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.
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
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
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.
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.
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
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
contractor is Honeywell Inc., Clearwater, Fla.