Seminar Report On
Nano Satellite
Submitted By
Shishu Priya Darshi
2007EEC08
SEM- VII, B.Tech
SMVD University, 2010
School of Electronics & Communication Engineering
College of Engineering
Shri Mata Vaishno Devi University
Katra, J&K
2010
Approved
Director- SECE
pg. 1
CERTIFICATE
TO WHOM IT MAY CONCERN
This is to certify that the Seminar entitled “NANO-SATELLITE” has been
submitted by SHISHU PRIYA DARSHI (2007EEC08) under my guidance in
partial fulfillment of the degree of Bachelor of Technology in Electronics
and Communication Engineering Shri Mata Vaishno Devi University Katra, J&K
during the academic 7th Semester.
Date:
Place:
Coordinator: Director-SECE:
Mr. Ashish Suri Dr. Vipan Kakar
pg. 2
INDEX
Chapter 1.1- Introduction ………………………….…..pg4
Chapter 1.2- Motivation…………………………………pg5
Chapter 2- Technologies………………………………..pg6
Chapter 2.1- Propulsion…………………………………..pg6
Chapter 2.2- Guidance Navigation & Control……………pg9
Chapter 2.3- Command and Data Handling…………… ..pg11
Chapter 2.4- Power System………………………………pg13
Chapter 2.5- Thermal……………………………………..pg15
Chapter 2.6- RF Communication…………………………pg17
Chapter 2.7- Mechanical/ Structure………………………pg19
Chapter 2.8- Instruments…………………………………pg21
Chapter 2.9- Ground Systems……………………………..pg22
Chapter 2.10- Autonomy………………………………….pg24
Chapter 3- Technology Transfer/ Spinoff…………………pg25
Chapter 4- Summary / Conclusion………………………..pg26
Chapter 5- References……………………………………..pg27
LIST OF IMAGES
1. STP nano-satellite Concept……………………….pg4
2. Orbit Concept……………………………………..pg5
3. MCM Concept…………………………………….pg12
4.Nano-satellite Ground Concept……………………pg22
LIST OF TABLES
1.RF Communication Specification…………………pg18
pg. 3
CHAPTER 1
1.1 INTRODUCTION
1.1.1 NanoSatellite?
The term "nanosatellite" or "nanosat" is an artificial satellite with a weight
mass between 1 and 10 kg. including propellant mass.
This type of satellite works together with another nanosatellite or require a
larger "mother" satellite for communication with ground controllers or for launching
and docking with nanosatellites.[1]
[9]
1.1.2Motives of nanosatellite
The main need of nanosatellite is to reduce the cost of satellite. Because of use
of advanced technology, the size of satellite can be dramatically reduced which
ultimately results in manufacturing and maintenance cost. The life of satellite is of 2
years.[9]
1.1.3Working of nanosatellite
These are spin stabilized with its spin axis normal to the ecliptic plane. This
configuration maximizes sunlight exposure on its solar cells, which are mounted
around its circumference
spacecraft operations alive during eclipse periods. At least 5 Watt of power can
be generated by multi junction solar cells, and batteries will keep spacecraft operation
alive during eclipse periods. Nanosatellites are placed in several highly elliptical
orbits. Every orbit shares the same perigee radius of 3Re to 42Re in 3Re increments.
Initially the two nanosatellites per orbit wiil be simultaneously deployed in opposite
directions. This aids in deployer-ship inertia consideration from the angular
momentum generated as a result of deployment. A constellation will require
pg. 4
simultaneous operation of multiple swarms of spacecraft. Perpetuation from
moon,earth and sun ,and other celestial bodies will eventually cause the nanosatellite
to become randomly distributed in space.
[9]
The deployer ship ejects the nanosatellite at 3Re with a minimum spin rate of
20rpm to ensure sufficient stabilization. Each nanosatellite boost itself to its particular
obit by firing its orbital insertion thruster when its spin axis is aligned with the
velocity vector. Pulsing of miniature thrusters will then be reused to orient the spin
axis of nano-satellite for optimum sunlight and communication effectiveness.
The low power available on the spacecraft for the communication subsystem
has created the requirement to send data to earth only during the portion of each nano-
satellite obit near perigee. This amounts to about 4.1 hours in duration for the 12 Re
apogee orbit, and 4.3 hours for 42Re apogee. As a consequence, the onboard memory
must be sufficient to hod a full orbit’s worth data. The longest orbit satellite must be
treated with the highest communications priority when they pass close to Earth.
