequently, There will be three variants; temperature, voltage and current. The current can be adjusted to be constant using an LM334 constant current source; as a result, only the voltage and temperature will vary. Hence, the changes in temperature can be measured by relating the changes in voltage with respect to temperature. Details of the calibration will be explained in calibrations. Figure 13: Schematic of the external temperature sensor The Figure 13 shows the schematic for the external temperature sensor. The U100 LM334 represents the constant current source. The AD822 operational amplifier is selected for the signal conditioning for the external temperature sensor. The ADS experiment will include a heater inside the payload to maintain the temperature within the operating temperature range of all the instruments. The internal temperature sensor (ITS) will allow monitoring the temperature; as a result, the microcontroller can determine the appropriate time at which the heater should receive the turn on and off control instruction. The SCA3000 accelerometer allows to obtain readings of temperature; therefore, the internal temperature measures will be provided by the accelerometer. In addition, the payload will incorporate five sun sensors using photodiodes; consequently, a photodiode will be positioned on all the sides, with the exception of the bottom side. Finally, GPS will be included in the payload. The Sun Sensor is a device or circuit that receives and detects visible light, infrared and/or ultraviolet energy to determine the direction of the payload with respect to the sun. This is obtained through an analysis of the measurements gathered through the difference of sensors e.g. photodiode [Ref.13]. The photodiode, the bipolar phototransistor and the photo FET (photosensitive field effect transistor) are the most common types of photo sensor. The difference between these devices and the diode, bipolar transistor and the field-effect transistor, is that the packages of the sensors Team EQUIS 18

26 previously mentioned have transparent windows that allow radiant energy to reach the junctions between the semiconductor materials inside. Furthermore bipolar and field-effect phototransistors provide amplification in addition to their sensing capabilities. Figure 14 Si S1336 Photodiode Figure 13 represents the possible photodiode that could be selected for the sun sensors system. This photodiode is a possible candidate because it has a UV to near IR photometry, it has a high sensitivity, it meets designing parameters and it is in a metal case which protects it from different conditions Sensor Interfacing The sensor connections with the Arduino are as shows the following table; Sensors Analog and Digital pins of the Arduino Mega Sun Sensor 1 (Photodiode) ADC0 pin 97 (ADCL) Sun Sensor 2 (Photodiode) ADC1 pin 96 (ADCL) Sun Sensor 3 (Photodiode) ADC2 pin 95 (ADCL) Sun Sensor 4 (Photodiode) ADC3 pin 94 (ADCL) Sun Sensor 5 (Photodiode) ADC4 pin 93 (ADCL) Internal Temperature Sensor ADC5 pin 92 (ADCL) External Temperature Sensor ADC6 pin 91 (ADCL) Not Connected ADC7 pin 90 (ADCL) Not Connected ADC8 pin 89 (ADCH) Not Connected ADC9 pin 88 (ADCH) Not Connected ADC10 pin 87 (ADCH) Not Connected ADC11 pin 86 (ADCH) Not Connected ADC12 pin 85 (ADCH) Not Connected ADC13 pin 84 (ADCH) Not Connected ADC14 pin 83 (ADCH) Not Connected ADC15 pin 82 (ADCH) Three Axis Accelerometer PA0 pin 22 (JP1 pin 16) Three Axis Gyroscope PA1 pin 23 (JP1 pin 15) Team EQUIS 19

27 Three Axis Magnetometer PA2 pin 24 (JP1 pin 14) Heater PA3 pin 25 (JP1 pin 13) Not Connected PA4 pin 26 (JP1 pin 12) Not Connected PA5 pin 27 (JP1 pin 11) Not Connected PA6 pin 28 (JP1 pin 10) Not Connected PA7 pin 29 (JP1 pin 9) Not Connected PC7 pin 30 (JP1 pin 8) Not Connected PC6 pin 31 (JP1 pin 7) Not Connected PC5 pin 32 (JP1 pin 6) Not Connected PC4 pin 33 (JP1 pin 5) Not Connected PC3 pin 34 (JP1 pin 4) Not Connected PC2 pin 35 (JP1 pin 3) Not Connected PC1 pin 36 (JP1 pin 2) Not Connected PC0 pin 37 (JP1 pin 1) Not Connected PD7 pin 38 (JP2 pin 16) Not Connected PG2 pin 39 (JP2 pin 15) Not Connected PG1 pin 40 (JP2 pin 14) Not Connected PG0 pin 41 (JP2 pin 13) Not Connected PL7 pin 42 (JP2 pin 12) Not Connected PL6 pin 43 (JP2 pin 11) Not Connected PL5 pin 44 (JP2 pin 10) Not Connected PL4 pin 45 (JP2 pin 9) Not Connected PL3 pin 46 (JP2 pin 8) Not Connected PL2 pin 47 (JP2 pin 7) Not Connected PL1 pin 48 (JP2 pin 6) Not Connected PL0 pin 49 (JP2 pin 5) Not Connected PB3 pin 50 (JP2 pin 4) MISO Not Connected PB2 pin 51 (JP2 pin 3) MOSI Not Connected PB1 pin 52 (JP2 pin 2) SCK Not Connected PB0 pin 53 (JP2 pin 1) SS GPS TX GPS RX PE0 pin 2 (PWML pin 0) RX0 PE1 pin 3 (PWML pin 1) TX0 Table 1 Pins for the Arduino Atmega All the sensors will have an interface between each one of the outputs and the Atmega The Atmega 1280 board has an internal analog-to-digital converter, which provides 16 analog inputs. This ADC is in charge of converting the analog signals that to digital signals that then can be process and stored on the Atmega 1280 as digital quantities. There are various types of sensors such as analog and digital sensors. The photodiode used falls in the category of analog, thus needing an analog to digital converter (ADC) which is internally in the Arduino. The difference between the analog and digital sensors is that the analog sensor has an output signal that is related to the angle between the sun and the payload. On the other hand, digital sensors produce a constant signal whenever the sun is in the field of view. Team EQUIS 20

