There are four accelerometer assemblies aboard the orbiter, each containing two identical single-axis accelerometers, one of which senses vehicle acceleration along the lateral (left and right) vehicle Y axis while the other senses vehicle acceleration along the vertical (normal, yaw and pitch) Z axis.

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.