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The Apollo spacecraft is guided and controlled by two interrelated subsystems. One is the guidance and navigation subsystem. The other is the stabilization and control subsystem.

    The two subsystems provide rotational, line-of-flight, and rate-of-speed information. They integrate and interpret this information and convert it into commands for the spacecraft's propulsion subsystems.

    The guidance and navigation subsystem contains three major elements. They are the inertial, optical, and computer subsystems.

    The inertial subsystem senses any changes in the velocity and angle of the spacecraft and relays this information to the computer. The computer digests the information and transmits any necessary signals to the spacecraft engines.

    The optical subsystem is used to obtain navigation sightings of celestial bodies and landmarks on the earth and moon. It passes this information along to the computer for guidance and control purposes.

    The computer subsystem uses information from a number of sources to determine the spacecraft's position and speed and, in automatic operation, to give commands for guidance and control. Data fed into the computer include: telemetry information from the ground regarding velocity, attitude, and position in space; a fixed memory which permanently stores navigation tables, trajectory parameters, programs, and constants; an erasable memory which stores intermediate results of computation, auxiliary program information, and variable data supplied by the guidance and control, and other subsystems of the spacecraft

The other of the two interrelated subsystems is the stabilization and control subsystem. In general, it operates in these three ways: it determines the spacecraft's attitude (its angular position); it maintains the spacecraft's attitude; it controls the direction of thrust of the service propulsion engine.

    Both of the subsystems are used by the computer in the command module to provide automatic control of the spacecraft. This is done electronically. Manual control of the spacecraft's attitude and thrust is provided mainly through the stabilization and control subsystem equipment.

    The space trajectory of Apollo is established essentially by the firing of engines, either its own or those of a launch vehicle. "Flying" Apollo, in the sense that an airplane is flown, is not possible because there is no atmosphere in space. The speed, altitude, and flight path angle of Apollo at the instant the last booster engine cuts off determines the characteristics of the spacecraft orbit: its apogee (maximum altitude), perigee (minimum altitude), velocity and position at any point, and the time to complete an orbit.

    Any additional engine firing automatically changes the flight path in space. To maneuver Apollo, the engines must be fired. The direction, amount, and duration of thrust of the service propulsion engine, for example, causes a change in the shape of the orbit about the earth, tilts the plane in which Apollo orbits the earth, or slows it down to permit entry and return to earth. During the flight to the moon, the firing changes the path to the moon (course correction), or slows the spacecraft at the proper time to permit the moon's gravity to pull Apollo into an orbit about the moon.

SPACECRAFT ATTITUDE

    Two indicators on the main display console show the spacecraft's attitude. The indicators show the angle, or angular position, in reference to star sightings. These two indicators are called flight director attitude indicators They tell the spacecraft's total angle (or position), the attitude errors, and rates of change.

    The term total attitude refers to a combination of information from two sources.

    One of the sources is the stable platform of the inertial measurement unit.

    This provides total attitude information by holding the gimbaled, gyro-stabilized stable platform to a fixed inertial reference (maintained by periodic star sightings). In other words, no matter how the spacecraft moves, the stable platform retains its fixed, star-sighted position.

    The second source is a gyro display coupler. This is a device which gives a reading of the spacecraft's actual attitude as compared with an attitude which the crew desires to maintain. A reading of the desired attitude is given by manually dialing roll, pitch, and yaw dials of the attitude set display in the console. The gyro display coupler then shows attitude errors by comparing total attitude information against the manually dialed attitude.

    Information about attitude error also is obtained by comparison of the inertial measurement unit's 9 imbal angles with computer-reference angles. Another source of this information is gyro assembly No. 1,which senses any spacecraft rotation about any of the three axes

    Total attitude information goes to the command module computer as well as to the attitude indicators on the console.

    Attitude control of the spacecraft is provided for the purpose of maintaining a certain angle or for changing it during maneuver. If a specific attitude or orientation is desired, attitude error signals are sent to the reaction jet engine control assembly. Then the proper reaction jet fires in the direction necessary to return the spacecraft to the desired position.

THRUST CONTROL

    The computer in the command module provides primary control of thrust. The flight crew pre-sets thrusting and spacecraft data into the computer by means of the display keyboard. The forthcoming commands include time and duration of thrust. Accelerometers sense the amount of change in velocity obtained by the thrust.

    Thrust direction control is required because of center of gravity shifts caused by depletion of propellants in service propulsion tanks. This control is accomplished through electromechanical actuators which position the service propulsion engine, which rides on gimbals. Automatic control commands may originate in either the guidance and navigation subsystem or the stabilization and control subsystem. There is also provision for manual controls

    The stabilization and control subsystem provides control and monitoring of the spacecraft's attitude, and control of the firing direction or thrust vector of the service propulsion engine. It also serves as a backup system for inertial reference. The subsystem may be operated either automatically or manually. It is produced by Honeywell, Inc., Minneapolis, Minn.

    All the control functions of this subsystem are backups to the guidance and navigation subsystem.

    This subsystem is divided into three basic elements. It provides attitude reference, attitude control,and thrust direction control.

    An electronic control assembly is used for attitude control and thrust direction control. This assembly has the circuits required for utilizing the rate and attitude error sensors. It also has the manual commands necessary to maintain backup stabilization and control in all axes (pitch, yaw, and roll ).

    An electronic display assembly provides the logic for establishing signal sources to be displayed and displays to be used. This assembly also provides for monitoring, isolation, and signal conditioning for telemetry of display signals.

    There are three hand controls. One is for line-of-flight thrusting. Two are rotation controls for angular thrusting. The control for line-of-flight (or rectilinear motion) of the spacecraft also enables the crew to initiate abort during launch. It can also be used to transfer spacecraft control from the guidance and navigation subsystem to the stabilization and control subsystem.

    The rotation control gives the crew manual control of spacecraft rotation in either direction about its three control axes. Also, it may be used for manual thrust direction control in pitch and yaw conditions when the service propulsion subsystem is thrusting.

    The flight data display associated with the stabilization and control subsystem includes the flight director attitude indicators. They show the spacecraft attitude, attitude error, and angular rate information. Also located on the main display console is a gimbal position and fuel pressure indicator which shows either the service propulsion engine pitch and yaw gimbal angles or the booster propel lent pressures. This provides a means for manually trimming the gimbals of the service propulsion subsystem engine.

EQUIPMENT

    Gyro Assemblies (Honeywell, Inc., Minneapolis) Two units, each 6 by 7 by 14-1/4 inches weighing 22 pounds 7 ounces, are on the left-hand side of the navigation station in the lower equipment bay. Each contains three body mounted attitude gyros mounted along body axes to sense attitude displacement. Components of each assembly are contained in a welded aluminum alloy enclosure designed to provide vacuum sealing and electrical continuity. In the attitude- hold mode, any displacement causes one gyro assembly to signal the reaction control system jets to restore the original attitude. The signal is displayed for the crew on the flight director attitude indicator. The other gyro assembly provides signals for display of body rates on either or both flight director attitude indicators, for rate damping, and for computation of inertial attitude changes.

    Rotation Controls (Honeywell) Two identical three- axes rotation controllers with control sticks provide proportional rate command signals for attitude or manual thrust vector maneuvers. Motion of the controls is analogous to rotation of the spacecraft about its axes. Direct angular acceleration command signals are provided when the direct reaction control system switch is actuated and the control is displaced to full stop position. Controls can be mounted on left and right couches and at the navigation position in the lower equipment bay. Each control is in a 3 by 5.5 by 7.26-inch metal box with a 5.25-inch aluminum stick handle. Each weighs 7 pounds, including cable and connector. A trigger-type push-to-talk switch is in the stick handle. There are redundant locking devices on each control. The control unit includes six breakout switches to provide on-off command and signals to the command module computer, stabilization control system minimum impulses, acceleration commands, caging of attitude gyros, and to enable proportional rate commands in the electronics; three transducers to command spacecraft rotation rates during proportional rate control and to command service propulsion engine gimbal position in pitch and yaw during manual thrust vector control; and direct switches in each axis for each direction of rotation. Direct switches produce acceleration commands through the direct solenoids on the reaction control system engines, bypassing stabilization control system electronics.

    Translation Control (Honeywell) This control with a T- handle stick provides a means of accelerating along one or more of the spacecraft axes. Motion of the control is analogous to translation motion of the spacecraft. The metal box measures 3.78 by 6.3 by 3.8 inches with a 3-inch rubber boot and aluminum T- handle. It weighs 5 pounds 11 ounces. The control is mounted with its axis approximately parallel to those of the spacecraft. Redundant switches close for each direction of control displacement. Clockwise switches transfer spacecraft control from the command module computer to the stabilization control system. One of the switches supplies 28 volts dc to a portion of the primary guidance navigation and control system. Redundant counterclockwise switches provide for initiation of manual abort during the launch phase. A dissrete signal from switch closure is fed to the master events sequence controller, which initiates other abort functions.