28 Figure #0 represents the basic circuit of the sun sensor. The sun sensor works by using a photodiode which will receive infrared waves that will allow current to flow in one direction from the collector to emitter. The resistor R1 is to control the current in the photodiode. The resistor RL represents the signal conditioner circuit to obtain a voltage in the range of the microcontroller s analog port. The output of the signal conditioner to the port will range from 0 to 5V. Figure #0 Basic Circuit of Sun Sensor To design the sun sensor a block diagram in Figure#1 is generated to obtain a general ideal of the process that will be developed and also to generate an ideal of what considerations need to be taken into account. 4.80V 5.0V Arduino Photodiodes Amplifier Internal ADC 1.20V Figure #1 Designing process of sun sensors Figure #1 illustrate what we will have for the sun senor circuitry, the component use to detect the light intensities is the photodiode these intensities will be obtain in volts. In this design it is assumed that the when the light intensity (LI) is High it will give of an output of 4.80 V and 1.20 when LI is low. This output will be passing through an amplifier that will condition the input to cover the range of the input of the internal ADC of the microcontroller. The Figure #2 allows an graphical understanding of the output obtain of the photodiode. In the y axis is represented by the voltage obtained from the sensor and the x axis is represented with the amount of light intensity. It can be observed that the output is linear. As the LI is higher also the output voltage will be higher. 0V Team EQUIS 21

29 6 LI vs. Vs 5 Vs (volts) L H 5 6 Light Intensity (LI) Figure #2 Light Intensity vs. Sensor Voltage (Vs) The graph in Figure#3 demonstrates the desired output and permits an idea of the mathematical analysis required to obtain the equations to develop the proper amplifier (signal conditioner). 6 Vs vs Vo Sensor Voltage (Vs) Output Voltage (Vo) Figure #3 Desired Output Voltage Range In Figure #3 its illustrated the desire output, which is that when the output is 1.20 it will be 0 and when it is 4.80 it will be amplified to 5V. To determine this, the equations for the amplifier circuit can be obtained by the use of the slope intercept equation Eq1., where m is the slope and b is the is the y intercept. Team EQUIS 22

30 Eq1. y mx b, Slope Intercept Solving for the slope equation Eq2 with the values from the Figure #3, results in a value of m=1.38. Eq2. m, Slope Using the slope intercept equation to obtain the output voltage (V o ) it is possible to obtain b, since V 0 = 5, m= 1.38, V s = 4.80, resulting that b= Eq3. V mv b Eq. 3 allows obtaining the equation for the signal conditioner in Eq.4 4. Where Eq5 M= 1.38 & Eq6. b= 1.67 After obtain Eq4. it is necessary to assume values of R f & R 1 to meet the values previously calculated and for the R o to obtain the all the values that will permit the proper circuit in Figure# 4. Figure #4 Signal Conditioner The signal conditioner in the Figure 4 a summing amplifier that has a reference voltage and a input voltage from the photodiode. Ro will be a potentiometer for calibration purposes, after obtaining proper calibration it will be replaced by a resistor with the value of the potentiometer s resistance. The amplifier chosen is and AD820, this is selected because it is single source and it has been used successfully in different experiments. The output of the signal conditioner will be connected to the internal ADC of the microcontroller. Five set of circuits of the Figure#4 will be developed, one for each sensor. It is possible that amplifier model could be change to a quad amplifier. Team EQUIS 23

31 4.3.3 Control Electronics The electrical design for the ADS experiment will include an interface between the each one of the instruments and the microcontroller. The microcontroller will retrieve data from each one of the instruments and will send the ON/OFF control instructions to each one of the three heaters. Consequently, the software code for the Arduino Atmega 1280 microcontroller requires including the control instructions for the heater and additional lines to receive the data from the instruments. The following figure shows the interface between the heater and the Arduino. Figure 15 Control Electronics A fuse is included at the inputs of each one of the DC-to-DC converters to isolate the circuit as a safety precaution. The ADS experiment will include three heaters as show the following figure. Each heater needs 20 VDC and HASP provides 30 VDC; therefore, a 20 V DC-to-DC converter is required to power the heaters. A DC-to DC converter is an electronic circuit that converts a source of direct current (DC) from one voltage level to another. The DC-to-DC converter is a class of power converter [Ref. 15]. The optical coupler subsystem that is shown in the previous figure represent three relays/switches that will be controlled by the Arduino microcontroller. The Arduino will send the ON/OFF control instructions to the relays/switches creating an open circuit when the switch is open by the OFF instruction that will turn off the heaters. When the Arduino send the ON instruction the relay/switch will close allowing that the power from the DC-to-DC converter reaches the Heaters to turn on the heaters. The Arduino will receive temperature readings from three internal temperature sensors. Each temperature sensor will be located next to a particular heater. The following figure relates each temperature sensor with a Heater by using the Area Label. Team EQUIS 24

32 Figure 16 Heaters Subsystem Team EQUIS 25

33 4.3.4 Power Supply Figure 17 Power System Diagram Figure 19 shows the power subsystem diagram. The power source will be the HASP platform and will provide VDC. The ADS experiment requires the use various voltage regulators, the selected is a LM317 due to its output capacity that ranges from 1.2 V up to 37 V. Also a voltage regulator will be used supply the voltage necessary for the heater. The arduino has an input voltage range from 6V to 20 V; however, the voltage regulator of the Arduino Board may overheat and damage the board when is using more than 12 V and if the supplied voltage is less than 7 V, the 5 V pin may supply less than five volts and the board may be unstable [Ref. 14]. Therefore, the Arduino will receive 9 V from a DC-to-DC converter as shows the figure above. The voltage regulator located at the Arduino Board will adjust the 9 V to the 5 V that the Arduino requires to operate. Team EQUIS 26

