Space Communication
Space Communication
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The design of space communication systems is not inherently
different from that of terrestrial microwave systems. However,
the unique environment of space cause a change of emphasis
in the design process. The major difference follow :
(A) There is no appreciable fading in space communication systems
because of the conditions to which terrestrial microwave systems
are subject. The only exception to this is rainfall, which affects
systems above 4 GHz. For this reason it is common to calculate
space communication systems with considerably more precision
than terrestrial systems and to design them with considerably less
margin. It is not unusual for system designers to argue about
discrepancies of the order of 0.1 dB in systems which may have
net losses of the order of 100 dB.
(B) In general, spacecraft equipment power output and
power supply are extremely expensive, as it is planetary
receiver sensitivity. Thus, the question of appropriate
economic tradeoffs is a basic one of space-system design
and will ordinarily control the design of the system.
(C) Because of the above , advantage is usually taken of
advance in detection and signal design as quickly as possible.
System Design
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There are three basic types of space communication systems , operating in
different environments and with decidedly different characteristics. On the
following are developed the characteristics, principles, and basic information
for design of each of these types.
(1) Planet-to-Spacecraft Type :
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A communication system working in the planet-to-spacecraft direction is
characterized , at least on Earth, by the relatively easy availability of
electrical power and supporting structures for large antennas. Such a
path will therefore have a relatively large available EIRP (Effective
Isotropic Radiated Power); a relatively high noise background (both
from the warm planet and from the man-made noise generated on
the planet), and a requirement for precise aiming information for the
planetary antenna.
(2) Spacecraft-to-Planet Type :
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In the spacecraft-to-planet direction, the receiving antenna looks into the
relative cold of space; it thus enjoys a relatively quiet background. Power
generation in the spacecraft is extremely expensive, because of both the
weight that must be put into orbit and the difficulty of disposing of waste
heat. Antenna gain is also very expensive, since increasingly large
antennas require increasingly precise stabilization of the spacecraft,
which eventually results in an increase in the amount of fuel necessary
for adjusting the position of the antenna and holding it within the
prescribed tolerance. Reliability is of surpassing importance.
(3) Spacecraft-to-Spacecraft Type :
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Spacecraft-to-spacecraft links enjoy tremendous freedom because
almost all frequencies are available for use, including optical
frequencies. Electrical design of these links is relatively
straightforward, the principal difficulty involving the maintenance
of track between the two spacecraft.
(1) Planet-to-Spacecraft Link
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1.1) System Noise :
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The antenna on the spacecraft receiver will "see" an environment that
contains a number of noise sources besides its intended signal of
y =2 arc sin { 1/ [ 1 + ( h/R ) ] } . There is a background comprising the
broad noise contribution of the galaxy as shown in Fig. 1, with
concentrated sources of noise. The most conspicuous of these
concentrated sources are the sun and any nearby planetary bodies in
the antenna's field of view. The sun's noise temperature is shown in Fig. 2.
The net background noise temperature can be found in a practical case
by simply averaging the noise temperature per unit solid angle, as follows.
When in the vicinity of a planet, the antenna of the spacecraft "sees" the
planet as a noise source having a noise temperature of T kelvins (when T
is the approximate average surface temperature of the planet). The planet
radiates radio noise at a level corresponding to this temperature. Table 1
lists approximate radio blackbody temperatures of bodies of planetary size,
plus other physical constants of interest.
The effective noise temperature of a spacecraft antenna's field of view
is approximately the blackbody radio temperature of the planet, so long
as the planet (as "seen" by the spacecraft antenna) subtends an angle
greater than the beamwidth of the satellite antenna.
