The flight controls consist of the rudder, ailerons and elevator which are actuated without power assistance via conventional control cables and pulleys. Electrically operated trim tabs are on all flight control axes The flaps are hydraulically actuated. The roll and pitch controls are redundant, physically separate and capable of being divided if a jam of the control path of surface should occur. Pitch and roll trim is also redundant and are electrically operated. Locking mechanisms on the primary controls prevent damage due to strong winds. The trim and flap controls and their related indications are located on the center pedestal and center instrument panel. To prevent over stressing of the aircraft structure at higher airspeeds, a rudder limiter system is installed which reduces the maximum available rudder angle to 15° each way above 150 kt and further restricts it to 5.7° each way above 200 kt. The rudder limiter stops are withdrawn automatically as the speed reduces allowing greater rudder travel.
Each aileron is controlled by an independent control system which is joined to the opposite system by the aileron interconnect unit. The autopilot servo is connected to the right aileron. A geared servo tab on each aileron is mechanically deflected as the aileron moves in order to reduce the aerodynamic forces associated with aileron deflection. For aileron trimming, the geared servo tabs are deflected by electric actuators. Normal aileron trimming is accomplished by the left tab while standby trimming is accomplished by the right tab.
The left geared servo tab actuator is supplied from the Right Battery Bus and controlled by the normal ROLL trim switch on the trim control panel. The right geared servo tab actuator is controlled by the STBY ROLL switch and is supplied from the Left Battery Bus for left trimming and from the Right Battery Bus for right trimming. Both elements of either switch must be operated simultaneously to activate the respective actuator.
Aileron Interconnect Unit
The independent aileron control systems are joined by an aileron interconnect unit which permits separation of the systems if jamming occurs. After disconnection, the pilot with the operable aileron controls roll attitude. A centering spring unit in each control system minimizes aileron rise during flight while the ailerons are disconnected. The ailerons are disconnected by either of the following methods:
• Manually: the pilots develop sufficient roll force on the control columns to overpower the friction unit. If jamming occurred in the aileron system, one control wheel resists the full range of rotation. The operable control wheel can be forced to rotate against a friction device to allow full deflection of the related aileron. Realigning the operable control wheel with the
jammed control wheel automatically resets the interconnect mechanism thereby re-engaging the jam.
• Electrically: the ailerons are electrically disconnected via power from the Left Battery Bus by pulling the ROLL disconnect handle located on the center pedestal. This method separates the aileron systems for the duration of the flight and roll control response is considerably diminished. System re-connection is accomplished only by Maintenance personnel on the ground.
Each elevator is controlled by an independent control system which is joined to the other system by the elevator interconnect unit. The left control system also integrates the autopilot and stickpusher systems. A mechanically operated geared servo tab on each elevator automatically reduces aerodynamic forces as the elevator is deflected. For elevator trimming, the servo tabs are uniformly deflected by separate electric actuators.
An airflow splitter mounted on the fuselage at the inboard leading edge of each horizontal stabilizer ensures proper airflow during all regimes of flight. Vortex generators located on the lower side of the right horizontal stabilizer ensure proper airflow over the stabilizer and elevator during flight in icing conditions when the flaps are fully extended. Normal pitch trim control is accomplished by the trim switches on either control wheel. When using normal pitch trim control, the left geared servo tab is directly controlled while operation of the right geared servo tab is actuated by the pitch trim synchronizer. If the pitch trim synchronizer is deactivated, trim tab synchronization is lost and the trim actuators are individually operated. Both trim actuators are supplied from the Right Battery Bus. Standby pitch trim is controlled by the STBY PITCH trim switch located on the center pedestal. When using the standby system, only the right geared servo tab is operated. The trim actuator is supplied from the Right Battery Bus with automatic backup from the Left Battery Bus. Use of the standby system automatically disengages the pitch trim synchronizer. Initial use of the STBY PITCH trim switch causes master caution alerting with flashing amber PITCH TRIM on the CWP. Subsequent use does not cause alerting.
The pitch trim synchronizer is reset by depressing the adjacent PITCH RESET pushbutton. The trim tabs must be set within one unit of each other prior to resetting the system. Both elements of either pitch trim switch must be operated simultaneously to activate the respective trim actuator. The Captain’s pitch trim switches override the first officer’s switches. The autopilot utilizes the normal trim system and activation of any pitch trim switch causes autopilot disengagement.
