FITSAT TECHNICAL PROPOSAL: A STUDENT SATELLITE INITIATIVE

REVISION B

FITSat Team

Project Manager: Jerry Tran Systems Engineer: Jason Barbour

Avionics Lead: Faye Tomimbang RF/Electrical: Michael Letsky

Launch Operations: Naim Torlakovic Mission Analysis: Christian Pederson

Submitted to: Dr. Paavo Sepri

Florida Institute of Technology 150 West University Boulevard

Melbourne, Fl 32905

Date: October 10, 2000

ii

REVISIONS

LETTER DESCRIPTION DATE APPROVAL

A

B

Submitted to Dr. Sepri for review

Minor corrections to schematic, schedule, and

references.

9/19/00

10/10/00

iii

Table of Contents

List of Authors ....................................................................................................................... iv Executive Summary ................................................................................................................1 Background ............................................................................................................................2 Problem Statement and Objectives ..........................................................................................4 Technical Approach ................................................................................................................5

Mission Analysis.................................................................................................................5 Case 1: Low Earth Orbit ..................................................................................................5 Case 2: Altitude of 12,000 ft.............................................................................................5

Launch Operations .............................................................................................................6 Mission Operations .............................................................................................................7

Parachute Deployment ....................................................................................................7 Mission ..........................................................................................................................7 Landing and Recovery ....................................................................................................7

Systems.............................................................................................................................7 Mechanical.........................................................................................................................9 Avionics ...........................................................................................................................10

Power ..........................................................................................................................11 RF/Electrical.................................................................................................................11

Ground Station .................................................................................................................12 Integration & Testing .........................................................................................................12 Schedule..........................................................................................................................13

Budget .................................................................................................................................16 Organization and Capabilities ................................................................................................17 Summary .............................................................................................................................19 References...........................................................................................................................20

List of Figures

Figure 1. FITSat-1 Preliminary Configuration ......................................................................9 Figure 2. Avionics Schematic...........................................................................................11 Figure 3. Preliminary Schedule ........................................................................................15 Figure 4. Personnel Chart................................................................................................18

List of Tables

Table 1. Mission Tradeoffs ...............................................................................................4

iv

List of Authors

Jerry Tran, Project Manager: Executive Summary, Background, Problem Statement and Objectives, Budget, Organization and Capabilities, Schedule, Summary, References. Jason Barbour, Systems Engineer: Systems. Faye Tomimbang, Avionics Lead: Avionics, Mechanical, Schedule. Michael Letsky, RF/Electrical: Power, RF/Electrical, Ground Station. Naim Torlakovic, Launch Operations: Launch Operations, Integration & Testing. Christian Pederson, Mission Analysis: Mission Analysis, Mission Operations

Executive Summary

FITSat Inc. is pleased to submit its Technical Proposal in regards to the Florida Institute of Technology’s (FIT) Announced Opportunity (AO) for a senior design engineering project. The purpose of FITSat Inc.’s proposed engineering project is the design, construction, and, if possible, the launch of a spacecraft. As the proposed contract limits FITSat Inc. to within a budget of $250, FITSat Inc. has thusly scaled its project to within the realistic constraints of budget, schedule, and manpower. To this end, FITSat Inc. is proposing the development of Florida Tech’s first student led spacecraft program with the aim of designing, building, and, if possible, the launching of FITSat-1, a picosatellite. This Technical Proposal will serve to answer not only how FITSat Inc. will accomplish the development of its spacecraft, but also the why. Detailed within the Technical Proposal is the background on the inception of FITSat, its objectives, the technical approach, budget, and capabilities.

Background To begin, it is necessary to define the scope of this project. A spacecraft is defined in Merriam-Webster’s Collegiate Dictionary as “a vehicle or device designed for travel or operation outside the earth's atmosphere.” A satellite is then “a manufactured object or vehicle intended to orbit the earth, the moon, or another celestial body.” A picosatellite is a classification of satellites based on mass. In this case, it has a mass of no more than one (1) kilogram. Thus, the overall problem becomes how to develop the smaller, faster, yet inexpensive approach of developing picosatellites for space applications for commercial exploitation or scientific study. At the industry level, picosatellites have the potential to fill in specific niches for plentiful and inexpensive space capabilities, albeit with a high risk factor. As Dr. Wiley J. Larson of Microcosm points out in Space Mission Analysis and Design,

Today the small-satellite industry focuses on special niches, to satisfy missions that conventional satellite technology cannot cover with large spacecraft and multiple payloads, developed in programs spanning 5 to 15 years.1

