305

multiple beam ant~nnas become more common. Systems such

as INTELSAT

V

and

DSCS-111

will have severe requirements

of

this kind.

The question arises of when the stabilization system

is

no

longer to be classified “dual-spin” but rather “three-axis with

spinning drum providing angular momentum.” One possible

definition

of

dual-spin stabilization

is

that the spinning portion

of

the satellite performs functions other than providing

angular momentum.

As

solar

arrays and antennas become very large (10 m on a

side, or in diameter), the problem of adequately balancing

solar

disturbing torques becomes difficult, and full three-axis

stabilization becomes necessary. More and more satellite

designs are

of

this type.

I)

Attitude Control:

A comparison

of

dual-spin versus

three-axis stabilization

is

instructive. The following three

points explore dual-spin advantages relative to three-axis

stabilization.

a) Simpler attitude-sensing system:

Scanning is provided

by the spinner, arid the spin momentum eliminates the need

for direct measurements

of

yaw angle.

b) Minimum number

of

jet thrusters:

The propulsion

system obtains ullage control (i.e., the feeding

of

propellents

to

the nozzles) from the centrifugal force of the spinner; a

minimum number

of

jet thrusters are required and the same

relatively high

thrust

level can be used for station keeping as

well as attitude control.

c) Attitude “Stiffness:”

The spinning momentum creates

attitude stiffness that reduces the effects

of

torques which are

created within the spacecraft and

also

prevents a rapid accu-

mulation

of

attitude error

as

a result

of

environmental torques.

Ground command thus has enough time

to

provide compen-

sation.

This

attitude “stiffness”

also

can

be used for attitude

control during

an

apogee motor burn (this

also

applies

to

a

three-axis system, but

to

a lesser degree).

The

following four points explore disadvantages of dual-spin

relative

to

three-axis stabilization.

d) Vulnerabih’fy:

A

single catastrophic bearing-failure

mode

can

cause a total telecommunications outage with dual-

spin stabilization. Vulnerable

sliprings

and brushes, and

binding of the despin bearings can cut

off

communications,

thus rendering the satellite useless. Furthermore, power losses

associated with transferring RF

sjgnals

increase with fre-

quency, and redundant encoders/decoders may have

to

be

used on both sides

of

the mechanical despin mechanism.

e) Spacecraft diameter limitations:

A

spinning body, to be

stabilized about a, desired axis, should have a higher stable

shape than a pencil, for example. If the spacecraft diameter

is

limited by the launch vehicle fairing, then this constraint

is

very serious.

f)

Nutational Instability:

Mechanical damping is needed

on the despun platform

to

compensate for nutational instabil-

ity (i.e., “coning”) that results from an unfavorable ratio

of

spin-to-lateral moments of inertia and by energy dissipation

from fuel ‘sloshing

in

the tanks in the spinning portion of the

spacecraft.

g)

Power:

More solar cells are needed for a given power

when mounted on a rotating drum, resulting in a weight and

cost penalty. This factor

is

increasing in importance because

of

the need for more RF power output from any single

antenna, the need for more channels, the use

of

higher fre-

quencies with their lower efficiency transmitters, and more

onboard data processing and automation.

Some general considerations are: reliability of the three-axis

design

is

decreased by the more complex attitude-sensing

system which it requires, but the sensing system can be made

redundant; and dual-spin reliability

is

degraded by the plat-

form despin system, which cannot easily be made redundant.

Spacecraft costs for the two design approaches appear

to

be comparable.

2)

Primary Power

a)

Solar

Cells:

Primary power for communication satellites

mostly

is

obtained by the

use

of silicon

solar

cells. They may

be fixed to the spacecraft body, or mounted

so

that they

can

be

oriented continuously for maximum

solar

energy.

During the equinox seasons, a geostationary satellite

will

be

eclipsed by the earth.

This

means that the satellite

will

be in

the dark for up

to

70

min

per day, depending on the incli-

nation

of

the orbit and the number

of

days before or after

equinox. To maintain operation during such periods, com-

munication satellites depend upon internal batteries, usually

consisting

of

nickel-cadmium cells, although nickel hydrogen

and other technologies are improving swiftly. The batteries

represent a major tradeoff among weight, power, and

performance.

To avoid the

solar-cell

battery limitations, consideration has

been given to the

use

of

nuclear

cells

for powering satellites.

Either radio isotope thermoselectric generations (RTG) or

nuclear reactor powered turbines

can

be used.

A

kg of

UB5

could supply

2.5

MWh

of

energy even at a 10 percent con-

version efficiency. With a half-life

of

lo8

years, it would

outlast most spacecraft.

The advantage of the nuclear supply over

solar

power

systems

is

that no

solar

orientation is required nor

is

any

battery needed. However, heavy shielding

is

required

to

protect the payload from radiation. This disadvantage has

caused solar

cells

to continue to be the preferred primary

power source for communications satellites. Nuclear fuel

handling continues

to

present safety problems both during

manufacture and in the event

of

launch malfunctions. The

safer fuels, such

as

Plutonium, Curium (CmN4), etc., are very

expensive. Strontium (Sr9’), although much cheaper and with

a convenient half-life

of

25

years, is very dangerous to handle.

b) Propulsion:

After launch, one or two types of propul-

sion are required. Satellites launched by Thor-Delta or Atlas-

Centaur launch vehicles inject into transfer orbit only and

require the use

of

an apogee kick motor for injection of the

satellite into geostationary orbit. The weight of this apogee

motor and its propellant

is

typically equal

to

that

of

the

weight of the rest of the spacecraft.

Spacecraft launched by Tital III-C “direct injection” launch

vehicles do not require a separate apogee kick motor, the

functions of orbit circulation and inclination removal being

performed by the launch vehicle itself.

Because of anomalies

in

the earth’s gravitational field and

the perturbing effects

of

the sun and moon,

all

spacecraft

require a small propulsion system for station keeping. Changes

in longitudinal position may be desired from time to time and

also require propulsion.

Hydrazine

is

very popular as a monopropellant because it

has

high

density for storage, low molecular weight and

high

specific impulse. It

is

dense, storable, and catalytic; i.e., it

needs no oxidizer but dissociates on its own.

The change in velocity of a spacecraft

Ar

that can be

achieved (e.g., for station keeping or apogeekick purposes)

is

AV

=

~eln

Mo/Mb

(7)