    Flight Director Attitude Indicator (Honeywell) There are two indicators, each in a metal box 6.9 inches in diameter and 9.33 inches long weighing 9 pounds. They have glass cover plates. One is on Panel 1 in front of the commander and the other on a panel to the right and directly above it, just over the display and keyboard. The display is a ball with circles showing pitch attitude, yaw attitude, and roll attitude, and three scales one above and below the ball and a third to the right of it showing pitch, yaw, and roll rate with needles showing pitch, yaw, and roll error. Attitude and attitude error signals to the displays are supplied by the stabilization and control system or the guidance and navigation system. Rate signals to the displays are supplied by the stabilization and control system only.

    Attitude Set Control Panel (Honeywell) This 3.5 by 4.8 by 4.4-inch panel weighs 3 pounds 6 ounces and is in the lower left-hand corner of the main display panel facing the commander. It has three windows to display roll, pitch, and yaw in degrees and three thumbwheels to the left of the windows. The panel provides the means for manually inserting desired attitude information into the stabilization control system in the form of three angles. It receives signals that represent the actual attitude of the spacecraft relative to an arbitrary inertial (fixed) reference frame. Output signals are provided, which represent the attitude error or the difference between the actual and desired total attitude of the spacecraft. These signals can be used to drive the attitude error needles on a flight director attitude indicator, thereby providing the astronaut with a visual indication of the spacecraft attitude error and to align the gyro display coupler to a fixed reference frame.

    Gimbal Position and Fuel Pressure Indicator (Honeywell) The indicator is in a 4.25 by 4.5 by 4.82-inch metal box with a single window showing four meter movements. The unit weighs 2 pounds 14 ounces and is directly below the flight director attitude indicator in front of the commander. It contains redundant indicators for both the pitch and yaw channels. During boost phases it displays second-stage fuel pressure on the redundant pitch indicators and third-stage fuel pressure on the two yaw indicators. The gimbal position indicator consists of two dual servometric meter movements mounted within a common hermetically sealed case. For a stabilization control system velocity change mode, manual service propulsion system engine gimbal trim capability is provided. Desired gimbal trim angles are set in with the pitch and yaw trim thumbwheels.

DETAILED DESCRIPTION

CONTROLS AND DISPLAYS

    Stabilization and control subsystem controls and displays consist of two rotation controls, a translation control, an attitude set control panel, a gimbal position and fuel pressure indicator, two flight director attitude indicators, and two gyro assemblies.

    Two identical rotation controls give the astronauts control of the spacecraft's rotation in either direction around all three axes. The controls are connected in parallel so that they operate in a redundant fashion without switching. Each axis of control performs three functions:

    1. Breakout Switches - A switch closure occurs whenever the control is moved 1.5 degrees from its rest position. Separate switches are provided in each axis and for each direction of rotation. These six switches are used to provide command signals to the CM computer, stabilization and control minimum impulses, acceleration commands, attitude gyro caging, and proportional rate commands in the electronics.

    2. Transducers - These produce alternating current signals proportional to the rotation control displacement from the null position which are used to command spacecraft rotation rates, and to command service propulsion engine gimbal position in pitch and yaw during manual thrust vector control. All three transducers can be used simultaneously.

    3. Direct Switches - A switch closure occurs whenever the control is moved a nominal 11 degrees from its null position (hardstops limit control movement to +/-11.5 degrees from null in all axes). Separate switches are provided in each axis and for each direction of rotation. These switches are enabled by placing the DlRECT RCS switch to ON. Direct switch closure will produce acceleration commands through the direct solenoids on the RCS engines. Al I SCS electronics are bypassed.

    The rotation control has a tapered female dovetail on each end of the housing which mates with mounting brackets on the couch armrests and at the navigation station in the lower equipment bay. When attached to the armrests, the input axes are approximately parallel with spacecraft body axed The control is spring-loaded to null in all axes. A trigger-type push-to-talk switch also is located in the control grip. Redundant locking devices are provided on each control.

The translation control gives the astronauts a means of accelerating the spacecraft in either direction along any of the three axes The control is mounted with its axes approximately parallel to those of the spacecraft. Redundant switches close for each direction of control displacement. These switches command the CM computer and the reaction jet and engine control. A mechanical lock inhibits the commands.

    The controls T-handle may be rotated in either direction about the centerline of the shaft on which it is mounted. Hardstops for these rotations are +/- 17 degrees from null with detent positions encountered at a nominal +/-12 degrees. In the detent position the hand can be removed and the T-handle will not return to null.

    The clockwise switches will transfer spacecraft control from the CM computer to the stabilization and control subsystem. It may also transfer control between certain submodes within the stabilization and control subsystem. The redundant counterclockwise switches initiate abort during the launch phase. A discrete signal from switch closure is fed to the master events sequence controller which initiates other abort functions.

    The attitude set control panel provides thumb-wheels to position resolvers for each of the three axes. The resolvers are mechanically linked with indicators to provide a readout of the dialed angles. The signals to these attitude set resolvers are from the inertial measurement unit or the gyro display coupler.

The panel counters indicate resolver angles in degrees, and allow continuous rotation from 000 through 359 to 000 without reversing direction. There are graduation marks every 0.2 degree. Pitch and roll are marked continuously between 0 and 359.8 degrees. Yaw is marked continuously from 0 to 90 degrees and from 270 to 359.8 degrees; it is also marked with 0.2-degree graduation marks from 270 to 0 to 90 degrees and is numbered at 180 degrees. Readings increase for an upward rotation of the thumb-wheels. One revolution of the thumbwheel produces a 20-degree change in the resolver angle and a corresponding 20-degree change in the counter reading.

The gimbal position and fuel pressure indicator contains redundant indicators for both the pitch and yaw channels. During the boost phases the indicators display fuel and oxidizer pressures for the Saturn second and third stages. Second stage fuel pressure (or third stage oxidizer pressure depending on the launch vehicle configuration) is on the redundant pitch indicators while third stage fuel pressure is on the two yaw indicators. The gimbal position indicator consists of two dual servometric meter movements mounted within a common hermetically sealed case. Thumbwheels enable crewmen to set service propulsion engine gimbal angler for stabilization and control subsystem velocity change maneuvers. Desired gimbal trim angles are set in with the pitch and yaw trim thumbwheels. The indicator displays service propulsion engine position relative to actuator null and not body axes. The range of the engine pitch and yaw gimbal displays is +/-4.5 degrees. This range is graduated with marks at each 0.5 degree and reference numerals at each 2-degree division. The range of the fuel pressure scale is 0 to 50 psi with graduations at each 5-psi division and reference numerals at each 10- psi division.

    The body rate (roll, yaw, or pitch) displayed on either or both flight director attitude indicators is derived from the body-mounted attitude gyros. positive angular rates are indicated by a downward displacement of the pitch rate needle and by leftward displacement of the yaw and roll rate needles. The angular rate displacements are related to the direction of motion required by the rotation control to reduce the indicated rates to zero. The angular rate scales are marked with graduations t null and full range, and at 1/5, 215, 3/5, and 4/5 of full range. Full-scale deflection ranges are obtained with the FDAI SCALE switch and are 1, , and 10 degrees per second in pitch and yaw, and l, 5, and 50 degrees per second in roll. Servometric meter movements are used for the three rate indicator needles.

    The indicator's attitude error needles show the difference between the actual and desired spacecraft attitude. Positive attitude error is indicated by a downward displacement of the pitch error needle, and by a leftward displacement of the yaw and roll error needles. The attitude error needle displacements also are related to the direction of motion required by the rotation control to reduce the error to zero. The ranges of the error needles are 5 degrees or 50 degrees for full-scale roll error, and 5 degrees or 15 degrees for pitch and yaw error. The error scale factors are selected by a switch that also establishes the rate scales. The pitch and yaw attitude error scales contain graduation marks at null and full scale, and at 1/3 and 2/3 of full scale. The roll attitude scale contains marks at null, 1/2, and full scale. The attitude error indicators also use servometric meter movements.

    Spacecraft orientation with respect to a selected inertial reference frame is also displayed on the attitude indicator ball. This display contains three servo control loops that are used to rotate the ball about three independent axes. These axes correspond to inertial pitch, yaw, and roll. The control loops can accept inputs from either the inertial measurement unit gimbal resolvers or gyro display coupler resolvers.

    Pitch attitude is represented on the ball by great semicircles. The semicircle displayed under the inverted wing symbol is the inertial pitch at the time of readout. The two semicircles that make up a great circle correspond to pitch attitudes of 0 and 0+180 degrees.

    Yaw attitude is represented by minor circles. The display readout is similar to the pitch readout. Yaw attitude circles are restricted to the intervals of 270 to 360 degrees (0) and 0 (360) to 90 degrees.

    Roll attitude is the angle between the wing symbol and the pitch attitude circle. The roll attitude is more accurately displayed on a scale attached to the indicator mounting under a pointer attached to the roll (ball) axis.

    The last digits of the circle markings are omitted. Thus, for example, 3 corresponds to 30, and 33 corresponds to 330. The ball is symmetrically marked about the O- degree yaw and 0/180-degree pitch circles. Marks at 1- degree increments are provided along the entire yaw Degree circle. The pitch 180-degree semicircles have the same marking increments as the 0-degree semicircle. Numerals along the 300- and 60-degree yaw circles are spaced 60-pitch degrees apart. Numerals along the 30- degree yaw circle are spaced 30-pitch degrees apart. Red areas of the ball indicate gimbal lock.

