A redundant set of four orbiter general-purpose computers forms the primary avionics software system; a fifth GPC is used as the backup flight system.
The GPCs interface with the various systems through the orbiter's flight forward and flight aft multiplexers/demultiplexers. The data buses serve as a conduit for signals going to and from the various sensors that provide velocity and attitude information as well as for signals traveling to and from the orbiter propulsion systems, orbiter aerodynamic control surfaces, and displays and controls.
The GN&C system consists of two operational modes: auto and manual (control stick steering). In the automatic mode, the primary avionics software system essentially allows the GPCs to fly the vehicle; the flight crew simply selects the various operational sequences. The flight crew may control the vehicle in the control stick steering mode using hand controls, such as the rotational hand controller, translational hand controller, speed brake/thrust controller and rudder pedals. The translational hand controller is available only for the commander, but both the commander and pilot have a rotational hand controller.
In the control stick steering mode, flight crew commands must still pass through and be issued by the GPCs. There are no direct mechanical links between the flight crew and the orbiter's various propulsion systems or aerodynamic surfaces; the orbiter is an entirely digitally controlled, fly-by-wire vehicle.
During launch and ascent, most of the GN&C commands are directed to gimbal the three space shuttle main engines and solid rocket boosters to maintain thrust vector control through the vehicle's center of gravity at a time when the amount of consumables is changing rapidly. In addition, the GN&C controls SSME throttling for maximum aerodynamic loading of the vehicle during ascent-referred to as max q-and to maintain an acceleration of no greater than 3 g's during the ascent phase. To circularize the orbit and perform on-orbit and deorbit maneuvers, the GN&C commands the orbital maneuvering system engines. At external tank separation, on orbit and during portions of entry, GN&C controls commands to the reaction control system. In atmospheric flight, GN&C controls the orbiter aerodynamic flight control surfaces.
Functions of GN&C software include flight control, guidance, navigation, hardware data processing and flight crew display. Specific function tasks and their associated GN&C hardware vary with each mission phase.
Vehicle control is maintained and in-flight trajectory changes are made during powered flight by firing and gimbaling engines. During atmospheric flight, these functions are performed by deflecting aerosurfaces. Flight control computes and issues the engine fire and gimbal commands and aerosurface deflection commands.
Flight control includes attitude processing, steering, thrust vector control and digital autopilots. Flight control receives vehicle dynamics commands (attitudes, rates and accelerations) from guidance software or flight crew controllers and processes them for conversion to effector commands (engine fire, gimbal or aerosurface). Flight control output commands are based on errors for stability augmentation. The errors are the difference between the commanded attitude, aerosurface position, body rate or body acceleration and the actual attitude, position, rate or acceleration.
Actual attitude is derived from inertial measurement unit angles, aerosurface position is provided by feedback transducers in the aerosurface servoamplifiers, body rates are sensed by rate gyro assemblies, and accelerations are sensed by accelerometer assemblies. In atmospheric flight, flight control adjusts control sensitivity based on air data parameters derived from local pressures sensed by air data probes and performs turn coordination using body attitude angles derived from IMU angles. Thus, GN&C hardware required to support flight control is a function of the mission phase.
The guidance steering commands used by the flight control software are augmented by the guidance software or are manually commanded by the hand controller or speed brake/thrust controller. When flight control software uses the steering commands computed by guidance software, it is termed automatic guidance; when the flight crew is controlling the vehicle by hand, it is called control stick steering. The commands computed by guidance are those required to get from the current state (position and velocity) to a desired state (specified by target conditions, attitude, airspeed and runway centerline). The steering commands consist of translational and rotational angles, rates and accelerations. Guidance receives the current state from navigation software. The desired state or targets are part of the initialized software load and some may be changed manually in flight.
The navigation system maintains an accurate estimate of vehicle position and velocity, referred to as a state vector. From position, attitude and velocity, other parameters (acceleration, angle of attack) are calculated for use in guidance and for display to the crew. The current state vector is mathematically determined from the previous state vector by integrating the equations of motion using vehicle acceleration as sensed by the IMUs and/or computed from gravity and drag models. The alignment of the IMU and, hence, the accuracy of the resulting state vector deteriorate as a function of time. Celestial navigation instruments (star trackers and crewman optical alignment sight) are used to maintain IMU alignment in orbit. For entry, the accuracy of the IMU-derived state vector is, however, insufficient for either guidance or the flight crew to bring the spacecraft to a pinpoint landing. Therefore, data from other navigation sensors-air data system, tactical air navigation, microwave scan beam landing system and radar altimeter-is blended into the state vector at different phases of entry to provide the necessary accuracy. The three IMUs maintain an inertial reference and provide velocity changes until the microwave scan beam landing system is acquired. Navigation-derived air data are needed during entry as inputs to guidance, flight control and flight crew dedicated displays. Such data are provided by tactical air navigation, which supplies range and bearing measurements beginning at 160,000 feet; the air data system provides information at about Mach 3. Tactical air navigation is used until the microwave scan beam landing system is acquired or an altitude of 1,500 feet is reached if MSBLS is not available.
During rendezvous and proximity operations, the onboard navigation system maintains the state vectors of both the orbiter and target vehicle. During close operations (separation of less than 15 miles), these two state vectors must be very accurate in order to maintain an accurate relative state vector. Rendezvous radar measurements (range and range rate) are used for a separation of about 15 miles to 100 feet to provide the necessary relative state vector accuracy. When two vehicles are separated by less than 100 feet, the flight crew relies primarily on visual monitoring (aft and overhead windows and closed-circuit television).
In summary, GN&C hardware sensors used by navigation include IMUs, star trackers, the crewman optical alignment sight, tactical air navigation, air data system, microwave scan beam landing system, radar altimeter and rendezvous radar. The GN&C hardware sensors used by the flight control system are accelerometer assemblies, orbiter rate gyro assemblies, solid rocket booster rate gyro assemblies, controllers and aerosurface servoamplifiers.
A sensor, controller or flight control effector cannot be assigned or rerouted to another MDM during flight, but the MDM it is wired to can be assigned to a different GPC. Each multiple unit of each type of GN&C hardware is hard-wired to a different MDM. For example, there are four accelerometer assemblies on board the orbiter. AA 1 is wired to forward flight MDM 1 and is part of string 1, AA 2 is wired to FF MDM 2 on string 2, AA 3 is wired to FF MDM 3 on string 3, and AA 4 is wired to FF MDM 4 on string 4.
Although flight could be accomplished with only one, three IMUs are installed on the orbiter for redundancy. The IMUs are mounted on the navigation base, which is located inside the crew compartment flight deck forward of the flight deck control and display panels. The navigation base mounting platform is pitched down 10.6 degrees from the orbiter's plus X body axis. The navigation base provides a platform for the IMUs that can be repeatedly mounted with great accuracy, enabling the definition of transformations that relate IMU reference frame measurements to any other reference frame.