34 4.3.5 Power Budget The HASP- High Altitude Student Platform provides power for small and large payloads. For small payloads the supplied voltage that HASP platform provides is VDC. Sensor Required Required Current Voltage (V) Three Axis Gyroscope ITG µa Standy & 6.5mA Photodiodes S Each µA Three Axis Accelerometer SCA µa Three Axis Magnetometer Micromag µa Temperature Sensor 1N ma Heater A Arduino Mega 9 50mA GPS Lassen iq ma LM334 3 terminal adjustable current source 20 5mA SD card circuit mA Max current consumption of components mA Table 2 Power Requirements (Amperes) Table 2 demonstrates the current and voltage required for each of the electronic components. Since the currents supplied is 500mA and none of the components require more than what is supplied by HASP, therefore the current consumption criteria is meet. 4.4 Software Design The data will be retrieved from the sensors and thus the software will be designed to store the data in the SD card and send it to ground by telemetry Data Format & Storage The total bytes required for the flight will be determined by the quantity of the sensors and the size of the integer and character values that will be used times the hours the payload may remain in flight. The integer size on the Arduino is of two bytes and the character size is of one byte. The system has eleven sensors; three axis magnetometer, three axis gyroscope, a GPS, three axis accelerometer, five photo diodes and two temperature sensors (internal and external). Each sensor with three axes will have six bytes, two bytes for each photo diode, with the total of ten bytes, ninety bytes for the GPS, and two bytes for each of the temperature sensors (external and internal) with a total of four bytes. The time stamp consists of six bytes, two bytes for the hours, two bytes for the minutes and two bytes for the seconds. In total we have 128 bytes for the whole system. The sample rate will be of 5 seconds. The data will be received 12 times each minute. The flight will be of approximately of 20 hours, the total bytes of data are 1,843,200. A high capacity SD card will be more than enough to store the data. Team EQUIS 27

35 The GPS data that will be stored in the 90 bytes is global positioning system fix data (GGA) string, time stamp, latitude, North/South position, longitude, East/West position, fix quality, number of satellites being tracked, horizontal dilution of position, altitude meters above mean sea level, height of geoids, and checksum data. The total bytes required for the system is: Byte Description 1 6 Time Stamp 7 96 GPS Accelerometer: X axis Accelerometer: Y axis Accelerometer: Z axis Gyroscope: X axis Gyroscope: Y axis Gyroscope: Z axis Magnetometer: X axis Magnetometer: Y axis Magnetometer: Z axis Sun Sensor Temperature Sensor (external) Temperature Sensor (internal) Table 3: Bytes Description The data will be saved on the SD card. The SD card will be formatted FAT 16 before use. That will be done on a computer. The program will use a library to create an object to initialize the SD card, set the volume of where the file will be located, check for errors, etc. The data will be saved as a comma separated values (CSV) format. Shown below is a code example of how the file will be created and saved. Saving to SD card (code example): Create a new file (CSV) // the files won t be erased Char file[] = TEST000.CSV ; // last three characters are zeros int j = 0; For(i = 0; i < 1000 ; i++) // creates up to 999 files, will be changed if necessary { // file[4], file[5] and file[6] are the position of the zeros, if the previous file name exists //the zeros will be changed file[4] = i/ ; // changes zero, if file created before exists if (j >= 100) // up to 100, to change the zero on the fifth position Team EQUIS 28

36 { j = 0; } If(i >= 100) // use j variable that will only get to 100 { File[5] = j/ ; j++; } If( i < 100) // if i is lower than 100 use i instead of j { File[5] = i/ ; } file[6] = i % ; // changes zero, if file created before exists, constantly changing } Write header // runs once Print(,HEADER ); // comma before Write data // loop Print(, ); // commas before to use the CSV format Print(data); Flight Software The system will have a data rate of approximately 204bps. The Arduino uses a USB cable or the receive (RX) and transmit (TX) lines. The software will use the Serial library that is included with the Arduino environment. An example of the serial functions and their use: Serial.begin(); // sets the baud rate Serial.available(); // Reads from serial port and stores it on a buffer Serial.write(); // writes binary data to the serial port. The data is sent as a byte or a series of //bytes Serial.read(); // reads incoming serial data The commanding will also use the serial library, specifically the serial.available() function which checks if data has been received through serial, the RX line. Team EQUIS 29

37 An example of the code: Char command[] = ; // initialize the variable to save the data received Void setup() // a must have function, initializes { Serial.begin(1200); // determine the baud rate Serial.println( Enter command ); // an outpt in ASCII } Void loop() // a must have function, infinite loop { If (Serial.available() > 0) // if data is received proceed { Delay(5); // delay for 5 milliseconds before saving data to variable For(int i = 0; i < 1; i++) // receives only one character { Command[i] = Serial.read(); // Read from the serial buffer, and save Delay(10); // delay for 10 milliseconds after saving data // this delay is needed if receiving more than 1 characters } Serial.flush(); // reset the serial buffer If((command[0] == S ) (command[0] == s )) // compares { Serial.println( Turn OFF ); // Output in ASCII } Else if ((Command[0] == o ) (command[0] == O )) // compares { Serial.println( Turn ON ); // Output in ASCII } Else if ((command[0] == r ) (command[0] == R )) // compares { Serial.println( RESET ); // Output in ASCII } Else // if the value is not equal to any of the above { Serial.println( Do nothing ); // Output in ASCII } } } Team EQUIS 30