Satellite in smaller orbits can hold several orbits of data for the same onboard memory
size, and thus need to download to earth less often.[1,6,9,11]
1.2 MOTIVATION
By seeing that our Indian student of IIT Kanpur has made a nano-satellite of weight 1
kg from University Nanosat program, I get inspired to know about nanosatellies. As
satellites are of great importance in electronics and communication engineer life and
this program is complete use of related advanced technology, it was necessary to
know nanosat is based on which technologies.[3,8]
pg. 5
CHAPTER2
Technologies
Nanosatellite require technologies that radically reduces the mass and power of
components without compromising performance. In addition to miniaturizing
components, we have to integrate similar function across subsystems. For example all
subsystem electronics, including instruments, can be integrated within the C&DH
subsystem. Simple, effective methods of thermal control are essential to keep nano-
satellite operational during extreme temperature variations. Autonomy is a critical
technology that impacts every subsystem. The nano-satellite ground system must be
kept inexpensive, simple, and made inter-operable with other missions.
2.1 Propulsion
There are various technology for propulsion- Chemical propulsion, Electronic
propulsion (pulsed plasma and field emission EP) etc. Chemical propulsion is best
suited because of relevant impulse bits and versatile nature. Each nano-satellite must
raise its orbit apogee to the appropriate radius (from 12 to 42Re). Then it must
reorient the axis of the spinning nano-satellite from the velocity direction (within the
orbit plane) to its science attitude (perpendicular to elliptical plane). These ΔV and
ACS thruster can have independent or shared feed system depending on whether a
single type of propellant can be used for both application.
The following products are most desirable for our applications, and are
actively being persued for development: miniaturized, solid propellant ΔV motors
with a low cost/mass ignition system; miniaturized liquid propellant thrusters
(hydrazine or advanced monopropellant); ultra low power cold gas micro-thrusters;
low cost tanks and other feed system components; low power gas generators for
liquid-storage cold gas feed systems; and micro-machined solid propellant motors for
attitude control firings.
A solid propellant motor is an attractive option to provide the necessary ΔV
for injection into the final mission orbit. Because the initial mission apogees of the
nano-satellites are not tightly constrained, the small ΔV errors typical of a solid motor
are acceptable. However, many challenges remain in development of an acceptable
motor. The motor must be able to accommodate a range of ΔV requirements without
pg. 6
incurring costly changes to nano-satellite’s mechanical interface. The thrust level
must be limited to ensure that the baseline spin rate of 20rpm is adequate to maintain
the nano-satellite attitude.
Miniaturized liquid propellant thrusters are another promising technology.
Liquid propellant offers storage density and performance comparable to solid
propellants, but with the addition capability to restart the engines for multiple burns. It
appears that the performance of hydrazine is inadequate for the ΔV portion if the
autonomous nano-satellite weight is to be kept below 10kg. However, Hydrazine
could be used with lower ΔV requirements, or for attitude control on spin-stabilized
or three-axis-stabilized nano-satellites.
One potentially near term technology is ultra low power cold gas thruster.
Because of the low specific impulse of cold gas thrusters, they cannot be used for any
substantial ΔV on a nano-satellite but their simplicity and multiple-pulse capability
make them a good choice for attitude control.
Propellant in a cold gas subsystem could also be stored as a liquid. The gas
might be generated by choosing a liquid with a very high vapour pressure.
Solid propellant gas generators could be used as ACS thrusters. Forty-eight 50
mN-sec pulses are required to reorient the nano-satellite after it achieves the required
altitude. Although this could be achieved either by a monopropellant or a cold gas
thruster, it could also be achieved using an array of gas generators. Such generators
are currently under development at NASA’s Lewis Research Center.[9]
2.11 CHALLENGES FOR PROPELLANTS
The power required to operate valve must be reduced by an order of magnitude.
For three axis stabilized applications, the thrust level must be reduced by two or three
order of magnitude.
Additionally, smaller thrusters will require novel thermal design approach to prevent
flow chocking or premature combustion.