When the angle subtended by the planet is smaller than the beamwidth
of the satellite antenna, the satellite sees a combination of galactic noise
and planetary radiation proportional to the relative areas of the planetary
distance. If the spacecraft is at a distance "h" from the surface of a planet
of radius R (in consistent units), the planet will subtend an angle "y" at
the spacecraft.
y = 2 are sin { 1 / [ 1 + ( h/R) ] }
The effective distance of the planet is h+R, and the area of the zone
bounded by the pattern of the spacecraft-antenna beam at that
distance is :
4π (h+R)2/G
where G is the gain of the spacecraft antenna over an isotropic antenna.
The effective noise temperature the antenna sees is :
At distances of three or more planetary radii, [ ( h/R ) = 3 ]. and with gains
of 3 or more, a further simplification can be made to
Te = GTp /4[ (h/R) + 1 ]2
with an error of less than 10 percent at (h/R) = 3. The error decrease rapidly
with increasing G or h/R. If the spacecraft antenna sees only part of the
planetary surface (as in communication satellites with high-gain antennas)
the noise contribution ideally should be found by vector integration over the
visible surface of the planet or, more practically, by estimating the ratios of
areas involved and their effective distances.
The noise density at the spacecraft receiver (in dBW of noise per hertz of
receiver bandwidth) is 10 log k( Te + TR ), where Te is the
effective external noise temperature, TR is the noise temperature of the receiver
without antenna, and k is Boltzmann's constant. This can be stated as
n = -228.60 + 10 log (Te + TR , in dBW / Hz.
(In the path calculation equations, decibels are used throughout where possible).
Because of the relatively noisy receivers often provided in spacecraft, especially
when intended doe use close to a planet, as in communication satellite service ,
it is sometimes more convenient to use the receiver noise figure. If F is the receiver
noise figure (a number; the noise figure expressed in dB is N = 10 log F ), it may be
converted to noise temperature by the relation
TR = 290 (F-1)
which assume the conventional 290 K reference temperature. The use of the
noise figure is declining in space applications, since the standard temperature
of 290 K is essentially meaningless away from a terrestrial environment. In
addition , low-noise receivers produce negative standard noise figures when
expressed in decibels, leading to confusion. If a space-borne receiver had a
"5-dB noise figure", its effective noise temperature by TR = 290 (F-1) would be
about 627 K.
Figures 1 & 2 and Table 1 as reference to post #2 :
1) Planet-to-Spacecraft Link
1.1) System Noise :
(1) Planet-to-Spacecraft Link
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1.2 ) Path Losses :
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If the planet has a substantial atmosphere, there will be losses in signal
strength from absorption by molecular-dipole resonances; in addition,
there will be attenuation from precipitation 4π (h/R)2/G. The amount of
this attenuation varies with the climate of the area to which communication
is intended, as well as with the angle at which the transmitted signal leaves
the surface of the planet (thus determining the distance the signal must
travel through the atmosphere). Because of the resonant nature of some
of the absorption, the attenuation is strongly frequency-dependent.
Figure 3 gives terrestrial measurements of clear-air attenuation at 30o and at
the zenith. Figure 4 gives information on millimeter-wave attenuation caused by
rainfall, about which information is particularly sparse. However, measurements
are continuously being taken, and designers should examine the latest literature
for statistical data if in a terrestrial path. The use of frequencies above 8 GHz is
contemplated.
In the following equations the atmospheric attenuation is given simply as La (in dB).
It is assumed that the designer of space communication systems will use the
available data from space probes for the composition of the atmospheres of
unfamiliar planets. The free-space loss is
LFS = 92.45 + 20 logf + 20 logd
where LFS is in decibels , d in kilometers, and f in gigahertz. Alternatively
LFS = 96.58 + 20 logf + 20 logd
where d is in statute miles.
The total path loss is
Lp = 92.45 + 20 logf + 20 logd + La
where Lp and La are in decibels, d in kilometers,
and f in gigahertz.
The rf carrier-to-noise ratio in the receiver is then
C/N = P - LFS -La + G - n - 10 logB
where P is the transmit EIRP in dBW, and B is the receiver noise bandwidth
in Hz.