    Each gyro assembly contains three body-mounted attitude gyros mounted so that the input axis of one gyro is parallel to one of the spacecraft body axes. Gyro assembly No. 2 provides signals for display of body rates on either or both attitude indicators; rate damping for stabilization and control subsystem control configurations (excluding acceleration command and minimum impulse), and computation of inertial attitude changes.

    The gyro assembly gyros provide signals equivalent to body attitude errors (deviations). These signals can be used for attitude control or display on the attitude indicator. The gyros also can provide backup rate signals for the functions of gyro assembly No. 2.

    The Block II stabilization and control subsystem uses a switching concept as opposed to the "mode select" switching used in Block I Apollo spacecraft. Functional switching means the manual operation of a number of independent switches to configure the subsystem for various mission functions (e.g., course correction, velocity changes, entry, etc.). Mode switching would, for example, use one switch labeled "entry" to accomplish automatically all the necessary system gain changes, etc., for that mission phase. Mode selection simplifies crew tasks but limits system flexibility. Functional switching offers flexibility in selection of subsystem elements and allows part of a failed signal path to be turned off without affecting the total signal source.

    There are two types of controls selectable from the main display console: stabilization and control subsystem or CM computer. The computer is the primary method of control and the stabilization and control subsystem is the backup. Attitude control is obtained from the reaction control engines and thrust vector control from the service propulsion engine.

ATTITUDE REFERENCE

    The gyro display coupler provides signals to either of the attitude indicators for display of spacecraft total attitude and attitude errors. Angular velocity for display will always be supplied from either of the two gyro assemblies. The spacecraft total attitude display requires a connection between either gyro assembly and the attitude ball. This combination provides a backup attitude reference system for accurate display of spacecraft position relative to a given set of reference axes. Spacecraft attitude errors can be developed using the attitude set control panel connection between the gyro display coupler and the attitude indicator error needles. This combination provides a means of aligning the attitude reference system to a fixed reference while monitoring the alignment process on the error needles; it can also be used in conjunction with manual maneuvering of the spacecraft.

The gyro display coupler can be operated in the following configurations:

    Alignment - Provides a means of aligning the coupler to a given reference.

    Euler - Computes total inertial attitude and body attitude error from body rate signal inputs.

    Non-Euler - Computes digital body rate signals from dc body rate signal inputs

    Entry = 0.05 G - Computes roll axis attitude about the entry roll stability axis from body rate signal inputs.

    The alignment function is used to align the coupler Euler angles (shafts) to the desired inertial reference selected by the attitude-set thumbwheels (resolvers). This is done by interfacing the coupler resolvers with the ASCP resolvers in each axis to generate error signals that are proportional to the difference between the resolver angles. These error signals are fed back to the gyro display coupler to drive the resolver angular difference to zero. During this operation all other functions for the coupler are inhibited.

    In the Euler configuration, the coupler accepts pitch, yaw, and roll dc rate signals from either gyro assembly and transforms them to Euler angles to be displayed on either attitude ball. The coupler Euler angles also are sent to the attitude set control panel to provide an Euler angle error, which is transformed to body angle errors for display on either attitude error indicator.

    Non-Euler pitch, yaw, and roll dc body rate signals from either gyro assembly are converted to digital body rate signals and sent to the CM computer. Power is removed from both attitude baIl-drive circuits when this configuration is selected.

    In the entry < 0.05 G operation, the coupler accepts yaw and roll dc body rate signals from either gyro assembly and computes roll attitude with respect to the stability axis to drive the roll stability indicator on the entry monitor system or from gyro assembly No. 1 and computes roll attitude with respect to the stability axis to drive either attitude ball in roll only.

    The purpose of displaying total attitude, attitude error, and rate is to monitor the spacecraft and system functions. Since a single source can supply more than one type of information, the mission function required at a particular time will normally dictate what type of information is required from the source.

    The flight director attitude indicator may be modified by an orbital rate display-earth and lunar unit. This unit is inserted electrically in the pitch channel between the electronic display assembly and attitude indicator to provide a local vertical display in the pitch axis. Local vertical is attitude with respect to the body (earth or moon) the spacecraft is orbiting. Controls on the unit permit selection of earth or lunar orbits and orbital attitude adjustment.

ATTITUDE CONTROL

    The attitude and translation control portion of the stabilization and control subsystem uses the reaction control subsystem. The reaction control subsystem and stabilization and control equipment are described only in relation to attitude control. Stabilization and control equipment used for attitude control includes:

    1. Gyro assembly No. 1, where three body mounted attitude gyros provide pitch, yaw, and roll attitude error signals for use when automatic attitude hold is desired.

    2. Gyro assembly No. 2, whose three body mounted attitude gyros provide pitch, yaw, and roll rate damping for automatic control and proportional rate maneuvering.

    3. The rotational controllers which enable crewmen to control the spacecraft attitude simultaneously in three axes.

    4. The translation controller which enables the crew to command simultaneous accelerations along all three spacecraft axes and is also used to initiate several transfer commands.

    5. The electronics control assembly which contains the electronics used for automatic, proportional rate, and minimum impulse capabilities.

    6. The reaction jet engine control which contains the automatic reaction control subsystem logic and the solenoid drivers that provide commands to the automatic coils of the reaction control engines.

    The reaction control subsystem provides the rotation control torques and translational thrusts for all attitude control functions. The reaction control engine is operated by applying excitation to a pair (fuel and oxidizer) of solenoid coils Each engine has two pairs of solenoid coils, one automatic and the other direct.

    Commands to the reaction control engines are initiated by switching a ground through the solenoid driver to the low side of the automatic coils. The solenoid drivers receive commands from the automatic reaction control subsystem logic circuitry contained in the reaction jet engine control. The automatic reaction control subsystem logic activates the command source selected and commands the solenoid drivers necessary to perform the desired attitude control function. The logic receives reaction control commands from the CM computer (for rotation and translation), the electronics control assembly (rotation for either automatic, proportional rate, or minimum impulse control), the rotation controls (for continuous rotational acceleration), and the translation control (for translational acceleration).

    Commands to the direct coils have priority over those to the automatic coils. The direct coils receive commands from the rotation control when the "Direct RCS" switch is actuated and the control is deflected 11 degrees about one or more of its axes. The direct coils are used to command an ullage maneuver before a service propulsion engine firing of normal ullage methods are not available; this is controlled through a pushbutton on the main display console. The master events sequence controller also can command an ullage maneuver to enable separation of the CSM from the third stage. Direct coils on the SM reaction control engines are activated by the SM jettison controller for SM-CM separation. CM direct coils are activated by a switch on the main display console for reaction control propellant dumping during the final descent on the main parachutes.

    The stabilization and control subsystem can be placed in various configurations for attitude control, depending on crew selection of control panel switches. The configuration desired is selected independently for each axis. Both automatic and manual control can be selected.

    Automatic control involves rate damping and attitude hold. In rate damping large spacecraft rates are reduced to a small range (rate deadband) and held within the range. In attitude hold, angular deviations about the body axes are kept within certain limits (attitude headband). If attitude hold is selected in pitch, yaw, and roll, the control can be defined as maintaining a fixed inertial reference. Rate damping is used in the mechanization of the attitude hold configuration.

    Attitude hold uses the control signals provided by the body-mounted attitude gyros. These signals are summed in the electronics control assembly switching amplifier. The switching amplifier has two output terminals that provide commands to the automatic reaction control subsystem logic: one terminal provides positive rotation commands and the other negative commands. If the magnitude of the signal input is smaller than a specific value, neither output is obtained. The input level required to obtain an output is referred to as the switching amplifier deadband. This can be interpreted as a rate deadband or an attitude (minimum) deadband. The deadband limits are a function of the control loop gains which depend on the position of the " Rate" switch: low is +/-0.2 degree per second and high is +/- 2 degrees per second. An additional deadband can be selected for attitude control: +/-0.2 to 4.2 degrees in low and +/-4 to 8 degrees in high.

    When the summation of rate and attitude signals exceeds the switching amplifier deadband, a rotation command is sent to the reaction control subsystem and the engines are automatically fired for the duration needed to correct the deviation.

    Manual attitude control in solves proportional rate, minimum-impulse, acceleration command and direct control. These commands are initiated by operation of either rotation control. With the exception of direct, the rotation control commands go through the reaction control subsystem automatic coils.

    Proportional rate provides the ability to command a spacecraft rate that is directly related to the amount of rotation control stick deflection. This capability is obtained by summing the control's transducer output with the body-mounted attitude gyro signal in the electronics control assembly; when the stick is deflected, an error is developed at the switching amplifier input that results in an acceleration command. The command is present until the gyro signal is large enough to reduce the error to less than the deadband. The spacecraft will then coast at a constant rate until the rotation control input is removed.

