Characteristics of the Soyuz attitude control system
The approximate location of the attitude control system sensors and their
boresight axes of early Soyuz spacecraft are shown in the figure on the
right and their functions described below. A detailed
description of the sensor locations is also provided.
system was (in early Soyuz models) equipped with the following actuators:
An infrared vertical sensor
an opto-electronic device, which measures the angular misalignment of the
-YC4 axis and the local vertical.
The sensor uses the Earth's and earth's atmosphere's radiation and converts
it into command signals in pitch and roll.
ionic sensor is an electronic device, which measures the misalignment
of the spacecraft +XC4 axis from
the orbital velocity vector. The sensor uses the incident ion flow on the
spacecraft and converts it into pitch and way command signals. There are
three ionic sensors: Two along the +XC4
axis and one along the -XC4 axis.
The sun sensor (designation
45K) is an opto-electronic instrument, producing a signal corresponding
to the solar attitude misalignment of the spacecraft +YC4
axis. The central zone of the sensor generates a signal indicating that
the sun is within +/- 6 degrees of the sensor line-of-sight axis. The crew
can monitor the attitude by means of the shade gauge. If the sun sensor
fails, solar orientation is performed manually.
The gyro system (designation
KI-38) consists of two gyro units. Each gyro unit has a free two-degree-of-freedom
gyro, caging mechanism, angle sensors, and a programming device. Both gyros
are used at the same time. the outer gimbals of the two gyros are aligned
to the spacecraft roll and yaw axes. The gyro system provides for selected
inertial attitude hold of the spacecraft and performs 0-360 degree maneuvers
about the roll and yaw axes.The maximum allowable angular deviation of
the spacecraft relative to a reference value about any axis is +/- 8 degrees.
If this limit is exceeded an "emergency" (Russian: 'Avaria') signal
The rate gyro unit contains
three rate gyros and puts out signals proportional to the projections of
the spacecraft angular rate vector on the XC4,
YC4, and ZC4
axes. The cosmonaut can switch the unit into an integrating mode where
the rate signals are integrated over time. In this mode the rate gyro unit
can provide inertial or orbital attitude hold and make it possible to make
programmed attitude maneuvers. An "emergency" signal is generated if the
angular deviation of the spacecraft from the reference direction exceeds
+/- 6 degrees.
The sketch on the right shows
both attitude control and translational thruster locations.
a set of hydrogen peroxide thrusters
with 10 N thrust.
a set of hydrogen peroxide thrusters
with 100 N thrust.
The system can operate in
the following modes.
The orbital orientation mode
is perforemd using the ionic and infrared sensors. The infrared sensor
is used to orient the -Y axis to the nadir vector. The earth is found from
an arbitrary by acquiring the earth through a roll search maneuver. Aftre
the earth has been sighted by the infrared vertical sensor, the -Y axis
is guided toward the earth center using the infrared sensor signals. The
orinetation in yaw is provided by the ionic senor which provides anlog
error signals. A rate command in pitch is provided to the rate gyros to
compensate for orbital rotation. The deadbands of the orientation engine
thrusters are 1-2 degrees in angle and 0.07 degrees/second in angular rate.
The accuracy in aligning the -Y axis with the local vertical is +/- 3 degrees
in pitch an roll and +/- 5 degrees in yaw. When the required accuracy in
yaw is achieved the instrument panel electroluminescent light "ORIENTATION"
is lit. The signal from the ionic sensors are then displyed to the pilot.
The inertial attitude hold
mode is performed in two ways. The first uses attitude signals from
the gyro system and rate signals from the rate gyro unit. The second uses
attitude and rate signals bith generated by the rate gyro system. The transfer
to the inertial orientation mode is made from the orbital orientation mode.
At a specified time, either the free gyros are uncaged or the integration
of the rate gyro datat is zeroed. If the spacecraft deviates in attitude
in any axis by more than 6 degrees when using rate gyros or 8 degrees using
the gyro system and emergency signal "Avaria" is generated. This
is also true during program attitude maneuvers. Such maneuvers are performed
with 0.45-0.66 degrees/second depending on which sensor (gyro or integrating
rate gyro) is used for angle input. Program maneuver can be commanded from
The sun orientation mode
uses the sun sensor. Sun acquisition is automatic by maneuvering the spacecraft
about the pitch axis. After the sun appears in the sun sensor field of
view, the spacercraft is maneuvered until the Sun is in the central zone
of the sensor. The attitude control system can maintain this orientation
within +/- 6 degrees. Alternately a rotation can be performed around the
Y axis at a rate of 2.5 degrees/second (rotation time 144 seconds) which
ensures a spin stabilization of the Y axis to the Sun which makes it possible
to have the solar panels recharge the batteries with the attitude control
system switched off.
The manual orientation
mode can either be made with the hand controller movement providing
a rotation rate as long as it is deflected or giving rate increments every
time it is deflected. The rate increments are 0.01-0.03 degrees/second
using 10 N thrusters or 0.1 degrees/second using the 100 N thrusters.
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