The gain
of the receiving antenna is G (in dB over isotropic). Expressions for the
gains of various types of antennas ar given in Table 2.
C/N Calculation :
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A complete equation for C/N (in dB) is
C/N = P + 136.15 - 20 logf - 20 logd - La
+ G - 10 log (Te + TR ) - 10 logB
Because the frequency term cancels, it is sometimes simpler to work with
the effective area A of the spacecraft antenna. The equation is then becomes
C/N = P + 157.61 + 10 logA - 20 logd - La
- 10 log (Te + TR ) - 10 logB
where A is in square meters.
Figures 3 & 4 and Table 2 as reference to post #4 :
1) Planet-to-Spacecraft Link
1.2) Path Losses :
(2) The Planetary Receiver
[ Spacecraft-to-Planet Type ]
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2.1 ) General Considerations
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Except for initial probes planetary stations are characterized by high available
transmitting powers and relatively high-gain antennas. Heavy cryogenic systems
can be provided for extremely low-noise receivers, and the relatively low cosmic
noise temperature makes their use practical. The effect is to reduce the power
requirements of the spacecraft transmitter. Also, the stable platform afforded by
the planetary surface allows precise antenna positioning.
2.2) Convention of Planetary Station Design
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Because of the very large antennas often used, antenna gains are often hard to
measure precisely, and in any event these gains may vary with dish position
because of deformation. Similarly, because of the extremely low noise
temperature of the receivers, precise in-place measurements of their noise
temperatures are very difficult to make. It has therefore become customary to
measure and specify the signal-to-noise ratio (which is relatively easier to
measure) with a known received signal strength. A convenient way to express
this is the ratio G/T, usually expressed in dBW/K . It is the antenna gain divided
by the system noise temperature , or
G/T = G - 10 logT
where G is expressed in decibels, and T in K.
Similarly, the received carrier-to-noise ratio is often expressed as
C/T (in dB), which is
C/T = 10 logPR - 10 logT
and is related to the true carrier-to-noise ratio by
C/N = C/T + 228.60 - 10 logB
where B is the receiver noise bandwidth in hertz, and the other values are
in decibels.
Because of the relatively broad patterns of spacecraft antennas, and for
convenience in comparing systems, the available signal is often expressed
as a "flux density" Ψ (PSI) in watts/square meter or more usually in dBW/m2.
This flux density is given by
Ψ = PT - 70.99 - 20 logd - La
where PT is the EIRP of the spacecraft in dBW, Ψ is in dBW/m2, and La is in dB.
With a known flux density and (G/T), the (C/T) can easily be found by
C/T = Ψ + G/T - 20 logf - 21.46
where f is in GHz, Ψ in dBW/m2, and the other values are in dB.
2.3) Noise Sources at a Planetary Station :
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It is much more complex to calculate the effective noise temperature of
a planetary station than that of a spacecraft, because of the greater
variety of noise sources and the greater sensitivity of the receiving
system. The galactic noise contribution must be considered, along with
noise radiation from the atmosphere and surrounding surfaces. In the
region between 1 to 10 GHz, all noise sources must be considered; the
greatest noise contribution is usually radiation from the planetary surface.
To some extent the design of the antenna must be compromised, in that
side-lobe levels must be kept lower than would otherwise be desirable,
to avoid pickup from the ground. This often results in a compromise of the
optimum illumination for the highest gain.
The general way to arrange a noise budget for such a station is to
consider the lineup of equipment between the antenna feed and
the low-noise pre-amplifier. This might look like the diagram show
in Fig. 5. The system noise temperature is obtained by adding the
noise contributions listed in Table 3.
Tf , Tp , and TFL are the actual physical
temperatures of the respective devices shown in Fig. 5. The blackbody
noise contributions of these devices will be significant if high transmit
powers and cooled pre-amplifiers are involved. Usually it is helpful to cool
as much of these assemblies as is possible.