The commercial and scientific niches of these small satellites might include low-cost imaging,

supplementing existing global digital communications, and observing scientific phenomena. These low-cost, high-risk small satellites can fulfill these space applications by deploying multiple and therefore redundant satellite constellation. In addition, a picosatellite becomes an ideal candidate for a student led spacecraft program. It limits the scope of a project within a manageable budget, realistic timeframe, and available manpower while serving to educate students in the design life of a “typical” spacecraft. It gives students the experience of working on a spacecraft at the multiple levels that are required to build a spacecraft bus, integrate its payload, provide mission operations, as well as the programmatic requirements. Typical student led spacecraft programs can be found at Stanford University, University of Santa Clara, Arizona State University, Dartmouth, University of Texas at Austin, as far away as the University of Tokyo, and much more. In addition, picosatellites have been launched into space on Minotaur rockets. The Aerospace Corporation achieved launch with its mini-constellation of picosatellites, the University of Santa Clara2, and even a group of amateur radio enthusiasts in StenSat (a CubeSat). Stanford University also developed OPAL (Orbiting Picosatellite Automated Launcher) which launched many of these picosatellites.3

Professor Bob Twiggs of Stanford has been a pioneering proponent of student led spacecraft

programs, like the CanSat programs at Stanford, University of Tokyo, and CubeSat programs at Dartmouth and California Polytechnic (to name but a few). A CanSat is defined as a picosatellite that meets the mechanical design of its namesake, namely a soda can. More specifically, it has a cylindrical body with a diameter of 66mm by 123mm in length.4 A CubeSat, in turn, is a picosatellite that is designed like a cube and is 10cm by 10cm by 10cm.5

CubeSats are typically more technically challenging, requiring a multidiscipline effort from

aerospace, electrical, and computer engineers. It also requires a large budget (approximately $5000) and a schedule of at least one (1) full school year. Of special note is that CubeSats have been launched into space and there is a developed and tested carrier.7 On the other hand, a CanSat can be considered a scaled down CubeSat. It is less technically challenging, does not require an extensive budget, and usually accomplished within one (1) whole school year. Unfortunately, CanSats have never been launched into space although the University of Texas at Austin has designed a CanSat carrier should a launch ever be acquired.8

Problem Statement and Objectives The overall objective of FITSat Inc.’s proposed project is to emulate the design lifecycle of a

spacecraft. The lifecycle of FITSat-1 will follow that of a “typical” spacecraft. Namely, it will go through a preliminary design, assembly and integration, integration and testing, and finally on-orbit operations.

Moreover, the project objectives necessary for completion is to design and build the FITSat-1

bus, do this within one (1) school year and hopefully establish a legacy in a student led spacecraft program at F.I.T.

Shown below in Table 1. Mission Tradeoffs are the factors that were involved in deciding FITSat

Inc.’s project approach. The CanSat bus is defined as the spacecraft without the payload while CanSats and CubeSats are picosatellites with their payloads.

Table 1. Mission Tradeoffs

Option Description Probability of Success

Cost to Build

Cost to Launch

Total Costs

A CanSat bus 100% $75 $0 $75 B CanSat bus w/

Launch @ 12K ft 95% $125 $2000 $2125

C Space-rated CanSat bus

90% $750 $0 $1000

D Space-rated CanSat @ 12K ft

85% $1000 $2000 $3100

E Space-rated CanSat w/ LEO

Launch

1% $1200 $15M $15M

F CubeSat 25% $5000 $0 $5000 G CubeSat w/ LEO

Launch 5% $7000 $15M $15M

As one can see, FITSat Inc. chose Option D. This incorporates both the CubeSat and CanSat

program. FITSat Inc. will raise the technical level of typical CanSat programs so that in the case of a definite launch vehicle, it may be selected as one of twelve payloads in the University of Texas at Austin’s carrier and be capable of operating in space environment. In such a case that there is no launch, FITSat Inc. has already taken steps to acquire a terrestrial level launch vehicle in the way of amateur rockets capable of going to 12,000 feet.

Thus, FITSat-1’s primary mission objective will be to build a CanSat bus that will be able to

withstand the environment up to 12,000 feet, the general space environment, the launch environment required to take it up there, and carry a primary payload in the form of a CCD/CMOS camera. As secondary mission objectives, FITSat-1 will also be recoverable and reusable. If possible, FITSat-1 will either be launched on a terrestrial level or in space. It may also be that FITSat Inc. will develop a prototype, barebones CanSat bus as a testbed.

Technical Approach

I. Mission Analysis

There are several important factors involved in the mission analysis for this project. Namely, they are the environmental conditions in which the flight takes place, the deployment of the satellite’s parachute, and the environmental conditions in which the satellite will land. These factors all play critical roles in the design of the satellite. Case 1: Low Earth Orbit

Even though the FITSat design team has found it to be unfeasible to launch the picosatellite into orbit, it is still important to keep this objective present in the background of the design process as the ideal objective of this spacecraft project. As such, it is important to briefly mention some conditions of the near-Earth environment that would impact the design of any satellite constructed to inhabit this region.