The IMU consists of a platform isolated from vehicle rotations by four gimbals. Since the platform does not rotate with the vehicle, its orientation remains fixed, or inertial, in space. The gimbal order from outermost to innermost is outer roll, pitch, inner roll and azimuth. The platform is attached to the azimuth gimbal. The inner roll gimbal is a redundant gimbal used to provide an all-attitude IMU while preventing the possibility of gimbal-lock (a condition that can occur with a three-gimbal system and cause the inertial platform to lose its reference). The outer roll gimbal is driven from error signals generated from disturbances to the inner roll gimbal. Thus, the inner roll gimbal will remain at its null position, orthogonal to the pitch gimbal.
The inertial sensors consist of two gyros, each with two degrees of freedom, that provide platform stabilization. The gyros are used to maintain the platform's inertial orientation by sensing rotations of the platform caused by vehicle-rotation-induced friction at the gimbal pivot points. The gyros output a signal that is proportional to the motion and is used by the gimbal electronics to drive the appropriate gimbals to null the gyro outputs. Thus, the platform remains essentially undisturbed, maintaining its inertial orientation while the gimbals respond to vehicle motion. One gyro-called the vertical gyro-is oriented so its input axes are aligned with the X and Y platform axes; its input axes provide IMU platform roll and pitch stabilization. The second gyro is oriented so that one input axis lies along the platform's Z axis and the other lies in the X-Y plane. This gyro-the azimuth gyro-provides platform yaw stabilization with the Z input axis, while the second input axis is used as a platform rate detector for built-in test equipment. Each gyro contains a two-axis pick-off that senses deflection of the rotating wheel. The gyro also contains a pair of two-axis torquers that provide compensation torquing for gyro drift and a means to reposition the platform.
The spin axis of a gyro is its axis of rotation. The inertial stability of the spin axis is a basic property of gyroscopes and is used in stabilization loops, which consist of the gyro pick-off, gimbals and gimbal torquers. When the vehicle is rotated, the platform also tends to rotate due to friction at the gimbal pivot points. Since the gyro casing is rigidly mounted to the platform, it will also rotate. The gyro resists this rotation tendency to remain inertial, but the resistance is overcome by friction. This rotation is detected by the pick-offs as a deflection of the rotating gyro wheel. A signal proportional to this deflection is sent to the gimbal electronics, which routes the signals to the appropriate torquers, which in turn rotate their gimbals to null the pick-off point. When the output is nulled, the loop is closed.
Four resolvers in an IMU are used to measure vehicle attitude. A resolver is located at one of two pivot points between adjacent gimbals. The IMU resolvers are electromechanical devices that apply the principle of magnetic induction to electrically measure the angle between two adjacent gimbals. This electrical signal is then transformed into a mechanical angle by the IMU electronics. There are two resolvers on each gimbal: one-speed (1X) and eight-speed (8X). The 1X electrical output represents a coarse measurement of the true gimbal mechanical angle. For greater resolution, the 8X electrical output represents a measurement eight times that of the true angle. These outputs are converted to an angle measurement in the IMU electronics and are sent to the GPCs, where they are combined into a single gimbal angle measurement and are used to determine vehicle attitude. Attitude information is used by flight control for turn coordination and steering command guidance. An attitude director indicator displays attitude and navigation data.
Two accelerometers in each IMU measure linear vehicle accelerations in the IMU inertial reference frame; one measures the acceleration along the platform's X and Y axes, the other along the Z axis. The accelerometer is basically a force rebalance-type instrument. When the accelerometer experiences an acceleration along its input axes, it causes a pendulum mass displacement. This displacement is measured by a pick-off device, which generates an electrical signal that is proportional to the sensed acceleration. This signal is amplified and returned to a torquer within the accelerometer, which attempts to reposition the proof mass to its null (no output) position.
The velocity data measured by the IMU are the primary sources that propagate the orbiter state vector during ascent and entry. On orbit, a sophisticated drag model is substituted for IMU velocity information, except during large vehicle accelerations. During large on-orbit accelerations, IMU velocity data are used in navigation calculations.
Platform attitude can be reoriented by two methods: slewing or pulse torquing. Slewing rotates the platform at a high rate (72 degrees per minute), while pulse torquing rotates it very slowly (0.417 degree per minute). Platform reorientation relies on another property of gyroscopes: precession. If a force is applied to a spinning gyroscope, the induced motion is 90 degrees from the input force. In each IMU, a two-axis torquer is located along the input axes of both gyros. Commands are sent to the torquers from the GPC to apply a force along the input axes. The result is a deflection of the gyro spin axis that is detected and nulled by the stabilization loops. Since the gyro spin axis is forced to point in a new direction, the platform has to rotate to null the gyro outputs.
The three IMUs have skewed orientations-their axes are not coaligned and are not aligned with the vehicle axes. This is done for two reasons. First, gimbaled platforms have problems at certain orientations. This skewing ensures that no more than one IMU will have an orientation problem for a given attitude. Skew allows resolution of a single-axis failure on one IMU by multiple axes on another IMU since the possibility of multiple-axis failure is more remote. Second, skewing is also used by redundancy management to determine which IMUs have failures.
The IMU platform is capable of remaining inertial for vehicle rotations of up to 35 degrees per second and angular accelerations of 35 degrees per second squared. Each IMU interfaces with the five onboard GPCs through a different flight forward multiplexer/demultiplexer of the data bus network. Under GPC control, each IMU is capable of orienting its platform to any attitude, determining platform alignment relative to a reference and providing velocity and attitude data for flight operations.
Very precise thermal control must be maintained in order to meet IMU performance requirements. The IMU thermal control system consists of an internal heater system and a forced-air cooling system. The internal heater system is completely automatic and is powered on when power is initially applied to the IMU. It continues to operate until the IMU is powered down. The forced-air cooling consists of three fans that serve all three IMUs. Only one fan is necessary to provide adequate air flow. The IMU fan pulls cabin air through the casing of each IMU and cools it in an IMU heat exchanger before returning it to the cabin. Each IMU fan is controlled by an individual on/off switch located on panel L1.
Each IMU is supplied with redundant 28-volt dc power through separate remote power controllers when control bus power is applied to the RPCs by the IMU power switch. The IMU 1 , 2 and 3 on/off power switches are located on panels O14, O15 and O16, respectively. Loss of one control bus or one main bus will not cause the loss of an IMU.
