INTRODUCTION
Communication links between space crafts is an important element of
space infrastructure, particularly where such links allow a major reduction in
the number of earth stations needed to service the system. An example of an
inter orbit link for relaying data from LEO space craft to ground.
Link between
a low earth orbiting (LEO) space craft and a geostationary (GEO) space craft
for the purpose of relaying data from the LEO space craft back to the ground in
real time. The link from the GEO Satellite to ground is implemented
using microwaves because of the need to communicate under all weather
conditions. However, the interorbit link (IOL) can employ either microwave or
optical technology. Optical technology offers a number of potential advantages
over microwave.
I.
The antenna can be much
smaller. A typical microwave dish is around 1 to 2m across and requires
deployment in the orbit, An optical antenna (le a telescope) occupies much less
space craft real estate having a diameter in the range of 5 to 30 cm and is
therefore easier to accommodate and deploy.
II.
Optical beam widths are much
less than for microwaves, leading to very high antenna gains on both transmit
and receive. This enables low transmitter (ie laser) powers to be used leading
to a low mass, low power terminal. It also makes the optical beam hard to
introsept on fan leading to convert features for military applications,
consequently there is a major effort under way in Europe, USA and Japan to
design and flight quality optical terminals
SOUT
The European Space Agency (ESA) has programmes underway to place
Satellites carrying optical terminals in GEO orbit within the next decade. The
first is the ARTEMIS technology demonstration satellite which carries both
microwave and SILEX (Semiconductor Laser Intro satellite Link Experiment)
optical interorbit communications terminal. SILEX employs direct detection and
GaAIAs diode laser technology; the optical antenna is a 25cm diameter
reflecting telescope. The SILEX GEO terminal is capable of receiving data
modulated on to an incoming laser beam at a bit rate of 50 Mbps and is equipped
with a high power beacon for initial link acquisition together with a low
divergence (and unmodulated) beam which is tracked by the communicating
partner. ARTEMIS will be followed by the operational European data relay system
(EDRS) which is planned to have data relay Satellites (DRS). These will also
carry SILEX.
Once
these elements of Europe’s space Infrastructure are in place, these will be a
need for optical communications terminals on LEO satellites which are capable
of transmitting data to the GEO terminals. A wide range of LEO space craft is
expected to fly within the next decade including earth observation and science,
manned and military reconnaissance system.
The
LEO terminal is referred to as a user terminal since it enables real time
transfer of LEO instrument data back to the ground to a user access to the DRS
s LEO instruments generate data over a range of bit rates extending of Mbps
depending upon the function of the instrument. A significant proportion have
data rates falling in the region around and below 2 Mbps. and the data would
normally be transmitted via an S-brand microwave IOL
ESA
initiated a development programme in 1992 for LEO optical IOL terminal targeted
at the segment of the user community. This is known as SMALL OPTICAL USER
TERMINALS (SOUT) with features of low mass, small size and compatibility with
SILEX. The programme is in two phases. Phase I was to produce a terminal flight
configuration and perform detailed subsystem design and modelling. Phase 2
which started in september 1993 is to build an elegant bread board of the
complete terminal.
The link from LEO to ground via the GEO
terminal is known as the return interorbit link (RIOL). The SOUT RIOL data rate
is specified as any data rate upto 2 Mbps with bit error ratio (BER) of better
than 106. The forward interorbit link (FIOL) from ground to LEO was a nominal
data rate of (34 K although some missions may not require data transmissions in
this directions. Hence the link is highly asymmetric with respect to data rate.
The
LEO technical is mounted on the anti earth face of the LEO satellite and must
have a clear line of sight to the GEO terminal over a large part of the LEO
orbit. This implies that there must be adequate height above the platform to
prevent obstruction of the line of sight by the platform solar arrays, antenna
and other appertages. On the other hand the terminal must be able to be
accommodated inside the launcher fairing. Since these constraints vary greatly
with different LEO platforms the SOUl configurations has been designed to be
adaptable to a wide range of platforms.
The in-orbit life time required for a LEO mission in typically 5
years and adequate reliability has to be built into each sub-systems by
provision of redundancy improved in recent years. and GaAIAs devices are
available with a projected mean time to failure of 1000 hours at 100 MW output
power.
The
terminal design which has been produced to meet these requirements includes a
number of naval features principally, a periscopic coarse pointing mechanism
(CPA) small refractive telescope, fibre coupled lasers and receivers, fibre
based point ahead mechanism (PAA), anti vibration mount (soft mount) and
combined acquisition and tracking sensor (ATDU). This combination has enabled a
unique terminal design to be produced which is small and lightweight These
features are described in the next sections.
