SEDSAT-2 Mission Feasibility Study

From SEDSWiki

A .pdf version of this document is also available.

Contents 1 Introduction 2 Subsystems 2.1 Payload 2.1.1 Optical observation 2.1.2 Detecting radar sources on earth 2.1.3 Maintaining constant exposure to sunlight 2.1.4 Power beaming to the surface 2.1.5 Interferometry in space 2.1.6 Surveying pollution 2.1.7 Wireless sensor network(WSN) 2.2 Power 2.3 Structure 2.4 Data handling 2.5 Communications 2.5.1 Antenna 2.5.2 Transceivers 2.5.3 Terminal Node Controller 2.5.4 RF (Radio Frequency) Amplifier 2.5.5 Amateur Radio Communication. 2.6 Ground control 3 Project Coordination 3.1 Communication 3.2 Cost 3.3 Schedule 4 Future Staff Plan 5 Contributors 5.1 Mission Concept Team 6 Appendix A: MCR Board Comments 7 Appendix B: Response to MCR Board Comments

Introduction

SEDSAT-2 is a CubeSat project that will be designed and built by members of the organization Students for the Exploration and Development of Space (SEDS). SEDSAT-2 is the second satellite mission designed by SEDS. The first mission, SEDSAT-1, was designed by members of SEDS-USA and launched as a secondary payload aboard the Deep Space 1 mission in 1998. As a result of that project, many members finished their university degrees and were placed in positions in the space industry.

The purpose of this document, the SEDSAT-2 Mission Feasibility Study (MFS), is to summarize the current work and discussions done by the Mission Concept Team in the initial development of SEDSAT-2. This is not a design document; the MFS considers the challenges that are expected in the design of SEDSAT-2 and an initial plan for overcoming the challenges. The majority of the work done was in researching concepts for major satellite subsystems. The MFS is not only remarkable for what it contains but what it does not contain: the experiences encountered by six students in four countries learning how to communicate, collaborate, and build a functional satellite. These are not skills that can be expressed in text, but will be useful as more students are recruited to assist with the design and manufacture of SEDSAT-2.

Subsystems

Payload

Optical observation

The type of device required for measuring light in the visible or ultraviolet/infrared spectrum is dependant upon the application we want to use it for (i.e., wavelength). For observation in the visible spectrum, a standard CMOS or CCD chip is a very likely candidate, for example, the Kodak kac-00400. Special consideration should be taken when looking into the optics, since it must be able to withstand extended temperature ranges. The AAU Cube built by Aalborg University in Denmark used a custom built lens. A standard camera implementation is a very typical CubeSat payload. Looking at a specific set of wavelengths might be more innovative. Good attitude determination and control would be important to ensure that we are able to observe our chosen target.

In the event of an imaging payload, compression might need to be looked at and either implemented in hardware or software. Important factors here are power consumption, cost, and potential image degradation. One candidate codec is the JPEG2000 lossless codec (http://www.jpeg.org/jpeg2000/).

One possible subject for observation is the Sun, an unusual mission for a CubeSat. Infrared (IR) and ultraviolet (UV) are the most interesting bands. More energetic radiation might be considered if we have the possibility of using detectors that can record it. Special filters would probably be required. A typical CubeSat polar orbit would probably not be very well suited for such a mission.

Observing celestial objects would require advanced/costly instruments with high sensitivity. We would probably be looking at light in the non-visible region or receiving radio signals, potentially even looking for laser signals similar to the optical SETI experiment although this is extremely advanced. A polar orbiting satellite would be less than ideal for this kind of mission. Observing the oceans of the Earth could possibly be done using a standard imager, since one could detect this by studying wavelengths in the visible light region. Several larger satellites are already looking at this. Looking at a specific type of algae would be unique for a CubeSat mission (http://www.esa.int/esaEO/SEMH7W2VQUD_planet_0.html).

Total cost for a standard camera system with jpeg compression would be around US$50 to US$150. This figure is based on the usage of a standard CMOS image sensor and other commercial components. This does not include the price of the optics, as this may greatly vary in cost. Spark Fun Electronics produces an example of a very small camera module which also integrates JPEG compression and a microcontroller (http://www.sparkfun.com/commerce/product_info.php?products_id=7906). If implemented with a standard CMOS imager approximately 35-100mW, depending on resolution and compression method.

