CCSDS Concept Paper
Future Spacecraft Systems Designs
Neptune and Uranus
- or -
Interstellar Missions of Multi Generational
Much has been
learned from the Pioneer Program, Voyager Program,
Galileo and Cassini-Huygens spacecraft on what spacecraft designs
most optimal for deep space scientific research.
subsystem minimal requirements
Absolute must requirements
There must be absolutly 3 command and control computers, for each
kind of command and control.
Sadly, there are missions in deep space that do not have 3 redundant
Cassini is a
three-axis-stabilized spacecraft and not a spin stabilized
craft. There is a set of four Reaction Wheel Assemblies (RWAs).
three RWAs are needed for spacecraft control, with one spare.
the RWAs is important for spacecraft stabilization and it is
accomplished by firing the RCS thrusters while changing the RWA
rotation speeds, all while the spacecraft attitude (pointing)
fixed. RWA biases are done quite frequently and are relatively
users of hydrazine, but they are absolutely necessary to keep the
speeds within safe ranges.
- For such a long duration mission, it is probable that 3 axis
stabilization designs will be mandatory.
- It is assumed that Cassini's level of redundancy (3, +1) in
end is not adequate.
- It is assumed that (3, +3) might be adequate, similar to
- A level of gyro redundancy of (3, +4) must be considered.
(RTGs) are absolutely necessary as no solar power is really
such great distances from the Sun.
It is assumed
that the Multi-Mission Radioisotope Thermoelectric
Generator (MMRTG) are a type of developed for NASA space missions
as Mars Science Laboratory and may be used on the Jupiter Europa
- The MMRTGs are produced by Boeing for NASA, but any space
for any deep space mission could use them.
- The GPHS-RTG used SiGe thermoelectric elements but these are
longer in production and the MMRTG will use PbTe/TAGS
(from Teledyne Energy Systems).
- The MMRTG is designed to produce 125 W electrical power per
at the start of mission, falling to about 100 W per module after
- At least 3 MMRTGs should be used on the primary craft.
- At least 1 MMRTG (of lesser power and weight) should be used
any ancillary craft.
- The bare minimum craft electrical power at the beginning of a
mission should be 350w or if ancillary craft shares power 400w.
- If RTGs become available that are capable of providing 1.0 kw
the beginning of a mission, they must be considered.
systems worth consideration
- To conserve heat and mass, spacecraft and instrument
housed together in IEMs (Integrated Electronics Modules).
- There should be redundant IEMs for each computer system.
- CPU speeds should be no higher than 50 MHz to conserve power,
with 250 nm processes.
- CPU dormancy speed should be 2.5 MHz or 1.5 MHz, as it is
expected that 2/3 rds of the total mission time will be spent in
The RAD750 is
radiation-hardened single board computer, based on
IBM's PowerPC 750. The RAD750 is manufactured by BAE Systems. It
intended for use in high radiation environments such as
board satellites and spacecraft. The RAD750 was released for
in 2001 and the first units were launched into space in 2005.
systems worth consideration
- The CPU has 10.4 million transistors which is nearly ten times
more than the RAD6000 which had 1.1 million transistors and it
manufactured using either 250nm or 150 nm photolithography and
die area of 130 mm².
- The RAD600 has a core clock of 110 to 200 MHz and can process
266 MIPS or more.
- The CPU can include an extended L2 cache to improve
- The CPU itself can withstand 2,000 to 10,000 grays and
temperature ranges between –55 °C and 125 °C and requires 5
watts of power.
- The standard RAD750 single-board system (CPU and motherboard)
withstand 1,000 grays and temperature ranges between –55 °C and
°C and requires 10 watts of power.
- The RAD750 system has a price that is comparable to the
which is US$200,000 per board.
spacecraft carries two computer systems, the Command
and Data Handling system and the Guidance and Control processor.
of the two systems is duplicated for redundancy, giving a total of
computers. The processor used is the Mongoose-V, a 12 MHz
radiation-hardened version of the MIPS R3000 CPU.
32-bit microprocessor for spacecraft onboard computer
applications is a radiation-hardened and expanded 10–15 MHz
the MIPS R3000 CPU.
- The Mongoose was developed by Synova, Inc. of Melbourne,
USA, with support from the NASA Goddard Space Flight Center.
- The Mongoose-V processor first flew on NASA's Earth Observer 1
(EO-1) satellite launched in November 2000 where it functioned
main flight computer. A second Mongoose-V controlled the
solid-state data recorder.