    Minimum impulse provides the ability to make small changes in the spacecraft rate. When minimum impulse is enabled in an axis, the output of the switching amplifier in that axis is inhibited. Thus, the spacecraft (attitude) is in free drift in the axis where minimum impulse is enabled if direct control is not being used. Minimum impulse is commanded by the rotation control breakout switch. When minimum impulse is selected in the roll axis, one-half of the roll solenoid drivers are inhibited for minimum impulse commands. A roll minimum impulse command is executed by two reaction control engines unless one of the "Channel Roll" switches is turned off, which reduces the command to a single engine.

    When acceleration command is activated and a breakout switch is closed, continuous commands are sent to the appropriate reaction control subsystem automatic coils. The selection of acceleration command in an axis inhibits all other inputs to the automatic reaction control subsystem logic for that axis. This differs from minimum impulse selection in that translation control commands are available during minimum impulse control.

    Direct rotation control is available as backup to any other control, including CM computer when the "Direct RCS" switch is on. When the rotation control stick is deflected to hard stop, a direct switch is closed and the voltage is routed to the direct coils on the appropriate reaction control engines. The control's direct switch also routes a signal to the automatic reaction control subsystem logic that inhibits all automatic coil commands in the axis under direct control.

    Commands from the translation control can be initiated simultaneously in the three axes and appear as logic inputs to the automatic reaction control subsystem logic. The logic signals are obtained from switch closures in the control. Translation control is not available after CM/SM separation.

THRUST VECTOR CONTROL

    Spacecraft attitude is controlled during a velocity change by positioning the engine gimbals for pitch and yaw control while maintaining roll attitude with the attitude control subsystem. The stabilization and control electronics can be configured to accept attitude sensor signals for automatic control or rotation control signals for manual control. Manual thrust vector control can be selected to use vehicle rate feedback signals summed with the manual signals. A different configuration can be selected for each axis; for example, one axis can be controlled manually while the other is controlled automatically.

    In automatic thrust vector control, spacecraft angular rates and attitude errors are sensed by the body-mounted attitude gyros. Attitude error, gimbal position, and gimbal trim signals are summed at the input to an integrator amplifier. The integrator output is then summed with rate, attitude error, gimbal position, and gimbal rate at the servo amplifier input.

Steady-state operation is obtained when the gimbal is positioned so that the thrust vector is aligned through the vehicle center of gravity and the error at both summing points is a constant-zero. The integrator input error is zero when gimbal position minus gimbal trim is equal to the negative of the attitude excursion sensed by gyro assembly No. 1. This is the spacecraft/ gimbal orientation necessary to obtain and maintain the desired thrust direction. Transients due to center-of-gravity uncertainty errors or shifts during thrusting are forced by the integrator to have the necessary steady-state solution. However, final pointing vector errors will be incurred because of the quadrature accelerations induced during the transient phases Errors also will result from amplifier gain and component inaccuracies.

The gimbals are trimmed before thrust by turning the trim wheels on the gimbal position indicator. The trim wheel in each axis is mechanically connected to two potentiometers connected with the gimbal servomechanisms. It is desirable to trim before velocity change to minimize the transient duration time and the accompanying quadrature acceleration. The trim wheels also are set before a velocity change controlled by the CM computer so that the stabilization and control subsystem can relocate the desired thrust direction if a transfer is required after engine ignition.

    In manual thrust vector control, the signal from the rotation control is sent to a proportional plus integral amplifier. This circuit maintains a gimbal deflection after the rotation control is returned to rest and makes corrections with the control about its rest position, rather than holding a large displacement. Depending on how switches are set, it also can damp out spacecraft rate.

    There are two manual thrust vector control configurations: rate command and acceleration command. Rate command is similar to the proportional rate control in the attitude control sum system except there is no deadband. The thrust vector is under body-mounted attitude gyro control. If there is an initial gimbal center-of-gravity misalignment, an angular acceleration will develop. The gyros, through the proportional gain, will drive the gimbal in the direction necessary to cancel this acceleration. The rate feedback is inhibited in acceleration command; the rotation control input must be properly trimmed to position the thrust vector through the center of gravity. This drives the rotational acceleration to zero but additional adjustments are necessary to cancel residual rates and obtain the desired attitude and positioning vector.

ENTRY MONITOR SYSTEM

    The entry monitor system provides a visual display of automatic guidance navigation and control system entry and velocity change maneuvers. It also provides sufficient display data to permit manual entry in case of guidance and control malfunctions and automatic velocity cutoff commands for the stabilization and control subsystem when controlling the service propulsion engine. The velocity display also can be used to cut off thrust manually if the automatic commands malfunction.

    The system provides five displays used to monitor an automatic entry or perform a manual entry: threshold indicator, roll attitude indicator, corridor verification indicators, range displays, and the flight monitor.

    The threshold indicator, labeled .05G, displays sensed deceleration. The altitude at which this indicator is illuminated depends on entry angle (velocity vector with respect to local horizontal), the magnitude of the velocity vector geographic location and heading, and atmospheric conditions. Bias comparator circuits and timers are used to activate this indicator. The signal used to illuminate the indicator is also used in the system to start the corridor evaluation timer, scroll velocity drive, and range-to-go circuits.

    The roll attitude indicator displays the lift vector position throughout entry. During entry, stability axis roll attitude is supplied to the indicator by the gyro display coupler. There are no degree markings on the display, but the equivalent readout will be zero when the indicator points toward the top of the control panel and increases up to 360 in a counterclockwise direction.

    The corridor verification indicators are located on the roll attitude indicator. They consist of two lights which indicate the necessity for lift vector up or down for a controlled entry. (The indicators are valid only for spacecraft entering at velocities and angles that will be used on the return from the moon.) The corridor comparison test is performed approximately 10 seconds after the .05G indicator is illuminated. The lift vector up light (top) indicates greater than approximately 0.2G. The lift vector down light (bottom) indicates less than approximately 0.2G. An entry angle is the angle displacement of the CM velocity vector with respect to local horizontal at 0.05G. The magni ude of the entry angles that determines the capture and undershoot boundaries depend on the CM lift-to- drag ratio. Entry angle less than the capture boundary will result in noncapture regardless of lift orientation. Noncapture would result in an elliptical orbit which will re-enter when perigee is again approached. The critical nature of this would depend on CM consumables: power, control propellant, lift support, etc. The CM and crew would undergo excessive G force (greater than 10G's) with an entry angle greater than the undershoot boundary, regardless of lift orientation.

    The range display is an electronic readout of inertial flight path distance in nautical miles to predicted splashdown after 0.05G. The predicted range will be obtained from the guidance and control subsystems or ground stations and inserted into the range display before entry. The range display also shows velocity in feet per second during service propulsion engine thrusting.

    The flight monitor provides an entry trace of total G level versus inertial velocity. A Mylar scroll has printed guide lines which provide monitor (or control) information during aerodynamic entry. The entry trace is generated by driving a scribe in a vertical direction as a function of G level, while the Mylar scroll is driven from right to left proportional to the CM inertial velocity change. Monitor and control information for safe entry and range potential can be observed by comparing the slope of entry trace to the slope of the nearest guide lines.

    In addition to entry functions, the entry monitor system provides outputs related to delta velocity maneuvers during either service propulsion or reaction control engine thrusting. Displays include a lamp which lights any time the service propulsion engine fires and a counter which shows the velocity remaining to be gained or lost. The latter display can have a range of 14,000 to -1000 feet per second in tenths of a foot per second. The desired velocity change for all service propulsion engine thrusting maneuvers is set in the panel and the display will count up or down. During thrust controlled by the stabilization and control subsystem, the entry monitor system automatically turns off the service propulsion engine when the display reads minus values.

    The Mylar scroll in the entry monitor system flight monitor has ground and flight test patterns together with four entry patterns. Each entry pattern is preceded by two identical flight test patterns and entry instructions that are used to verify operation of the system's entry circuits and set initial conditions. Each entry pattern contains velocity increments from 37,000 to 4,000 feet per second as well as entry guidelines. The entry guidelines are called G on-set or G off-set and range potential lined During entry the scribe trace should not become parallel to either the nearest G on-set or G off- set lines. If the scope of the entry trace becomes more negative than the nearest G on-set line, the CM should be oriented so that a positive lift vector orientation (lift vector up) exists to prevent excessive G buildup. If the entry trace slope becomes more positive than the nearest G off-set line, the CM should be oriented to produce negative lift (lift vector down). The G on-set and off-set lines are designed to allow a 2-second crew response time and a 180-degree roll maneuver if the entry trace becomes parallel to the target of the nearest guideline.

    The range potential lines, shown in hundreds of nautical miles, are used by the crew during entry. They indicate the ranging potential of the CM at the present G level. The crew will compare the range displayed by the range so-go counter with the range potential indicated at the position of the entry trace. The slope and position of the entry trace relative to a desired ranging line indicates the need for lift vector up or down.

    The vertical line on the scroll corresponds to where the CM velocity becomes suborbital; that is, where the velocity has been reduced to less than that required to maintain orbit. The full positive lift profile line represents the steady-state minimum-G entry profile for an entry.