Elevator Interconnect Unit
The independent elevator control systems are joined by an elevator interconnect unit which permits separation of the systems if elevator jamming occurs. After disconnection, the pilot with the operable elevator controls pitch attitude The elevators are disconnected by either of the following methods:
• Manually: the pilots develop sufficient pitch force on the control columns to overpower the friction unit. If jamming occurs in the elevator system, one control column resists the full range of movement. The operable control column can be forced to move against a friction device to allow full deflection of the related elevator. Realigning the operable control column with the jammed column automatically resets the interconnect mechanism thereby re-engaging the jam
• Electrically: the elevators are electrically disconnected via supply from the Left Battery Bus by pulling the PITCH disconnect handle located on the center pedestal. This method separates the elevator systems for the duration of the flight and pitch control response is considerably diminished. System re-connection is accomplished only by Maintenance
Elevator Trim Monitoring
Both trim actuators are monitored by the warning annunciator system for trim settings outside the green takeoff range displayed on the pitch trim indicator. Master warning alerting with an intermittent horn are triggered with flashing red CONFIG on the CWP if takeoff is attempted with the pitch trim set outside of the takeoff range.
Pitch Trim Synchronizer
Failure of the pitch trim synchronizer results in master caution alerting and flashing amber PITCH TRIM on the CWP.
Stall Warning and Stickpusher Operation
Stall warning is provided by the following simultaneous indications:
• Aural clacker alert is generated by the WAS
• The stickshaker devices mounted on each control column activate
At the stall threshold, the stickpusher activates:
• A brisk, unmistakable forward movement of both control columns is made (approximately 80 pounds force)
• Both amber PUSH 1 and PUSH 2 lights illuminate on the each pilot instrument panel for as long as the stickpusher is active
Stall warning and stickpusher activation are controlled by two independent stall warning/ident computers supplied from the respective Battery Bus. Each stall warning/ident computer monitors flap position, operation of the wing deice system, and angle-of-attack signals received from the AOA probes. Either computer can independently activate stall warnings. However, to prevent unwarranted stickpusher activation the system requires that both computers:
• Detect excessive AOA
• Participate in activation of the electric motor and clutch comprising the stickpusher actuator
The Stall Warning System is designed to give an artificial stall warning (stick shaker and aural warning) and subsequent pusher at a preset angle of attack, for an ice free wing, depending on flap position and if the DE-ICE BOOTS are cycling. The stick shaker is activated at approximately 6-8 kt before stall (stick pusher activation). With ice on the wing stall might be encountered before, or at, stick pusher activation. In some adverse cases stall may even be encountered before the artificial stall warning is activated. Stall warnings and stickpusher activation are inhibited while on the ground, and the stickpusher is further inhibited for seven seconds after takeoff and anytime the load factor is less than .5g. The stickpusher actuator is supplied by the Right Battery Bus and installed on the left elevator control system. It transmits movement to the right elevator system through the elevator interconnect unit. When activated, the stickpusher actuator causes the elevators to move to four degrees in the nose-down direction which is sustained as long as AOA is excessive or until less than .5g load factor is encountered. Improper stickpusher operation causes master caution alerting with amber PUSHER SYSTEM illumination on the CWP. The stickpusher system is disabled by depressing either PUSHER DISARM switch located on each side panel. When depressed, stickpusher operation is disabled while full stall warning protection continues, and can be reset by maintenance only on the ground. Amber PUSHER SYSTEM may illuminate on the CWP during normal stickpusher operation without master caution alerting. Faults detected in a stall warning/ident computer result in master caution alerting with the associated flashing amber STALL FAIL on the CWP. The stall warning/ident computers are tested via the STALL switches on the overhead TEST 1 panel.
