NASA/Goddard Space Flight Center
Greenbelt, Maryland
Summary One of NASA’s newest Explorer class satellite programs,the Small
Astronomy Satellite (SAS), will provide much-needed information in the most recently
studied fields of astronomy, X-ray, Gamma-ray, UV and IR. This paper will describe the
basic spacecraft functions with emphasis on its key feature, the SAS control system, as
proposed for early sky surveys, and the changes needed in it for later flights which will
require a pointing capability of one arc-minute or better. Its flexibility and versatility for
application to many different types of astronomy experiments will be examined.
Introduction The objective of the SAS program is to study the celestial sphere above
the earth’s atmosphere and to search for sources radiating in the X-ray, Gamma-ray,
ultraviolet, visible, and infrared regions of the spectrum both inside and outside of our
galaxy. Surveys performed from such a small, Scout-launched satellite (Figure 1) will
provide valuable data on the position, strength, and time variation of sources radiating in
these regions of the spectrum and lead to the selection of the more interesting ones for
detailed study by later missions on this or more sophisticated spacecraft. It is anticipated
that one of these satellites will be launched each year, with an estimated lifetime of a
year, unless limited by the experiment sensors.
The planned orbital characteristics of the SAS series are a nominally circular 300
nautical mile (550 kilometer) equatorial orbit. The equatorial orbit is highly desirable for
an astronomy mission because the satellite will miss the South Atlantic anomaly where
the radiation belts come far down into the earth’s atmosphere (Figure 2). This will avoid
degradation of the spacecraft operation and maintain at a minimum the background count
which can affect the data returned from several different types of sensors applicable to
satellite astronomy. This orbit, with an inclination of 2.9 degrees, can be achieved by
launching from the San Marco platform located three miles east of Kenya in the Indian
Ocean. The maximum payload capability under these conditions is 333 pounds.
Basic Spacecraft Design The satellite is designed to have a clean interface between
the basic spacecraft structure and the experiment payload (Figure 3). This is intended to
minimize costs for follow-on programs. The basic spacecraft is a cylinder approximately
22 inches in diameter and 20 inches high, weighing about 180 pounds. This lightweight
cylindrical shell acts as the primary support for the experiment. Four solar paddles,
hinged to the outer shell, provide raw power to the spacecraft and experiment. At the tips
of the paddles are the command and telemetry antennas. Inside the shell (Figure 4), and
thermally isolated from it is a honeycomb deck which is itself a good thermal conductor.
On it are the systems to provide the basic spacecraft functions: a rechargeable battery
with its charge control and regulator systems; command receivers and decoders; a 1000
bit per second telemetry system with one orbit storage capability on tape; and a
magnetically torqued commendable control system which can point the spacecraft stably
to any point in the sky. The electronic circuitry is packaged in “books” mounted with the
mutation damper on the underside of the honeycomb deck. The battery, tape recorder,
transmitter, command receivers, and rotor are mounted on the top side.
One of the design requirements for SAS was to be able to point in any orientation, thus
complicating the thermal design as well as that of the power system. The surface of the
cylinder is the only part of the spacecraft portion which radiates the internal heat loads to
space. Since the experiment is isolated from the spacecraft section with a high thermal
resistance, we will consider the spacecraft section independently of the experiment
section. The most critical items are the tape recorder and the battery, which are colocated for efficient thermal design. The judicious use of passive thermal coatings and
multilayer insulation, with a small amount of active thermal control in the form of
heaters, permits operation in any random orientation. While most of the spacecraft
components can be operated satisfactorially from -20EC to +70EC, the battery prefers an
environment between 0E and +30EC. A separate array of solar cells on the base of the
spacecraft provides heater power for the experiment in the worst case cold condition.
Power Subsystem The power system consists of four paddles with solar cells on both
sides; a 6-ampere-hour 8-cell nickel cadmium battery; and a shunt regulator. Still
assuming a random orientation, the power system must guarantee 27 watts average
power over the entire orbit, both nighttime and daylight portions. This is particularly
important to astronomy experiments, which frequently get their most useful data at night,
when scattered sunlight cannot interfere with the sensors. The basic spacecraft functions
require 17 watts average power, allowing 10 watts for the experiment. From Figure 5 you
can see the variation in available power as a function of sun angle. Total power includes
operational needs plus enough additional power to charge the battery to provide the same
amount of power through the nighttime portion of the orbit.