Detecting radar sources on earth

Another payload concept is detecting radars such as ship radars or larger radar installations using relativistic Doppler measurements to “triangulate” their positions. This might be used to detect unauthorized presence in restricted waters or similar applications. Radars emit very strong pulses that might be picked up by the satellite with the proper receivers, however ship radars probably have a very narrow transmission angle so they might be more difficult to detect from orbit.

Maintaining constant exposure to sunlight

If the CubeSat were put into orbit around the Earth so that the plane of its orbit faced the sun, the attitude of the satellite could be controlled so that its same side always faced the sun, and the satellite would never enter the Earth’s shadow. In this way, power supply would be constant, and there would be no need for batteries or chargers, assuming the power output of the sun were constant. Also, with the same side of the satellite always toward the sun, shielding could be reduced by not needing it on all sides of the satellite, and deployable solar panels could easily be used to increase the available power.

However, the plane of the satellites orbit would not remain facing the sun. Every three months it would rotate by 90 degrees, and then the solar arrays would be useless. A possible solution to this problem could be to include a high-efficiency thruster on the satellite to allow its orbit to track the sun. Such a maneuver could be accomplished by firing the thruster incrementally at the point of it’s orbit highest above the plane of the ecliptic, pushing the satellite toward the sun (assuming the Earth orbiting counter-clockwise when viewed from above the plane of the ecliptic, and the satellite orbiting the Earth counter clockwise when observed from the sun). This action would result in the orbit of the satellite pivoting around the line connecting that point and the corresponding point of the orbit on the other side of the Earth. Firing the thruster a minute amount in this way each orbit would allow the plane of the orbit to maintain its attitude relative to the sun year-round.

Clearly, such a thruster would have to be extremely small, light weight, and most importantly, efficient. An ion engine (the exhaust being positively or negatively charged ions accelerated by a magnetic or electric field) would be perfect, from an efficiency perspective. Building and testing such an engine could be the goal of the mission. This mission would depend on the availability of an appropriate orbital insertion.

To be more specific: Assuming an Earth radius of transparency of 6500 km, a satellite orbit altitude of 400km (http://www.hit.ac.jp/~satori/hitsat/system/hiteps-e.htm), and an infinite Earth-sun separation, the maximum angle through which a satellite orbiting the Earth can see the sun on a continual basis is 0.6 degrees, which corresponds to 0.6 days. It also means that if the orbit control propulsion concept is to work, it must fire about twice per day. With a time per orbit of approximately 90 minutes, it must therefore fire about once every 9.6 or 10 orbits.

Power beaming to the surface

Power beaming would be another mission to test a new technology. If excess power could be generated by the satellite and successfully beamed back to Earth, the satellite would effectively become an orbiting solar power plant. Ideally one could combine this technology with the solar pointing payload design. As the efficiencies of solar panels increase (the current landmark being 40.1%), the desirability of this combination increases, assuming that it is not proven more feasible to place the solar panels at Lagrange points.

Interferometry in space

A remote sensing (RS) enabling technology that we could test with our CubeSat is interferometry. This is used commonly in astronomy where two telescopes are aimed at the same object, so that their combined data attains the resolution of a single telescope with the diameter of their separation and combined radii. (Resolution increases with the diameter of the telescope.) If we could implement such a mechanism in space, we could attain a much higher resolution.

The way to do this is to use an extendable antenna. The satellite is built in two pieces and connected with an extendable telescoping antenna. Each side could each have paraboloid faces with extendable antennae. Given the existence of an extension mechanism, it could be used to simultaneously deploy extra solar panels between the twin satellites, and deploy the real antennae as well. Apparently a tiny camera (<3 mm × <3 mm × <3 mm) is already in development, although the project has not yet been made public. If it was to be completed on time, it could be incorporated into the satellite, making the interferometer more feasible in terms of mass and volume requirements.