(3) The Spacecraft-to-Spacecraft Link
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At the present state of knowledge, it appears that the planner of space
communication systems has almost complete design freedom so long as
he is concerned with communication between two points that are
permanently in space. The net path loss (in dB) is simply.
Lp = 49.53 - 20 logf + 20 logd - 10 log Ar - 10 log AR
where f is in GHz, d in km, Lp in dB, and Ar and AR
(the areas of transmitting and receiving antenna, respectively) in m2.
As frequency is increased, the transmit power or the antenna sizes can be
reduced; it is clearly better to use the highest frequency for which generators
and receivers are available . Optical frequencies are a possible choice; the basic
difficulty of this approach is to maintain the precise track and platform stability
necessary for use of the narrow beamwidths available with an optical system.
Satellite Repeaters
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The earliest example of space communication involved a satellite repeater
used for communication between two terrestrial points. Of course, satellite
repeaters can be provided for communication between any two points at
which the satellite is mutually visible, and they can used in complex
arrangements involving switching, store-and-forward, and the like. They
quite useful in space exploration. The most practical arrangements have
involved active satellites. Passive satellite's relative immunity from jamming
makes it attractive for military applications.
Passive Satellite Repeaters
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A passive satellite repeater is really a radar system in which the receiver
and transmitter are not co-located. Radio-frequency system design is
relatively uncomplicated, involving the basic radar equation.
Expressed in dBW/m2, the received flux density is :
Ψ = Pr - 141.98 - 20 logd1 - 20 logd2 + 10 loga - L aU - LaD
where Pr is the transmitted power in dBW , d1 and d2 are the slant ranges
of the satellite to the transmitter and the receiver in km, "a" is the radar
cross section of the satellite in m2, and LaU and LaD are the
up- and down-path atmospheric losses, respectively, in dB. It is obvious
that a passive satellite system using a relatively high altitude requires
enormous transmitter EIRP for more than a few channels. For a
carrier-to noise ratio of 20 dB in a telephone channel (kTB = 141.7 dBm),
an 85-foot-diameter antenna and a noiseless receiver, a 6,000-km orbit,
and a 100-foot-diameter spherical reflector, (and ignoring atmospheric losses),
the transmitter power at 6 GHz must be about 150 watts per channel.
More-efficient reflectors can be made, of course. The most effective in terms of
reflecting cross section per unit weight is the type made of orbiting thin wire
dipoles (needles).
It has been determined that the dipoles have an average reflecting cross
section of 0.158 λ2 (per dipole), including the effects of random
dipole orientation. Exact calculations are complex, since, to be precise, a
way must be found to sum the contributions of each dipole in the scattering
volume.
The common scattering volume can be found geometrically. If the belt is much
narrower than the antenna beamwidth, the scattering cross section is
x = NlL (0.158 λ2 )
where Nl is the number of dipole per unit length and L is the length
of the common cylinder.
If the belt is considerably wider than the beamwidth, the common volume will
be an ellipsoid, the scattering cross section of which is
x = NvV ( 0.158 λ2 )
where Nv is the number of dipoles per unit volume and V is the
volume of the ellipsoid.
Natural satellites are occasionally used for passive repeaters. They have
the characteristics of large cross-sectional areas and relatively poor reflectivity.
In general, these rough bodies are far from specular reflectors. The Moon,
for instance, has an effective reflecting cross section of 9.48 x 1011 m2,
about 1/10 of its physical cross section. Similar
information is not now available for other natural satellites; it is reasonable
to assume, however, that essentially airless satellites have similar terrain
and thus similar ratios of radar to physical cross section.
The great disadvantage of these satellites is that they are not continuously in
usable positions; however, their extremely large reflecting cross sections make
them attractive for moving low-priority record traffic and for planetary
exploration and development.