The guidance and navigation subsystem gives astronauts the ability to navigate the spacecraft on a required course through space. It can be operated either semi-automatically or manually and performs the basic functions of guidance and navigation -similar to the navigation of an airplane or of a ship at sea. It can also be updated by the ground via telemetry, while an airplane or ship at sea is concerned with two-dimensional navigation (it is always on or near the surface of the earth), Apollo is faced with exacting three-dimensional navigation as it speeds through deep space. Sightings from the spacecraft of stars and pre-determined landmarks on the earth and moon are used to establish the location and position of the spacecraft in space. The guidance and navigation system is used in conjunction with the stabilization and control, service propulsion, reaction control, electrical power, environmental control, and telecommunications subsystems.

Massachusetts Institute of Technology, an associate contractor to NASA's Manned Spacecraft Center, is responsible for development and design of the subsystem. AC Electronics Division of General Motors is responsible for its production, operation, and integration. The guidance computer is manufactured by the Raytheon Co, and the optics are produced by Kollsman Instrument Co.

There are three main elements in the guidance and navigation subsystem. The inertial guidance subsystem measures changes in the spacecraft position and velocity and assists in generating steering commands. The optical subsystem is used to take precision navigational sightings and provide the computer with measured angles between the stars and landmarks. The computer subsystem consists of a digital computer which takes data from the inertial guidance and optical subsystems and calculates spacecraft position, velocity, and steering commands.

The components of the guidance and navigation subsystem, including the primary controls and displays, are located in the lower equipment bay of the command module. The main display console contains the switches and displays necessary for the astronauts to control the spacecraft while in their couches,

This subsystem measures changes in the spacecraft position, assists in generating steering commands, and measures spacecraft velocity changes. Its instruments sense changes in velocity and attitude in a manner similar to the balance system in the human ear.

The main part of this subsystem is an inertial measurement unit mounted on a navigation base. The key part of this mechanism is a device called a stable platform, suspended on gimbals which allow it to incline freely regardless of spacecraft position. It is aligned to star direction references and retains this alignment regardless of the rotational movement of the spacecraft and thus provides a reference against which the movements of the spacecraft can be measured.

Mounted on this stable platform are the actual sensors: three accelerometers and three gyroscopes. The gyros are used to keep the stable platform fixed with respect to some point in space. The accelerometers sense any change in the speed of the vehicle -- forward, backward, up, down, sideways.

Any change from the pre-determined flight path or attitude generates electronic signals which result in firing of the reaction control subsystem engines (for attitude change only) or positioning of the service propulsion engine for flight path changes. The service propulsion engine is fired on signal from the computer or astronauts.

Devices known as resolvers, mounted on the gimbal axes, measure how far the spacecraft has rotated with respect to the stable platform. These measurements are transmitted to the guidance computer. While the gyros are used to maintain the spacecraft in a required attitude, the resolvers are used primarily to orient the spacecraft when firing the service propulsion engine.

The inertial guidance subsystem is controlled automatically by the guidance computer through crew selection of a computer program.

This subsystem is used by the astronauts to take navigational sightings of the stars and earth or moon landmarks, It consists of a navigation base, a telescope, a sextant, and equipment to permit operation with the computer and inertial guidance subsystems,

The telescope and sextant can be operated independently but generally are used together to obtain precision navigational sightings. The telescope has a 60-degree field of view with no magnification and is used to obtain coarse sightings of the stars or landmarks. Because of the telescope's limited field of view, controls are provided so the astronaut can maneuver the entire spacecraft to point the instrument in the general desired direction. The sextant is used to take the precision sightings and has a much smaller field of view (1.8 degrees) but has a magnification of 28.

These two instruments are used to measure the angle between two targets such as stars and earth or moon landmarks. The telescope and then the sextant are manipulated by the astronaut to line up a sighting and enter the reading into the computer.

This subsystem consists of a digital computer which stores and uses signals from the inertial guidance subsystem and sightings by the optical subsystem with other data. This highly sophisticated computer uses this data to calculate necessary corrections to maintain course. The computer memory contains 38,912 words, or pieces of guidance data, and is divided into non-erasable and erasable sections. The non-erasable memory contains all the basic data necessary to achieve the round trip to the moon. The erasable memory is used by the astronauts when performing the various guidance and navigation computations.

The major functions of the computer are to calculate spacecraft position, velocity, and steering data; to calculate the signals for the main engine and attitude control jets necessary to keep the spacecraft on the required flight path and attitude; to position the stable platform in the inertial measurement unit as defined by precise optical measurements; to position the optical unit to celestial objects; to monitor the guidance and navigation subsystems for failure indications; and to supply information to the display and control panel.

The display of information, the results of computations, and the control of the computer by the astronauts is accomplished with two display and keyboard panels. One is located in the lower equipment bay and the other (a duplicate) is on the main display console. Lights on the panels also indicate the detection of malfunctions in guidance and navigation systems.

Operation of the guidance and navigation subsystem in specific flight phases, control modes, and critical maneuvers during a typical lunar mission is described here briefly.

To conserve electrical power and fuel, the subsystem is activated only about 20 percent of the time, for specific sightings, alignments, and engine-firing maneuvers. Each time the guidance and navigation subsystem is activated the stable platform must be aligned with respect to a predetermined reference. Before launch the platform is aligned with respect to the earth and during flight it is aligned to the stars.

Launch and Translunar Injection - Guidance of the launch vehicle is monitored during the ascent from earth and firing of the third stage which sends the spacecraft on a trajectory to the moon. At launch, the inertial measurement unit is switched from an earth reference to a space reference frame. The crew then uses the guidance and navigation subsystem to monitor the spacecraft flight profile.

Earth and Lunar Orbit - During these phases, the crew checks the spacecraft position and orbital path with optical sightings. The crew takes these sightings by identifying and tracking landmarks with the optical instruments. The computer records optical sighting data, spacecraft attitude, and the time of the optical sighting.

Midcourse Navigation - This may be performed several times during the translunar and transearth mission phases. Starlandmark sightings are taken and the computer records the angles and time of sighting and determines spacecraft position and velocity.

Course Correction - The spacecraft's course must be corrected during both the translunar and transearth journeys. The computer, after calculating spacecraft position and velocity from navigation sightings, determines the need for a course correction by comparing the actual to the required trajectory. If a course correction is necessary, the computer calculates the time of firing and velocity change needed, repositions the spacecraft, and controls the initiation and duration of thrusting of the service propulsion engine.

(P-257) Guidance and navigation equipment in test fixture simulating spacecraft installation.

Lunar Injection and Return to Earth -The guidance and navigation subsystem places the spacecraft into the attitude required for firing the service propulsion engine and controls the time of thrusting required.

Entry - The guidance and navigation subsystem controls the flight path of the command module during entry. The computer determines the proper trajectory and steers the command module by rolling it. This changes the lifting force acting on the command module and thereby varies its trajectory,

The guidance and navigation system occupies a space 4 feet high, 2 feet deep, and about 3 feet across the top and 2-1/2 feet across the bottom. It is in the command module lower equipment bay.

Inertial Measuring Unit (AC Electronics Division of General Motors) - This ball-shaped unit has a diameter of 12.6 inches and weighs 42.5 pounds. It consists of three gimbals of which the inner gimbal is the stable member, with three gyroscopes and three accelerometers, all can-shaped and mounted onto the stabilized inner member structure, or platform. The gimbals are connected to each other by drive motors and angle resolvers. The unit is pressurized in dry air for good heat transfer. When in operation, the unit requires 217 watts at 28 volts dc. It maintains an inertially referenced coordinate system for spacecraft attitude control and measurement and maintains three accelerometers in this coordinate system for accurate measurement of spacecraft velocity changes.

Navigation Base (AC Electronics) - This 27-by-22-by-4.5-inch unit weighs 17.4 pounds and is made of riveted and bonded anodized preformed sheet aluminum alloys. It is filled with polyurethane foam. It is a rigid supporting structure for the inertial measuring unit and optical equipment.

Power and Servo Assembly (AC Electronics) - It is 2.75 inches high, 23.1 inches wide, and 22.6 inches deep and weighs 49.4 pounds. It contains 37 modules, most of the electronic modules for the inertial and optical subsystem servo loops and power supplies.

Coupling Data Unit (AC Electronics) - It is 20 by 11.3 by 5.5 inches and weighs 36.5 pounds. This sealed unit contains modular packaged solid-state electronics necessary to provide five separate coupling data unit channels for use with inner, middle, and outer inertial measuring unit gimbal resolvers and the shaft and trunnion resolvers of the optical subsystem. It also contains failure detection circuitry for inertial subsystem and optical subsystem. It provides analog-to-digital conversion of the inertial measuring unit gimbal angles and optical subsystem trunnion and shaft angles, and digital-to-analog conversion of computer-derived data to control inertial subsystem and optical subsystem modes of operation using discretes issued by the computer. It also controls service propulsion system engine gimbaling, attitude error display, and, as a backup mode, it controls the attitude of the launch's third stage.