Development Boards with 32-bit Rad-Hard Mongoose-V
- MIPS R3000 Instruction Set
- MIPS R3010 Floating-point Unit
- On-Chip 2KB Data Cache
- On-Chip 4KB Instruction Cache
- Speed grades 10MHz, 15MHz
- Mongoose-V microprocessor fabricated using 1-MRAD
Silicon-On-Insulator (Silicon on Saffire) technology
- The board is immune to single event upsets
- Spaceflight qualified units are readily available
- Pricing for Mongoose-V Software Development Board (MV-SDB-01)
around $30,000 USD per board.
Each of the
systems is duplicated for redundancy, giving a total of
four computers. The processor used is the Mongoose-V, a 15 MHz
radiation-hardened version of the MIPS R3000 CPU -- that should be
capable of running at 2.5 MHz during sleep or dormant phases of
and timing routines should be implemented in hardware
as well as software to help prevent faults or downtime.
A minimum of
low-power solid-state recorders (one primary, one
backup, one interstellar backup -- not used until interstellar
begins) holding up to 8 gigabytes (64 gigabits) each. One should
that there is about 2% less available for data storage due to file
light levels on the 2 outer gas giants
As a general
rule, visible light cameras should be able to cope with 1
Watt per square meter
As a general
rule, infrared cameras should be able to cope with 40 K
blackbody imaging conditions.
- 10 second exposure times should allow for rings to become
visible, without excessive smearing due to dynamic motion
second exposure times should be typical and nominal
second exposure times should still produce usable images
cameras by megapixel capabilities and bit depths.
RULE : All
cameras should be autonomous to whatever extent is possible. The
Voyager Program arrangement of having camara controls linked to
flight subsystem should no longer be needed. Each camera should
its own 2 core (or 4 core) 32 bit CPU with Error Correcting Memory
an autonomous pointing system. Each camera should also have its
command and control system that is able to obtain spacecraft state
determine when to acquire images.
- 2 x 5 mpx x 16 bits, Narrow angle
- 2 x 3 mpx x 16 bits, Wide angle
- 2 x 1 mpx x 10 bits, for
navigation cameras with 8 bit fallback mode (broadband only,
polarization filtering for one or both is acceptable)
- 2 sun locators @ 256 kpx, 10 bit but with 8 bit fallback mode
- 2 star locators @ 256 kpx, 12 bit but with 8 bit fallback mode
- 2 x 4 mpx x 16 bits if feasible, Wide angle and Narrow angle
facing in the preferred targeting direction
- 2 x 1 mpx x 14 bits, spare Wide angle not facing in the
- It is assumed that to increase coverage the Wide and Narrow
angle cameras would not be perfectly centred
- 2 x 4 mpx x 16 bits, Wide and Narrow angle but offset by ~20
- 2 x 1 mpx x 16 bits, monochrome if visible
- The camera must have at least 4 megabytes of memory, its own
operating system and compression subsystem
of establishing a link between a spacecraft and a DSN
station is, to a first order, a function of the required data rate
the square of the distance over which the link is occurring.
simple measure of end-to-end link difficulty can be obtained by
the product of the data rate and the square of the distance. Note
this measure makes no assumptions about the telecommunications
capabilities of either the spacecraft or the DSN ground station at
end of the link. It is simply indicative of the inherent
the link itself. Over the next 25 years, this downlink difficulty
expected to increase by roughly two-and-a-half
uplink difficulty trends are, of course, similar to those of
the downlink difficulty trends and involve roughly the same driver
missions. However, because of the asymmetry between uplink and
rates for robotic missions discussed earlier, the effective
radiated power needed to support the uplink to such missions is
the capability of the current DSN (assuming appropriately sized
high-gain antennas onboard the spacecraft and forward-error
coding on the uplink when needed).
requirements and their associated link difficulties are
generally bounded by the problem of providing emergency uplink at
planet distances. The DSN’s current 70-meter, 20 kW, X-band
enables spacecraft at Jupiter’s maximum distance from the Earth to
receive a 7.8125 bps emergency transmission via their
enable emergency uplink into an omni antenna at greater distances,
equivalent of 10 to 20 times the current 20 kW capability on a 70
uplink dishes is needed, depending upon one’s spacecraft
(e.g., system temperature, receiver loop bandwidth, etc.).
using bandwidth efficient technique compatible with
Block V receiver structure in a deep space network. JPL is now
researching deep space exploring mission (eg. MRO) modulation
suitable for future high data rate (eg.10-100Mbps) by combining
efficient error correction code (eg. Turbo and LDPC code)
- Offset QPSK (O-QPSK) : For phase restriction, its peak average
power ratio (PAPR) is smaller than that of BPSK and standard
research keystones in O-QPSK are how to use rectangular and
raised cosine (SRRC) pulse shaping techniques.