A block diagram of the power subsystem is shown in Figure 6. The solar cell array is
connected into two separate sections, the main array and an auxiliary array. The auxiliary
array is used to provide redundant power for the command subsystem and to trickle
charge the battery if it has been removed from the main bus by the low voltage sensing
The six ampere-hour nickel cadmium battery is protected by two redundant charge
control devices, a coulometer and an enhanced zener. The coulometer senses the total
energy taken from and returned to the battery by integrating the voltage developed across
the battery telemetry resistor. When the energy which was taken from the battery on
discharge has been returned, the coulometer will reduce the battery charge current to a
“safe”, preset trickle charge level by shunting the remainder. This unwanted current is
shunted internally if the battery is cold, or externally if the battery is hot. The enhanced
zener is a voltage limiter. When the battery voltage reaches a preset level, the enhanced
zener will actuate a shunt, decreasing the battery charge current. Thus the battery charge
current will be reduced to a level which will cause the battery voltage to be equal to the
enhanced zener limit voltage. As with the coulometer, the shunted current may be
dissipated internally or externally depending on battery temperature. The power
subsystem will normally be operated with both the coulometer and the enhanced zener.
However, it can operate with either one alone.
The low voltage sensing switch continuously monitors the voltage of the main bus. If it
drops below 8.8 volts (normal voltage is 10.7 v.), power is switched from the main
converters and the battery in order to protect them from damage due to possible shorted
loads. In this “solar only” mode, the battery will be trickle charged by the auxiliary array,
while the only loads on the main array are the dual command systems and the heater
circuits. Each circuit can be switched out by command if it becomes defective.
Command Subsystem SAS employs the NASA standard PCM command system. The
spacecraft carries dual redundant receivers and decoders. The redundancy continues
through the relay coils themselves, but a single set of contacts is used. Figure 7 shows
the command system in block diagram form. Independent dipole antennas are mounted
on the tips of two adjacent solar paddles. The command signal, a pulse code modulated,
frequency shift keyed, amplitude modulated (PCM/ FSK-AM/AM) signal, is
demodulated by either or both command receivers and appears at the video output. Bit
synchronization is obtained from a signal that has been amplitude-modulated onto the
data subcarrier. The video signal is sent to a command decoder, which tests the tones for
authenticity by examining timing and frequencies. If the tones are proper, the decoder
converts the tones to digital levels and sends the levels and appropriately timed clocking
pulses to the logic circuitry. The decoder also applies power to the standby portions of
the logic, beginning with the first correctly received tone burst and remaining on long
enough to complete the command. The logic circuitry accumulates the decoded levels in
a shift register and, if the satellite address bits are proper, it activates the proper row and
column of the relay switching matrix to switch the desired relay or relay cluster.
The redundant interconnections show that only one receiver, one command logic, and
one relay switching matrix need be operating to send commands. One bit in the
command word (Figure 8) addresses the particular logic circuit and one bit the particular
relay switching circuit to be used, so that only one logic and one matrix are activated
during any given command. Each half system has its own power converter. The system’s
magnetic latching relays have two coils per relay, one for each system half, and consume
power only during a command. The direction of current flow in any given relay coil
determines the state in which the relay will latch. Because of these dual coils, actually
located in the relay switching matrix, the system’s redundancy continues up to the
magnetic flux circuit in the relay pole piece, and no single relay coil open or short can
disable the system.
The command system provides SAS with 35 on-off relay commands, 25 for the
spacecraft and 10 for the experiment. In addition, if bit 25 in the command word
specifies a “data command” instead of a “relay command”, then bits 33 through 56 are
routed, as specified by bits 27 through 32, to a location other than the relay matrix for
separate decoding. This permits an experimenter to increase his allocation of 10 on-off
commands by decoding up to 224 within the experiment.