Satellites in a space environment would be subject to severe temperature change. This would make it necessary to account for thermal fatigue and temperature extremes in the design process. This would not only create possible structural problems, it could also impact the functioning of various systems contained within the satellite. In addition to temperature changes, the satellite would also be subject to radiation degradation. Once again, this must be accounted in the structural design of the satellite. The possibility of impact with orbiting debris or micrometeoroids is also a very real problem that needs to be addressed from the standpoint of structural design. Other negative phenomena’s found in space is outgassing due to vacuum and cold metal welding due to vacuum and temperature. Case 2: Altitude of 12,000 ft.

Deployment of the satellite at this is altitude is currently the primary objective of FITSat’s design team (The specifics of which will be discussed under Launch Operations). Preliminary calculations indicate a minimal flight time of approximately twenty (20) minutes from deployment at 12,000 feet. At this height, it is necessary to consider the various properties of the air that would affect the flight and the functioning of the various systems of which the satellite is to be comprised.

The air density is a very important property that is directly involved in determining the duration of the flight, as it partially responsible for determining the drag associated with the parachute, and thereby indirectly related to the rate at which the satellite descends. At an altitude of 12,000 feet, the air density is approximately 0.849 kg/m3. Of course as the satellite descends, the air density increases to its sea level value of about 1.225 kg/m3.9 This density gradient must also be accounted for in estimations of the duration of the flight.

The temperature at 12,000 is approximately 264.4 K. Temperature varies linearly with altitude, and in this region of the atmosphere, the temperature will increase at a rate of 6.5 K/m until it reaches sea level temperatures. While this temperature range may not be sufficiently extreme to be problematic for the structural design, it may very well influence the design and selection of the components that will make up the internal systems of the satellite. All electronic devices must be capable of functioning at any temperature within the range that will be experienced during the flight.

The atmospheric pressure must also be considered in the design process. At a height of 12,000 feet, the atmospheric pressure is about 64.46 kPa. This pressure increases linearly until it reaches the sea level pressure of 101.3 kPa.

In addition to these environmental conditions that are relatively easy to determine or estimate,

there are two other factors, namely humidity and wind, that must be considered in the design process. These two factors are difficult to predict or estimate. Wind direction and speed will affect the flight time of the satellite as well as the location at which it lands. Humidity levels may possibly impede the proper functioning of the various electronic devices that are being considered for use in the satellite and must be considered when this equipment is selected.

II. Launch Operations

To launch picosatellites into orbit, previous universities have used the Minotaur Launch Vehicle, which has previously carried microsatellites into orbit. This launch vehicle is a combination of the Minuteman 2 ICBM system and the proven Pegasus launch vehicle. The first and second stages of the vehicle are the M-55A1 and SR-19 Minuteman 2 rocket motors respectively. The third and fourth stages are the Orion 50 XL and Orion 38 Pegasus rocket motors respectively. The basic configuration of the vehicle uses the Pegasus fairing to protect the payload. This vehicle can operate with two fairings allowing for the launch of oversized payloads. The standard configuration uses a slightly modified Pegasus fairing. It has a dynamic payload volume of about 46 inches in diameter by 88 inches long. The larger fairing has a dynamic payload volume of about 61 inches in diameter and is about 133 inches long. Minotaur can lift 750 pounds to a 400-nautical mile, sun-synchronous orbit.

This vehicle is designed to provide low cost, reliable space launch capability in support of U.S.

government small-satellite launch requirements. In July 2000, a Minotaur rocket carried U.S. Air Force Academy's FalconSat and the Air Force JAWSat. The JAWSat vehicle also serves as a multi-payload adapter carrying three micro satellites. The microsatellites included the Optical Calibration Spheres, Stanford University's OPAL carrier, and Arizona State University's ASUSat. Ideally Florida Tech’s satellite would be on the next trip, which does not have a set date yet.

Because the chance of getting onboard the Minotaur rocket is fairly slim, FITSat Inc. has

developed a launch vehicle descope plan. The rocketry club NEFAR (NorthEast Florida Assocation of Rocketry) stationed in Jacksonville, FL is building a M-Class rocket which will be able to fly to the altitude of 12000 ft. The members of the club, including its president, Greg Peebles, were very enthusiastic about FITSat Inc.’s payload and has preliminarily agreed to provide free launch services not including cost of payload integration. FITSat Inc. would need to design an additional deployment mechanism either in the way of a payload fairing or carrier. While NEFAR has carried much heavier payloads on their rockets, they have never before dealt with deploying a payload in-flight. The launch would take place from Palm Bay, Jacksonville or from NEFAR’s launch site in South Carolina.