Each IMU has two modes of operation: a warm-up/standby mode and an operate mode. When the respective IMU switch is positioned to on , that IMU is powered and enters the warm-up/standby mode, which applies power only to the heater circuits. It takes approximately 30 minutes for the IMU to reach its operating range, at which time the IMU enters a standby mode, when it can be moded to the operate mode by flight crew command in GN&C OPS 2, 3 or 9.
To mode the IMU to operate, the controlling GPC sends the operate discrete to the IMU through the forward flight multiplexer/demultiplexer. The IMU, upon receiving this command, initiates its run-up sequence.
The run-up sequence first cages the IMU-a process of reorienting the IMU gimbals and then mechanically locking them into place so that the gyros may begin to spin. When the IMU is caged, its platform orientation will be known when it becomes inertial. The caged orientation is defined as the point at which all resolver outputs are zero. This causes the IMU platform to lie parallel to the navigation base plane with its coordinate axes lying parallel to the navigation base's coordinate axes.
Once the IMU gimbals are caged, the gyros begin to spin and power is applied to the remaining IMU components. When the gyros have reached the correct spin rate, the stabilization loops are powered, and the IMU becomes inertial. At this time, the IMU returns an in operate mode discrete to the GPC, indicating that the run-up sequence is complete. This process requires approximately 38 seconds.
The IMUs are in operate by the time the flight crew enters the vehicle before launch and remain in that state for the duration of the flight unless powered down to minimize power consumption. While in the operate mode, the IMU maintains its inertial orientation and is used for calibrations and preflight, flight and on-orbit alignments.
Before preflight, the IMUs are taken through three levels of calibration to correct for hardware inaccuracies: factory calibration, hangar calibration and preflight calibration. Sixty-one IMU parameters are developed during this extensive calibration period. These parameters are stored in the orbiter GPC mass memory units and are used in the software to compensate for hardware inaccuracies.
At T minus two hours during the launch countdown, the IMU calibration is complete and the IMUs are ready for the preflight alignment. At T minus one hour and one minute, the Launch Control Center initiates this alignment by a display electronics unit equivalent. (A DEU equivalent is simply a ground command that looks to the GPCs like a crew keyboard input.)
Preflight alignment requires 48 minutes to complete and consists of two different operations: gyrocompass alignment and velocity/tilt initialization.
In the gyrocompass alignment, each IMU is oriented so that the desired relative skew is achieved when the platforms are at their alignment orientation. During this phase, the IMUs are placed in two orientations relative to the north-west-up coordinate system. These two orientations differ only in a 90-degree rotation about the up axis. Data are collected for 90 seconds by the accelerometers to remove any misalignment resulting from the reorientation. The accelerometers are used here because their accuracy is much better than that of the resolvers and the acceleration due to Earth rotation is definitely known. Therefore, any unexpected acceleration is due to IMU misalignment. Once this misalignment is nulled, the platform is torqued about the north axis to compensate for Earth rotation. Data are then collected for 10 minutes to measure platform drifts. This sequence of data collecting is repeated at the second orientation. The relative attitude errors for each IMU pair are also computed, first with resolver data and then with accelerometer data. The two values are subtracted and transformed into body coordinates. A factory-calibrated relative resolver error term is then subtracted, and a reasonableness test is performed to check the relative alignment between each IMU pair to assure a good preflight alignment. The velocity/tilt initialization mode is then entered, during which the drifts experienced while waiting for the OPS 1 transition are estimated. The compensation developed by these drifts is applied to the gyros from the OPS 1 transition to T minus 12 minutes and is also used to compute the current platform to the mean of 1950 reference stable member matrix at the OPS 1 transition. In addition, a level-axis test is performed on each platform three times a second; failure to pass this test requires the alignment to be repeated.
At T minus 22 seconds, a one-shot data transfer from the primary avionics software system to the backup flight system is commanded by display electronics unit equivalent. IMU compensation data computed by the PASS GPCs in OPS 9 are sent to the BFS GPC at this time so that it will have the same data for controlling the IMUs if it is engaged.
At the OPS 1 transition, the IMUs enter the ''tuned inertial'' drift compensation mode. It is tuned because a compensation factor computed in the velocity/tilt is applied to the IMU gyros. At T minus 12 seconds, this compensation is removed and the IMUs enter the ''free inertial'' mode. The IMUs are now flight ready, and all functions, both hardware and software, remain the same throughout the flight.
During the orbital flight phase, the IMUs provide GN&C software with attitude and accelerometer data.
On-orbit alignments are necessary to correct platform misalignment caused by uncompensated gyro drift.
The IMU can be safely powered off from either the warm-up/standby mode or the operate mode. If an IMU is moded to standby, an internal timer inhibits moding operation for three minutes to allow the gyros to spin to a stop so that the proper sequencing to the operate mode can occur.
In the event of an IMU failure, the IMU red caution and warning light on panel F7 will be illuminated. If temperatures are out of limits or if built-in test equipment detects a failure, a fault message and SM alert will be annunciated.
The accuracy of the IMU deteriorates with time. If the errors are known, they can be physically or mathematically corrected. Software based on preflight calibrations is used to compensate for most of the inaccuracy. The star trackers and crewman optical alignment sight are used to determine additional inaccuracies.
The IMU subsystem operating program processes the data from the IMUs and converts it to a usable form for other users. The following computations are performed in the IMU SOP: conversion of velocities to the mean of 1950 coordinates; conversion of resolver outputs to gimbal angles; computation of accelerations for displays; performance of additional software built-in test equipment checks; support of selection, control and monitoring of IMU submodes of the operate mode; and computation of torquing commands based on the misalignment determined by the star trackers, crewman optical alignment sight or another IMU. Misalignments are due to gyro drifts.
A new high-accuracy inertial navigation system will be phased in to augment the present KT-70 IMU during 1988-89. The HAINS will provide spares support for the inertial navigation system and will eventually phase out the KT-70 IMU design. Benefits of the HAINS include lower program costs over the next decade, ongoing production support, improved performance, lower failure rates, and reduced size and weight. The HAINS is 9.24 inches high, 8.49 inches wide and 22 inches long. The unit weighs 43.5 pounds. The HAINS also contains an internal dedicated microprocessor with memory for processing and storing compensation and scale factor data from the vendor's calibrations, thereby reducing the need for extensive initial load data for the orbiter GPCs. The HAINS is both physically and functionally interchangeable with the KT-70 IMU.
The IMU contractor is Singer Electronics Systems Division, Little Falls, N.J.
Alignment of the IMUs is required approximately every 12 hours to correct IMU drift, within one to two hours before major on-orbit thrusting duration or after a crewman optical alignment sight IMU alignment. IMU alignment is accomplished by using the star trackers to measure the line-of-sight vector to at least two stars. With this information, the GPC calculates the orientation between these stars and the orbiter to define the orbiter's attitude. A comparison of this attitude with the attitude measured by the IMU provides the correction factor necessary to null the IMU error.