OVERVIEW
OF THE SOUT TERMINAL
The SOUT terminal consists of two main parts: a terminal head unit
and a remote electronics module (REM). The REM contains the digital processing
electronics for the pointing acquisition and tracking (PAT) and terminal
control functions together with the communications electronics. This unit is
hard mounted to the space craft and has dimensions 200 by 200 by 150mm. The REM
will have the advantage of advanced packaging ASIC and technologies to obtain a
compact low mass design.
In the figure the SOUT configuration head unit is shown. The REM is
not shown and the supporting structure and terminal control hardware have been
removed for clarity. The terminal head performs the critical functions of
generating and pointing the transmit laser beam and acquiring and tracking the
received beacon and tracking beams.
There
is fixed head unit with a periscopic course pointing assembly (CPA) on top of
the telescope. The telescope with the CPA is referred to as the optical
antenna. The head unit is soft mounted to the satellite by a set of three anti
vibration mounts arranged in a triangular geometry. This fillers out high
frequency microvibrations, originating from the space craft. Inclusion of the
soft mount has a major impact on the terminal fine pointing loop design and
structural configuration as described below. All of the optical components and
mechanisms needed for transmit and receive functions except for the telescope
and CPA are mounted on the double sided optical bench. The head unit also
includes an electronics package (CPEM) which contains electronics required to
be in close proximity to the sensors and pointing mechanisms.
Key
elements of the head unit are the integrated transmitter comprising diode laser
and point ahead assembly (PAA) optical antenna comprising telescope and coarse
pointing assembly, fine pointing loop comprising acquisition and tracking
sensor (ATDU) and fine pointing assembly (FPA) and optical bench.
STRUCTURAL
CONFIGURATION OF SOUT
The SOUT has a novel structural and thermal design which satisfies
the unique demands imposed by the various sub-systems. The main structural
elements are a truss frame assembly which supports the optical antenna
orthogonal to the optical bench, a triangular plate which forms the lower truss
support and carries the soft mounts, optical bench and electronic units. Key
design drivers for the structure are the optical bench pointing stability, soft
mount constrains and base-bending moments associated with the telescope CPA.
There has to be a high degree of Coaligtnment between the transmit and receive
beam paths on the optical bench in order that the transmit beam can be pointed
towards the GEO terminal with an acceptably small pointing loss.
The height of the terminal above the
space craft depends upon the mounting interface; options include mounting
through a hole in the side wail of the space craft (Suitable for large
platforms), external mounting on a support frame, mounting on a deployment
mechanism. The head unit occupies an area of about 40 by 40cm depending upon
the platform interface.
Mass and
Power
The base-line SOUT has a total mass (including REM) of around 25 Kg
and a dynamic mass of 3.7kg due to the motion of the CPA. The maximum power
dissipation is around 65 W.
OPTICAL
ANTENNA
The optical antenna comprises the telescope and coarse pointing
assembly. The telescope is a refractive keplerian design which does not have
the secondary mirror obscurration loss associated with reflective systems. The
CPA uses stepping motors together with a conventional spur gear and planetary
gear. The total height of the optical antenna is a major contributor to the
height of the CPA above the platform which affects LEO and GEO link
obscurration by solar arrays, antennas and other space craft appendages.
Course pointing
assembly (CPA)
This
provides coarse pointing of the terminal over greater than hemispherical
coverage with high open-loop pointing
accuracy and very fine angular resolution.
It consists of two elliptical mirrors on zerodur
substrates which are driven by a stepper
motor and gearbox arrangement.
An external quartz window protects the mirrors from
the ingress of contamination in the
laboratory.
Telescope
This is a refractive Keplerian telescope with
70mm entrance pupil diameter and a
magnification of 8.75.
The breadboard design has an air-spaced triplet
objective and a triplet eyepiece. The lenses are housed in a steel tube with a
length of 42.8cm. This provides athermalised performance.
OPTICAL
BENCH
The diplexer, quarter wave plate and other lens system required too
acquisition and tracking are all placed in the optical bench. The diplexer has
a dietetric multilayer coating which provides efficient transmission of one
type polarised light at the transmit wavelength (848 nm) and rejects another
type poiarised light at the receive wavelength (800 nm). A quarter wave plate
(QWP) converts the transmit light to circular polarisation state prior to the
telescope. The PAA, lasers, and redundancy switching mechanisms are on one side
while the diplexer, receive paths and calibration path are on the other side of
the optical bench.
LINK
DESIGN
Wave length and polarization.