Telemetry Subsystem Due to the low equatorial orbit chosen for SAS, the spacecraft
can be seen regularly by only one of NASA’s Satellite Tracking and Data Acquisition
Network (STADAN) stations. For this reason, the spacecraft telemetry subsystem
(Figure 9) must include on-board storage of the data collected during the remaining 90%
of the orbit. Therefore, the pulse code modulated/phase modulated (PCM/PM) telemetry
system has two basic modes of operation - record and playback. In the normal, or record
mode, digital data from various sources are multiplexed and sent to be stored serially in
Manchester Code on an endless loop tape recorder and to phase-modulate a VHF
transmitter. This continuous output, radiated through a turnstile antenna at the tip of one
of the solar paddles, is used as a beacon for tracking the satellite and for the collection of
real-time data. As the satellite passes over its data acquisition station, however, the tape
recorder is commanded into playback mode and the data from the full 96 minute orbit
phase-modulates the same transmitter with a 30 kbps signal. Simultaneously, the
transmitter is switched into its high power (2 watt) mode. The tape recorder and the
transmitter automatically return to the record mode of operation after completion of the
3.4 minute playback cycle, or upon ground command. In a special operating mode, realtime data can be transmitted at 2 watts. This system meets the requirements of the
NASA/GSFC Aerospace Data Systems Standards for PCM telemetry systems. Phase
deviation is ±64E peak, with a realtime bandwidth allocation of ±15kHz and a tape
recorder playback allocation of ±45kHz. Linear phase filters, having their 3 db points at
1kHz and 30kHz respectively for record and playback mode, are used for premodulation
The 1000 bps real-time data rate was determined by optimizing experimenter and
housekeeping data requirements with a reasonable length of tape (300 feet) and a
reasonably conservative bit packing density for single track recording (1667 bits per
inch). The tape recorder, of flight-proven endless loop design, weighs only 7 pounds and
requires 1.5 watts in record mode and 3.0 watts in playback mode. Wow and flutter will
be kept below 2% peak-to-peak, with jitter below 3% peak-to-peak. Commands will
bypass the output reclocking circuitry and energize the erase head, as well as provide the
obvious record/playback and on/off control. Digital data recording was selected for its
better signal-to-noise properties and the fact that the data could be provided to the
telemetry system most easily in digital form.
The basic format is shown in Figure 10. The minor frame is composed of 32 words, each
having three 8-bit syllables, which permits easy handling of 8, 16, or 24 bit data words.
Of these 96 syllables, the first three are allocated to frame synchronization. The frame
synchronization word used is the optimum 24 bit code word determined by Maury and
Styles1, 11111011111001100100000. One syllable is allocated to frame identification,
three to 64 position analog subcommutators, two to digital subcommutators, one for a
parity word and the remaining 86 for data. The period of the minor frame is 0.768 second
and, for the major frame, 64 times that or 49.152 seconds. The basic timing source for
the system is a temperature-compensated 1.024MHz crystal oscillator, whose frequency
variation is less than one part in 106 per orbit. It serves as the source for all readout gates,
clock signals to control the multiplexing of spacecraft and experiment data, motor drive
frequencies for the tape recorder and rotor, and pulses for the 20 bit accumulator which
is incremented by the minor frame identifier every major frame, and sampled every eight
minor frames. Unambiguous time correlation for about 20 months is provided.
While most of the data originates in digital form, many of the housekeeping functions are
analog and must be converted. A dual-slope integrating analog-to-digital converter,
accurate to 8 bits, is used for this purpose. It is clocked at 16kHz and has an aperture of
8ms. While its normal inputs are 0 to 250mv, the telemetry system can provide a variety
of attenuators so that signal levels of ±0.5, ±1.25, ±2.5, ±5.0, or ±7.5 volts can be
accepted. An additional feature of the analog subcommutators is that it is possible to
command either of them to stop and sample continuously, at the minor frame rate of
0.768 second, any one of its 64 positions. After receiving the RUN command again, it
will automatically resynchronize itself at the beginning of the next major frame.
One of the digital subcommutators has 16 channels. Twelve of the channels accept 8-bit
parallel data while the other four accept 8-bit serial data. The second digital
subcommutator has 8 channels, each of which accepts 8-bit serial data. These digital
subcommutators provide us with command verification, the unambiguous time code, a
monitor of the rotor speed, and time correlation to 24 ms for the digital solar aspect
sensor data.