38 To run the code on the Arduino, the code must have two main functions: Void setup() o Will run once, when the Arduino starts o The baud rate is determined o The pins are set (input or output) Void loop() o Is an infinite loop, starts running when setup function finishes o Runs the rest of the code The flowchart explained briefly: INPUT Receive time stamp Receive data from sensors GPS Sun sensors Magnetometer Gyroscope Accelerometer Internal temperature External temperature PROCESSING If temperature LOW then Turn heater ON If temperature HIGH then Turn heater OFF OUTPUT Send through serial downlink Save to SD card Team EQUIS 31

39 Figure 18 Flight software control flow chart Team EQUIS 32

40 4.5 Thermal Design In order to maintain the electronic components inside the payload in their operative optimal temperature ranges and to reduce any measure errors (i.e. bias) that could appear due to temperature changes, a thermal control system will be adapted inside the payload. During this experiment the payload will be submitted mostly to the stratosphere thermal environment. This environment can reach temperature as low as -80 C and as high as 90 0 C and winds that fluctuate between 20mph to 100mph. These are the atmospheric conditions that would lead to drastic changes in the environment surrounding the equipment, producing damage in these devices. The thermal operating ranges of the main components inside the payload are presented in the following table: Electronic equipment Arduino Gyroscope Accelerometer Magnetometer Temp. Sensor GPS -40 to 85( 0 C) -40 to 85( 0 C) -40 to 85( 0 C) -20 to 70( 0 C) -65 to70( 0 C) -40 to 85( 0 C) The table illustrates that the device that will limit the temperature parameter inside the payload is the magnetometer. Considering the temperature range of the magnetometer, a temperature of 10 0 C is selected as the lowest temperature allowed for the medium inside the payload. A heater will be adapted inside the payload to compensate the heat loss from the payloads inside to the environment. An evaluation of the overall thermal dissipation from the inside to the external environment was made to determine which heater is appropriate for this application. The worst case scenario was considered in the analysis and the structure is study at 100,000ft above sea level. In addition is important to mention that the materials and their dimensions were determined before the analysis. This analysis was performed considering steady state conditions (i.e. no change at a point with time) and the coldest environment that the equipment can experience. The assumption of steady state conditions helps to solve the heat transfer problem without involving any differential equations or temperature distribution and permits the use of the thermal resistance concept. The concept of thermal resistance is used mainly for the case of steady heat conduction through walls or any solid material. Thermal resistances for the analysis of convection and radiation had been developed considering Newton s law of cooling and the Stefan-Boltzmann law, and the expressions obtained can be applied only at the boundary of the solid material where conduction occurs. The expressions that define the thermal resistances of the different heat transfer mechanism are the following: Conduction: Convection: Radiation: R cond = L/kAs R conv =1/h conv As R rad =1/h rad As To simplify even more the heat transfer analysis performed the overall heat transfer coefficient term is used so that the heat loss by the inner space can be expressed using only one expression Team EQUIS 33

41 instead of considering each heat transfer mechanism separately. This expression is analogous to Newton s law of cooling and has the following form: Q=UA T where: U = overall heat transfer coefficient UA = 1/R total Is important to mention that convection was examined in this analysis so that all the possible heat transfer mechanism that could occurred in the stratosphere can be considered. The stratosphere contains very dry air which is the medium that allows the presence of convection in that layer of the earth atmosphere (National Center for Atmospheric Research, 2009,. 4). Natural convection was considered in the payloads inside, but at the outside forced convection is the mean of heat transfer considered. For the analysis of natural convection at the payloads inside the following terms were determined: β = (1/T) ideal gas Gr L = ^ ѵ^ Pr = Nu = represents the variation of the density of a fluid with temperature at a constant pressure. represents the ratio of the buoyancy force to the viscous force acting on the fluid. (Grashof number) represents the ratio between the velocity boundary layer and the thermal boundary layer form in the contact surface between a fluid and a solid surface. (Prandtl number) Dimensionless convection heat transfer convection (Nusselt number) For the analysis of external forced convection at the payloads outside the following terms were determined: Nu = Re = ѵ represents the enhancement of heat transfer through a fluid layer as a result of convection relative to conduction across the same fluid layer. represents the ratio between the inertial forces and the viscous forces. Used to determine the fluid flow regime. (Reynolds number) Team EQUIS 34

42 The terms used in the convection analysis are mainly used to define the convection heat transfer coefficient h which can be determine for different cases once the Nusselt number had been evaluated. The radiation in this analysis is only considered in the payloads inside and this is because the structure in this analysis is examined in the coldest environment that can occur only during night. For the evaluation of radiation in the payload the following equation was used: Q rad eq = є FR4 σa eq (Ts 4 - T in 4 ) A summary of the analysis performed is presented in the appendix. The results obtained show that for this application a 5 W heater is recommended so that the loss encountered with this condition could be supply. The types of heaters considered for this application are thermo-foils heaters. These types of heaters are ideal for applications with space and weight limitations. To select the desire heater model the following parameters were determined: Maximum power provided: P=VI=(30V)(0.5A)= 15W Therefore the 5W of heat loss are within the maximum power provided by the HASP platform. Heater Selection Desired Temperature 0 0 C Power Required 5W at 20V Maximum Heater Size 11cmX11cm Ideal Resistance (20V) 2 /5W = 80Ω Mounting Method Acrylic Pressure Sensitive Adhesive Model Chosen HK5448R83.8L12B Effective Area 8.90in 2 Actual Power (20V) 2 /83.8 Ω=4.77W Watt Density 4.77W/8.9in 2 =0.54W/in 2 Max Watt Density 12W/in 2 at 10 0 C Wattage Ok? Yes(0.54<12) Leadwire Current 20V/83.8 Ω = 0.24A Current Ok? Yes(0.24<0.5) Table 3 Heater Parameters The heater evaluated in this analysis is fabricated by the MINCO Company. This heater is one of their standard kapton heaters. The dimensions of the model chosen in this analysis are illustrated in the following figure: Team EQUIS 35