Surveying pollution

If we were able to gain sufficient resolution with an interferometer, we might have the option of studying (the oceans) in greater detail. Such a study might contain valuable information about pollution level and type, and contribute to awareness of the extent of pollution of the Earth's oceans, and its distribution and origin.

Wireless sensor network(WSN)

By installing a network of small, autonomous sensors inside the satellite we could potentially get extremely accurate data on the internal conditions of the cube, such as temperature and pressure. This could be valuable for future CubeSat missions and could also prove a unique opportunity to test this kind of technology inside a satellite. These are usually very small MEMS-enabled devices that include two-way communications via bluetooth, laser or zigbee (or other RF standards), a programmable microcontroller, one or more sensors and a power source (http://robotics.eecs.berkeley.edu/~pister/SmartDust/)]. The feasibility of using these devices is largely dependent on the size of commercially available sensors modules (motes).

Power

Solar cells will clearly provide the main source of power. There are two kinds of possible orbits: those that involve the satellite traveling through the Earth's shadow for part of it's orbit and those where it does not. If it never enters the Earth's shadow, solar cells can be used as the satellites only power source. However, it has been found both that the opportunity for such an orbital insertion is very rare, and that maintaining such an orbit would involve a propulsion system that might not be feasible when taking into account budget, time, mass, and volume constraints. Much more common orbits involve the eclipsing of the satellite. One option could be to power down the satellite during these periods, simplifying it's design and construction by not requiring batteries and battery chargers. However, this might introduce other design problems, and might needlessly compromise the mission objectives.

The system in use on all satellites researched so far involves solar cells as the primary power supply, working in concert with lithium-ion batteries and battery chargers, and voltage regulators for power distribution. Space rated solar cells are mainly classified by their efficiency, which has varied roughly between 12% and 27%, the 27% efficiency cells being priced at approximately US$300 each (http://CubeSat.ece.uiuc.edu/Power.html). The latest technology has now produced an efficiency of 40.1% (http://www.energy.gov/news/4503.htm), although these may or may not be commercially available and their price would be correspondingly high. However, a possible mission objective might be to test the performance of this new technology in the space environment.

Another alternative power source is power beaming from the Earth, although this would apparently shorten the life of the satellite, and its feasibility has not yet been established. It has been concluded that for the moment, the most likely solution will be the solar cell battery combination, unless the mission objective is to specifically test another technology.

Batteries and battery chargers have been found in general to cost under US$100 each. However, no pricing information has been found on the specific costs of the batteries and chargers used to date on satellites, although two companies, Saft, and Nec-Tokin have been found, who can be officially contacted for cost information.

Saft has developed widely recognized cells and batteries for the satellite market using rechargeable lithium-ion technology. Li-ion batteries are used in all satellite applications such as GEO, MEO, LEO and Microsats. Recommended for all satellite batteries, the Li-ion offers:

• Smaller and lighter than conventional technology (30 to 50% weight reduction compared to Ni-H2)
• High specific energy: 125 to 165 Wh/kg
• Low thermal power and high energy efficiency leads to smaller solar panels and battery radiators
• Easier launch pad operations
• No memory effect
• An energy gauge given by the voltage
• Modular flexibility in designing batteries

Saft's Li-ion cells are specially adapted for space applications from field-proven designs. Saft's lithium-ion modular systems are built from the proven types of VES cylindrical cells and MP prismatic cells. The amount of power expected from the solar cells is 3.75 W, assuming incident 2 2, total effective surface area is 0.0103 m , and solar panel efficiency is solar energy is 1350 W/m 27%.

The battery/charger/converters setup introduces further energy losses in the system, depending on the statistics of the particular components that are used. Converters were found with efficiencies over 80% that cost less than US$10 (http://www.national.com/pf/LM/LM2670.html).

The average available power estimation taking into account time in eclipse and efficiency losses is approximately 2 W. The total estimated cost for power components is displayed in Table 1. Extensive documentation of similar satellite power subsystems was performed by the nCube CubeSat team at Narvik University College (http://www.ncube.no/project_documents/masterthesis/DiplomPower.pdf and http://www.ncube.no/project_documents/masterthesis/vedleggPower.pdf).