The GPC memory contains inertial information for 50 stars chosen for their brightness and their ability to provide complete sky coverage.
The star trackers are oriented so that the optical axis of the negative Z star tracker is pointed approximately along the negative Z axis of the orbiter and the optical axis of the negative Y star tracker is pointed approximately along the negative Y axis of the orbiter. Since the navigation base provides the mount for the IMUs and star trackers, the star tracker line of sight is referenced to the navigation base and the orbiter coordinate system; thus, the GPC knows where the star tracker is pointed and its orientation with respect to the IMUs.
To enable the star tracker doors to open, the star tracker power minus Y and minus Z switches on panel O6 must be positioned to on. The star tracker door control sys 1 and sys 2 switches on panel O6 control one three-phase ac motor on each door. Positioning the sys 1 switch to open controls motor control logic and drives the minus Y and minus Z star tracker door by electromechanical actuators to the open position. Limit switches stop the motors when the doors are open and control a talkback indicator above the sys 1 switch. The talkback indicator indicates when the doors are open. Setting the sys 2 switch to open controls a redundant ac motor and electromechanical actuators to open the minus Y and minus Z star tracker door, and limit switches stop the motors when the doors are open and control the talkback indicator above the sys 2 switch in the same manner as for system 1. Positioning the sys 1 switch to close drives the minus Y and minus Z door closed; the talkback indicator above the switch indicates cl. The door opening or closing time with two motors is six seconds; with one motor, it is 12 seconds. Setting the sys 2 switch to close drives the system 2 motors and closes the minus Y and minus Z door; the talkback indicator above the switch indicates cl . The indicators indicate barberpole when a door is between open or closed. The off position of the sys 1 or 2 switch removes power from the corresponding motor control logic circuitry.
The difference between the inertial attitudes defined by the star tracker and the IMU is processed by software and results in IMU torquing angles. If the IMU gimbals are physically torqued or the matrix defining its orientation is recomputed, the effects of the IMU gyro drift are removed and the IMU is restored to its inertial attitude. If the IMU alignment is in error by more than 1.4 degrees, the star tracker is unable to acquire and track stars. In this case, the crewman optical alignment sight must be used to realign the IMUs to within 1.4 degrees; the star trackers can then be used to realign the IMUs more precisely. The star tracker cannot be used if the IMU alignment error is greater than 1.4 degrees because the angles the star tracker is given for searching are based on current knowledge of the orbiter attitude, which is based on IMU gimbal angles. If that attitude is greatly in error, the star tracker may acquire and track the wrong star.
In addition to aligning the IMUs, the star trackers can be used to provide angular data from the orbiter to a target. This capability can be used during rendezvous or proximity operations with a target satellite.
The star tracker includes a light shade assembly and an electronics assembly mounted on top of the navigation base. The light shade assembly defines the tracker field of view (10 degrees square). Its shutter mechanism may be opened manually by the crew using an entry on the cathode ray tube display, or it can be opened and closed automatically by a bright object sensor or target suppress software. The bright object sensor reacts before a bright object, such as the sun or moon, can damage the star tracker (the sensor has a larger field of view than the star tracker shutter). The target suppress software reacts to a broad light source (such as the sunlit Earth), which may not trip the bright object sensor but could produce overall illumination large enough to cause photo currents larger than desired.
The electronics assembly contains an image dissector tube mounted on the underside of the navigation base. The star tracker itself does not move-the field of view is scanned electronically. The star tracker may be commanded to scan the entire field of view or a smaller offset field of view (1 degree square) about a point defined by horizontal and vertical offsets. An object is tracked when the proper intensity and the correct location are sensed. Star tracker outputs are the horizontal and vertical position within the field of view of the object being tracked and its intensity.
There is no redundancy management for the star tracker assemblies; they operate independently, and either can do the whole task. They can be operated either separately or concurrently.
The star tracker subsystem operating program supports the modes that are commanded manually: self-test, star track, target track, break track and term/idle. Self-test consists of software and hardware tests. In the star track mode, the star tracker does an offset scan search for the star, acquires it and tracks it. The star may be selected by the flight crew or GPC; in either case, field-of-view and occultation checks are made. Target track is the same as star track, but the flight crew must specify the target and its threshold. Break track forces the star tracker to stop tracking a star and to perform a search scan from the current location to track the next star it acquires. In the term/idle mode, the star tracker continues its operation, but all star tracker software processing ceases.
In addition, the star tracker SOP maintains the star table. When a star tracker has acquired and tracked a star and the data has passed software checks, the star identification, time tag and line-of-sight vector are stored. The identification and time elapsed since time tag are displayed in the star table. When two or three stars are in the table, the angular difference between their line-of-sight vectors is displayed. The difference between the star tracker and star catalog angular differences is displayed as an error. The star tracker SOP selects line-of-sight vectors of two stars in the star table for IMU alignment and outputs an align ena discrete. The software selects the star pair whose angular difference is closest to 90 degrees or the pair whose elapsed time of entry into the table is less than 60 minutes. The flight crew may manually override the SOP selection or clear the table if desired.
The SOP also determines and displays star tracker status.
The contractor for the star trackers is Ball Brothers, Boulder, Colo.
The COAS is mounted at the commander's station so the crew can check for proper attitude orientation during ascent and deorbit thrusting periods. For on-orbit operations, the COAS at the commander's station is removed and installed next to the aft flight deck overhead right minus Z window.
The COAS is an optical device with a reticle projected on a combining glass that is focused on infinity. The reticle consists of 10-degree-wide vertical and horizontal cross hairs with 1-degree marks and an elevation scale on the right side of minus 10 to 31.5 degrees. A light bulb with variable brightness illuminates the reticle. The COAS requires 115-volt ac power for reticle illumination. The COAS is 9.5 by 6 by 4.3 inches and weighs 2.5 pounds.
After mounting the COAS at the aft flight station, the flight crew member must manually maneuver the orbiter until the selected star is in the field of view. The crew member maneuvers the orbiter so that the star crosses the center of the reticle. At the instant of the crossing, the crew member makes a mark by depressing the most convenient att ref push button; the three att ref push buttons are located on panels F6, F8 and A6. At the time of the mark, software stores the gimbal angles of the three IMUs. This process can be repeated if the accuracy of the star's centering is in doubt. When the crew member feels a good mark has been taken, the software is notified to accept it. Good marks for two stars are required for an IMU alignment. The separation between the two stars should be between 60 and 120 degrees.