The primary challenge in using ultra low power gas thruster is decreasing the
requirement input power for the thruster value by an order of magnitude.
pg. 7
This choice of propellant in a cold gas subsystem, although simple, presents several
problems: the evaporation rate would be highly dependent on temperature; the low
thruster inlet pressure would result in poor performance; and the exhaust could
possibly condense on cold spacecraft surfaces.
Propellant selection, low power ignition, and thruster array packaging are some of the
challenges ahead for this technology.[2,7,9]
2.12 PROPOSAL OF NEW SOLLUTION
Advanced monopropellants, such as those based on hydroxyammonium nitrate (HAN)
and other chemicals, offer all thesity, non-toxicity, and lower freezing point.
Advantages of hydrazine with several additional benefits, including higher specific
impulse, higher density, non-toxicity, and lower freezing point.[7,9,11]
For ultra low power cold gas thruster one likely subsystem configuration is a
blowdown feed system with high pressure gas tank feeding the thruster directly,
thereby eliminating any need for pressure regulator.[7,9,11]
For cold gas subsystem, a gas generator could be used, although this would require
some power input.[9]
pg. 8
2.2 GUIDENCE, NAVIGATION & CONTROL
Guidance Navigation and Control (GN&C) subsystem key technologies and concepts
have been identified to enable the successful altitude determination of spin-stabilized
and three-axis-stabilized nano-satellite for future missions. They include
miniaturization of sun sensor and horizon crossing indicator. The miniature precision
‘fan’ sun sensor will pinpoint the sun virtually anywhere in the entire celestial sphere
with every satellite rotation. The sun sensor will be required to weigh less than 0.25
kg, draw less than 0.1 watts, operate on no greater than a 3.3 volt bus, and meet a 0.1̊
resolution requirement. The miniature horizon crossing indicator has a small bore-
sight FOV that is mounted at an angle off the spin axis. As the spacecraft rotates, a
cone of coverage is formed. The sensor must be capable of detecting Earth from 3 to
5Re with a pointing accuracy of 0.005̊. Total horizon crossing indicator weight and
power will be less than 0.2kg and 0.1 watt, respectively.
To presses the spacecraft spin axis from the orbit plane to the ecliptic normal
requires a nutation damper in conjunction with thrusters. The damper will reduce a 15̊
nutation angle in under a few hours.
There are four concepts of Navigation: Navigation using Magnetometer Data,
TDRSS Onboard Navigation System (TONS), Navigation using Ground stations, and
Navigation using GPS.
Navigation using Magnetometer Data assumes the spacecraft attitude is
known. As the spacecraft passes through a low altitude region of the orbit, the
magnetometer data can be compared to onboard magnetic field model. This
information is processed through a Kalman filter to produce an onboard ephemeris
solution.
TONS was successfully completed on the NASA Extreme Ultraviolet
Explorer (EUVE) mission. This system uses the Doppler shift of the communication
signal from TDRSS to generate onboard navigation solution.
The ground Onboard Navigation System (GONS) is currently being
developed. As with TONES, the GONS program will be evaluated for potential use in
nano-satellite missions.
pg. 9
Of particular interest to Constellation missions is the incorporation of GPS
onboard the nano-satellites, to eliminate ground-based ephemeris generation. This
allows for increased autonomy and simpler, more accurate time resolution onboard
the spacecraft. For GPS to fit within the constraints of a nano-satellite, the receiver
electronics need to be miniaturized into a layer within the C&DH module.[1,7,9,11]
pg. 10
2.3 COMMAND & DATA HANDLING
Developing the Command and Data Handling (C&DH) subsystem for a nano-satellite
presents some unique challenges, with low mass (0.25 kg) and low power (0.5 W)
requirements being the biggest drivers. Advanced microelectronic solutions must be
developed to meet these challenges. The microelectronics developed must be modular
and of scalable packaging to both reduce cost and meet the requirements of various
missions. This development will utilize the most cost effective approach, whether
infusing commercially driven semiconductor devices into spacecraft applications or
partnering with industry in the design and development of high capacity data
processing devices. The major technologies that will be covered in this section
include: lightweight, low power electronics packaging; radiation hard, l ow power
processing platforms; high capacity, low power memory systems; and radiation hard,
reconfigurable, field programmable gate arrays (RHrFPGA). The requirements of a
nano-satellite C&DH subsystem are included in Table 3.
In order to develop a low mass C&DH, a lightweight and low power
electronics packaging method must be used. The packaging method that will be
chosen must have a small volume and small footprint (6cm x 6 cm x variable height).