Inertial Reference Integrating Gyros (AC Electronics) - Three gyros are mounted mutually perpendicular to each other on the stable platform of the inertial measuring unit. Each is 2-1'2 inches in diameter. They sense displacement of the inertial measuring unit's stable platform and generate error signals. They are pressurized in an atmosphere of helium to provide good heat transfer.

Accelerometers (AC Electronics) - There are three pulse-integrating pendulous accelerometers mounted perpendicular to each other on the stable member of the inertial measuring unit. Each is can-shaped with a 1.6-inch diameter. They measure velocity changes along all axes of the three-axis inertial measuring unit.

(P-258) This tiny pendulum is heart of an accelerometer, a key component in guidance and navigation subsystem.

Sextant (Kollsman Instrument Co., Syosset, N.Y.) - This is a highly accurate dual line-of-sight electro-optical instrument with 28X magnification and 1.8 degrees field of view. It can sight two celestial targets simultaneously and measure the angle between them with 10 arc seconds accuracy to determine the position of the spacecraft. It is mounted on the navigation base. One line of sight is fixed along the shaft axis normal to the local conical surface of the spacecraft. It is positioned by changes in the spacecraft attitude. The other line of sight has two degrees of rotational freedom about the shaft axis (plus or minus 270 degrees) and trunnion axis (minus 5 to plus 50 degrees). The variation about the trunnion axis is represented by movement of an indexing mirror.

Scanning Telescope (Kollsman) - It is a single line-of-sight, refracting-type, IX magnification instrument with 60-degree instantaneous field of view. It is similar to theodolite (surveyor's instrument used to measure horizontal and vertical angles). It has a double-dove prism mounted in the head assembly. The operating power for the telescope, sextant, and associated electronic equipment is 94.5 watts at 28 volts dc. The telescope has two axes of rotational freedom, which are normally slaved to the sextant axis. The wide field of view is used for general celestial viewing and recognition of target bodies. It is also used to track landmark points during earth and lunar orbits.

Computer (Raytheon Co.) - This is 24 by 12.5 by 6 inches and weighs 70.1 pounds with six memory modules. It consumes 70 watts of power at 28 volts dc during normal operation. It is a digital computer with fixed and erasable memory. The erasable memory has a capacity of 2048 words; fixed memory has a capacity of 36,864 words. The fixed memory contains programs, routines, constants, star and landmark coordinates, and other pertinent data. The computer solves guidance and navigation problems, provides control information to optical and inertial subsystems as well as other spacecraft systems, provides pertinent information to astronauts and the ground on request, provides means by which astronauts or ground control can directly communicate with the primary guidance navigation control system, provides direct on-off control for reaction control jets and service propulsion engines, and monitors its own operation and other primary guidance navigation control system operations.

Display Keyboard (Raytheon Co.) - This 8-by-8-by-7-inch panel weighs 17.5 pounds. It is made up of a keyboard, power supply, a decoder relay matrix, status and caution circuits, and displays. It is a 21-digit character display and a 16-button keyboard through which crewmen can communicate in a coded numerical language, Crewman inserts data and commands the computer by punching numbers on the keyboard. They are then displayed to him in electroluminescent counter-type readout windows. The computer communicates with the crewman by displaying numbers in the same windows. When the computer requests the crewman to take some action, numbers flash to attract attention.

Signal Conditioner Assembly - This 3-by 5.7-by-14.3-inch unit weighs 5.8 pounds. It contains encapsulated electronic circuitry to condition primary guidance navigation control system signals so that they are acceptable to the spacecraft telemetry system.

The inertial subsystem provides a space stabilized inertial reference from which velocity and attitude changes can be sensed. It is composed of the inertial measurement unit, the navigation base, parts of the power and servo assembly, parts of the control and display panels, and parts of the coupling data unit. The navigation base is the rigid supporting structure on which the inertial measurement unit and optical instruments are mounted. It is manufactured and installed to close tolerances to provide accurate alignment of the equipment mounted on it. It also provides shock mounting for the inertial measurement unit and optics.

The inertial measurement unit is the main unit of the inertial subsystem. It is a three-degree-of-freedom stabilized platform assembly, containing three inertial reference integrating gyros, and three pulsed integrating pendulous accelerometers. The stable member is machined from a solid block of beryllium with holes bored for mounting the accelerometers and gyros.

The stable platform attitude is maintained by the gyros, stabilization loop electronics, and gimbal torque motors. Any angular displacement of the stable platform is sensed by the gyros which generate error signals. These signals are resolved and amplified at the inertial measurement unit and applied to stabilization loop electronics. The resultant signal is conditioned and applied to the gimbal torque motors, which restores the desired attitude.

The stable platform provides a space-referenced mount for the three accelerometers, which sense velocity changes. The accelerometers are mounted orthogonally (each at right angles to the other two) to sense the velocity changes along all three axes. Any translational force experienced by the spacecraft causes an acceleration or deceleration which is sensed by one or more accelerometers. Each generates an output signal proportional to the magnitude and direction of velocity change. This signal, in the form of a pulse train, is sent to the computer which uses it to update the velocity information.

Temperature is controlled by a thermostatic system that maintains the gyro and accelerometer temperatures within their required limits during inertial measurement unit standby and operating modes. Heat is applied by end-mount heaters on the inertial components, stable member heaters, and a temperature control anticipatory heater. Heat is removed by convection, conduction, and radiation. The natural convection used during inertial measurement unit standby modes is changed to blower controlled, forced convection during operating modes. Inertial measurement unit internal pressure is normally between 3.5 and 15 psia, enabling the required forced convection. To aid in removing heat, a water-glycol solution passes through coolant passages in the outer case of the inertial measurement unit. Therefore, heat flow is from the stable member to the case and coolant. The temperature control system consists of the temperature control circuit, the blower control circuit, and the temperature alarm circuit.

The coupling data unit is an all-electronic device used to convert and transfer signals among the inertial, computer, and optical subsystems and between the computer and various controls and displays. Each of the five channels has two sections: one acting as an analog-to-digital converter and the other as a digital-to-analog converter. These five channels are: one each for the inner, middle, and outer inertial measurement unit gimbals, and one each for the shaft and trunnion optical axes. Each channel converts inertial or optical gimbal angles from a resolver analog form to a digital form, supplies the computer with this information, and converts digital signals from the computer to either 800-cycle per second or direct current analog signals. The coupling data unit also controls the modes of the inertial and optical subsystems through logical manipulations with the computer.

The power and servo assembly provides a central collection point for most of the guidance and navigation subsystem power supplies, amplifiers, and other electronic components. It is located in the lower equipment bay directly beneath the inertial measurement unit. It consists of 42 modules mounted to a header assembly. Connectors and harnessing are integral to the construction of the header assembly, and guidance and navigation harness branches are brought out from the power servo assembly header. A thin cover plate is mounted on the assembly to provide a hermetic seal for the interior. During flight this permits pressurization of the assembly to remain at 15 psi. Connectors are available for measuring signals at various system test points.

The optical subsystem is used to take precise optical sightings of celestial bodies and landmarks. These sightings are used to align the inertial measurement unit and to determine the position of the spacecraft. The system includes the navigational base, two of the five coupling and data units, parts of the power and servo assembly, controls and displays, and the optics, which include the scanning telescope and the sextant.

The optics consist of the telescope and the sextant mounted in two protruding tubular sections of the optical base assembly. The scanning telescope and sextant line of sight may be offset depending on the mode of operation.

The sextant is a highly accurate optical instrument capable of measuring the included angle between two targets and the direction of any single target with respect to the navigation base. Angular sighting between two targets is made through a fixed beam splitter and a movable mirror located in the sextant head. The sextant lens provides 1.8-degree true field-of-view with 28X magnification. The movable mirror is capable of sighting a target to 57 degrees line of sight from the shaft axis. This is reduced to approximately 45 degrees when installed in the spacecraft, however, because of interference from structure. The mechanical accuracy of the trunnion axis is twice that of the line-of-sight requirement due to mirror reflection which doubles any angular displacement in trunnion axis.

The scanning telescope is similar to a theodolite in its ability to measure elevation and azimuth angles of a single target accurately using an established reference. The lenses provide 60-degree true field of view at 1 X magnification. This is reduced to about 40 degrees when installed because of obstruction by vehicle structure.

The coupling data unit used in the inertial subsystem is also used as a part of the optical subsystem. Two channels of the unit are used, one for the sextant shaft axis and one for the sextant trunnion axis. These channels repeat the sextant shaft and trunnion angles and transmit angular change information to the computer in digital form. The angular data transmission in the trunnion channel is mechanized to generate 1 pulse to the computer for 5 arc-seconds of movement of the sextant trunnion which is equivalent to 10 arcseconds of star line-of-sight movement. The shaft channel issues 1 pulse for each 40 arc-seconds of shaft movement. The location of the sextant shaft and trunnion axes are transmitted to the coupling data units through 16X and 64X resolvers, located on the sextant shaft and trunnion axes. This angular information is transmitted to the units in the form of electrical signals proportional to the sine and cosine of 16X shaft angle and 64X trunnion angle. During the computer mode of operation, the unit provides digital-to-analog conversion of the computer output to generate an ac input to the sextant shaft and trunnion servos. This analog input to the sextant axes will drive the star line of sight to some desired position. In addition, the optical subsystem channels of the coupling data unit perform a second function on a time-sharing basis. During a thrust vector control function these channels provide digital-to-analog conversion of the service propulsion engine gimbal angle command between the computer and the service propulsion subsystem gimbals.