NEFAR members are using six (6) inch tubing for their rockets, which limits the size of the FITSat

carrier design. The M-Class rocket can carry the twelve (12) oz of payload with commercial rocket motor whose size is determined based on the vehicle’s weight. The members of a NEFAR will develop a

costume made rocket using a computer simulation to make sure FITSat-1 will reach the desired height. The M-Class rocket meets the requirements as set by ARLISS (a Rocket Launch for International Student Satellites) and Professor Twiggs. It was the ARLISS rockets which launched previous CanSats at an altitude of 12,000 feet in order to mimic a LEO orbit.10

In this case, FITSat would have to integrate the parachute into the CanSat, which will bring the

satellite down slowly enough for planned operations. The M-Class rockets use two-stage parachute deployment mechanism. The first parachute, smaller one of two, is deployed once the maximum altitude is reached. FITSat-1 will probably deploy at this time. The reason behind this, other than the highest altitude, is that the launch vehicle is suspended on the fairly small parachute and it will descend much faster than the payload, which eliminates the chance for collision. FITSat-1 would have only one large parachute, unlike the launch vehicle. The combination of altimeter, accelerometer, and timer electronics are the primary, secondary and tertiary systems respectively, designed to assure deployment and activation of pyrotechnic charges. FITSat Inc. plans to use similar mechanism to deploy and activate Can-Sat.

III. Mission Operations Parachute Deployment

The initial concept of the manner in which the parachute will be deployed is quite simple. The parachute will be tied to the top of the satellite and will open naturally after the satellite has been deployed from the rocket. It is hoped that the simplicity of this design will eliminate the need for any firing mechanism to deploy the parachute, thereby reducing the chances of a faulty parachute deployment. Mission Once FITSat-1 has been deployed from its launch vehicle, it will be in state of flight. Mission Analysis has calculated a time of twenty (20) minutes of flight. This mimics the time it takes for a satellite to complete one (1) LEO orbit. During these twenty minutes, the primary payload of FITSat-1 will be activated by a deployment switch which in itself will be activated by the release from the rocket. The CCD/CMOS camera will begin to take images while transmitting its data to the ground station. Landing and Recovery

As it is one of the stated objectives to recover the satellite upon its landing, the location and conditions of the landing environment are quite important. Given the possibility that the launch site may be in fairly close proximity to the Atlantic Ocean and/or Indian River, it is necessary to acknowledge that this landing may be a “splash down”. Obviously, in this instance recovery would not be possible. The landing location will very probably be determined by the wind conditions on the launch date, something that unfortunately cannot be predicted or estimated with great accuracy until much closer to the launch date. In such case, it may be possible to launch at a differing launch sites in northern Florida or South Carolina.

IV. Systems

The objective of FITSat Inc.’s Senior Design Project is to build and launch a CanSat into orbit. However, this is a very optimistic goal for FITSat Inc. since the chances of getting launch space on either

a shuttle or rocket are slim. FITSat Inc.’s realistic goal is to build a space worthy CanSat so that if an opportunity does arise to launch the satellite into orbit then FITSat-1 can still carry out that objective. The design of FITSat-1 will generally follow that of any small satellite design. Namely, the use of low cost parts and proven approach combined with a sense of ruthless minimalism. The Artemis picosatellite of the Santa Clara University epitomizes such an approach and it is this successful approach that FITSat Inc. hopes to incorporate. The Artemis approach can be stated thusly:

Maximal efficiency is essential in the design of the picosatellites due to the volume constraints. Only the most fundamental subsystems can be considered a part of the design process. Unnecessarily complex subsystems could result in inflated budgets and prolonged project lifetimes. Developments on a pico-class spacecraft need to place an emphasis on minimum missions and the simplest way of meeting a set of specifications.11

To build this satellite, it will cost approximately $1000 (of which a general breakdown can be found in the Budget section of this proposal). Should funding be unavailable to build FITSat-1, FITSat Inc. will pursue the design, development and launch of an unrated CanSat bus, which will cost 10% that of the FITSat-1. The picosatellite will still be able to be sent up on the twelve thousand foot capable rocket, but it will not be space worthy. Also, the frozen structural design size will greatly affect the payload that FITSat-1 as well as all spacecraft subsystems like power and communications. There will be four main systems that make up the satellite: power, the primary payload, RF/electrical, and mechanical structure. The primary payload, transmitter, flight board computer, and general avionics will each be connected to the power supply separately. Each system will then be connected to each other so that the data from the camera can be converted to transferable data in the avionics and then transmitted to the ground station via the transmitter. The key necessity for each of these components is that they operate at the same voltage or regulated voltage so that each component is able to feed off the same power supply. If not, it could lead to a short circuit, which would render the satellite useless. It will also be necessary to establish the data rate through which the camera will pass its data to the flight board computer and then the transmitter. The camera itself will determine the amount of power necessary to send with the satellite. While there are still power design configurations to be figured out, it will be necessary to determine what initial power will be needed to power the camera and transmitter while generating enough power to run the general avionics. The primary payload will also be the driver in the configuration of the transmitter and thusly the ground station. The ground station must be able to accept the data rate at which the camera downlinks as well as have a powerful enough receiver to receive it in the first place. FITSat-1 would be required to have the appropriate power supply to feed its camera, power the general avionics, and enough power to supply a transmitter powerful to reach the ground.