By knowing the star being sighted and the COAS's location and mounting relationship in the orbiter, software can determine a line-of-sight vector from the COAS to the star in an inertial coordinate system. Line-of-sight vectors to two stars define the attitude of the orbiter in inertial space. This attitude can be compared to the attitude defined by the IMUs and can be realigned to the more correct orientation by the COAS sightings if the IMUs are in error.
The COAS's mounting relative to the navigation base on which the IMUs are mounted is calibrated before launch. The constants are stored in software, and COAS line-of-sight vectors are based on known relationships between the COAS line of sight and the navigation base.
COAS can also be used to visually track targets during proximity operations or to visually verify tracking of the correct star by the minus Z star tracker.
COAS data processing is accomplished in the star tracker SOP. This SOP accepts and stores crew inputs on COAS location, star identification or calibration mode; accepts marks; computes and stores the line-of-sight vectors; enables IMU alignment when two marks have been accepted; and computes, updates and provides display data.
The orbiter is equipped with three TACAN sets that operate redundantly. Each TACAN has two antennas: one on the orbiter's lower forward fuselage and one on the orbiter's upper forward fuselage. The antennas are covered with reusable thermal protection system tiles.
The onboard TACAN sets are used for external navigation and for the orbiter during the entry phase and return-to-launch-site abort. Normally, several ground stations will be used after leaving L-band communications blackout and during the terminal area energy management phases. TACAN's maximum range is 400 nautical miles (460 statute miles).
Each ground station has an assigned frequency (L-band) and a three-letter Morse code identification. The ground station transmits on one of 252 (126X, 126Y) preselected frequencies (channels) that correspond to the frequencies the onboard TACAN sets are capable of receiving. These frequencies are spaced at 63-MHz intervals.
The TACAN ground station beacon continuously transmits pulse pairs on its assigned frequency. The orbiter TACAN receivers pick up these pulse pairs, and the TACAN data processors decode them to compute bearing. The onboard TACAN sets detect the phase angle between magnetic north and the position of the orbiter with respect to the ground station. The ground beacon is omnidirectional; when the orbiter is over the ground station, or nearly so, it is in a cone of confusion. Within this cone, bearing is unusable.
Periodically, the onboard TACAN sets emit an interrogation pulse that causes the selected TACAN ground station to respond with distance-measuring equipment pulses. The slant range (orbiter to ground station) is computed by the onboard TACAN sets by measuring the elapsed time from interrogation to valid reply and subtracting known system delays. As the orbiter approaches a ground TACAN station, the range decreases. After a course has been selected, the onboard TACAN sets derive concise deviation data.
The range and bearing data are used to update the state vector position components after the data are transformed by the TACANs in the entry phase (or return to launch site) by navigation and for display on the horizontal situation indicators on panels F6 and F8, as well as for display of raw TACAN data on the cathode ray tube.
Each of the onboard TACANs has an ant sel switch on panel O7. In the auto position, the onboard GPCs automatically select the upper L-band antenna or lower L-band antenna for that TACAN. The upper and lower positions of each TACAN ant sel switch allow the flight crew to select the upper or lower L-band antenna manually.
Each of the onboard TACANs is controlled by its mode rotary switch on panel O7. The modes are off, receive, transmit and receive, and GPC. In the GPC mode, the onboard GPCs control TACAN ground station channel selection automatically, and both bearing and range are processed by hardware and software. In the transmit and receive mode, both bearing and range are processed by hardware and software, but TACAN ground station channels are selected manually using the four thumbwheels for that TACAN on panel O7. The first three thumbwheels (left to right) select the channel (frequency), and the fourth selects the X or Y. In the receive mode, only bearing is received and processed by the hardware; the thumbwheels for that TACAN would be used to select the channel.
Approximately every 37 seconds, the selected ground TACAN station transmits its three-letter identification to the onboard TACAN. In order for the Morse code identification to be verified by the commander and pilot, TACAN ID audio controls are located on panel O5 for the commander and panel O9 for the pilot. The TACAN on/off switch is positioned to on to transmit the TACAN identification. The TACAN 1 , 2 and 3 switch selects the onboard TACAN that will transmit the TACAN identification code, and the TACAN on/off switch is positioned to on to transmit the code to the orbiter's audio system, thus the commander and pilot. Volume TACAN thumbwheels on panels O5 and O9 control the volume setting of the TACAN identification code to the commander and pilot.
In the GPC mode, 10 TACAN ground stations are programmed into the software and are divided into three geometric regions: the acquisition region (three stations), the navigation region (six stations), and the landing site region (one station).
During orbital operations, landing sites are grouped into minitable and maxitable programs. The maxitable programs provide data sets that support a broad range of trajectories for contingency deorbits and enable reselection of runway and navigation and data sets for those deorbits. The minitable consists of three runways determined by the flight crew, one of which is initialized as a primary runway. The minitable is transferred from entry operations and becomes unchangeable. Entry guidance is targeted from one of the three runways selected by the crew, initialized with the primary runway for the well-defined trajectory and nominal end-of-mission data sets. Since the TACAN units are placed in groups of 10 and 10 TACAN units from one group (primary) form the TACAN half of the minitable, the secondary and alternate runways should be from the same group as the primary runway to assure TACAN coverage.
The acquisition region is the area in which the onboard TACAN sets automatically begin searching for a range lock-on of three ground stations at approximately 160,000 feet. After one TACAN acquires a range lock, the other two will lock on to the same ground station. When at least two TACAN sets lock on, TACAN range and bearing are used by navigation to update state vector until microwave scan beam landing system selection and acquisition at approximately 18,000 feet.
When the distance to the landing site is approximately 120 nautical miles (138 statute miles), the TACAN begins the navigation region of interrogating the six navigation stations. As the spacecraft progresses, the distance to the remaining stations is computed. The next-nearest station is automatically selected when the spacecraft is closer to it than to the previous locked-on station. Only one station is interrogated when the distance to the landing site is less than approximately 20 nautical miles (23 statute miles). Again, the TACAN sets will automatically switch from the last locked-on navigation region station to begin searching for the landing site station. TACAN azimuth and range are provided on the horizontal situation indicator. TACAN range and bearing cannot be used to produce a good estimate of the altitude position component, so navigation uses barometric altitude derived from the air data system probes, which are deployed by the fight crew at approximately Mach 3.
If the microwave scan beam landing system is not acquired, TACAN data can be used until an altitude of 1,500 feet. When runways with MSBLS are acquired, MSBLS operation can be automatic. The flight crew is provided with the controls and displays necessary to evaluate MSBLS performance and take over manually if required. The runways with MSBLS must be in the primary or secondary slot in the minitable for the minitable to copy the MSBLS data. The maxitable is an initial-loaded table of 18 runway data sets and MSBLS data for runways and 50 TACAN data sets. In orbital operations, the landing site function provides the capability to transfer data from the maxitable to the minitable.