43 x Heater x=6.35cm y=10.2cm Figure 19 Heater Dimensions The model of the heater was chosen considering the maximum heater size, the mounting method and the ideal resistance value. All the model s parameters were provided by the MINCO Company via web. 4.6 Mechanical Design y The first step in the mechanical design was to determine which type of aluminum alloy was the appropriate for this application. This alloy was selected considering that the structures have to withstand thermal and mechanical loads during the various mission phases. Weight was another important parameter considered for the material selection, since there is a maximum value of mass that can t be exceeded in this project (i.e. 3kg). The materials examined are presented in the following table: Aluminum ρ (Mg/m³) (Mpa) (Mpa) K (W/m C) 6066 T T T Table 4 Alloy Selection The aluminum 2014-T4 was the alloy selected for this application because of its advantages compared with the other materials. As can be observed from the table this material posses the best combination of thermal and mechanical properties. In addition Kapton was the insulating material selected for this project. Kapton is a plastic film that has the ability of maintaining its physical, electrical, and mechanical properties over a wide temperature range. This characteristic and its low conductivity is what make this material a good selection for heat transfer problems, specifically for applications with space limitations. The Kapton thermal conductivity is of 0.12 W/m 0 C. The payloads geometry selected is basically a cube; this symmetrical geometry facilitates the structural and thermal analysis of the payload as well as the computation of any unknown dimension. The cube structure made of aluminum will be form from a solid cube of material to construct a monoblock structure. The monoblock structure is ideal to maintain the structure stresses within a considerable range during the payloads ascent External Structure Team EQUIS 36

44 The external structure is basically a cube used to provide the first thermal and mechanical protection to the equipment inside the payload. It s a box built with Kapton insulation in its core and aluminum skin at its outside. The external aluminum structure is constituted of a monoblock arrangement that includes the sides and bottom of the cube. The top cover of the external structure is a separate part of the rest of the structure. An illustration of the payloads external structure is the following: Solid structure Connector for the cover Mounting Plate Figure 20 Isometric view of the external structure As can be observed from the figure to attach the cover to the rest of the external structure four L shape couplers will be used. Also, is important to mention that five sun sensors will be attached to the external wall of the payload s structure. These sun sensors will be positioned in the sides and top walls of the external structure and they will be specifically place in the center of the walls. In addition, to attach the payload to the mounting plate provided, four L-shape couplers will be used. Figures that illustrate in more detail these two points are the following: Sun Sensor L-Shape Couplers Figure 21 Sun sensor positioning and L-shape couplers The mounting plate provided, include wiring for the electrical and data connections. For that reason a hole will be made in one of the sides, so that all the electrical cabling can be route to the Team EQUIS 37

45 payload s inside. All the important dimensions and drawing views of the external structure are provided in the following figure: Figure 22 External Structure Drawing Internal Structure The internal aluminum case provides support for the internal components which are in this case electronic devices. This internal aluminum case is surrounded by an insulation of Kapton material. The internal structure has the same monoblock -cover arrangement than the external structure, but they differ in their dimensions. An illustration of the payload structure and its electronic devices is the following: Team EQUIS 38

46 External Structure Kapton Insulation Internal Structure GPS with antenna integrated. Other Sensors Figure 23 Top view payload structure Is important to mention that a set of holes will be made in the internal structure to route the electrical cables to the payload s inside. These cables are necessary for the electrical and data connections. The following figure illustrates the position of these holes in the internal structure: Cover hole Side holes Figure 24 View of punctures on structures The internal structure will be protected by the external structure which withstand all the loads directly and by the Kapton material that enclose this part of the payload. All the important dimensions and drawing views of the external structure are provided in the following figure: Team EQUIS 39

47 Figure 25 Internal Structure Drawing Mass Budget In this project the maximum mass allowed for a small payload is 3kg. The approximate value of the total mass of our payload is g. This total mass was determined using values obtained from the Solidworks (CAD software) mass properties feature and considering approximated values for the PCBs and microcontroller (i.e. Arduino). The masses obtained directly from the software were the kapton insulation, the external structure and the internal structure. The value of the PCBs were base in past experiment of this kind. To eliminate any bias introduced by the presumed mass values of the PCBs and to work in conservative range of values, the mass value of the PCBs observed in previous experiment was increased from 95g to 100g for each board. Also the mass value of the microcontroller board was approximated to 100g. The following table is used to show the payload s main components and their masses. Team EQUIS 40