Table 1:Estimated cost for power components

Component Cost(US$)
Three solar cells	900
One lithium-ion battery	200
Battery Charger	10
Five voltage converters	50
Total	1160

Structure

Two main issues were considered in the structure: material and machining. If we use either Al-7075-T73, 6061-T6, these are the costs in America and UK. The team undertaking the construction of the structure should ensure that the aluminum types are available locally. We can also use a material similar to the ones mentioned above. Extra material will need to be purchased to consider margin for material that might be wasted.

The machining costs depend on which university undertakes it. The cost for the structure on a whole should cost approximately US$500 to $1000. This is a cost of not only building outer structure but also building internal support structures. It has been over-estimated by at least $100 to include time overruns.

Data handling

The plan for SEDSAT-2 is to use all gratis software with source code available on the internet (if not developed in-house), or a vendor-provided SDK. For hardware, important points to consider are the component temperature ranges of operation, form factor, and power consumption. With worthwhile trade-offs, important limitations can be increased or alleviated. Hardware researched for use on SEDSAT-2 includes:

• H8S 2674R CPU, manufactured by Renesas Technology, Inc.
• MC68332 microcontroller, manufactured by Freescale Semiconductor, Inc.
• C8051F123 microcontroller, manufactured by Keil
• Atmega8 flash program memory, manufactured by Atmel Corporation

Communications

SEDSat-2, being a CubeSat, will operate in the amateur radio frequencies. It will do this via the AX.25 protocol, a data link layer protocol used extensively for amateur packet radio networks. Further, as shown by a study of previous CubeSat missions as well as taking into account the possibility that the satellite will fly in low Earth orbit (LEO), the best possible ranges for operation would be the VHF and the UHF frequencies.

Very high frequency (VHF) is the radio frequency range from 30 MHz to 300 MHz. VHF is also commonly used for terrestrial navigation systems (VOR in particular), Marine Communication, and aircraft communications.

Ultra high frequency (UHF) designates a range (band) of electromagnetic waves whose frequency is between 300 MHz and 3 GHz. UHF is widely used for two-way radio communication (usually using narrowband frequency modulation, but digital services are on the rise) by both public service agencies and the general public.The main advantage of UHF transmission is that its high frequency means it has a physically short wave. Since the size of transmission and reception equipment (particularly antennas) is related to the size of the wave, smaller, less conspicuous antennas can be used than with VHF or lower bands.

The UHF range is recommended for two reasons. First, the antenna dimensions are reduced, as shown by the following equation: ? = c /f, or wavelength equals the speed of light divided by the frequency. This implies that the higher the frequency, the smaller the wavelength, upon which the size of the antenna is dependent. Second, less power is consumed. Supposing that power is mainly acquired through solar cells, we have a total of [number] W. Looking at the past CubeSat missions, it is seen that the communications system on average consumes power in the range of 1.5 to 2 W. Thus minimizing power consumption should factor into the design considerations.

In terms of hardware, the communication subsystem is generally composed of the following components: antenna, transceiver, terminal node controller and RF (radio frequency) amplifier.

Antenna

Quarter-wave monopole is a monopole antenna (formed by a single conductor) of a length corresponding to 1/4 the wavelength, that behaves like a dipole antenna. It does this by using a conductive surface as a reflector; the current in the reflection has the same phase and direction as that of the real antenna. Adding the effect of the two, we get a half-wave dipole antenna. The conductive surface that acts as a reflector is usually the Earth. Since in space, it is not possible to use the ground as a reflector, one can use two quarter-wave monopoles to realize a half-wave dipole, such as done by the ION team.

A half-wave dipole antenna is formed by two conductors whose total length is half the wavelength. The current in this antenna is described as having a sinusodial distribution with a maximum in the center. The gain of this antenna, the ratio of the surface power radiated by the antenna and the surface power radiated by a hypothetical isotropic antenna, is 2.14 dBi.