The fore-aft position of each set of rudder pedals is adjustable by a lever located forward of the control column below the instrument panel. The rudder pedals move AFT under spring tension when the lever is pulled, and the pilot pushes the pedals to the desired position. Rudder deflection is accomplished by rudder pedal or autopilot inputs acting directly on a servo tab and, through an opposing spring, the rudder itself. A damper installed between the rudder and servo tab minimizes tab oscillations caused by airflow disturbances. Vortex generators installed on each side of the vertical stabilizer ensure laminar airflow over the stabilizer and rudder during high speed flight. Rudder trimming is accomplished by an electric actuator which moves the servo tab. The trim actuator is supplied from the Right Battery Bus and controlled by the YAW trim switch. Both elements of the YAW trim switch must be operated simultaneously to activate the trim actuator.
Rudder Limiter System
The rudder limiter system limits rudder travel at high speeds in order to prevent rudder/stabilizer overload. The system is supplied from the Left Avionics Start Bus and controlled by the RUDDER LIMIT switch. The rudder limiter mechanism consists of an electric actuator which blocks rudder travel according to indicated airspeed. The system monitors airspeed obtained from the air data computer and the standby airspeed indicator and restricts rudder travel as follows:
• Full rudder travel (to 30°) is permitted at speeds up to 150 knots
• Intermediate travel (to 15°) is permitted at speeds between 150 to 200 knots
• Minimum travel (to 5.7°) is permitted at speeds above 200 knots
Discrepancies between the monitored airspeeds or an improperly positioned rudder limiter result in master caution alerting with flashing amber RUDDER LIMIT on the CWP. Selecting the RUDDER LIMIT switch to OVRD enables full rudder travel by retracting the rudder limiting mechanism. If an airspeed vs. rudder limiter position discrepancy caused the alert, then the RUDDER LIMIT light remains illuminated until airspeed is reduced to a value appropriate for the actuator position, as listed above. The override function is supplied from the Right Battery Bus.
The trim indicating panel, supplied from the Right Battery Bus, displays all trim tab positions. The left aileron tab position is shown on the MAIN side of the ROLL trim scale and right aileron tab position is shown on the STBY scale. Elevator trim tab position is displayed on the corresponding PITCH trim scale. The rudder servo tab position is displayed on the YAW trim scale.
The aircraft is equipped with a single-panel Fowler type flaps. Each flap is operated by a single hydraulic actuator which incorporates fail-safe design to prevent flap retraction if hydraulic pressure fails, but will permit blow-back during excessive airspeeds in order to prevent flap damage. Flap asymmetry is prevented by mechanical interconnection between the flap panels.
Flap Control and Indications
Flap position is controlled by the FLAP handle which can be set in five detented positions: 0°, 7°, 15°, 20°, and 35°. The FLAP handle operates an electrical switch and must be lifted from each detent in order to make another selection. For flap retraction, there are gates at the 20° and 7° positions which require the handle be pushed into the detent before the handle movement can continue forward. Each flap detent position transmits a control signal to the flap control unit. The flap control unit, supplied from the Left Battery Bus, monitors FLAP handle selection and the position of the left flap panel. An appropriate control signal (extend or retract) is directed to the flap control valve when disagreement occurs between the two inputs. Flap control system unit failure or loss of electrical supply results in closure of the flap control valve and the flaps subsequently become hydraulically locked in position without further control. Electrical faults in the flap control system or sustained disagreement between flap handle selection and left flap position triggers master caution alerting with flashing amber FLAPS on the CWP. Each flap panel is monitored by a separate position transmitter which supplies signals to the FLAPS indicator via the flap control box. The indicator is supplied from the Right Battery Bus and incorporates L and R pointers. Normally, only the L pointer is visible since both flaps are at the same position.
The gust lock is controlled by the GUST LOCK handle located on the center pedestal. When pulled AFT, the elevators and ailerons are mechanically locked while the rudder is electrically locked by an actuator supplied from the Left Essential Bus. Before the GUST LOCK lever can be moved, the GUST LOCK RELEASE knob located on the center pedestal must be actuated in order to release the handle. The control columns must be pushed forward and the ailerons and rudder must be centered in order to engage the gust locks. When the GUST LOCK handle is engaged, power lever movement is limited to prevent takeoff with the controls locked. All gust locks fail to the disengaged position if a disconnection occurs in the gust lock control system. Master caution alerting is triggered with flashing amber GUST LOCK on the CWP if the rudder gust lock remains engaged after the GUST LOCK handle is released.