The transmit and receive wavelengths are determined by the need for
interoperability with future GEO terminals such as SILEX which are based on
GaAIAs laser diodes. Circular polarisation is used over the link so that the
received power does not depend upon the orientation of the satellite. The
transmit and receive beams inside the terminal are arranged to have orthogonal
linear polarisation and are separated in wave length. This enables the same
telescope and pointing system to be used for both transmit and receive beams
since the optical deplexing scheme can then be used.
Link
budgets for an asymmetric link
The requirement to transmit a much higher data rate on the return
link than on the forward link implies that the minimum configuration is one
with a large telescope diameter at GEO ie maximise the light collection
capabilities and a smaller diameter telescope at Leo. A smaller telescope at
LEO has the disadvantages of reduced light collection hut the advantage of
reduced pointing loss due to wider beam width.
The smaller
telescope on LEO facilitates the design of a small user terminal. For SILEX the
telescope diameter in 25 cm but it is highly desirable k a telescope with less
than 10 cm aperture in the user terminal. The design process begins with the link
budgets to ensure that adequate link margins is available at end of life too
the chosen telescope diameters and laser powers.
Pointing,
Acquisition and Tracking
The narrow optical beam width gives rise to a need to perform the
following critical pointing factions.
Pointing
The LEO terminal must be able to point in the direction of the GEO
terminal around a large part of the LEO orbit. Pointing error do occur some
time and it is determined by the accuracy with which the transmitting satellite
can illuminate the receiving satellites. This is turn depends on
1.
accuracy to which one satellite
knows the location of the other
2.
accuracy to which it knows its
own attitude and
3.
accuracy to which it can aim
its beam knowing the required direction.
Acquisition
The transmitted beam cannot be pointed at the communicating pointer
in the open loop made with sufficient accuracy because of uncertainties in the
attitude of the space craft, pointing uncertainties in the terminal and
inadequate knowledge of the location of the other satellite. Consequently
before communication can commence, a high power beam laser located on GEO end
has to scan over the region of uncertainty until it illuminates the GEO
terminal and is detected. This enables the user terminal to lock on to the
beacon and transmit its communication beam back along the same path. Once the
GEO terminal receives the LEO communication beam it switches from the beacon to
the forward link communication beam. The LEO and GEO terminals then track on the
received communication beams, thereby foaming. a communication link between the
LEO and GEO space craft.
Tracking
After successful acquisition, the LEO and GEO terminals are
operating in tracking mode In this mode the on-board disturbances which introduce
pointing fitter into the communication beam are alternated by means f a fine
pointing control loop (FPL) to enable acceptable communications to be obtained.
These disturbances are due to thruster firings, solar arrays drive mechanisms,
instrument harmonics and other effects.
FINE POINTING LOOP
The fine pointing loop (FPL) is required to attenuate external
pointing disturbances so that the residual mispoint angle is a small fraction
of the optical beam width. The closed loop tracking subsystem consists of a
tracking sensor which determines the direction of the incoming communications
beam with an angular resolution around 5% of the optical beam width and a fine
pointing mirror assembly (FPA) which compensates beam mispointing effects. The
SOUT FPL is used to compensate for frequencies upto 80 HZ.
A
three point antivibration mount (soft mount) acts as a low pass filter to form
an isolating interface between the satellite microvibration environment and the
SOUT thereby reducing the bandwidth requirements of the FPL. This also removes
any concerns about uncertainities in the vibration spectrum of the user space
craft. The EPA is implemented by a pair of orthogonal mirrors. The EPA for the
SOUT is based on a dual axis tilting mirror mechanism. This employs a single
mirror and a permanently excited DC motor.
Point ahead
This is needed because of the relative orbital motion between the
satellites which calls for the transmitted beam to be aimed at a point in space
where the receiving terminal will be at the time of arrival of the beam. The
point ahead angle is calculated using the equation
Point ahead angle 2Vt /c where
Vt = transverse
Velocity component of the satellite.
C = Speed
of light
The
point ahead angle is independent of the satellite cross link distance.
CONCLUSION
Optical intersatellite communications promises to become an
important element in future space infrastructure and considerable development
effort is currently underway in Europe and elsewhere. There will be a need for
small optical terminals for LEO space craft once Europe’s data relay satellites
are in orbit within the next five years. The small official user terminal
(SOUT) programme funded by ESA seeks to fill this need for data rate around
2Mbps.
Detailed
design and modelling of the SOUT fight configuration has been carried out and
has provided a high confidence level that the unique terminal design can be
built and qualified with a total mass around 25 Kg. The next phase of the
programme will be to integrate and test a bread board terminal which is
representative of the flight equipment. This breadboard will be used to test
the performance of the PAT subsystem and to verify the structural and optical
configuration for the SOUT.
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