TACAN data is processed in the TACAN subsystem operating program, which converts range and bearing to units of feet and radians.
TACAN redundancy management consists of processing and mid-value-selecting range and bearing data. The three TACAN sets are compared to determine if a significant difference is detected. When all three TACAN sets are good, redundancy management selects middle values of range and bearing. If one of the two parameters is out of tolerance, the remaining two will average that parameter. If a fault is verified, the SM alert light is illuminated, and a cathode ray tube fault message occurs for the applicable TACAN set.
The TACAN contractor is Hoffman Electronics Corporation, Navigation Communication System Division, El Monte, Calif.
The air data system senses air pressures related to spacecraft movement through the atmosphere to update navigation state vector in altitude; provide guidance in calculating steering and speed brake commands; and update flight control law computations and provide display data for the commander's and pilot's alpha Mach indicators, altitude/vertical velocity indicators and CRTs. The AMIs display essential flight parameters relative to the spacecraft's travel in the air mass, such as angle of attack (alpha), acceleration, Mach/velocity and knots equivalent airspeed. The altitude/vertical velocity indicators display such essential flight parameters as radar altitude, barometric altitude, altitude rate and altitude acceleration.
Each probe is independently deployed by an actuator consisting of two ac motors connected to rotary electromechanical actuator and limit switches. Each probe is controlled by its air data probe switch on panel C3. To deploy the air data probes, the left and right switches are positioned to deploy. The redundant motors for each probe drive the probe to the deployed position. When the probe is fully deployed, limit switches remove electrical power from the motors. Deployment time is 15 seconds for two-motor operation and 30 seconds for single-motor operation. The deploy position deploys the probe without electrical heaters. The deploy/heat position also deploys the air data probes with heaters powered.
The air data probe stow left and right switches on panel C3 are used during ground turnaround operations to stow the respective probe. Positioning the respective switch to enable and positioning the corresponding air data probe switch to stow stows the corresponding air data probe. The air data probe stow inhibit position opens the ac motor circuits, disables the stow and protects microswitches.
Each air data probe has four pressure-port sensors and two temperature sensors. The pressures sensed are static pressure, total pressure, angle-of-attack upper pressure and angle-of-attack lower pressure. The four pressures are sensed at ports on each probe: static pressure at the side, total pressure at the front and angle-of-attack lower near the bottom front. The probe-sensed pressures are connected by a set of pneumatic lines to two air data transducer assemblies. The two temperature sensors are installed on each probe and are wired to an ADTA. The pressures sensed by the left probe are connected by pneumatic tubing to ADTAs 1 and 3. Those sensed by the right probe are connected to ADTAs 2 and 4. Temperatures and sensed pressure from the probes are sent to the same ADTAs.
Within each ADTA, the pressure signals are directed to four transducers, and the temperature signal is directed to a bridge. The pressure transducer analogs are converted to digital data by digital-processor-controlled counters. The temperature signal is converted by an analog-to-digital converter. The digital processor corrects errors, linearizes the pressure data and converts the temperature bridge data to temperatures in degrees centigrade. These data are sent to the digital output device, which converts the signals into serial digital format, and then to the onboard computers to update the navigation state vector. The data are also sent to the commander's and pilot's altitude/vertical velocity indicators, alpha Mach indicators and CRT.
The ADTA SOP uses ADTA data to compute angle of attack, Mach number (M), equivalent airspeed (EAS), true airspeed (TAS), dynamic pressure (q), barometric altitude (h) and altitude rate ( . h ).
The altitude/vertical velocity indicators and alpha Mach indicators are located on panels F6 and F8 for the commander and pilot, respectively. The information to be displayed on the commander's AVVI and AMI is controlled by the commander's air data switch on panel F6, and the pilot's AVVI and AMI data are controlled by the pilot's air data switch on panel F8. When the commander's or pilot's switch is positioned to nav , the AVVIs and AMIs receive information from the navigation attitude processor. When the air data probes are deployed, the commander's and pilot's switches can be positioned to left or right to receive information from the corresponding air data probe.
The AVVIs display altitude deceleration ( alt accel ) in feet per second squared, rate in feet per second, navigation/air data system (nav/ADS) in feet and radar ( rdr ) altitude in feet. The AVVI's altitude acceleration indicator remains on the navigation attitude processor from the IMUs. In addition, the radar altimeters on the AVVIs will not receive information until the orbiter reaches an altitude of 5,000 feet.
The AMIs display angle of attack ( alpha ) in degrees, acceleration (accel) in feet per second squared, Mach number or velocity (M/vel) in feet per second and equivalent airspeed ( EAS ) in knots. The AMI's acceleration indicator remains on the navigation attitude processor from the IMUs.
All but the alpha indicators (a moving drum) and the altitude acceleration indicators (a moving pointer displayed against a fixed line) are moving tapes behind fixed lines. The AMI's angle-of-attack indicator reads from minus 18 to plus 60 degrees, the acceleration indicator from minus 50 to plus 100 feet per second squared, the Mach/velocity indicator from Mach zero to 4 and 4,000 to 27,000 feet per second, and equivalent airspeed from zero to 500 knots. The AVVIs read altitude acceleration from minus 13.3 to plus 13.3 feet per second squared, altitude rate from minus 2,940 to plus 2,940 feet per second, altitude from minus 1,100 to plus 400,000 feet and then changes scale to plus 40 to plus 165 nautical miles (barometric altitude), and radar altitude from zero to plus 9,000 feet.
Failure warning flags are provided for all four scales on the AVVIs and AMIs. The flags appear in the event of a malfunction in the indicator or in received data. In the event of power failure, all four flags appear.
The four computers compare the pressure readings from the four ADTAs for error. If all the pressure readings compare within a specified value, one set of pressure readings from each probe is summed, averaged and sent to the software. If one or more pressure signals of a set of probe pressure readings fail, the failed set's data flow from that ADTA to the averager is interrupted, and the software will receive data from the other ADTA of that probe. If both probe sets fail, the software operates on data from the two ADTAs connected to the other probe. The best total temperature from all four ADTAs is sent to the software. A fault detection will illuminate the air data red caution and warning light on panel F7, the backup caution and warning alarm light, and the master alarm and will also sound the audible tone and generate a fault message on the CRT. A communication fault will illuminate the SM alert light.
The four ADTAs are located in the orbiter crew compartment middeck forward avionics bays and are convection cooled. Each is 4.87 inches high, 21.25 inches long and 4.37 inches wide and weighs 19.2 pounds.