The packaging technique must provide data on programmable substrates to accelerate
the process of “prototype to flight” with less cost. The packaging technique must also
provide data on compliant interconnects for space use. Figure 6 illustrates one such
electronics package, a multi-chip module (MCM) made by Pico Systems Inc. [Banker
et al., 1998].
This stackable MCM technique enables modularity and scalability for
flexibility in design to meet the needs of multiple missions. The approach shown in
Figure 6 allows for rapid custom designs, fast design iterations and moderate design
costs, while allowing high performance working over required temperature ranges
with radiation tolerance.
A combined effort in the reduction of mass, power, size, and cost is underway to
produce optimal electronics. The CMOS Ultra Low Power Radiation Tolerant
(CULPRiT) system on a chip, and “C&DH in your Palm” are technologies that will
pg. 11
enable the power reduction required for nano-satellites. The goals of these
technologies are a 20:1 power reduction over current 5 Volt technology, foundry
independence of die production, and radiation tolerance.
Every three years memory technology advances enable a doubling of memory
capacity and a halving of silicon area. Memory trends starting in 1996 are toward a
3.3 V core and a 3.3 V I/O, reducing by 1/3 the power for Gbit size solid state
recorders. Trends in packaging technology are enabling denser 3-D stacking in
smaller volume packages for multi-bit stacks in the next three to five years. This will
be accomplished by incorporating Chip Scale Packaging technology where the
package is less than 1.2 times the area of the silicon. DRAM memory will be at the
128 Mbit per die level within the next three years. With these current trends, it
appears promising that an off-theshelf solution is viable for the C&DH subsystem of a
nano-satellite.
Another technology enabling a decrease in volume is the radiation hard,
reconfigurable, field programmable gate array. The RHrFPGA reduces volume by
replacing many logic functions/circuits with one die. The RHrFPGA also allows
concurrent design by decoupling the logic design from the module, shortens the
design schedule, lowers the part count, and eases rework.
The above technologies allow for higher levels of electronic integration,
effectively combining spacecraft subsystem electronics and instrument electronics
into the smallest possible mass, power and volume.[9]
[9]
pg. 12
2.4 POWER SYSTEMS
Total spacecraft power is limited by the small satellite size. The Sun s power density
is 1.35 kW/m2. Assuming 15% conversion efficiency for a 0.3m x 0.1m disk shaped
spacecraft (cross section of 0.03m2), with a 67% area coverage, this results in a total
electric power of only 4.0 watts. Lightweight, efficient solar array panels that
minimize the effective array mounting area are needed. Dual or triple junction GaAs
solar cells that give 18% conversion efficiency at end of life (EOL), and assuming a
more optimistic area factor of 85%, will result in only 6.2 W at EOL. Small satellites
that do not have extended solar panels simply do not intercept a large solar power
density and must use the available power very efficiently. For a small spinning
satellite, it is expected that three solar cells will be connected in series along the spin
axis, and groups of three will be connected in parallel around the circumference. Each
section will generate 3.3 V and rotate into and out of sunlight as a unit. Voltage drops
at 3.3 volts, bus regulation, circuit protection (e.g. fuse or circuit breaker) and LiIon
battery discharge characteristics are being studied.
Highly elliptical orbits in the ecliptic plane where the apogee velocity is very low will
cause a several hour eclipse during part of the year. Spacecraft batteries to cover this
eclipse period presents a significant mass impact. However, only a 10° orbit plane
inclination relative to the ecliptic, will reduce the maximum eclipse period to about
one hour. Inclusion of spacecraft batteries is then justified. Passive thermal control
will be used to keep the spacecraft electronics within 10° C of ambient temperature,
and hence will not require electric power for heating. Using such a scenario, a battery
requirement of about 2 amp hours at 3.3 volts will allow full spacecraft functionality
during an eclipse. Twelve AA size LiIon batteries meet the requirement and only
weigh 480 grams.