The data rate of the camera will determine the type of transmitter and avionics that FITSat-1 will incorporate. If the transmitter or the avionics cannot keep up with the camera then data will be lost. If the camera takes pictures at a rate of one every ten seconds, the transmitter transmits a pictures at a rate of one per twenty seconds then the avionics must be set up to either be able to store data until it can be transmitted or to slow the speed of the camera to the speed of the transmitter. This can be done designing the avionics to give power to the camera every twenty seconds, which would give the transmitter the time it needs to transmit.

This also relates to the ground station as well. The ground station must be able to process and store the data quick so that the back up in data does not cause a system failure in the ground station or the satellite, which would result in the failure of the experiment and possible loss of the satellite itself. If the satellite is unable to send and the ground station is unable receive then it would not be possible to retrieve the satellite upon landing. There are two ways for the ground station to keep up with the transmitter and that is with either a fast processor or by placing enough memory in the receiver so the information can be stored until the computer can process the information.

As part of integration and testing (I&T), it will be necessary to test each components so that they

meet the necessary requirements for a space launch. Before this satellite is built, proper planning, design and testing must be exercised so that any shorts, faulty wiring can be found and the proper equipment is used in processing the pictures that will be sent from the satellite.

V. Mechanical

The CanSat bus will house the electrical components, flight board computer, power supply, and primary payload. FITSat will adhere to the specifications for CanSats provided by Professor Twiggs, the originator of the CanSat projects. The structure size is 66 mm in diameter and 123 mm in length.4 The rule of thumb for utilization of the volume is roughly 1/3 batteries, 1/3 payload, and 1/3 for the other electronics. The total volume of the structure will be dependent on the electrical power system, the payload, and the flight board computer. The preliminary arrangement for FITSat-1 (drawn in ProEngineer2000) is presented in Figure 1. FITSat-1 Preliminary Configuration

Figure 1. FITSat-1 Preliminary Configuration

STRUCTURE

PAYLOAD(CAMERA)

FLIGHT BOARD COMPUTER

ELECTONICS

POWER SUPPLY

UPPER PANEL

Although the FITSat is expected to travel at an altitude range of 12,000 to 15,000 feet above sea level, one of the objectives of the FITSat Inc.’s mission is building a space-rated CanSat. In conjunction with Mission Analysis, analyses and tests will be performed to ensure that the CanSat will survive the FITSat mission, and to some extent, be assigned a space rating. A space rating, defined as the ability of the satellite to survive in a space environment, is a secondary, but desirable, objective of FITSat.

FITSat is considering using AL-6061, an aluminum alloy that is widely used for aircraft. Past

CanSat teams, in addition to the picosatellites built by the Artemis group, have also used Aluminum 6061.11 The Artemis group of the University of Santa Clara provided three picosatellites for the Stanford’s OPAL. The picosatellites have gone into orbit and have withstood a harsh space environment and thus has an established space heritage.

Several configurations for the CanSat casing are possible, but the optimum design is a

compromise of strength and modularity. The shell configuration is the simplest and most efficient design. Previous CanSat projects attest that a simple tubing with no stiffeners or props have been the most successful in withstanding vibration tests and structural forces applied on to it. FITSat Inc. is capable of performing vibration tests on FITSat-1 with their components at the Fluids Lab. FITSat will investigate different configurations and different testing facilities where available.

The most practical structure for the CanSat bus is a tube with welded upper and lower panels.

The upper panel will act as a partition between the parachute packing and the contents of the CanSat. Holes will be drilled onto the upper panel for the chute strings as well as to eliminate some of the panel’s weight without compromising its role to protect the CanSat’s contents.

Another objective of the team is capturing pictures in real-time. Instead of shooting the camera

straight down as most CanSats have done, FITSat proposes to angle the lens of the camera to get an oblique view of the landscape. The resulting photographs will show landscapes in three dimensions (as opposed to a two dimensional and constant view of the ground). Such a configuration will allow FITSat to observe the realistic path that the CanSat will take and provide an estimate of the CanSat’s rotation and descent.