The air data probe sensor contractor is Rosemount Inc., Eden Prairie, Minn. The air data transducer assembly contractor is AirResearch Manufacturing Co., Garrett Corp., Torrance, Calif. The contractor for the air data probe deploy system is Ellanef, Corona, N.Y.
The orbiter is equipped with three independent MSBLS sets, each consisting of a Ku-band receiver/transmitter and decoder. Data computation capabilities determine elevation angle, azimuth angle and orbiter range with respect to the MSBLS ground station. The MSBLS provides highly accurate three-dimensional navigation position information to the orbiter to compute state vector components for steering commands that maintain the orbiter on its proper flight trajectory. The three orbiter Ku-band antennas are located on the upper forward fuselage nose. The three MSBLS and decoder assemblies are located in the crew compartment middeck avionics bays and are convection cooled.
The ground portion of the MSBLS consists of two shelters: an elevation shelter and an azimuth/distance-measuring equipment shelter. The elevation shelter is located near the projected touchdown point, with the azimuth/DME shelter located near the far end of the runway. Both ends of the runway are instrumented to enable landing in either direction.
The MSBLS ground station signals are acquired when the orbiter is close to the landing site and has turned on its final leg. This usually occurs on or near the heading alignment cylinder, about 8 to 12 nautical miles (9 to 13 statute miles) from touchdown at an altitude of approximately 18,000 feet.
Final tracking occurs at the terminal area energy management ''autoland'' interface at approximately 10,000 feet altitude and 8 nautical miles (9 statute miles) from the azimuth/DME station.
The MSBLS angle and range data are used to compute steering commands until the orbiter is over the runway approach threshold, at an altitude of approximately 100 feet. If the autoland system is used, it may be overridden by the commander or pilot at any time using the control stick steering mode.
The commander's and pilot's horizontal situation indicators display the orbiter's position with respect to the runway. Elevation and azimuth are shown relative to a GPC-derived glide slope on a glide slope indicator; course deviation needles and range are displayed on a mileage indicator. When the orbiter is over the runway threshold, the radar altimeter is used to provide elevation (pitch) guidance. Azimuth/DME data are used during the landing rollout.
The three orbiter MSBLS sets operate on a common channel during the landing phase. The MSBLS ground station transmits a DME solicit pulse. The onboard MSBLS receiver responds with a DME interrogation pulse. The ground equipment responds by transmitting a return pulse. A decoder in the onboard MSBLS decodes the pulses to determine range, azimuth and elevation. Range is a function of the elapsed time between interrogation pulse transmission and signal return. Azimuth pulses are returned in pairs. The spacing between the two pulses in a pair identifies the pair as azimuth and indicates which side of the runway the orbiter is on; spacing between pulse pairs defines the angular position from runway centerline. The spacing between the two pulses in a pair identifies the pair as elevation, and the spacing between pulse pairs defines the angular position of the orbiter above the runway.
The elevation beam is 1.3 to 29 degrees high and 25 degrees to the left and right of the runway. The azimuth/DME beam is zero to 23 degrees high and 13.5 degrees to the left and right of the runway.
Each RF assembly routes range, azimuth and elevation information in RF form to its decoder assembly, which processes the information and converts it to digital words for transmission to the onboard GN&C via the multiplexers/demultiplexers for the GPCs.
Elevation, azimuth and range data from the MSBLS are used by the GN&C system from the time of acquisition until the runway approach threshold is reached. After that point, the azimuth and range data are used to control rollout. Altitude data are provided separately by the orbiter's radar altimeter.
Since the azimuth/DME shelters are at the far ends of the runway, the MSBLS can provide useful data until the orbiter is stopped. Azimuth data give position in relation to the runway centerline, while the DME gives the distance from the orbiter to the end of the runway.
Each MSBLS has an on/off power switch on panel O8 and on the channel (frequency) selection thumbwheel on panel O8. Positioning the MLS 1, 2 and 3 switch provides power to the corresponding microwave scan beam landing system. MSBLS 1 receives power from main bus A, MSBLS 2 from main bus B and MSBLS 3 from main bus C. Positioning the channel 1 , 2 and 3 thumbwheels selects the frequency (channel) for the ground station at the selected runway for the corresponding MSBLS.
Redundancy management mid-value-selects azimuth and elevation angles for processing navigation data. The three MSBLS sets are compared to identify any significant differences among them.
When data from all three MSBLS sets are valid, redundancy management selects middle values of three ranges, azimuths and elevations. In the event that only two MSBLS sets are valid, the two ranges, azimuths and elevations are averaged. If only one MSBLS set is valid, its range, azimuth and elevation are passed for display. When a fault is detected, the SM alert light is illuminated, and a CRT fault message is shown.
Each MSBLS decoder assembly is 8.25 inches high, 5 inches wide and 16.16 inches long and weighs 17.5 pounds. The RF assembly is 7 inches high, 3.5 inches wide and 10.25 inches long and weighs 6 pounds.
The MSBLS contractor is Eaton Corp., AIL Division, Farmingdale, N.Y.
The RAs constitute a low-altitude terrain-tracking and altitude-sensing system based on the precise time it takes an electromagnetic energy pulse to travel from the orbiter to the nearest object on the ground below and return during altitude rate changes of as much as 2,000 feet per second. This enables tracking of mountain or cliff sides ahead or alongside the orbiter if these obstacles are nearer than the ground below and warns of rapid changes in absolute altitude.
The two independent RAs consist of a transmitter and receiver antenna. The systems can operate simultaneously without affecting each other. The four C-band antennas are located on the lower forward fuselage. The two receiver/transmitters are located in the middeck forward avionics bays and are convection cooled.
Each RA transmits a C-band (4,300 MHz modulated at 8.5 kHz) pulse through its transmitting antenna. The signal is reflected by the nearest terrain, and the leading edge of the return radar echo is locked on by the RA through its receiving antenna. The altitude outputs by the RA are analog voltages that are proportional to the elapsed time required for the ground pulse to return, which is a function of height or distance to the nearest terrain. The range output of the RA is from zero to 5,000 feet. The RA will not lock on if the orbiter has large pitch or roll angles.
The onboard GPCs process the data for the autoland mode and touchdown guidance after the orbiter has crossed the runway threshold from an altitude of 100 feet down to touchdown. If the autoland mode is not used, the GPCs process the data for display on the commander's and pilot's altitude/vertical velocity meters from 5,000 feet.
The commander and pilot can select RA 1 or 2 for display on their respective AVVI. The commander's radar al tm 1 and 2 switch is located on panel F7, and the pilot's switch is located on panel F8. The radar altimeter on/off 1 and 2 power switches are on panel O8. Positioning radar altimeter 1 to on provides electrical power to RA 1 from main bus A; positioning radar altimeter 2 to on provides electrical power to RA 2 from main bus B.