Circuits that have high current demands, such as thruster solenoids and fuses,
need to be augmented with components that have a lower power density than
pg. 13
batteries, but also have lower internal resistance. Ultra capacitors are a candidate for
this application.[2,9,11]
2.41 CHALLANNGES
Miniaturization of the power system electronics (PSE) to meet the weight and
size requirements of the nano-satellites is a considerable challenge.[9]
2.42 SOLUTION OF CHALLANGE
The ideal approach is to eliminate the PSE completely, by having a fixed
electrical load and batteries provide the needed bus regulation. This yields a
simplified system consisting of the solar cells, batteries, and minimal circuitry. A
more immediate approach to miniaturization is to produce hybrid modules that
measure approximately 2"x 1.25" x 0.5" and weigh about 100 grams for each PSE
component, namely the solar array regulator, battery regulator, and low voltage power
converter. The combination of these three components into one module will reduce
the size and weight another order of magnitude.[9,11]
pg. 14
2.5 THERMAL
Although an inclination change by 10° renders maximum shadows below 2 hours, we
hereby study the case of a maximum 8 hour shadow for the purpose of generality. We
investigated several thermal control strategies from the viewpoint of design
robustness and the effect of the long earth shadow on each design.
Three thermal configurations were considered: (1) top and bottom of the
spacecraft are insulated, the inside of the cylindrical solar array is not insulated
allowing internal heat transfer between the internal equipment and the solar array; (2)
the entire spacecraft is insulated, top and bottom as well as inside the solar arrays,
except for a radiator on top, sized to radiate the internal electrical dissipation; and (3)
w serving as the only thermal coupling between the equipment and a radiator on the
outside surface.
The key advantage of configuration (1) is its reliability, or robustness. Since
the temperature of the spacecraft is set by a high energy balance (heat in = heat out)
dominated by the absorbed solar energy, the operational temperature of the spacecraft
is relatively insensitive to top and bottom multilayer insulation (MLI) properties, or,
largely, to internal heat dissipation. However, the feature that yields the operational
reliability, i.e., the high energy balance, also results in a rapid drop in temperature
when the solar load disappears during the earth shadow. During the maximum 8 hour
eclipse used for study purposes, internal temperatures dropped by about 60° C, which
would result in internal temperatures in the range of -30 to -40° C. At the same time,
the solar arrays dropped to a temperature of about –60° C. Based on past experience,
these end-of-eclipse temperatures are reasonable.
Because configuration (2) has a much smaller overall energy balance than
configuration (1), it is much more sensitive to MLI properties and to internal power
dissipation. However, eclipse performance improves. During the ~8 hour eclipse,
internal temperatures drop by only about 20° C, a marked improvement, with end-of-
eclipse temperatures well within the range of most spacecraft components. It should
be noted that the solar arrays, since they are now isolated from the body of the
spacecraft, drop to temperatures of about 110° C. Even these solar array temperatures
should not pose a problem. For example, the solar arrays of many geosynchronous
pg. 15
satellites drop routinely to temperatures of about -150° C during the 72-minute eclipse
experienced by these spacecraft at each equinox season.
The key feature of configuration (3) is that the equipment is coupled to an
external radiator only with a two-phase heat transport device, such as a capillary
pumped loop (CPL) or loop heat pipe (LHP). Operational temperatures are again
maintained to temperatures of about 20° C nominal with a properly sized radiator.
However, the temperature is also totally dependent on the proper operation of the two-
phase loop”. The two-phase heat transport device can be made redundant by the
addition of a second loop if single fault tolerance is desired. Note that redundancy is
not a consideration for the other two configurations studied. During the ~8 hour
eclipse, further improvement is realized, with internal temperatures dropping by as
little as 6° C if the internal payload is well insulated from the exterior of the
spacecraft. As in configuration (2), the solar array temperatures drop to about -110°
C. For certain equipment or science instruments, the temperature control afforded by
this type of “active” design may be necessary.
A moderate amount of technology development will be necessary to enable a
two-phase heat transport system for use in a nano-satellite. The small size and low
heat transport requirements of the nano-satellite will necessitate significant
downsizing of today’s flight qualified two-phase systems. This reduction will be
accomplished by leveraging recent successful tests of a small, cryogenic two-phase
CPL.[9]
pg. 16
2.6 RF COMMUNICATIONS
The onboard RF subsystem must be small, light, and low power. The tracking system
should be coupled with this communications subsystem, to maximize efficiency in
mass and power.