The MEROPE satellite chose the dipole over the monopole for the following reasons:

• it allows for enhanced directional gain over an omni antenna since a dipole does not require a ground plane and is not shadowed by the spacecraft
• the use of one antenna for both up-/downlink frequencies avoids the addition of duplexer hardware
• it is relatively simple to tune to both uplink and downlink frequencies
• it reduces the number of potential points of failure, and will radiate if only one element deploys.
• A quarter-wave monopole antenna system is recommended because of its greater directionality and power density.

Transceivers

In choosing the transceivers, one must take into account the operating frequencies of the satellite. Because the uplink frequency and downlink frequency must of necessity be different, it is probable that two transceivers will be used, as the CUTE-1 mission did. [It is, however recommended that a transceiver that can cater to both frequencies be used, such as the blank mission did].

The transceivers used by the CUTE-1 mission were the DJ-C1T and DJ-C4T. A more recent version is the DJ-C5T, which costs about US$160. Another option would be the Microhard Corporation modems, of which the price ranges from US$179 to US$595.

Another factor that might come into play in choosing a transceiver is the modulation. Most transceivers have a specific modulation schemes and the choice of the modulation and eventually the transceiver itself may depend on the efficiency of the modulation and power available. In particular, the following digital modulation schemes are noted either for their efficiency or for their low power consumption:

GMSK is a kind of continuous phase frequency-shift keying. The baseband modulation is generated by starting with a bitstream 0/1 and a bit-clock giving a timeslice for each bit. This is the type of modulation used in Global System for Mobile Communications (GSM) - although that system includes differential encoding of the bitstream, which makes the overall modulation behave like a kind of continuous phase Binary Phase Shift Keying with constant envelope. GMSK has high spectral efficiency, but it needs higher power level than for instance Quadrature phase-shift keying (QPSK) to reliably communicate the same amount of data. QPSK is a form of phase shift keying (PSK) modulation, the QPSK can encode two bits per symbol.

MSK is a particularly spectrally efficient form of coherent frequency-shift keying. In MSK the difference between the higher and lower frequency is identical to half the bit rate. As a result, the waveforms used to represent a 0 and a 1 bit differ by exactly half a carrier period. This is the smallest FSK modulation index that can be chosen such that the waveforms for 0 and 1 are orthogonal.

The transceivers used by the CUTE-1 mission were Alinco’s DJ-C1T and DJ-C4T. A more recent version is the DJ-C5T, which costs about US$160. Another option would be the Microhard Corporation modems, with a price range from US$179 to US$595; the MHX2400 model in particular, while expensive, would give optimal performance.

Terminal Node Controller

The terminal node controller is necessary in order to be able to connect to and communicate through the AX.25 protocol. The terminal node controller contains a microprocessor and an implementation of the protocol in firmware.

RF (Radio Frequency) Amplifier

Most of the previous missions have included this onboard. Considering that this is a remote sensing satellite that will most probably be placed in low Earth orbit, an amplifier for the signal will more likely be necessary than not.

Amateur Radio Communication.

CubeSats generally communicate via amateur radio frequencies and as such, rules and regulations that govern the use of this must be strictly observed. Information on this subject can be found at the sites outlined at the CubeSat web site, http://CubeSat.atl.calpoly.edu/pages/documents/developers.php.

There are three key points that must be considered:

• The data cannot be encrypted. All communication with the satellite must be in the clear and available to all.
• There must be at least one member of the SEDSat-2 team that has an amateur radio operator license. Without this, the team’s communication with SEDSat-2 will be considered illegal.
• There must be no pecuniary interest in the satellite and/or its mission. This means that while the satellite may be sponsored by corporate entities, all support must be freely given, i.e., without any expectations of return whatsoever.

The estimated cost for communications components are displayed in Table 2.

Table 2: Estimated cost for communications components

Component	Cost(US$)
Transceiver	160
Antenna	20
Terminal node controller	150
RF amplifier	220

Ground control

A Ground Control subsystem was also considered as an important aspect of the SEDSAT-2 mission and will be studied in further detail in future reports. Attitude control was identified as a potential task and subsystem on the mission, but not as a necessary task unless required by payload or electrical power needs.