The display scale on the commander's and pilot's AVVI raw data recorder indicators ranges from 5,000 to zero feet. Altitude is displayed on a moving tape. Above 9,000 feet, the scale will be pegged. At 1,500 feet, the raw data recorder indicator changes scale. The RA off flag will appear if there is a loss of power, loss of lock, data good-bad or after three communications faults.
Each radar altimeter receiver/transmitter measures 3.13 inches high, 7.41 inches long and 3.83 inches wide and weighs 4.5 pounds.
The radar altimeter contractor is Honeywell Inc., Minneapolis, Minn.
The AAs provide feedback to the flight control system concerning acceleration errors, which are used to augment stability during first-stage ascent, aborts and entry; elevon load relief during first-stage ascent; and computation of steering errors for display on the commander's and pilot's attitude director indicators during terminal area energy management and approach and landing phases.
The lateral acceleration readings enable the flight control system to null side forces during both ascent and entry. The normal acceleration readings indicate the need to relieve the load on the wings during ascent. During entry, the normal acceleration measurements cue guidance at the proper time to begin ranging. During the latter stages of entry, these measurements provide feedback for guidance to control sink rate. In contrast, the accelerometers within the IMUs measure three accelerations used in navigation to calculate state vector changes.
Each accelerometer consists of a pendulum suspended so that its base is in a permanent magnetic field between two torquer magnets. A lamp is beamed through an opening in one of the torquer magnets; photodiodes are located on both sides of the other torquer magnet. When acceleration deflects the pendulum toward one photodiode, the resulting light imbalance on the two photodiodes causes a differential voltage, which increases the magnetic field on one of the torquer magnets to return the pendulum to an offset position. The magnitude of the current that is required to accomplish this is proportional to the acceleration. The polarity of the differential voltage depends on the direction of the pendulum's movement, which is opposite to the direction of acceleration. The only difference between the lateral and normal accelerometers is the position in which they are mounted within the assembly. When the acceleration is removed, the pendulum returns to the null position. The maximum output for a lateral accelerometer is plus or minus 1 g; for a normal accelerometer, the maximum output is plus or minus 4 g.
The accelerations transmitted to the forward MDMs are voltages proportional to the sensed acceleration. These accelerations are multiplexed and sent to the GPCs, where an accelerometer assembly subsystem operating program converts the eight accelerometer output voltages to gravitational units. This data is also sent to the CRTs and attitude director indicator error needles during entry.
The accelerometer assemblies provide fail-operational redundancy during both ascent and entry. The four AAs employ a quad mid value software scheme to select the best data for redundancy management and failure detection.
Accelerometer 1 is powered from main bus A through the accel 1 circuit breaker on panel O14. Accelerometer 2 is powered from main bus B through the accel 2 circuit breaker on panel O15. Accelerometer 3 is controlled by the accel 3 on/off switch on panel O16. When the switch is positioned to on, power from control buses controls remote power controllers, which supplies main bus A and main bus C to accelerometer 3. The accel 4 on/off switch on panel O15 operates similarly, except that accelerometer 4 receives power from main bus B and main bus C. The accelerometers are turned off once on orbit and on again before entry.
An RGA/accel red caution and warning light on panel F7 will be illuminated if an accelerometer fails.
The four AAs are located in crew compartment middeck forward avionics bays 1 and 2. The AAs are convection cooled and require a five-minute warm-up period.
The accelerometer contractor is Honeywell Inc., Clearwater, Fla.
The RGAs sense roll rates (about the X axis), pitch rates (about the Y axis) and yaw rates (about the Z axis). These rates are used by the flight control system to augment stability during both ascent and entry.
Each RGA contains three identical single-degree-of-freedom rate gyros so that each gyro senses rotation about one of the vehicle axes. Thus, each RGA contains one gyro-sensing roll rate (about the X axis), one gyro-sensing pitch rate (about the Y axis) and one gyro-sensing yaw rate (about the Z axis).
Each gyro has three axes. A motor forces the gyro to rotate about its spin axis. When the vehicle rotates about the gyro input axis, a torque results in a rotation about the output axis. An electrical voltage proportional to the angular deflection about the output axis-representing vehicle rate about the input axis-is generated and transmitted through the flight aft MDMs to the GPCs and RGA SOP. This same voltage is used within the RGA to generate a counteracting torque that prevents excessive gimbal movement about the output axis. The maximum output for roll rate gyros is plus or minus 40 degrees per second; for the pitch and yaw gyros, the maximum output is plus or minus 20 degrees per second.
The RGA SOP converts the voltage rate into units of degrees per second.
The RGA 1, 2, 3 and 4 on/off power switches are located on panels O14, O15, O16 and O15, respectively. The redundant power supplies for RGAs 1 and 4 prevent the loss of more than one rate gyro assembly if main bus power is lost.
The RGAs remain off on orbit except during flight control system checkout to conserve power.
The RGA/accel red caution and warning light on panel F7 will be illuminated to inform the flight crew of an RGA failure.
The RGAs are located on the aft bulkhead below the floor of the payload bay. They are mounted on cold plates for cooling by the Freon-21 coolant loops. The RGAs require a five-minute warm-up time.
The RGA contractor is Northrop Corp., Electronics Division, Norwood, Mass.
The SRB RGAs sense pitch and yaw rates, but not roll rates, during the first stage of ascent. Because the SRBs are more rigid than the orbiter body, these rates are less vulnerable to errors created by structural bending. They are thus particularly useful in thrust vector control.
The three RGAs in each SRB are mounted on the forward ring within the forward skirt near the SRB-external tank attach point.
The SRB RGA SOP converts the 12 voltages representing a rate into units of degrees per second. These rates are used by the flight control system during first-stage ascent as feedback to identify rate errors, which are used for stability augmentation. The pitch and yaw axes and a combination of rate, attitude and acceleration signals are blended to provide a common signal to the space shuttle main engines and SRB thrust vector control during first stage. In the roll axis, rate and attitude are summed to provide a common signal to the SSMEs and SRB thrust vector control.
Each of the SRB RGAs is hard-wired to a flight aft MDM to the GPCs through flight-critical buses 5, 6 and 7. In the GPCs, the SOP applies the rate compensation equation to each of the left or right pitch and yaw rates. The compensated rate signals are sent to redundancy management, where the mid value software scheme selects the best data for use and failure detection.
The SRB RGAs are commanded to null and switched out of the flight control system two to three seconds before SRB separation; SRB yaw and pitch rate data are then replaced with orbiter pitch and yaw RGA data.
The RGA/accel red caution and warning light on panel F7 will be illuminated to inform the flight crew of an RGA failure.
The RGA contractor is Northrop Corp., Electronics Division, Norwood, Mass.