The modes of operation for the optical subsystem are selected by the astronaut using the controls located on the indicator control panel. There are three major modes: zero optics, manual control, and computer control.

During the zero optics mode, the shaft and trunnion axis of the sextant are driven to their zero positions by taking the output of the transmitting resolvers OX and 64X in trunnion, and 1/2X and 16X in shaft) and feeding them through the twospeed switches to the motor drive amplifier. This in turn drives loops to null positions as indicated by zero output from the resolvers. The telescope and trunnion axes follow to a zero position. The zero optics mode can be selected either manually by the flight crew or by the computer after the computer control mode of operation has been manually selected. The manual mode can be selected to operate under either direct hand control or resolved hand control. Independent control of the telescope trunnion is possible in both of these variations.

In direct hand control the hand controller outputs are applied directly to the sextant shaft and trunnion motor drive amplifiers. Forward and back motion of the hand controller commands increasing and decreasing trunnion angles, and right and left motion of the hand controller commands increasing and decreasing shaft angles. The apparent speed of the image motion can be regulated by the flight crew by selecting either low, medium, or high controller speed on the indicator control panel. This regulates the voltage applied to the motor drive amplifier and therefore the shaft and trunnion drive rates.

The slave telescope modes provide for alternate operation of the telescope trunnion while the sextant is being operated manually. Three alternate modes can be selected by a telescope trunnion switch: the telescope trunnion axis slaved to the sextant trunnion (the normal operating position), the telescope trunnion locked in a zero position (by applying fixed voltage to the telescope trunnion 1 X receiving resolver, causing this position loop to null in a zero orientation. Therefore, this means the centerline of the telescope 60-degree field-of-view is held parallel to the landmark line of sight of the sextant), and the telescope trunnion axis is offset 25 degrees (the centerline of the 60-degree field of view is offset 25 degrees from the landmark line of sight of the sextant). This last position of the telescope trunnion will allow the landmark to remain in the 60-degree field of view while still providing a total possible field of view of 110 degrees if the telescope shaft is swept through 360 degrees.

In manual resolved operation, the hand controller outputs are put through a matrix transformation before being directed to the shaft and trunnion motor drive amplifiers. The matrix transformation makes the image correspond directly to the hand controller motion. That is, up, down, right, and left motions of the hand controller commands the target image to move up, down, right, and left, respectively, in the field of view.

Buttons on the indicator control panel are used to instruct the computer that a navigation fix has been made and that sextant shaft and trunnion position and time should either be recorded or rejected. The "mark" command is generated manually by the flight crew which energizes the mark relay. The mark relay transmits a mark command to the computer. If an erroneous mark is made, the "mark reject" button is depressed.

The computer-controlled operation is selected by placing the moding switch in computer position. The mechanization of this loop is chosen by the computer program that has been selected by the flight crew. The computer controls the sextant by completing the circuit from the coupling data unit digital-to-analog converters to the shaft and trunnion motor drive amplifier. The computer can then provide inputs to these amplifiers via a digital input to the coupling data unit which converts it to an 800cycle signal that can be used by the motor drive amplifier. This mode is used when it is desired to look at a specific star for which the computer has the corresponding star coordinates. The computer will know the attitude of the spacecraft from the position of the inertial measurement unit gimbals and will, therefore, be able to calculate the position of the sextant axes required to acquire the star.

The computer subsystem consists of the command module computer and two display and keyboard panels. The computer and one keyboard are located in the lower equipment bay. The other keyboard is located on the main display console. All computer controls and displays are located on the keyboards.

The computer is a core memory digital computer with both fixed and erasable memory. The fixed memory permanently stores navigation tables, trajectory parameters, programs, and constants. The erasable memory stores intermediate information. The computer processes data and issues discrete control signals, both for guidance and control and for other spacecraft subsystems. It is a control computer with many of the features of a general-purpose computer. As a control computer, it aligns the stable platform of the inertial measurement unit in the inertial subsystem, positions the optical unit in the optical subsystem, and issues control commands to the spacecraft thrusters. As a general-purpose computer, it solves guidance problems required for the spacecraft mission. In addition, it monitors the operation of the guidance and control and other spacecraft subsystems.

The computer computes information about the flight profile that the spacecraft must assume in order to complete its mission. This position, velocity, and trajectory information is used by the computer to solve the various flight equations. The results of various equations can be used to determine the required magnitude and direction of thrust required. Corrections to be made are established by the computer. The spacecraft engines are turned on at the correct time, and steering and engine control signals are generated by the computer to re-orient the spacecraft to a new trajectory if required. The inertial subsystem senses acceleration and supplies velocity changes to the computer for calculating the total velocity. Drive signals are supplied from the computer to coupling data unit and stabilization gyros in the inertial subsystem to align the gimbal angles in the inertial measurement unit. Error signals are also supplied to the coupling data unit to provide attitude error display signals for the flight director attitude indicator. Coupling data unit position signals are fed to the computer to indicate commanded changes in engine gimbal angles. The computer receives mode indications and angular information from the optical subsystem during optical sightings and uses it to calculate present position and orientation and to refine trajectory information. Optical subsystem components can also be positioned by drive signals supplied from the computer.

The computer is functionally divided into seven blocks: timer, sequence generator, central processor, memory, priority control, input-output, and power.

The timer generates the synchronization pulses to assure a logical data flow from one area to another within the computer. It also generates timing waveforms which are used by the computer in alarm circuitry and other areas of the spacecraft for control and synchronization purposes.

The master clock frequency is generated by an oscillator and is applied to the clock divider logic. The divider logic divides the master clock input into gating and timing pulses at the basic clock rate of the computer. Several outputs are available from the pulses at the basic clock rate of the computer. Several outputs are available from the scaler, which further divides the divider logic output into output pulses and signals which are used for gating, to generate rate signal outputs, and for the accumulation of time. Outputs from the divider logic also drive the time pulse generator which produces a recurring set of time pulses. This set of time pulses defines a specific interval (memory cycle time) in which access to memory and word flow takes place within the computer. The start-stop logic senses the status of the power supplies and specific alarm conditions in the computer and generates a stop signal which is applied to the time pulse generator to inhibit word flow. Simultaneously, a fresh-start signal is generated which is applied to all functional areas in the computer.

The sequence generator directs the execution of machine instructions. It does this by generating control pulses which sequence data throughout the computer. The control pulses are formed by combining the order code of an instruction word with synchronization pulses from the timer, The sequence generator contains the order code processor, command generator, and control pulse generator. The sequence generator executes the instructions stored in memory by producing control pulses which regulate the data flow of the computer. The manner in which the data flow is regulated among the various functional areas of the computer and between the elements of the central processor causes the data to be processed according to the specifications of each machine instruction.

The order code processor receives signals from the central processor, priority control, and peripheral equipment (test equipment). The order code signals are stored in the order code processor and converted to coded signals for the command generator. The command generator decodes these signals and produces instruction commands. The instruction commands are sent to the control pulse generator to produce a particular sequence of control pulses, depending on the instruction being executed. At the completion of each instruction, new order code signals are sent to the order code processor to continue the execution of the program.

The central processor performs all arithmetic operations required of the computer, buffers all information coming from and going to memory, checks for correct parity on all words coming from memory, and generates a parity bit for all words written into memory. It consists of flip-flop registers, the write, clear, and read control logic, write amplifiers, memory buffer register, memory address register and decoder, and the parity logic.

Primarily, the central processor performs operations indicated by the basic instructions of the program stored in memory. Communication within the central processor is accomplished through the write amplifiers. Data flows from memory to the flip-flop registers or vice versa, between individual flip-flop registers, or into the central processor from external sources. In all instances, data is placed on the write lines and routed to a specific register, or to another functional area under control of the write, clear, and read logic. The logic section accepts control pulses from the sequence generator and generates signals to read the content of a register onto the write lines, and write this content into another register of the central processor or to another functional area of the computer. The particular memory location is specified by the content of the memory address register. The address is fed from the write lines into this register, the output of which is decoded by the address decoder logic. Data is subsequently transferred from memory to the memory buffer register. The decoded address outputs are also used as gating functions within the computer.

The memory buffer register buffers all information read out or written into memory. During readout, parity is checked by the parity logic and an alarm is generated in case of incorrect parity. During write-in, the parity logic generates a parity bit for information being written into memory. The flip-flop registers are used to accomplish the data manipulations and arithmetic operations. Each register is 16 bits or one computer word in length. Data flows into and out of each register as dictated by control pulses associated with each register. The control pulses are generated by the write, clear, and read control logic.

