IRAS Explanatory Supplement
III. The IRAS Mission
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- Attitude Control
- Solar Radiation
- Earth Radiation
- Moon and Planets
- South Atlantic Anomaly
- Station Passes
- Constant Sun Angle
- Eclipse Operations
B. 1 Introduction
Figure III.B.1 A schematic drawing of the orbital geometry. The orbital
altitude, 900 km, and inclination, 99°, combined with the Earth's
equatorial bulge lead to a precession of the plane of the orbit about 1°
per day. As a result, the orbit normal always pointed towards the
Sun as the satellite orbited above the Earth's terminator. By pointing
the satellite radially away from the Earth, the cold telescope was shielded
from the heat loads from the Sun and Earth while providing natural scanning
motion across the entire sky in about six months. A sequence of
hours-confirming scans on the celestial sphere is also shown.
IRAS was successfully launched into its planned 900 km altitude,
99° inclination Sun-synchronous polar orbit with a period
of 103 minutes. This orbital altitude was low enough to be below
most particles in the Earth's trapped particle belts yet high
enough that only a negligible amount of residual atmospheric gasses
would build up on the cold mirror surfaces during the mission.
With the telescope pointing radially outwards from the Earth
and perpendicular to the Sun vector, no Earth or sunlight could
enter the telescope and all ecliptic latitudes would be swept
out during one orbit while, as the line of nodes precessed at
a rate of about 1° per day to remain perpendicular to the
Sun vector, all ecliptic longitudes would be covered in a period
of six months (Fig.III.B.1). Such a simple
solution would, however,
have allowed no flexibility. The attitude control system and
telescope were designed to allow pointing away from the local
vertical within certain constraints which are described below.
B.2 Attitude Control
The first constraint was designed into the spacecraft attitude control system and required the telescope to point no further away from the Sun than 120°. At greater angles the fine Sun sensor could no longer see the Sun well enough to function.
B.3 Solar Radiation
Figure III.B.2 With IRAS at the center of the celestial sphere,
the solar constraint on solar radiation and visibility to fine Sun
sensors prohibited viewing closer than 60° from the Sun and
farther than 120°, the area shown shaded in the figure.
The second constraint was that the sunshield design did
not allow the telescope to point closer than 60° towards
the Sun without solar radiation falling onto the inside of the
sunshade. These first two constraints limited the celestial sphere
available to IRAS at any one epoch as shown schematically in
B.4 Earth Radiation
Figure III.B.3 The Earth infrared radiation constraint is shown in
telescope coordinates. The telescope boresight is at the original with
the +X axis out of the page. The scanning direction is toward -Y.
The satellite position vector had to remain within the shaded region
at all times.
The third constraint arose from prohibiting any infrared
radiation from the Earth from falling upon the inside of the sunshield
or the top of the telescope baffle system. This constraint is
shown in telescope coordinates in Fig. III.B.3
and the corresponding
area of the sky available on any one orbit is shown in
The Earth radiation constraint together with the spacecraft orbital
rate determined the maximum time of 15.6 min. during which the
satellite could point at a given fixed celestial position, if
no other constraints interfered. Because of the varying solar
declination during the mission, the joint effect of these constraints
changed during the mission; Fig. III.B.5 shows
schematically the combined constraints at two epochs six months apart.
Figure III.B.4 The shaded region depicts the portion of the celestial
sphere not available for viewing during any given orbit due to the Earth
Figure III.B.5 Basic viewing window (unshaded) on the celestial sphere
for two different dates during the survey.
B.5 Moon and Planets
Infrared radiation from the Moon and the planet Jupiter was sufficiently strong to affect the performance of the detectors for a significant time after passage over the focal plane. Of the other planets, only Venus was bright enough to have this effect but it was always too close to the Sun to be observed. An avoidance radius of 1° from Jupiter was set within which the telescope did not point. For the moon, an avoidance radius of 25° was used during the first two months of the survey but was lowered to 20° after April 3 except between August 26 and September 9 where it was lowered to 13°; at 25° significant "moon glints" were introduced into the data-stream (see Section IV.C).
B.6 South Atlantic Anomaly
Figure III.B.6 Sample orbital tracks through the South Atlantic
Anomaly (SAA) at 900 km and the contours used for SAA avoidance are
given. The contour (A) was determined during in orbit check-out;
the less conservative contour (A) was used after May 9 to help reduce
the SAA effects on the survey scans.
Another constraint was the depression in the Van Allen belts known as the South Atlantic Anomaly (SAA). Proton hits in the detectors when passing through the SAA increased the noise to such an extent that it was impossible to continue observations. Data were not taken whenever the satellite entered a geographically fixed flux/energy contour. As a result of analysis of the effects of radiation on the detectors, the contour shown in Fig. III.B.6 was adopted. In May, it was slightly reduced in an attempt to minimize the adverse effects on the survey scans. Both contours are shown in Fig. III.B.6.
A second effect of a passage through the SAA was a long
term enhancement, by as much as factors of ten, in the responsivity
and noise caused by the large radiation dosage. As described in
Section II.C.5, these large changes
could be erased by increasing
the bias voltage on the detectors, a technique referred to as
B.7 Station Passes
No observations could be carried out during the prime pass over the ground station (Bevan et al. 1983), as during this time (typically 10 minutes every 10-14 hours) data from the preceding 10-14 hour observation period were being transmitted from the on-board tape recorders to the ground and the commands for the next 10-14 hours of observations were being sent to the satellite (Mount 1983; MacDougall et al. 1984).
B.8 Constant Sun Angle
Figure III.B.7 The cone and clock system were used to define the
scan geometry. The cone angle
() is the angle
between the satellite-Sunline and the boresight. The clock angle
is measured in the plane perpendicular to the Sun-line, clockwise
as viewed from the Sun. Sometimes
= 360° -
is also used.
Figure III.B.7 shows the attitude-control
coordinate system of the spacecraft. Although the spacecraft had gyros
for 3-axis control, it normally used only one, the z-axis gyro
II.B.1). Control of the other two axes was maintained by the
fine Sun sensor which ensured that the y-axis was always perpendicular
to the satellite-Sun vector see Section
II.B.2). The satellite
scanned with a fixed cone-angle
between the telescope
boresight and the Sun vector.
Figure III.B.8 The curvature of scans taken at
90° is called the
"banana effect". See Section III.B.8.
A consequence of the constant Sun cone angle constraint, together with the fact that the ecliptic longitude of the Sun constantly increases, is that it was never possible to repeat exactly the coverage obtained by any scan, except at the ecliptic plane, or after an interval of six months For example, a pole-to-pole scan executed at cone angle = 90° and passing through a specified longitude in the ecliptic plane would be rectangular in the ecliptic coordinate system. A pole-to-pole scan through the same specified longitude in the ecliptic plane at a later or earlier time ( = 90°) would be an arc whose curvature increased with |90- | (Fig. III.B.8). This "banana effect" meant that overlapping scans had to be made at almost the same time with only small differences.
B.9 Eclipse Operations
Towards the end of the mission, the solar declination became such that the Sun was eclipsed by the Earth during part of the orbit. Because the satellite could not use its Sun sensor, gyros were used to control all three axes during this period. This resulted in a considerable loss of control accuracy and in particular, meant that slewing maneuvers ended in unpredictable positions. Consequently, it was decided not to continue survey scans or other normal observations during eclipses.