48 Component Mass (grams) External Structure Internal Structure Kapton Insulation Two PCB (approximately) 200 Arduino 100 Total Mass Table 5 Mass Budget It can be observed considering the total mass value, that even with the increment of mass in the PCB values we still have g that can be used for any component that is not considered in this moment. Additional Mechanical Feature (optional). The Boston University is working in cooperation with scientists and engineers from the United States Air Force Research Lab, in the designing of an extendable boom for a cube-satellite, as well as the satellite housing. The boom and its mechanism is used to deploy a magnetometer at a specified distance with respect to the structure and this magnetometer will be measuring the electric field in that zone. In the case of satellites the electric field measurements help scientists to have a better understanding of space weather. The University of Boston will like to test their design in a real working environment and this project will represent a good opportunity for that. Consequently, we are considering adding their design in our structure to help them with the testing of their design if that modification can be included within our time limitation. The total cost of this feature is approximately $1, Payload Development Plan Prior to the payload fabrication in the development phase various task should accomplished such as verifying calculations, designs, prototyping and testing. This is done to resolve issues that can only be fulfilled through testing results made to the prototype. The areas that will be mostly focused on will be software, electrical and mechanical. On the software, tests and measurements will be obtained to identify the size of files developed when storing data to the SD card. It is necessary to establish the processing time of the program which is obtained by prototyping. Verify that all interfaces, connections, and programming will work as plan, otherwise apply the suitable contingency plan previously develop. Simulate or develop a model to ensure that the microcontroller can communicate through proper connections that will be used for telemetry purpose. The electrical section will involve prototyping the signal conditional for the external temperature sensor and the sun sensor to verify the output response and perform calibrations. Develop the complete design of the print circuit board and develop a prototype. Perform the prototyping and calibration for the temperature sensor. Proving that the degrees calculated with the sun sensor data are approximately identical to the sun positioning measured. Team EQUIS 41

49 Mechanical prototyping will be performed to ensure optimal quality in the payload s structure. Test will be performed to ensure the thermal characteristics calculated are similar to the characteristics and behavior physically. Determine the period of time that the heater takes to reach the desired temperature. For the prototyping parts are necessary to be order with a degree of urgency thus requiring a organize list of parts with sequence. The order the part must be order is: 1. Arduino 2. Phototodiode 3. GPS with antenas 4. Magnetometer 5. Gyroscope 6. Accelerometer 7. Slot and SD Memory Card SDHC compatible 8. Payload Structure materials More detail on order or purchase of parts can be observe in the table in section Payload Construction Plan When we have the component dimensions we will start sketching for a box that meets the size and making an estimate of the weight. After the sketching we will start making possible real size models and testing then. When we finish with the result we will make the final box and test it with all of the components. 6.1 Hardware Fabrication and Testing The hardware fabrication will be worked in parallel since the measurement of the payload enclosure is known, allowing the members to start their respective sections. There will be various tests done in the mechanical section such as thermal, impact, stress, strain and vibration. On the electrical part there will be test in current and power dissipation, verifying that circuits work properly. In the software development there will be a test of all sensors to make sure it is giving and storing the proper data. The parts and components will be ordered as soon as possible to prevent shipping delays. If a delay were to happen the team will work on other areas accordingly until the parts are received to reduce downtime. 6.2 Integration Plan The payload will be integrated in steps. The first is to place the heaviest parts at the lowest part of the box to prevent or minimize the swing of the box and preventing the components to come loose inside the box. Next assembling the components with the internal structure outside the box and the put it inside is intended. This is done to make sure every cable is no interfering with anything and that the components are place properly. The board with the gyroscope will be place in the middle of the rack to have an excellent reading with it. Team EQUIS 42

50 6.3 Software Implementation and Verification The first step is to design the program to meet the Arduino MEGA needs and run it for testing purposes. This phase will be executed as follow: Determine the required bytes Flowchart analysis Sensor calibration Display sensor data during testing Verify Arduino s processing time 6.4 Flight Certification Testing The testing will be done by: Styrofoam to hold the sensors in place Full system test o Payload shock testing o Payload vacuum testing Electrical system testing o PCB board continuity test o Cable connection testing o Electronic components test (IC- integrated circuit testing) 7.0 Mission Operations Equipment final checkup Complete system integration Place the equipment on the launch site Ground platform tracking Data analysis o Data retrieval o Data conversion o Data plots o Plot analysis 7.1 Pre-Launch Requirements and Operations A successful flight testing in simulated environment (similar to that were the payload is going to be, such as vacuum and thermal testing) will be done to make sure that the sensors are working as expected. The Arduino also needs to boot up and run the pre-flight software to ensure the memory will be clean, in other words with 0 data in the SD card. Team EQUIS 43

51 7.1.1 Calibrations Proper sensor calibration, with the recommended procedure suggested by the manufacturer, needs to be done to all of the sensors in the ADS payload. The SCA3000 Three axis accelerometer sensor will have an output response when a capacitance change occurs due to acceleration. Once the accelerometer sensor arrives the prototype for the accelerometer sensor will be implemented to perform the required tests. The accelerometer prototype will be exposed to at least three different acceleration positions on each one of its axes to determine the output response of the sensor at that particular acceleration. The next step is to assign a digital value to each one of the selected outputs response of the accelerometer to obtain a linear equation that will relate the outputs with the digital values. The ITG-3200 Three axis gyroscope sensor will measure the rate of rotation; as a result, the sensor will have an output response when sense a change in rotation in each one of the axis. Each axis will be rotated at approximately the same rate of rotation to have a similar calibration for each axis. After rotating each axis at a particular rate of rotation and obtain the output response, a linear equation can be created to relate the digital values with the gyroscope output response. The SEN Thee axis magnetometer can be calibrated by measuring a generated magnetic field Pre-Launch Checklist Make sure all the instruments are working properly Realize a final check of the prototype comparing it with the schematics and the simulations Revise the list of test, such as thermal and vacuum test to ensure that all the testing has been accomplish Run all the program and make sure are working properly Bring all the necessary equipment for any emergency 7.2 Flight Requirements, Operations and Recovery The payload will ascend up to an altitude of 120,000 ft for approximately 2 hours and remain there for 16 hours total. The payload and components will have to support temperature conditions of an expected -40 degree of temperature. The system needs the sensors to read all the time to get all the data during the flight. The batteries must be able to support the flight with the minimum required power. If the batteries are not in the minimum required power then the circuit would be in equilibrium with the sensor voltage required for they to work. We need a GPS or any radiofrequency tracking device of high power so we can track the altitude in which the payload is located and to track the retrieval of the payload in the landing site. Team EQUIS 44