Jesse L. Maury, Jr. and Frederick J. Styles, “Development of Optimum Frame Synchronization
Codes for Goddard Space Flight Center PCM Telemetry Standards”, NTC ’64 Proceedings
This telemetry system is less than 0.2 cubic foot, weighs 14 pounds, and consumes 4.5
watts. It is built in modular form for easy expansion, and the design is such that the
present hard-wired format could be replaced by an in-flight programmable unit if these
requirements occur in the future.
Control Subsystem The most unique feature of the Small Astronomy Satellite is its
control system. An evolutionary system is planned which will permit the present
scanning system to progress to an accurately-pointed, three-axis stabilized spacecraft, a
most valuable tool in astronomy.
First, a look at the basic control system. The prime contractor for SAS, the Applied
Physics Laboratory of the Johns Hopkins University, has a well-established reputation in
the field of magnetic control. An obvious advantage of the use of the earth’s magnetic
field for torquing near-earth satellites is the elimination of the need for expendables, and
their inherent life limitations.
Experiments can look along the Z-axis of the SAS spacecraft or perpendicular to it. If
one is looking out perpendicular to the Z-axis, and the spacecraft is rotated slowly about
the Z-axis, then a swath is swept out of the celestial sphere. The vector made by the
Z-axis in space can be moved, and another swath swept. Thus, ultimately, the entire
celestial sphere can be scanned.
This sweeping technique is valid even if the Z-axis wobbles, but data reduction can be
made much easier if it is stable. One way of achieving stability is by spinning the
spacecraft. This is quite satisfactory for many experiments looking along the Z-axis.
However, a high spin rate will reduce the effective time-on-target for perpendicularlooking experiments and, consequently, the signal-to-noise ratio of the celestial sources.
An alternative method of adding the required angular momentum along the Z-axis is to
include a momentum wheel whose axis is parallel to the Z-axis.
Figure 11 shows the various components of the control system. Controlling the Z-axis
orientation is accomplished by energizing a torquing coil. This electromagnet, acting like
a compass needle, attempts to align the spacecraft with the earth’s magnetic field. At the
maximum rate, using a magnetic dipole of 5 x 104 pole-centimeters, it is possible to
move the Z-axis of the spacecraft 1.72 degrees per minute.
Figure 12 shows how a maneuver might be performed to point the Z-axis to a new
direction as requested by an experimenter. This particular sequence is designed to result
in no change in right ascension, but only in declination. In normal operation a combined
motion in both right ascension and declination will be accomplished with the aid of a
computer. The computer must store the values of the earth’s magnetic field in each
predicted satellite position. Knowing this, the dipole moment of the satellite, and the
desired change of position, the computer will determine at what point in the orbit the
torquing coil is to be turned on and for how long. Time delays of up to an hour are
available in the satellite so that the maneuver can begin at the optimum time in the orbit,
not just over the command station. Time increments for this delayed maneuver are 2.5
minutes, both for the delay and for the maneuvering period. Finer control can be
obtained by turning the torquing command on and off while within sight of the station.
Once the position is achieved, it is held primarily due to the momentum generated by the
rotor, (Figure 13), which is 2.12 X 107 gm-cm2/sec. Drifting from this position is caused
by gravity gradient, aerodynamic, and magnetic torques, the latter resulting from
uncompensated dipoles. The total of these drift rates should not exceed 5 degrees per day
on the average. To guarantee that the residual magnetism of the spacecraft is minimized,
it is provided with a degaussing coil and a chargeable trim magnet system to compensate
in orbit for residual magnetism along each axis. It is possible to use this system to
produce a specific dipole moment for controlled spin axis drift.
The nutation damper is used to dissipate the lateral components of satellite angular
velocity created by the magnetic torquing. It consists (Figure 14) of a torsion wire
suspended arm with an end mass and a copper damping vane. If mutations occur, the arm
oscillates causing the copper vane to swing back and forth through the gap in a
permanent magnet, thus inducing eddy currents in the copper vane. With the rotor on, the
residual nutation should be less than 0.2 degree; with the rotor off, it may be a few
degrees. The motion of the nutation damper, detected optically by a coded digital mask
mounted on the movable arm, is telemetered every 8 seconds via a 7 bit binary word.
The spin rate control system is shown in Figure 15. Signals sensed by the X- and Y-axis
magnetometers are amplified and applied in quadrature to the Y- and X-axis coils. This
system, in conjunction with the earth’s magnetic field, operates essentially as a motor to
increase or decrease the speed of rotation about the Z-axis. The maximum rate of change
of spin rate is 10 rpm per day.