Project Coordination

Communication

Operating a project with participants in various universities, countries, and continents presents a unique challenge. One important lesson learned in SEDSAT-2 is the use of Greenwich Mean Time for scheduling meetings. For communications between the mission concept team members, an email list (sedsat2@seds.org) was created. This medium is used for team discussions and announcements. Online meetings typically occur every weekend using MSN Messenger to communicate via text chat.

The mission concept team maintains a blog at http://blogs.seds.org/sedsat2. The function of the blog is for discussing selected items of the mission design in public in order to engage the public and display team news.

Information for the mission concept phase was managed at SEDSWiki, http://wiki.seds.org. The use of wiki software allowed all members of the team to add, edit, and manage information for the project, including the simultaneous development of this report. In addition to SEDSWiki, ProtoForge (http://www.protoforge.org), a tool created by SEDSAT-2 adviser Aaron Schultz, will be used to manage mission requirements during the next phase of mission development. During the next phase of the mission, http://www.seds.org/sedsat2 will be introduced as the public face of SEDSAT-2 information.

The first in-person meeting is scheduled for 22 September 2007 at Vellore Institute of Technology in Vellore, Tamil Nadu, India. This will coincide with the 2007 SEDS International Conference hosted by the SEDS chapter at VIT and the International Astronautical Congress hosted in Hyderabad, Andhra Pradesh, India. This meeting will constitute the Preliminary Design Review of SEDSAT-2.

Cost

The total cost of the SEDSAT-2 program will be calculated in US dollars for this study. Program costs will include flight hardware; flight software; facilities and equipment; travel costs for team members to workshops, conferences, and meetings; launch systems. Subsystem leads will responsible for raising funds for their subsystem; there will be no central bank account for this project. The Program Manager(s) and Advisers will assist with fundraising.

Schedule

• 16 Jan 2007: Mission Feasibility Study (MFS).
• 17 to 24 Jan 2007: Mission Concept Review (MCR). The MCR will be the first program milestone, summarizing work-to-date and a plan for future development. The MCR board will comments and recommendations on MFS. Comments from the MCR board will be added to the MFS as an appendix.
• 28 Jan 2007: Open team and payload applications. After receiving comments from the MCR board, the team will respond to the comments. This response will be added as an appendix to the MFS. Then applications will be distributed to SEDS email lists and to chapter contacts.
• 25 Feb 2007: Close applications. After applications are closed, the SEDSAT-2 advisers will choose and distribute team members to various subsystems.
• 11 Mar 2007: Team chosen and distributed.
• 8 Apr 2007: Mission Definition Review (MDR). The MDR will detail trade studies for choosing the SEDSAT-2 payload and the basic requirements for supporting the payload.
• 3 June 2007: System Definition Review (SDR). The SDR will be the milestone for defining mission requirements and allocating them to subsystem tasks. After the SDR, the system and operation will be developed enough to begin designing and acquiring system components.
• 22 Sept 2007: Preliminary Design Review (PDR). The PDR will demonstrate that the preliminary design meets system requirements. The PDR will verify that the preliminary design is acceptable and that detailed design can proceed. The PDR is scheduled to coincide with the 2007 International Astronautical Congress in Hyderabad, India.
• 31 Jan 2008: Critical Design Review (CDR). The CDR will verify that the detailed technical design has been completed or nearly completed and that manufacture of the satellite can proceed.
• Sept 2008: System Acceptance Review (SAR). The SAR will verify that the completed system is ready for integration and testing. The SAR is scheduled to coincide with the 2008 International Astronautical Congress in Glasgow, Scotland.

Future Staff Plan

After the MCR, an advertisement will be sent electronically to various SEDS and other student-interest groups to apply to participate on the SEDSAT-2 design team. A team of at least 25 students will be needed to fully staff the SEDSAT-2 program. This allows 3 students per subsystem (Power, Payload, Structure, Data Handling, Communications, and Ground Control), 5 students to manage Systems and Integration, and 2 students to maintain information and educational programs. Students may apply to be team members, subsystem leads, or program manager(s).

Preference for design team will be given to members of SEDS or students that apply to join SEDS as part of the SEDSAT-2 application process. Subsystem Leads must demonstrate technical knowledge and have an adviser from their university that can participate on the project.