52 7.3 Data Acquisition and Analysis Plan The Arduino MEGA requires an SD card memory module for the storage of the data during the flight then we going to download the data using an Arduino MEGA and a serial/usb cable in which is going to be connected from the payload to the computer software but to process the receiving data we need to convert the measurement in ADC counts requiring some equations for the convection of ADC to physical values then we analyze the data of the atmosphere and flight of the payload Ground Software The ground software requires the following components; a microcontroller (Atmega 1280), a serial/usb cable, the code to download the data and a computer to process the information and download the data from the payload. Graphical analysis 3.2 for plot the data will also be needed Data Analysis Plan The first step on the ground station after the land will be to use the post flight program to take the data from the payload. The software will be prepared using the Arduino IDE. We will implement some software for the calibration of the sensors. The data will be store using ADC counts to represent the measurements from the sensors. In the process of receiving the data, the ADC will convert the measurements into ADC counts that will be stored in the SD card. This requires some equations to convert from ADC counts to physical values. Some uncertainties during the launch are expected as well as in the balloon cut off and in the landing. Expecting those irregularities, make some extra calibrations looking for the most precise measurements. 8.0 Project Management To ensure documentation version control a single team member (A.M. Espinal Mena) will be the team member that will have the latest version of all the documentation. It has been agreed that when a team member do a modification in the PDR, CDR and FRR they have to send an of the modification to this team member, so she can add it to the proper document and send a notification to the rest of the team so they can now which parts has been updated. In addition to the days of class regular meetings in the week to monitor the progress of the project will be scheduled. 8.1 Organization and Responsibilities The EQUIS Team consists of three students of the Inter American University of Puerto Rico, Bayamón Campus. The responsibilities of these students are to work on the following part of their experiment; the electrical/mechanical designs, prototype development, fabrication, integration and testing of the payload that will be launch with a balloon. The name of the students of the EQUIS Team and their tasks are as follows: A. M. Espinal Mena, anaespinal@gmail.com, Calibration Electrical Design, Prototype and Team EQUIS 45

53 E. M. Portilla Matías, Software Design and Implementation F. O. Rivera Vélez, Mechanical and Thermal Design J. I. Espinosa Acevedo, Sun Sensors Subsystem, Risks and Management plan 8.2 Configuration Management Plan Every time a design is made we consult with the team member so we can make a decision of the design and to analyze de advantage and disadvantage of the design before getting approved. 8.3 Interface Control Interface control allows us to maintain a constant monitoring of the Preliminary Design Review (PDR) document at all times. It was established in a group meeting to assign one person in charge of the PDR, this person will be in charge of constantly adding new information to the document and uploading its final version, this way every member will allow access to the file. Also an online group section was established where all document will be uploaded. It was notified to the members that every time a document is uploaded or any change will be made to notify Ana Espinal through a call, which is the person in charge of the PDR, doing so she will be aware of any modifications. Finally a group meeting will be held to verified the complete document and ensure it is the final version before submitting. 9.0 Master Schedule We establish to divide the tasks into different member of the team but some task will be in parallel with the other task, because this will makes us to meet the experiment goal. The amount of days assigned to each task is shown in the Figure 26. Following the work breakdown schedule is essential to realize the mission of this experiment. 9.1 Work Breakdown Structure (WBS) Team EQUIS 46

54 Figure 26 WBS- Work Breakdown Schedule Figure 26 shows the WBS of the EQUIS team to complete the EQUIS payload and instruments on time. 9.2 Staffing Plan A. M. Espinal Mena- Electrical design, Prototype and Calibration E. M. Portilla Matías- Software Design and Implementation F. O. Rivera Vélez- Mechanical and Thermal Design J. I. Espinosa Acevedo- Sun Sensors Subsystem, Risks and Management plan Advisor: Dr. H. B. VO/ E. G. Delgado 9.3 Timeline and Milestones The PDR, CDR and FRR documents have to be submitted by the following deadlines: PDR- Preliminary Design Review due March 5, 2010 PDR Revision- Preliminary Design Review Revision due April 12, 2010 PDR Defense- Preliminary Design Review Defense June 4, 2010 CDR- Critical Design Review due June 16, 2010 FRR- Flight Readiness Review due July 23, Master Budget Team EQUIS 47

55 The maximum capital cost budget required for this project is $5,000. It is also required to have a reserve contingency fund of 10%, which is $500 in this case; therefore, the capital cost budget is $4, Expenditure Plan Integrating the instruments for the ADS experiment requires several components. Table 7 shows the list of electronic components necessary to integrate the instruments, their price and delivery status. Lead time Sensors Part Sales Quantity Status Price ($) Number Company 8 Three Axis Accelerometer SCA3000 Sparkfun 2 - $89.98 ($44.99 each) Three Axis Gyroscope Three Axis Magnetometer Temperature (Small Signal Diode) Sun Sensor ( five photodiodes) GPS (Lassen IQ) Lassen IQ SMD Mating Header Antenna GPS Ultra Compact for Lassen IQ ITG-3200 Sparkfun 2 - $99.90 ($49.95 each) Micromag3 Sparkfun 2 - $ ($59.95 each) 1N4148 Sparkfun 4 - $0.60 ($0.15 each) QP TO5 GPS GPS GPS Mouser Electronics 10 - $ ($17.02 each) Sparkfun 2 - $ ($56.95 each) Sparkfun 2 - $3.90 ($1.95 each) Sparkfun 2 - $37.90 ($18.95 each) 16 Heater 2 - San Disk 32GB SDSDX3- Amazon 2 - $ ($ G-A31 each) 1 Arduino MEGA DEV Sparkfun 2 - $ ($64.95 each) 11 Arduino MEGA shield kit DEV Sparkfun 2 - $35.90 ($17.95 each) Team EQUIS 48