- Pre-encoded GMSK : Pre-encoding technique is used to
differential coding of MSK modulator. Consequently, it can avoid
performance reduction in modulator and demodulator system
combination of differential coding at sender and differential
- Trellis-coded OQPSK : By using two-state convolutional encoder
introduce memory among sending data and raised cosine pulse
OQPSK signal of improved envelope performance can be obtained.
- FQPSK : By introducing a controllable correlation and adopting
given signal waveform between in-phase and quadrature arms (its
is similar to constant), the FQPSK modulator can be regarded as
16-state joint I-Q TCM modulator.
- In order to further improve bandwidth efficiency, more
fluctuations bandwidth efficient TCM technique is now being a
of deep space communications
- Long Distance : A lot of planets in deep space are several
hundred million kilometres away from the earth. Such long
results in very low signal to noise ratio (SNR).
- High Signal Propagation Delays : This is due to the enormous
distances involved between the communicating entities and the
relativistic constraint restricting signal transmissions to the
of light. For example, one-way signal propagation delays for the
Cassini mission to Saturn are in the range of 1 hour and 8
minutes to 1
hour and 24 minutes.
- High Data Corruption Rates : Extremely long distances cause
signals to be received at extremely low strengths at the
thereby increase the probability of bit-errors in the channel
random thermal noise errors, burst errors due to solar flares,
- Disruption Events : Since communicating entities in deep-space
tend to be in motion relative to one another, the communication
between them is prone to disruption. A planetary probe on the
of Saturn’s moon Titan, for example, could experience disruption
the rotation of Titan on its own axis (when it goes to the night
of Titan), when Titan passes under Saturn’s shadow during its
revolution around the planet, and when other moons / planets/or
itself block the line of sight to the destination.
- Primary communication with the spacecraft should be via X band
- Secondary communications and radio science experiments should
possible via Ka band (32–34 GHz), with data rates similar to X
- Primary telemetry communication should be via S band.
- Due to the vast distances involved, error correction systems
be chosen that offer the most reliable coding.
- The backup engineering telemetry communication system should
band, but it should be possible to allocate up to 50% of the
stream to science data.
- There should be a target X band downlink rate of 40 kbit/s at
- There should be a target X band downlink rate of 20 kbit/s at
- There should be a target X band downlink rate of 10 kbit/s at
- The spacecraft should have a minimum 100% redundancy of
transmitters and receivers.
- The spacecraft should use Right-Hand Circular Polarization and
Left-Hand Circular Polarization (RHCP & LHCP).
- The downlink signal is amplified by dual redundant 15-watt
(travelling wave tube amplifiers) mounted on the body under the
- The primary X band system should permit control to power both
TWTAs at the same time. This option should permit a
downlink signal, with a 90% increase in the downlink datarate.
Paradigm" with 3 tunable parameters of control for
applications, namely Immediacy, Probability of Delivery and
system should be based on using Immediacy and "Probability
of Delivery" but "Mission Value" should act like a 'gamma
factor to tweak up the delivery of some data streams or image
- Immediacy : Immediacy is a notion of how urgently a unit of
application data (called a "job" in the subsequent discussion)
be received. For example, a message from a science instrument
a critical state implying status such as \Instrument too hot",
battery condition" or commands such as "Stop! Don't go down
from Earth, might need to be reported ahead of all other
experimental results / commands. We say that such a job has
immediacy requirements over the others.
- Probability of Delivery (PoD) : In the space environment where
there is always a non-trivial bit-error rate on the channel, the
communication channel cannot guarantee 100% reliable data
Some application jobs may seek higher PoD guarantees compared to
however. For example, the picture of a microbe found on the
Titan may be of much more importance to the mission compared to
house-keeping telemetry data, and thus may need higher
- Mission Value : How valuable is the data to the mission.
Value" should act like a 'gamma correction' factor to tweak up
delivery of some data streams or image files.
- Jobs in Cloud A [high immediacy (low actual value), low PoD]
could include spacecraft location data and certain classes of
data which must be delivered within a very small timespan or the
would not be valid or valuable any more for the receiver. This
classes of telemetry data updates that are sent often enough
loss of a single update is not critical.