Contributors

Mission Concept Team

• Geoff Hilton.
  • Geoff is Canadian, he began playing with computers at 8 years old, and started programming in elementary school. Now a programmer by profession, he is interested in such things as Artificial Intelligence, avionics and complex systems in general such as large-scale server/client software frameworks or as the case may be -- satellites, SEDSat-2 being a prime example. His plans include getting a University degree in Computer Sciences, and eventually ideally having a career in the space industry. Geoff is also involved in a project with the Canadian Space Society.

• Mike Jensen.
  • Michael is a Canadian with an Honours Bachelors degree in theoretical physics, currently in training at the Quebec school of circus arts. His twin passions in life are circus arts and aviation/aerospace engineering, and his life's goal to find a way to pursue them actively and in parallel. He would be thrilled to be a pilot for one of the first human missions to the moon and mars. He also loves physics, languages, and programming, and is actively looking for ways to travel and contribute all of his experience and passion to international development and the inspiration and empowerment of youth, and people everywhere. His current projects include building a glider along with a hangar (and safety system as he has no piloting experience and very little glider building experience), web development for SEDS Canada, and SedSat2. He has applied for graduate scholarships in engineering.

• Pradeep Nair.
  • Pradeep is an Indian from the commercial capital of India, Mumbai. He's currently pursuing undergraduate studies in mechanical engineering in a college along the edges of Mumbai. Although interested in space, he has not been able to take up any major project till now. This is his first space project and a great learning experience.

• Tom Nordheim.
  • Tom is a Norwegian studying for a bachelors in space science and robotics at the University of Wales, Aberystwyth. He graduated from the Norwegian space engineering high school at Andøya in 2005.While at Narvik University College in 2005/2006 he became involved in the HiNcube satellite program, and ended up on the payload design team, working on the imaging system. In the summer of 2006 Tom became involved in the cansat concept and after winning a national competition, he and the two other members of Team XR demonstrated their cansat at the IAC 2K6 in Valencia,Spain. His technical interests include embedded electronics, imaging systems, rocketry and robotics. Tom was enslaved by UKSEDS in a lunch break at a summer university course in Kiruna, Sweden. He also has plans to start a local chapter of SEDS at his university. Tom likes challenges.

• Lavina Parwani.
  • Lavina is an Indian citizen living in the Philippines, where she was born and brought up. Although her original focus in high school was the liberal arts, her dream to become an astronaut brought her to the thorny path of the engineering major. She started out as with Aerospace Engineering at Embry-Riddle Aeronautical University--where she first met SEDS--but financial reasons compelled her to return home, where she is majoring in Mechanical Engineering. Apart from working on SEDSat-2, she is also currently working with a university team to build a solar-powered car, which they hope to race in Australia at the World Solar Challenge. Her other interests are astrophysics, astronomy, writing and chocolate.

• Amrut Yalagi.
  • Amrut Yalagi is an Indian from the historical place called 'Bagalkot'. He's currently pursuing undergraduate studies in BSC-III year with Physics as a major. He is also the Secretary of The Planetary Society of Youth. Last year he received prestigious award i.e. 'Young Achiever' from the Karnataka Govt. for achieving excellence in the Space Education. He is currently working on Weather Remote Sensing and Village Resource Centre Project with The Planetary Society of Youth.

The Mission Concept Team was assisted by four SEDS alumni advisers:

• Kirk Kittell, Systems Engineer, Orbital Sciences Corporation
• Chris Lewicki, Senior Systems Engineer, Jet Propulsion Laboratory
• Aaron Schultz, Systems Engineer, Honeywell
• Purvesh Thakker, PhD Candidate, University of Illinois at Urbana-Champaign

Appendix A: MCR Board Comments

Please see the SEDSAT-2 Mission Concept Review page.

Appendix B: Response to MCR Board Comments

• Pradeep: I believe that the comments that we received showed that though we were organising data we were not able to present it as such as information. Some reviewing of the way we add information to our documents has also been changed accordingly. I, personally, writing my first Study contribution feel I tried to get too technical and thank the reviewers for correcting me.