56 12 microsd shield DEV Sparkfun 2 - $29.90 ($14.95 each) 15 GPS shield GPS Sparkfun 2 - $33.90 ($16.95 each) 14 Material for housing $ Extra PCB board $600 Total Cost: - $2, Table 6 Materials Acquirement & Costs 11.0 Risk Management and Contingency Risk Type Electrical Mechanical Software Risk Exceed component s temperature Faulty connection between cables Impact Severity Assessment of Risk Likelihood Probability Detention Difficulty HIGH HIGH MEDIUM HIGH HIGH MEDIUM Short circuit HIGH MEDIUM MEDIUM SD card failure HIGH LOW MEDIUM Temperature Leakage into payload HIGH MEDIUM MEDIUM Structure Failure HIGH LOW MEDIUM Loosening of nuts and bolts due to vibration Fall into wet terrain MEDIUM LOW LOW HIGH MEDIUM LOW Data Corruption HIGH HIGH HIGH Loss of Data due to Overwriting Improper Logic Programming HIGH HIGH MEDIUM HIGH MEDIUM HIGH Risk Control Measures Provide a heating system Ensure connections and strap cable to prevent movement Review the PCB population process Transmit through Telemetry Pressure and temperature testing Minimize apertures, use alloy 2014 and mono block design Use pressure bolts Ensure proper sealing of Payload Store Data in SD memory card Divide SD in sectors of 512bytes Test program before flight Team EQUIS 49

57 Scheduling Insufficient memory storage Not meeting PDR deadline HIGH HIGH LOW HIGH MEDIUM MEDIUM Errors in PDR HIGH MEDIUM MEDIUM Table 7 Risk Management & Contingency Obtain memory card with greater storage Getting more hours of work and more days in the lab Review document various time Through a risk management it is possible to reduce the likelihood of unexpected events. By having a contingency plan it is possible to reduce costs and severity of a risk. This is done by identifying, assessing and developing a strategy to response for each risk, while monitoring for additional risks as shown in Figure 1. Table 8 shows the potential risks that the EQUIS experiment may encounter and the contingency plan that the ADS team has for each of these risks. As illustrated in the Table 8, the potential risks have several categories. It is important to keep constant cycle of risk management in order to control possible variables that may affect project performance. Figure 27 Risk Management Cycle Team EQUIS 50

58 12.0 Glossary CDR Critical Design Review FRR Flight Readiness Review PDR Preliminary Design Review TBD To be determined TBS To be supplied WBS Work breakdown structure HASP High Altitude Student Platform ITS Internal Temperature Sensor ETS External Temperature Sensor ADS Attitude Determination System EQUIS Experiment with Quality United In Science NEU North-East-Up MRI Magnetic Resonance Imaging GPS Global Positioning Satellite SD Secure Digital USB Universal Serial Bus EEPROM Electrically Erasable Programmable Read-Only Memory ADC Analog to Digital Converter FET Field Effect Transistor IC Integrated Circuits PCB Printed Circuit Board Team EQUIS 51

59 Appendix Appendix A I. Data. T out = C Q emt eq = 1W (computed) T in = 0 0 C (design parameter) k alm = 134W/m 0 C k ins = 0.12 W/m 0 C L alm = m L ins = 0.005m σ = 5.67X10-8W/m 2 K 4 (Stefan-Boltzmann const.) є FR4 = 0.8 є Alm = 0.07 (rough surface) ѵ air@100000ft = 8.012X10 4 m 2 /s V mean@100000ft = 26.82m/s ρ air@100000ft = 1.841X10-2 kg/m 3 μ air@100000ft = 1.475X10-5 kg/m s II. Find Q heater = Q loss =? Team EQUIS 52

60 III. Diagrams Figure 28 Heat Loss Analysis IV. Analysis Heat loss: Q loss =UA T T= (Tin Tout), UA=, R total = R eq1 +R A1,1 + Ri ns +R Al2 +R eq2 Thermal Resistances: R eq2 = h comb = h rad,in + h conv,in h conv,in =? ; h rad,in =? Equipment power radiation: Q rad eq = є FR4 σa eq (Ts 4 - T in 4 ) Ts=450K=182 0 C; Aeq= m 2 Q rad eq = h rad A eq (Ts- T in ) h rad,in = Ts =Tin + є = 0.27W/m2 0 C Internal Convection: -0.5 Team EQUIS 53

61 Nu L = s=0.06m =,., (distance between boards). ; L =0.1m ; R a,s = ; B = Ideal ; 2 gas T f = = 15 0 C = 288K ; B = /K; Pr=0.7362; k= w/m 0 C R a,s = Ra L = R a,s (L 3 /s 3 ) = Nu L =1.79 = Therefore: R eq2 = C/W h conv =0.71W/m 0 C h comb,in 1W/m 2 0 C R eq1 = h comb = h rad,out + h conv,out h rad,out =? h conv,out =? Q rad out = є Alm σa s (Ts 4 - T out 4 ) = 84.8mW, h rad,out = = 0.15W/m2 0 C For external convection: Walls modeled as: V mean T out T s Re = L` Figure 29 External Convention where Lc = L Re = 4686 = 4.69X10 3 Re cr =5X10 5 Re < Re cr (laminar flow regime) Therefore: Team EQUIS 54