- Jobs in Cloud B [high immediacy (low actual value), high PoD]
could carry data representing critical instrument status seeking
immediate attention (Instrument too hot" or "Battery draining
fast") or urgent commands sent on the uplink that need to be
on ahead of all other ongoing jobs ("Stop by that interesting
a while"-command being sent to a rover on Mars).
- Jobs in Cloud C [low immediacy (high actual value), high PoD]
could include important results of scientific experiments,
interesting images that need to be sent reliably, but not
- Jobs in Cloud D [low immediacy (high actual value), low PoD]
could include relatively less important scientific data and bulk
that may be sent just on a best-efforts basis. Note that while
typically is a large solid-state data store available on
its capacity is finite, and an implementation of our priority
paradigm would store jobs in Cloud C ahead of jobs in
Cloud D, and under severe storage space constraints may even
to store Cloud D jobs for transmission.
consider various mechanisms that could be used to guarantee
the priority requirements of application jobs.
- Adapting the error correction mechanisms We might choose to
increase the quality of Forward Error Correction (FEC)
use for the higher priority jobs and thereby improving the
transmitting them successfully, with the additional overhead
- Modulating the number of redundant transmissions In a scenario
is typical in current deep-space missions, the FEC mechanisms
sizes tend to be used for a mission phase. In such a case, the
option we may have to guarantee the job priority requirements
to transmit the frames redundantly in original transmission. We
that such a simple mechanism is being used in current missions
command / data to be sent is extremely critical.
- Adapting the frame size. We might choose to decrease the size
datalink frames for the higher priority application tasks,
the chances of successful transmission of each frame. This would
introduce additional overhead in data transmission as the header
to application data ratio would increase.
Lessons from the Voyager Program and
spacecraft was only able to transmit mission data through a
low-gain antenna because the high-gain antenna on board the
has refused to deploy properly and was essentially useless. This
failure scenario should be avoided at all costs by using fixed,
antennas -- but this failure scenario risk cannot be eliminated.
The data rate
from this antenna was not designed to exceed 100 bits per
second. To offset some of the perform ante loss, the spacecraft’s
computer had to be extensively reprogrammed to include new data
compression and coding algorithms.
coding system for the low gain antenna mission uses a
Reed-Solomon code of block length 255 concatenated with a (14,
convolutional inner code, and interleaves the Reed-Solomon symbols
depth 8. The convolousionally encoded symbols are decoded by a
likelihood (Viterbi) decoder. Each Reed-Solomon decoded codeword
then decoded algebraically.
In order to
the problem of only being able to use the low gain
antenna 2 types of decoding enhancements were proposed. These
enhancements involve “re-decoding” of some of the data.
of re-decoding is confined to the Reed-Solomon decoder
and utilizes information from neighbouring codewords within the
interleaved block to erase unreliable symbols in un-decoded words.
second type involves re-decoding by the Viterbi decoder, using
information fed back from codewords that have been successfully
de-clocked by the Reed-Solomon decoder.
conclusions were drawn from the analysis and delivered to the
(Galileo mission planners. These comparisons are valid for the
system using a (14, 1/4) convolutional coded depth-8 interleaving
Reed-Solomon symbols, and achieving a final decoded bit error rate
x 10–7. A second stage of Viterbi decoding without any
erasure declarations is worth about 0.37 dB relative to the
stages of Viterbi decoding is worth an additional 0.19
dB for a total gain of 0.56 dB. The marginal additional
from utilizing erasure declarations was shown to be around 0.19 dB
one-stage decoding (no Viterbi re-decoding), but only 0.02 dB for
two-stage decoding and essentially nil (0.00 dB) for four-stage
Communication System (downlink baudrate > 300 bps)
- The Galileo high gain antennae failure forced the low gain
antenna into becoming the primary antenna for the mission --
failure scenario should be avoided at all costs by using a
fixed dish antenna capable of 10000 bps at Neptune's radius.
- Antenna polarization : LCHP (Left Hand Circular Polarization)
RCHP (Right Hand Circular Polarization) should be implemented,
provides better noise filtering vs the typically Horizontal and
Vertical polarizations found from natural sources.
- Three transmitters
Telemetry System (downlink baudrate < 300 bps)
- There is a possibility of failure of the primary antenna, so
Engineering Telemetry System must be ready to serve as a backup
- The Engineering Telemetry System need to have some backup
capability for the primary antenna system, that is to say it
capable of 300 bps at Neptune's radius if the primary antenna
- Normal Engineering Telemetry System downlink rates should be
bps, but 50 bps should be available for fallback.
|15 April 2007
|11 November 2009
|25 November 2013
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