The IMUs consist of an all-attitude, four-gimbal, inertially stabilized platform. They provide inertial attitude and velocity data to the GN&C; software functions. Navigation software uses the processed IMU velocity and attitude data to propagate the orbiter state vector. Guidance uses the attitude data, along with state vector from the navigation software, to develop steering commands for flight control. Flight control uses the IMU attitude data to convert the steering commands into control surface, engine gimbal (thrust vector control) and reaction control system thruster fire commands.

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, theLaunch 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 ascent, the IMUs provide accelerometer and resolver data to the GN&C; software to propagate the state vector, determine attitude and display flight parameters.

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.

During entry, IMU operation differs only in the manner in which accelerometer data are used by navigation.

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.

The IMU software scheme is designed to select the best data for GPC use and to detect system failures. This scheme is referred to as redundancy management.

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.

Each KT-70 IMU is 10.28 inches high, 11.5 inches wide and 22 inches long and weighs 58 pounds.

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.