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from DADC #2.
Most subsystems with dual installations normally source DADC data
according to a standard coupling arrangement: the left (L), #1 or pilot side
sourced to the #1 DADC, and the right (R), #2 or copilot side connected to
OPERATING MANUAL
2A-34-00 PRODUCTION AIRCRAFT SYSTEMS
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October 11/01
Revision 5
the #2 DADC. For some, but not all installations, the data source is
selectable to provide redundancy in case of DADC failure. Transponders
(ATC) #1 and #2 are normally sourced to their respective DADC, but may
be selected to the alternate DADC. Angle of Attack (AOA) indicators are
referenced to the DADC selected to the respective PFD. EPR sensors are
normally paired with engine #1 to DADC #1, engine #2 to DADC #2, but
failure of a DADC will prompt an automatic switch to the remaining DADC
for EPR computation. The DADC source for cabin pressure control is
selectable on the cockpit overhead panel.
DADC Failure Modes (Flagged, Unflagged)
(1) A “flagged” DADC failure is one where the failure is readily apparent,
because of the blue DADC 1 (or 2) FAIL advisory CAS message,
and red “X’s” through all four air data scales (airspeed, altitude,
AOA, and vertical speed) of the PFD using the failed DADC as its air
data source, as selected on the display controllers. Other
confirmation of failure is as follows:
• Transponder indications
• AOA indexer failure
• Automatic cabin pressurization control problems and faulty
guidance panel indications (if operating on the failed DADC)
• EICAS message indicating that EPR is receiving pressure
information from the opposite DADC (EPR 1 - DADC 2, or
EPR 2 -DADC 1)
The solution is to select the opposite (good) DADC as the air data
source, on display controllers, guidance panel, cabin pressure
control panel, etc.
OPERATING MANUAL
PRODUCTION AIRCRAFT SYSTEMS 2A-34-00
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October 11/01
Revision 5
(2) An “unflagged” DADC failure will produce a blue DADC
MISCOMPARE advisory CAS message, and the failure may not be
readily apparent. The autopilot and yaw damper will disconnect and
will not re-engage until the faulty DADC has been identified and
isolated by pulling its circuit breaker. Pitch trim will remain operative.
The flight crew may be able to identify the faulty DADC by looking
for an amber IAS and / or ALT comparator warning annunciation to
left of the horizon in each PFD. The IAS indication means a split of
20 or more knots exists between air data systems; the ALT
indication means a split of 200 feet or more exists between systems.
To determine which system is correct requires reference to an
independent data source, in this case standby flight instrumentation.
Since the standby flight instruments show large errors because they
are uncorrected for static source error, it is recommended that
standby altimeter be set so as to read the same as the cruising flight
level. Once stable cruise speed is attained, the settable airspeed
bug should be aligned with the standby airspeed pointer. Thus,
reference can be made to the standby altimeter and airspeed
indications, as “voters” in helping to determine which DADC outputs
are more nearly correct. Then, check the other DADC outputs, the
pressurization system, AOA indexers and transponders, for
indications of faulty operation. If observation leads to a
determination of which DADC is faulty, select the “good” DADC to
both PFDs, guidance panel, transponders, and the cabin
pressurization system. Then isolate the faulty DADC by pulling its
circuit breaker, and after at least a one minute wait, re-engage the
autopilot.
3. Controls and Indications:
(See Figure 1.)
NOTE:
A description of the SPZ-8000 (or SPZ-8400) Digital
Automatic Flight Control System (DAFCS) can be
found in Honeywell’s SPZ-8000 (or SPZ-8400) Digital
Automatic Flight Control System Pilot’s Manual for the
Gulfstream IV. A description of the Integrated
Automatic Tuning System (Collins RTU-4200 Series
Radio Tuning Unit (RTU)) can be found in Section
2A-23-40, Integrated Automatic Tuning System and
Collins’ RTU-4200 Series Pilot’s Guide.
A. Circuit Breakers (CBs):
Circuit Breaker Name CB Panel Location Power Source
TOTAL TEMP PROBE HTR CP L-10 MAIN 115V AC φB
TOTAL TEMP VALVE CP M-10 R MAIN 28V DC
L PITOT HT PWR CP L-11 ESS 115V AC φA
R PITOT HT PWR CP M-11 MAIN 115V AC φA
L PITOT HT CONT CP L-12 ESS 28V DC
R PITOT HT CONT CP M-12 MAIN 28V DC
OPERATING MANUAL
2A-34-00 PRODUCTION AIRCRAFT SYSTEMS
　
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