The communications subsystem is further complicated by constellations
requiring spin-stabilized nano-satellites. A spinning nano-satellite cannot easily point
an antenna toward Earth. Therefore, a low gain Omni antenna is assumed and
communications must take place near perigee, when the range is 3-5 Earth radii. A
large ground antenna and high compression must be used to achieve reasonable data
rates with minimum power. This places an additional burden on the ground stations
for both sensitive receivers/bit synchronizers and advanced decoders. These same
considerations limit the data rate for satellite-to-satellite communication.
Although the inclusion of an onboard command receiver is highly desired, it puts an
additional strain on an already challenging nano-satellite mass and power budget. For
this reason, the concept of a totally autonomous, receiver less nano-satellite design
appears most attractive. However, “receiver on a chip” technology is advancing
quickly enough that including a receiver onboard looks feasible. The biggest
disadvantage of a receiver now becomes the ground personnel and software needed to
support the ability to command the nano-satellite. Command actions taken onboard
will of course be limited to basic functions such as “transmit data” because of the lack
of redundancy and mechanical functions. Although scenarios have been defined to
allow the nano-satellites to autonomously determine when to transmit their stored
data, utilizing a receiver to control the telemetry downlink from the ground still has
value. The capability of uploading flight software changes, as well as sending a
master reset if necessary, would also exist with such an onboard command receiver.
[2,9,10,11]
pg. 17
[9]
pg. 18
2.7 MECHANICAL/STRUCTURE
The nano-satellite mechanical system will be kept as simple as possible. The
ideal nano-satellite mechanical design should consist of a one-piece structure to which
all other components are mounted.
Multifunctional structures can provide thermal control, shielding and serve as
substrates for printed circuit boards. For example, diamond facesheet honeycomb
panels can serve as a structure, thermal conductor and radiator, and printed circuit
board substrates. The diamond facesheet provides 10 times greater thermal
conductivity than aluminum and can dissipate heat from high power density
electronics modules with a low mass comparable to carbon fiber composites.
Another example is the structural battery system. It consists of a honeycomb panel
whose core is filled with the cells of a nickel-hydrogen battery (or other flight
qualified cell technology).
Concurrent engineering and fabrication techniques will be used to create a
single computer model for the design, analysis (structural, thermal, and dynamic), and
fabrication of the nano-satellite and its components. Dynamic modeling capabilities to
simulate nano-satellite deployments will provide faster designs and a reduction in the
amount of deployment testing required. This approach will significantly lower
development costs by reducing duplication of effort, chances for errors, the number of
drawings and paperwork required.
Mass production techniques not traditionally used for spaceflight hardware
will be used, such as casting and injection molding. Options being considered for the
nano-satellite structure material are: cast aluminum; cast aluminum-beryllium alloy;
injection molded plastic; fiber reinforced plastic; and flat stock composite
construction. The material will be selected based on mass, cost, manufacturability,
ease of assembly and integration, and suitability for the space environment.
Streamlined testing is needed for up to 100 nano-satellites per mission.
Performing a complete test program on each unit would be prohibitively expensive
and time consuming. We need to reduce the quantity of testing required while
assuring product quality to meet program cost and schedule goals.
pg. 19
The deployer-ship carries the nano-satellites and deploy them into their proper
transfer orbits. The deployer-ship will be a conventional spin-stabilized or three-axis
stabilized spacecraft. The deployer ship release system will impart the required
minimum spin rate of 20 rpm to the nano-satellites. The following innovative
deployer-ship designs, nano-satellite packaging, and deployment techniques help
accomplish these goals.
A spinning deployer-ship with simple “Let-Go” deployment: In this case, the
deployer-ship is spinning at 20 rpm with the nano-satellite spin axis aligned with the
deployer-ship spin axis. The deployer-ship spin axis is then oriented in the desired
direction and the nano-satellite is released by a simple mechanism which lets the
nano-satellite go while imparting no additional spin. The nano-satellites are released
in opposing pairs to maintain the balance of the deployer-ship.[2,9,10]
pg. 20
2.8 INSTRUMENTS
Instruments for in-situ and remote measurements must be miniaturized to fit within
the mass and volume constraints of a nano-satellite. Power consumption must also be
scaled down accordingly. Instrument sensitivities cannot be compromised in the
process. Instrument electronics need to be combined with spacecraft subsystem
electronics to achieve higher degrees of integration, yielding reduced mass and
volume. Instrument software will be designed to evaluate the onboard data and adjust
instrument data rates and modes to efficiently capture the data of highest priority.