This summarizes the basic SAS control system with which many astronomy experiments
can be performed well. It is ideally suited for sky surveys when used in conjunction with
a good attitude determination system. Data telemetered from the magnetometers and a
digital sun sensor on board the spacecraft can be used to determine attitude to ±5E or
Experiments and Future Development On SAS-A, the principal of Dr. Riccardo
Giacconi of American Science of American Science and Engineering, Inc., is flying star
and sun sensors accurately aligned with the experiment X-ray collimators so that attitude
may be determine to a minute of arc. This experiment will provide us with an accurate
map of X-ray sources throughout the celestial sphere. After such a survey which will
determine source strength, position, and spectral composition, it is obvious that later X-
ray experiments will require more accurate pointing capability for detailed study of
individual sources. The SAS design permits building on this basic control system by
adding a star tracker or star camera system, to provide three axis stabilization to hold the
attitude of the satellite to one minute of arc or better. This would permit useful
experiments in the UV and IR regions of the spectrum as well.
Rough positioning of the spacecraft would be accomplished by the basic control system
as described above, and the exact attitude determined by ground computer. Then, for
example, a program could be transmitted for storage in the spacecraft defining a pattern
that the star camera should see and fine torquing would be used to find and maintain the
proper position Two sets of optics with about 100 fields of view located at 900 to each
other could permit three-axis attitude detection to a resolution of 10 seconds of arc, and
control to I minute of arc. Error signals would control the speed of the basic rotor whose
excess momentum would be dumped into the earth’s magnetic field - again, no
expendables. Small gyros or a worm gear tilting mechanism would be suitable for fine
control of the X- and Y-axes. A system of this general type is planned for SAS-C and
later flights. SAS-B will be a sky survey of Gamma-Ray sources, with Dr. Carl Fichtel of
the Goddard Space Flight Center as principal investigator. The basic control system is
adequate for SAS-B as the Gamma Ray Spark Chamber has a 45E field of view and does
not require very fine pointing.
Summary To summarize the capabilities of the Small Astronomy Satellite: It is a
small, inexpensive, Scout-launched spacecraft specifically designed to meet the
requirements of astronomy missions in a 300 n. mi. circular equatorial orbit. The
spacecraft portion requires about 4.4 cubic feet, weighs 180 pounds and needs 17 watts
average power to perform all the basic spacecraft functions in any random orientation.
Depending on the Scout envelope chosen (Figure 16), the volume available for an
experiment is either 6.4 or 12.5 cubic feet, the available weight is 150 pounds, and power
is 10 watts. If the fine pointing system is desired, these experiment allocations are
decreased by the amount required for it, about 20 pounds, 2 watts, and less than one
cubic foot. One orbit storage is provided for a data rate of 1000 bps, more than 90% of
which is available for experiment data. Although most astronomy experiments are large
and will require the entire spacecraft capability, it is possible to fly several small
compatible experiments simultaneously. SAS can provide for astronomers a much
needed, versatile platform for studying sources radiating in the X-ray, Gamma-Ray,
visible, infra-red, and ultra-violet regions of the spectrum.
Fig. 1 - Artist’s Conception of the Small Astronomy Satellite (SAS)
Fig. 2 - Plot of Radiation Belts Provided by Dr. James Vette, GSFC
Fig. 3 - SAS in Launch Configuration on Scout with Standard Shroud
Fig. 4 - Cross-section of Basic Spacecraft Structure
Fig. 5 - SAS Solar Cell Array Current
Fig. 6 - SAS Power System
Fig. 7 - SAS Command System
Fig. 8 - SAS Command Word
Fig. 9 - SAS Telemetry System
Fig. 10 - SAS PCM Telemetry System Frame Format
Fig. 11 - SAS-A Spin Axis Orientation Control System
Fig. 12 - SAS-A Declination Maneuver Sequence
Fig. 13 - Simplified Drawing of Angular Momentum Flywheel
Fig. 14 - Simplified Drawing of SAS-A Nutation Damper
Fig. 15 - Block Diagram of Spin Control System
Fig. 16 - Outline Drawing of SAS Experiment Package
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