Characteristics of the Soyuz attitude control system

• An infrared vertical sensor is 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.
• The 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 is generated.
• 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.

• 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 ground.
• 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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