External inputs through the write amplifiers included the content of both the erasable and fixed memory bank registers, all interrupt addresses from priority control, control pulses which are associated with specific arithmetic operations, and the start address for an initial start condition. Information from the input and output channels is placed on the write lines, and routed to specific destinations either within or external to the central processor.

Memory provides the storage for the computer and is divided into two sections: erasable memory and fixed memory. Erasable memory can be written into or read from; it is destroyed when it is read out (displayed); therefore, information required for later use must be restored. Fixed memory cannot be written into and its readout is nondestructive. Erasable memory stores intermediate results of computations, auxiliary program information, and variable data supplied by the guidance and control and other subsystems of the spacecraft. Fixed memory stores programs, constants, and tables. There is a total of 38,912 sixteen-bit word storage locations in fixed and erasable memories. It should be noted that the majority of the memory capacity is in fixed memory (36,864 word locations). The erasable memory uses planes or ferrite (iron) cores as storage devices. A core is a magnetic storage device having two stable states. It can be magnetized in one or two directions by passing a sufficient current through a wire which pierces the core. The direction of current determines the direction of magnetization. The core will retain its magnetization indefinitely until an opposing current switches the core in the opposite direction. Wires carrying current through the same core are algebraically additive. Sense wires which pierce a switched core will carry an induced pulse. The fixed memory is high-density core rope tiny nickel-iron cores woven together with thousands of copper wires and encapsulated in plastic. Each core functions as a transformer and storage does not depend on magnetization. The advantages of encapsulated core rope for fixed memory are indestructibility, permanence of data, and storage of a vast amount in a small volume. The technique requires, however, that programs for classes of missions be developed and verified before the rope is woven, because the program determines the wiring sequence.

Priority control establishes a processing priority of operations that must be performed by the computer. These operations are a result of conditions which occur both internally and externally. Priority control consists of counter priority control and interrupt priority control. Counter priority control initiates actions which update counters in erasable memory. Interrupt priority control transfers control of the computer to one of several interrupt subroutines stored in fixed memory.

The start instruction control restarts the computer after a hardware or program failure. The counter instruction control updates the various counters in erasable memory upon reception of certain incremental pulses. The interrupt instruction control forces the execution of the interrupt instruction to interrupt the current operation of the computer in favour of a programmed operation of a higher priority.

The input-output section routes and conditions signals between the computer and other areas of the spacecraft. In addition to the counter interrupt and the program interrupts previously described, the computer has a number of other inputs derived from its interfacing hardware. These inputs are a result of the functioning of the hardware, or an action by the operator of the spacecraft. The counter interrupts, in most cases, enable the computer to process inputs representing such things as changes in velocity. The program interrupt inputs are used to initiate processing that must be done a relatively short time after a particular function is present. The other inputs to the computer, in general, enable it to be aware of conditions which exist in its environment. These inputs are routed to computer and are available to its programs through the input channels.

The outputs of the computer fall in one of the following categories: data, control, or condition indications. Some of these outputs are controllable through the computer program while others are present as a function of computer circuitry. All of the outputs which are controlled by the computer programs are developed through the computer output channels.

The power section provides voltage levels necessary for the proper operation of the computer. Power is furnished by two switching-regulator power supplies: a +4-volt and a +14-volt power supply which are energized by fuel cell powerplants in the electrical power subsystem.

The power supply outputs are monitored by a failure detector consisting of four differential amplifiers. There are two amplifiers for each power supply, one for overvoltage and one for undervoltage detection. If an overvoltage or undervoltage condition exists, a relay closure signal indicating a power fail is supplied to the spacecraft.

The display and keyboard panels provide communication between the flight crew and the computer. They operate in parallel, with the main display console keyboard providing computer display and control while crewmen are in their couches.

The exchange of data between the flight crew and the computer is usually initiated by crew action; however, it can also be initiated by internal computer programs. The exchanged information is processed by the keyboard program. This program allows the following four different modes of operation:

1. Display of internal data-both a one-shot display and a periodically updating display (called monitor) are provided.

2. Loading external data-as each numerical character is entered, it is displayed in the appropriate display panel location.

3. Program calling and control-the keyboard is used to initiate a class of routines which are concerned with neither loading nor display; certain routines require instructions from the operator to determine whether to stop or continue at a given point.

4. Changing major mode-the initiation of large scale mission phases can be commanded by the operator.

The data involved in both loading and display can be presented in either octal or decimal form as the operator indicates. If decimal form is chosen, the appropriate scale factors are supplied by the program. Decimal entries are indicated by entering a plus or minus sign.

The basic language of communication between the operator and the computer is a pair of words known as verb and noun. Each of these is represented by a two-character decimal number. The verb code indicates what action is to be taken (operation); the noun code indicates to what the action is applied (operand). Typical verbs are those for displaying and loading. Nouns usually refer to a group of erasable registers within the computer memory. The program, verb, and noun displays provide two-digit numbers which are coded numbers describing the action being performed. The register 1, 2, and 3 displays show the contents of registers or memory locations. These displays are numbers which are read as decimal numbers if a plus or minus sign is present and octal numbers if no sign is used. The register displays operate under program control unless the contents of a specific register or memory location is desired. The crew may request display of the contents of a specific register or memory location by commanding the display from the keyboard. The only other displays are the "activity" lights which indicate whether the computer is computing or accepting telemetry from the ground and status lights.

The keyboard consist of ten numerical keys (pushbuttons) labeled 0 through 9, two sign keys (+ or -) and seven instruction keys: Verb, Noun, Clr (clear), Pro (proceed), Key Rel (key release), Entr (enter) and Rset (reset). Whenever a key is depressed, 14 volts are applied to a diode encoder which generates a unique five-bit code associated with that key. There is, however, no five-bit code associated with the proceed key. The function of the keys is as follows:

Pushbutton	Function
0 through 9	Enters numerical data, noun codes, and verb codes into the computer
+ and -	Informs the computer that the following numerical data is decimal and indicates the sign of the data
Noun	Conditions the computer to interpret the next two numerical characters as a noun code and causes the noun display to be blanked
Clear	Clears data contained in the data displays; depressing this key clears the data display currently being used; successive depressions clear the other two data displays
Proceed	Commands the computer to proceed on to the standby mode when depressed, if in standby mode, depression commands the computer to resume regular operation (operate mode)
Key Release	Releases the keyboard displays initiated by keyboard action so that information supplied by the computer program may be displayed
Enter	Informs the computer that the assembled data is complete and that the requested function is to be executed
Reset	Extinguishes the lamps that are controlled by the computer
Verb	Conditions the computer to interpret the next two numerical characters as a verb code and causes the verb display to be blanked

The standard procedure for the execution of keyboard operations consists of a sequence of seven key depressions:

Pressing the Verb key blanks the two verb lights on the keyboard and clears the verb code register in the computer. The next two numerical inputs are interpreted as the verb code. Each of these characters is displayed by the verb lights as it is inserted. The Noun key operates similarly with the keyboard noun lights and computer noun code register. Pressing the Enter key initiates the program indicated by the verb-noun combination. Thus, it is not necessary to follow a standard procedure in keying verb-noun codes; it can be done in reverse order or a previously inserted verb or noun can be used without re-keying it. No action is taken by the computer in initiating the verb-noun-defined program until the Enter key is actuated. If an error is noticed in either the verb or noun code before actuation of the Enter key, it can be corrected simply by pressing the corresponding Verb or Noun key and inserting the proper code.

If the selected verb-noun combination requires data from the operator, the Verb and Noun lights flash on and off about once per second after the Enter key is pressed. Data is loaded in five-character words and is displayed character-by-character in one of the five-position data display registers. The Enter key must be pressed after each data word. This tells the program that the numerical word being keyed in is complete.

The keyboard also can be used by internal computer programs for subroutines. However, any operator keyboard action (except error reset) inhibits keyboard use by internal routines. The operator retains control of the keyboard until he wishes to release it. This assures that the data he wishes to observe will not be replaced by internally initiated data displays.

A noun code may refer to a device, a group of computer registers, or a group of counter registers, or it may simply serve to convey information without referring to any particular computer register. The noun is made up of 1, 2, or 3 components, each entered separately as requested by the verb code. As each component is keyed, it is displayed on the display panel: component 1 in Register 1, component 2 in Register 2, and component 3 in Register 3. There are two classes of nouns: normal and mixed. Normal nouns (codes 01 through 39) are those whose component members refer to computer registers which have consecutive addresses and use the same scale factor when converted to decimal. Mixed nouns (codes 40 through 99) are those whose component members refer to non consecutive addresses or whose component members require different scale factors when converted to decimal, or both.

A verb code indicates what action is to be taken. It also determines which component member of the noun group is to be acted upon. For example, there are five different load verbs. Verb 21 is required for loading the first component of the selected noun; verb 22 loads the second component; verb 23 loads the third component; verb 24 loads the first and second component; and verb 25 loads all three components. A similar component format is used in the display and monitor verbs. There are two general classes of verbs: standard and extended. The standard verbs (codes 01 through 39) deal mainly with loading, displaying, and monitoring data. The extended verbs (codes 40 through 99) are principally concerned with calling up internal programs whose function is system testing and operation.