Space Communication


Space Communication


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.

November 21, 2020 6:46 PM


System Design


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 :


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 :


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 :


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.

November 21, 2020 6:48 PM


(1) Planet-to-Spacecraft Link


1.1) System Noise :


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.

November 21, 2020 6:49 PM


Figures 1 & 2 and Table 1 as reference to post #2 :

1) Planet-to-Spacecraft Link

   1.1) System Noise :


November 21, 2020 6:50 PM


(1) Planet-to-Spacecraft Link


1.2 ) Path Losses :


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


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 :


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.

November 21, 2020 6:52 PM


Figures 3 & 4 and Table 2 as reference to post #4 :

1) Planet-to-Spacecraft Link

  1.2) Path Losses :

November 21, 2020 6:52 PM


(2) The Planetary Receiver

   [ Spacecraft-to-Planet Type ]


2.1 ) General Considerations


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


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 :


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.

November 21, 2020 6:55 PM


(3) The Spacecraft-to-Spacecraft Link


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.

November 21, 2020 6:56 PM


Satellite Repeaters


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.

November 21, 2020 6:57 PM


Passive Satellite Repeaters


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


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. 

November 21, 2020 6:58 PM


Active Satellite Repeaters


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


(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.

November 21, 2020 7:00 PM


Effects of Motion


November 21, 2020 7:00 PM


Spacecraft in an Atmosphere


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.

November 21, 2020 7:01 PM


Spacecraft in an Atmosphere ........continued...


November 21, 2020 7:03 PM


Orbit Mechanics


November 21, 2020 7:03 PM


Idealized Case


November 21, 2020 7:04 PM