A highly integrated spacecraft will result, reducing both time and cost for final
spacecraft integration and testing.
Nano-satellites for in-situ measurements, such as those baselined for the STP
Constellation missions, will carry a low energy particle detector (electrons and ions)
and a magnetic field instrument.
One of the targets for reducing the mass of the particle detectors is the
miniaturization of the high voltage power supply. The magnetometer consists of a tri-
axial fluxgate sensor and an electronics module. The instrument sensor is mounted on
a deployable boom, while the electronics module is placed inside the spacecraft
structure. The sensor can be made small enough today to be used on a nano-satellite.
The challenge for magnetometers as well as particle nanosatellite instruments remains
to reduce the electronics modules to a fraction of the C&DH unit, while maintaining
the sensitivity and accuracy of present day, larger-size designs.[9]
pg. 21
2.9 GROUND SYSTEMS
Figure 7 shows the ground system concept for a nano-satellite constellation.
[9]
The large number of spacecraft in a constellation is a challenge to the ground system
in getting all of the data to the users. In the baseline mission, there are times when up
to nine spacecraft would be within communications range of a ground station at a
single time. We have modeled the ground station contacts and can support the
constellation with only two stations, located on opposite sides of the earth.
The schedulers will prioritize the contacts, with the spacecraft in the higher period
orbits getting priority. Spacecraft in the lower period orbits have more opportunities
to dump their data, and therefore can have lower priority without risking any data
loss.
Since the nano-satellites are autonomous, the operations concept for a mission
requires only a few operators to determine the orbits of the spacecraft, schedule the
ground stations, and to investigate anomalies on the spacecraft. Automated systems
will monitor the housekeeping data from the spacecraft and they will flag problems
for the spacecraft engineers to investigate. The large number of spacecraft allows the
risk management to be different for this mission than for single spacecraft missions.
Except for commands to initiate the data downlink, the ground system will not
command the nano-satellites for normal operations. The only commands that the
ground system sends would be program loads to resolve or work around problems and
failures.
The large number of spacecraft is a configuration control challenge for the
data tracking, the schedules, the command loads, the science data, and the engineering
pg. 22
data. The ground system will use IDs, color coded user interfaces, and other
techniques to ensure that the operators and users can keep track of the data associated
with a particular satellite. Constellations that fly in a close formation can benefit by
the use of inter-satellite communications to reduce ground station contention. The
data would flow from a single spacecraft to the ground, instead of coming from every
spacecraft. Communications protocols for inter-satellite communications will be
investigated in the future.[9]
pg. 23
2.10 AUTONOMY
Support costs are high if single-satellite mission operations and data analysis
practices are scaled to a constellation mission. Autonomy onboard the spacecraft and
on the ground is therefore required to ensure that science objectives are efficiently and
inexpensively met.
Nano-satellite autonomy will make use of onboard and ground-based remote
agents, with the overarching goal of maximizing the scientific return from each
satellite during the mission lifetime. The remote agents achieve this goal by
monitoring and appropriately controlling spacecraft subsystems. Additionally, the
onboard agent monitors the full complement of spacecraft sensors and instruments to
heuristically separate scientific events of interest from background events, thereby
intelligently fitting the science data within allocated spacecraft storage resources.
Spacecraft subsystems could be compromised if faults occurring during this
blackout period were not readily addressed. An unacceptable loss of scientific data
could also occur. Therefore, the onboard agent will incorporate the capability to
detect, diagnose and recover from faults.
Certain failure scenarios may not be correctable by the onboard agent. These
faults will be deferred to the ground agent for handling. Each spacecraft will include
data in its telemetry stream on the health and status of each subsystem and a history of
commands autonomously issued since the last ground contact. The ground system will
then attempt to diagnose problems based on this data. Additionally, collective
knowledge of actions taken by all satellites in the constellation will reside within the
ground system by virtue of the data dumps made during each contact. From this data
the agent can detect trends and systematic conditions not otherwise observable
onboard the spacecraft.[9]
2.10.1 CHALLANGES
These highly autonomous systems will present a unique set of challenges not
only to the system designers, but also to those involved in spacecraft testing. Careful
consideration must be given to the design of the test program to ensure that the state-
space of the remote agents is validated and verified. It is equally important to
implement this program in a cost-effective manner. However, we could likely justify
exerting considerable resources to address this issue since the methods developed to
solve these challenges can be applied to numerous missions.[2,9,11]
pg. 24
CHAPTER 3
TECHNOLOGY TRANSFER & SPINOFF
In addition to enabling a wide array of scientific space missions, nano-satellite
technology will have applications to a variety of industries. Such technology transfers,
or spinoffs, have been and will remain an important link between NASA and other
organizations.