A secondary goal for FITSat Inc. is set for Fall 2000 to deploy a prototype testbed from a common

NEFAR Rocket to an altitude of 6,000 feet to test the structure of the CanSat bus, but more importantly, the electronic configuration that is the heart of FITSat’s mission.

VI. Avionics

The primary goal of FITSat is to design a satellite that produces valuable science in space at a

low cost while giving students a hands-on learning experience. Avionics is the most challenging and exciting aspect of FITSat-1’s mission. The Avionics Team will configure FITSat-1’s electronics such that the satellite will send data to the ground station. Data of interest include altitude reading, state of health, location, and real-time pictures from the CCD/CMOS Camera.

The technology FITSat Inc. hopes to incorporate in the CanSat includes a flight board computer,

transmitters and antennas to relay information to the Ground Station, and a CCD/CMOS camera that will provide the CanSat with pictures illuminating the CanSat’s descent. Engineers who work extensively with aircraft and satellite communication will be providing Workshops on Avionics to aid FITSat Inc. in understanding avionic systems. The Avionics team will coordinate closely with RF/Comm and I&T to ensure that data is received and recorded appropriately.

The main Avionics configuration being considered is shown in Figure 2. Avionics Schematic.

Data from the GPS Unit and Payload (the camera and any other secondary payload) will be piped into the Flight Board Computer via RS-232 data buses. That data in turn will be sent to a Video Imaging Overlay device that will combine it into one data stream and sent to the transmitter. The transmitter sends the signal to a receiver from the ground station. The signal will be demodulated at a more appropriate frequency and the data is stored at the Ground Station computer.

Figure 2. Avionics Schematic

Power

The primary bus system of FITSat-1 consists of an onboard flight computer, digital converter, primary communications array, and the payload. The onboard flight computer has three primary functions. One function of the onboard flight computer, is receiving data from the payload (CCD Camera), sending it through the digital converter, through the primary communications array, and to the ground station. The second function of the onboard flight computer is to interface with the tracking system, and send that information to both the primary and secondary communications array. The third and final function of the onboard flight computer is power management and state-of-health of the CanSat, which is transmitted through the primary communications array.

The planned operational life cycle for FITSat’s primary bus system is estimated at thirty (30) minutes, this time period was determined from previous CanSat projects as well as calculations from Mission Ops and Analysis. Because the planned operation lifecycle is thirty (30) minutes, the power system is looking to draw its energy from a number of batteries and/or solar cells. This combination will be used according to power requirements, and weight restrictions of the CanSat. In the event of a complete power failure, the onboard flight computer will receive enough power from the secondary communications array to support all tracking capabilities for a maximum period of four (4) hours.

RF/Electrical

The preliminary design of FITSat-1’s communications subsystem consists of a primary and a

secondary communications array. The primary communications array consists of an antenna, transmitter,

Ground Station

Demodulator

Receiver GPS Unit

Payload

Flight Board Computer

Video Image Overlay

Transmitter

Power

RS-232

RS-232

multiplexor and amplifier. Its function is to transmit the payload data to the ground station in real time, as well as provide the ground crew with its location. The primary communications array will receive a total of 15VDC from the primary power source; 6VDC for the transmitter and 9VDC for the amplifier.

The secondary communications array consists of an antenna, transmitter, amplifier, and an independent power source. The function of the secondary power source is to transmit the CanSat’s location in the event the primary communications array is damaged or fails. The secondary communications array will operate similar to the primary communications array receiving 15VDC from an independent power source. The prospective outlook for the independent power supply is a minimal amount of batteries accompanied by a minimal amount of solar cells strategically placed for the best reception of solar energy, despite the CanSat’s three-dimensional orientation.

To keep FITSat-1’s design simple, lightweight, and cost effective, FITSat Inc. has minimized the amount of components the CanSat will require for operation. For this purpose, the CanSat’s communications arrays do not have a receiver, because the CanSat is fully autonomous and does not need to receive any flight commands.

VII. Ground Station

Preliminary studies suggest that FITSat-1’s ground station needs to be completely compact and

portable in order to ensure complete recovery of the flight unit as well as transportation to the launch site. The ground station is equipped to receive imaging and telemetry data from the CanSat. The ground station is composed of a receiver, decoder, amplifier, antenna, computer, and software, and is powered by 18VDC from an internal power source (batteries), which makes it powerful enough to receive signals from the CanSat at a large distance.

VIII. Integration & Testing The structural testing for the shell of the CanSat will consist of impact, temperature durability, and

vibration testing. Impact testing will determine the degree of damage caused to the CanSat during the landing

procedure. FITSat Inc. will subject the potential designs to free fall from different heights as well as descent with the parachute attached. These tests will give an idea about the magnitude of descent speed and material properties, which can lead to the weight reduction.