• Tom: The usefulness of an optical observation payload is definitely something that needs looking into, with regards to resolution and compression, but I feel that at such an early stage it is understandable that we don't have any data on many of these things since we don't actually know the orbit: Therefore we do not know any parametres for bandwidth, power, radiation exposure or opportunities for observing various targets(such as the sun). I still feel however that compression is neither completely unjustified or required, this would have to be decided when more data on bandwidth and the actual target of the observation is available. With regards to the WSN, I believe this section should have been dropped from the MFS, since we came up with many of the same comments when discussing it as a group, and more or less came to the conclusion that it would be pointless.I do agree with the fact that it might be the best option to avoid complex payloads, and rather go with something quite standard(like a camera for taking "pr" pictures for no perticular scientific use). Having a successful cubesat is more than what most achieve, so this should be a our primary goal. Finally I would like to thank everyone who submitted their comments and especially to Ed Chester for his critical and detailed evaluation of our report.

• Geoff: I have no special comments going beyond Tom's, but would also like to thank you of the MCR board for all the excellent comments on our works.

• Lavina: I'd like to thank the reviewers for their comments. In writing the MFS, I had only a vague idea of what was required in it, and their comments have helped give me a clearer definition. Perhaps, if we had more time, we could edit the document to include vital information that is missing from it.
  • Response to comments by Gregg Maryniak:
    • Wireless Power Transmission: he is absolutely correct. It is not practical considering the limitations we have to work with. However, I do like his suggestion of beaming between two parts of the satellite. We may even have a 1x2 CubeSat to carry this out.
    • Suggestions: If the subsequent payload team can take up the challenge of developing such an engine, then it might be best to make this the mission of SEDSat-2, given Gregg's insight as to its usefulness. After all, our challenge in the payload discussions have been to figure out what could be most feasible as well as useful.
  • Response to comments by J.F. Gauthier:
    • While it is true that the actual mission takes a back seat to the experience gained by the students, I still believe that we need to come up with an interesting, practical and useful mission; this is actually part of the learning process. As engineers, this is precisely our challenge, and the more experience we have with that part of it, the better.
    • It is also true that the range of payloads given in the Mission Feasibility Study makes it seem like it lacks focus. Perhaps what we should have done was to pick one that we felt was most feasible, and list the others as backup. After all, the payload team that will eventually get picked may not have the capabilities for our specific mission outline.
    • Heritage, while providing security, also provides limitations. If the real goal of this is for us to learn, this means that we also have to learn about the risks that come with designing and engineering something like this. Both failure and success give us entirely new knowledge, so it is never a waste. Not to mention, the success would be infinitely sweeter. Knowing the members of the team that exist now, it'd be a shame not to capitalise on the creativity that exists.
  • Response to comments by Ed Chester:
    • The antenna system only has one conclusion, though I understand why it may seem like it has two. What I was doing was presenting the arguments for the other side before giving my conclusion.
    • I haven't gotten a chance to talk to AMSAT yet. However, I am under the impression that choice for modulation, etc, will be affected (at least in part) by the choice of payload. It is for this reason that I've outlined the possible options without a real conclusion.

• Michael Jensen:Here are some of my preliminary coments on the MCR: First, it comes out as abundantly clear that we need to decide as early as possible on a specific mission to perform so that we can spend as much of our time and effort as possible on actually designing, building, and testing it. Also, that we should choose a simple mission, and go about performing it in the simplest way possible, maximizing our chances of success. A third thing that seems generaly agreed on is that having an attitude control system would be highly desierable. I agree on all of these points :)

For some specifics, Gregg Maryniak of the X-prise foundation identifies small high efficiency micro-thrusters as a payload that would be of great value to the space exploration community, and the world. This is exciting and encouraging to say the least. However, the belief has also been voiced that an accomplishment of this magnitude might not be possible within the time frame of the mission. With these views in mind, the possibility of such a payload will be throughly investigated :)

Finally, I would like to thank Ed. Chester for his extensive coments and feedback. He has put a great many things clearly and preciselly, and I believe following his advice, which is echoed in many ways in the coments of the others, will produce much higher quality, more informative, and more convincing documents in the future, along with a much better defined mission and operating procedure, and a more successful satellite :)