Two of the more versatile propulsion technologies are miniaturized ignition
systems and ultra low power control valves. The former will increase the efficiency of
many gas-generating or explosive devices, from air bags to pneumatic hand tools. The
latter will enable the incorporation of precise and reliable fluid control into an ever-
increasing number of medical devices, automotive systems, and aircraft systems.
The C&DH subsystem requires rugged, radiation tolerant, low power, and
lightweight electronics. Once developed, this technology can improve many types of
remote and mobile devices. Portable medical devices, advanced aircraft systems, and
mobile communications equipment all can benefit from the C&DH characteristics.
The miniaturized two-phase heat transfer technology described in the thermal
section has several potential terrestrial and commercial applications. A patent has
been awarded for a “bio-CPL”, which can be applied to utilize excess body heat to
warm appendages such as hands and feet in medical applications as well as for
recreational equipment. Additional commercial possibilities exist in energy
management for a variety of process and equipment applications.
Some of the more aggressive communications coding techniques, those with
gains similar to turbo code, will become more routinely accepted and incorporated
into commercial ground stations. Along with these coding techniques that allow all”
errors to be corrected in very weak signals, improvement in the quality of bit
synchronizers is expected, which convert noisy analog inputs into clean, digital
outputs. These advances will improve ground-to-ground as well as space-to-ground
communications.[9]
pg. 25
CHAPTER4
SUMMARY / CONCLUSION
Each nano-satellite will be an autonomous, highly capable miniaturized
satellite with a maximum mass of 10 kg, and designed for a two year mission life.
Provisions for orbital maneuvers, attitude control, onboard orbit determination, and
command and data handling will be included. Fully capable power and thermal
systems, RF communications, multiple sensors, and scientific instruments will be
integrated on an efficient structure. Nano-satellites developed for in-situ
measurements will be spin-stabilized, and those developed for remote measurements
will be three-axis-stabilized. Autonomy both onboard the nano-satellites and at the
ground stations will minimize the mission operational costs for tracking and
managing a constellation.
Key technologies being actively pursued include miniaturized propulsion
systems, sensors, electronics, heat transport systems, tracking techniques for orbit
determination, autonomy, lightweight batteries, higher efficiency solar arrays, and
advanced structural materials. Deployer-ships will carry and deploy a constellation of
up to 100 nano-satellites, delivered to space by one launch vehicle. This initiative is
scheduled to produce the first generation of mature technologies by 2004, with the
launch of the first nano-satellite constellation in 2008.
Partnerships with other NASA centers, other government agencies, private
industry, universities, and foreign institutions are currently being established in the
areas of manufacturing and testing of up to 100 nano-satellites per mission, the
development of multifunctional structures, and integration of instrument sensors and
electronics with the spacecraft subsystems.
pg. 26
CHAPTER5
REFERENCES
1. http://en.wikipedia.org/wiki/Miniaturized_satellite
2. www.wisegeek.com/what-is-a-nanosatellite.htm
3. www.timesofindia.indiatimes.com
4. www.mscweb.gsfc.nasa.gov/543web/.../tsld011.html
5.www.spacedaily.com/.../ISRO_To_Build_Nano_Satellite_Platform_Eyes_Overseas
_Business_999.html
6. www.ssdl.stanford.edu/ssdl/images/stories/papers/1999/ssdl9907.pdf
7. www.tethers.com/Nanosats.htm
8. www.iitk.ac.in/dord/research_news/nano.pdf
9. www.plasma2.ssl.berkeley.edu/.../panetta.pdf
10.www.epubs2.cclrc.ac.uk/bitstream/765/nsat_paper_final. pdf
11. Microsoft student Encarta 2007
IMAGE REFERANCE:
www.plasma2.ssl.berkeley.edu/.../panetta.pdf
TABLE REFERANCE:
www.plasma2.ssl.berkeley.edu/.../panetta.pdf
pg. 27