Temperature durability will determine how will design act under the exposure to sudden change in

ambient temperatures. As FITSat Inc. is limited by testing facilities, it is possible to use liquid nitrogen to reach low temperatures and observe the design for different structural deformations such as cold welding phenomenon. This will be the thermal stress screening that will weed out infant mortality and test the operating range of the avionics.

FITSat-1 will be placed in the upper level of the rocket so it will not be subjected to the extreme

high temperature but high temperatures will be tested as well. Test for extreme highs will give an estimation of chance for recovering the CanSat in case of explosion or misfire.

Vibration testing will be performed on the vibration table available in FIT laboratory. This test will

simulate the vibrations produced by the rocket engine during takeoff and flight. The tests other than structural will include wind tunnel use and testing of avionics for proper functioning.

The wind tunnel may be used in order to determine the drag coefficient of parachute. This test

will help determine the shape and size for the parachute needed for desired time of descent. A series of functional testing will be conducted on the payload, power, transmitter, flight board

computer, and general avionics to insure their proper functioning in all types of environment. A major test for the whole project will be the prototype testbed. The CanSat will be on the test

flight on a smaller rocket to the altitude of 6000 ft. This test will show if design is really working or not. The smaller rocket is another fallback plan that can be executed in case NEFAR M-Class rocket plan fails or it is delayed.

IX. Schedule

In Figure 4. Preliminary Schedule, FITSat Inc. has outlined some milestones for this space program. The launch of the prototype testbed is scheduled around the December/January timeframe while actual launch is projected to be during the March/April months.

FITSat Inc. has already begun a great deal of research as indicated on the schedule as well as

an internal Feasibility Assessment Study. Although the schedule for FITSat-1 appears to be aggressive, this can be considered the norm of any small spacecraft program as longer spacecraft programs translate into higher program costs.

Slack in the schedule is estimated to be one month. Further slack can be achieved by eliminating

the prototype testbed launch as it is not necessary, but is very useful. Finally, the main launch itself may be scrubbed in order to achieve another two months. While this is extremely undesirable, building FITSat-1 at an acceptable engineering standard must take priority. It also acknowledges that launch services is a capricious mistress in that no launch date is ever really certain.

15

Figure 3. Preliminary Schedule

FITSat* SEPT OCT NOV DEC JAN FEB MAR APR Comments Group Formation Research Feasibility Assessment Proposal Tech/Sponsors Sponsor Letters w/ Proposal Pending Account NEFAR Launch Midterm Exam Preliminary Design Review Order/ Purchase Prototype Materials Prototype Avionics Build Prototype Structure Build Instrument Integration Prototype Launch Date not exact Final Course Presentation (FA2000) 1 month prep Final Exams FAA Waivers for 12,000 Launch Winter/Spring Break Order/Purchase CanSat Materials CanSat Avionics Configure CanSat Structure Build CanSat Individual Component Test CanSat Integration CanSat Test Space Rating Qualification CanSat Launch CanSat Final Report (SP2001) 1.5 month prep Graduation

*KEY

Each cell = 1 week period Completed milestone

Mandatory

Suggested time

16

Budget

The average cost of a CanSat bus is $75. This estimate is based on a plethora of experience and parts list from Stanford University and University of Texas at Austin.

FITSat-1’s payload, the CCD camera, costs on upwards of $300 while the descope

payload, the CMOS camera, is priced at $100. A GPS receiver would cost $180. Avionics, including transmitter, power supply, wiring, etc., should cost around $100. The aluminum used for the spacecraft structure would cost $75 while special modifications at a machine shop would cost an approximate $100. Transportation, depending on launch site and vehicle, may cost upwards of $150. Ground station equipment will cost $300. Lab equipment and testing will be done at existing Florida Tech facilities.

The total proposed cost of this spacecraft is approximately $1000. This does not include

launch services. If included, it would bring the total cost, including payload integration, to $3100.

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Organization and Capabilities