Active Satellite Repeaters
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The simplest form of an active satellite is one in which a signal is received,
amplified , and re-radiated. The design constraints on such a device are
set by the following considerations :
(a) When the satellite is used as a relay between planetary stations,
up-link power is usually adequate to permit the use of uncooled receiver
pre-amplifiers.
(b) The system performance is therefore set by the satellite EIRP. This is ,
in the final analysis, a function of the orbited weight of the satellite; it is
determined by the specific weight of available primary power sources and
by the lifting capacity and cost of available boosters.
(c) The operating cost of the satellite itself is simply the cost of replacing
the satellite divided by the satellite's life expectancy.
operating cost/year = launch cost / (MTBF) x P
where MTBF is the mean time, before failure, and P is the possibility
that the launch will be successful and that the satellite will work
properly when in place.
For systems involving relatively small numbers of satellites, such as
synchronous systems, conservative accounting demands that at
least one launch failure be assumed at the outset. In-place standby
equipment is not unusual in such systems. Therefore, reliability is of
almost overwhelming importance.
The block diagram of a simple satellite repeater is given in Fig. 6.
Because efficiency is paramount, systems of this type are usually
operated quite close to the level at which limiting occurs. The
limiting characteristics of most practical , efficient amplifiers are
rather "hard". This characteristic is helpful when the repeater is
used to amplify pulsed signals, and the amplifier in these cases
is normally run in a limiting condition.
Some systems, however, use arrangements in which each of a
number of rf carriers handles multi-channel telephony, frequency
modulated on the carrier. As carriers are added , each shares the
available output power of the satellite. However, when the
satellite's amplifiers are operated close to the limiting point, some
of the power is dissipated as distortion products. This reduces the
expected output power slightly and also causes cross modulation
and distortion among he carriers.
Measurements have been made on a satellite using traveling-wave
tubes. Table 4 gives the effect of added carriers on the available
output power of the satellite , relative to its saturated power out.
Figure 7 gives the inter-modulation performance under the same
conditions, using equipment for 340 voice channels.
This type of fm multiple-access system is quite susceptible to
jamming and interference, since a signal strong enough to drive
the amplifier into hard limiting will monopolize essentially all the
available power of the satellite transponder.
Effects of Motion
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Spacecraft in an Atmosphere
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If a spacecraft is to pass through an atmosphere at high speed during
some part of its flight, it may generate a plasma sheath that has drastic
effects on both the attenuation of the atmosphere and on the patterns
and impedances of the spacecraft antennas. The plasma is formed
behind a shock front , the shape and position of which are determined
by aerodynamic considerations.
The electron density of the plasma sheath is the primary determinant of
the electrical properties of the plasma. This electron density (in electrons
per cubic centimeter) can be obtained from Fig. 8 for the Earth's
atmosphere, for a point immediately behind the normal shock, as a
function of altitude and spacecraft velocity. For cylindrical spacecraft,
this electron density drops to about 1/100 of this value at 3 spacecraft
radii behind the normal shock, and roughly linearly thereafter at 1/10 per
10 radii.
It is usual to assume the plasma thickness L, above the antenna,
as 1/2 of the spacecraft radius.
Figure 8 also can be used to find the collision frequency "g"
(collisions per second), which , at 3 spacecraft radii behind
normal shock, drops to about 1/10 of the value at normal
shock and decreases linearly thereafter at 1/10 per 10 radii.
The resulting plasma will behave much like a high-pass filter,
with a definite cutoff at approximately the "plasma frequency" wp.
wp = 5.642 x 10-2 (Ne)1/2
where Ne is the electron density from Fig. 8. Above cutoff,
there will be several absorption lines from dipole resonances. These
absorption lines usually extend well into the infrared region.
A general idea of the resulting behavior of the plasma can be gained
from Fig. 9 , where the absorption coefficient and refractive index of
the plasma are plotted for an electron density of 1012 electrons
per cubic centimeter.
Spacecraft in an Atmosphere ........continued...
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Idealized Case
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