The FITSat Team is composed of Jerry Tran, Jason Barbour, Faye Tomimbang, Michael Letsky, Naim Torlakovic, and Christian Pederson. As is the norm of many small spacecraft programs, departmental responsibilities are divided amongst the small team, but this also requires a certain amount of flexibility, as all team members must wear multiple hats. This will ensure that all subsystems of the spacecraft are being supervised as well as allowing extensive peer reviews and interaction on a personal, team level. Project Manager Jerry Tran has had some spacecraft experience working as an intern at SpaceDev Inc. where he was involved in CHIPSat, a UNEX NASA spacecraft. He also participated in the design study of Hektor, a deep space asteroid rendezvous mission, and the SpaceDev Lunar Orbiter. He is familiar with the life process of a spacecraft and worked on the programmatic and systems level of a spacecraft. He is also attending class at the Florida Space Institute at the Kennedy Space Center for Small Satellite/Payload Integration project which will serve as the FITSat-0. Systems Engineer Jason Barbour is also involved with FITSat-0. In addition, Mr. Barbour has extensive knowledge with aircraft as well as civil engineering and survey work. His previous experience includes designing and competing in the Odyssey Mechanical Car competition. Avionics Lead Faye Tomimbang works as an intern at Rockwell Collins as part of the Manufacturing & Engineering Department. She has experience working with avionics at Rockwell as well as extensive CAD work experience. RF/Electrical Engineer Michael Letsky has interned with NASA at the Kennedy Space Center. He worked on the Mars Life Sciences chamber for simulating a Mars environment as well as general NASA flight hardware. His previous experience includes designing and competing in the Moon Buggy competition. Launch Engineer Naim Torlakovic is currently employed at the Florida Tech Fluids Lab for the past two years and has an appreciable knowledge in manufacturing and tooling. His previous experience also includes working on solar oven competition. Mission Analyst Christian Pedersen has done survey work at Pedersen & Pedersen. FITSat Inc. will rely on his strength as an analyst. Intern Tammy Stuart has worked at SpaceDev as an intern doing research as well as market analyses.

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The organizational hierarchy and breakdown of responsibilities can be seen in Figure 3. Personnel Chart.

Figure 4. Personnel Chart Consultant work, to ensure the effective performance of FITSat-1, will be taken up by multiple individuals. Namely, NEFAR’s 30 to 50 members have volunteered their expertise for launch operations and payload integration. In addition, Energy Laboratories has volunteered their assistance in the realm of FITSat-1’s power system. Mark Bentley, avionics test engineer at Rockwell Collins, has agreed to provide outside consulting for the FITSat team. Florida Tech faculty will provide further “graybeard” consulting.

Tammy StuartPR Intern

Letsky/TomimbangPublic Relations

Christian PedersonPayload Specialist

Jerry TranPayload

Naim TorlhkovicTesting

Tran/BarbourI&T

Christian PedersonThermal Control

Christian PedersonMission Analysis

Naim TorlhkovicTechnician

Faye TomimbangMechanical

Michael LetskyRF/Electrical

Christian PedersonAvionics

Naim TorlhkovicAvionics

Jason BarbourAvionics

Letksy/TranPower

Faye TomimbangAvionics Lead

FITSat TeamFlight Software

FITSat TeamSoftware

Michael LetskyRF/Comm

Torlhkovic/PedersonMission Ops

Barbour/TorlhkovicLaunch Ops

Barbour/TranSystems Engineer

Jerry TranProgram Manager

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Summary

In response to Florida Tech’s AO for a Senior Engineering Capstone Project, FITSat Inc. is proud to present this Technical Proposal for review. This technical proposal has shed some light on the design approach of FITSat-1.

As this is a preliminary design, FITSat-1 may undergo several revisions in order to best

meet the contractual obligations of Florida Tech’s AO. However, the basic approach still applies. FITSat Inc. will emulate the design lifecycle of an actual spacecraft program. This will involve the development of all systems required for a normal spacecraft. Namely, these program components include launch operations, mission operations, mission analysis, mechanical, power, communications, integration and testing.

The expected result of FITSat Inc. is the successful acquisition of data from the primary

payload of the space-rated FITSat-1. Failing that, FITSat Inc. will design and build a spacecraft bus at a lower standard and cost.

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References

1 Larson, Wiley J. 1992 Space Mission Analysis and Design, 3rd edition, Torrance, California: Microcosm, Inc.

2 University of Santa Clara: Artemis Program - http://screem.engr.scu.edu/artemis/ 3 Stanford OPAL - http://ssdl.stanford.edu/opal/

4 Stanford CanSat - http://ssdl.stanford.edu/cansat/

5 Stanford CubeSat - http://ssdl.stanford.edu/cubesat/ 6 StenSat - http://users.erols.com/hheidt/ 7 California Polytechnic State University (PolySat) - http://www.calpoly.edu/~aero/polysat/ 8 CanSat Delivery System - http://www.ae.utexas.edu/~cansat/usss99/ 9 John David Anderson, 1989. Introduction to Flight, Third Edition, New York, New York: McGraw

Hill College Division. 10 ARLISS (a Rocket Launch for International Student Satellites) – http://ssdl.stanford.edu/arliss/ 11 Valdez, Adelia. “The Artemis Project: Picosatellites and the Feasibility of the Smaller, Faster,

Cheaper Approach,” Santa Clara University. 12 University of Tokyo: CanSat Imaging Data

http://www.space.t.utokyo.ac.jp/cansat/arliss/analysis/ana3.html