TM 55-1510-219-10 LANDING    GEAR    WARN,    located    on    the    overhead circuit breaker panel (fig.  2-26). f. Landing Gear Warning Horn Test Switch. The landing gear warning horn may be tested by the test switch  on  the  right  subpanel.    The  switch,  placarded STALL  WARN  TEST  OFF  LDG  GEAR  WARN  TEST, will sound the landing gear warning horn and illuminate the landing gear position warning lights when moved to the momentary LDG GEAR WARN TEST position.  The circuit    is    protected    by    a    5-ampere    circuit    breaker, placarded  LANDING  GEAR  WARN,  on  the  overhead circuit breaker panel (fig.  2-26). g. Landing Gear Safety Switches.  A safety switch on   each   main   landing   gear   shock   strut   controls   the operation  of  various  aircraft  systems  that  function  only during  flight  or  only  during  ground  operation.    These switches are mechanically actuated whenever the main landing  gear  shock  struts  are  extended  (normally  after takeoff),  or  compressed  (normally  after  landing).    The safety   switch   on   the   right   main   landing   gear   strut activates     the     landing     gear     control     circuits,     cabin pressurization  circuits  and  the  flight  hour  meter  when the strut is extended.  This switch also activates a down- lock hook, preventing the landing gear from being raised while  the  aircraft  is  on  the  ground.    The  hook,  which unlocks   automatically   after   takeoff,   can   be   manually over-ridden    by    pressing    down    on    the    red    button, placarded    DN    LOCK    REL    located    adjacent    to    the landing  gear  switch.    If  the  over-ride  is  used  and  the landing   gear   control   switch   is   raised,   power   will   be supplied  to  the  warning  horn  circuit  and  the  horn  will sound.  The safety switch on the left main landing gear strut   activates   the   left   and   right   engine   ambient   air shutoff valves when the strut is extended. CAUTION Continued   pumping   of   handle   after GEAR  DOWN  position  indicator  lights (3)  are  illuminated  could  damage  the drive mechanism, and prevent subsequent gear retraction. h. Manual Landing Gear Extension System. Manual  landing  gear  extension  is  provided  through  a manually powered system as a backup to the electrically operated system.  Before manually extending the gear, make certain that the landing gear switch is in the down position    with    the    LANDING    GEAR    RELAY    circuit breaker     pulled.          During     a     manual     landing     gear extension,  the  landing  gear  motor  must  be  disengaged from the landing gear drive mech- anism.  This is accomplished through use of an alternate engage   handle   located   adjacent   to   the   landing   gear alternate  extension  handle.    To  disengage  the  landing gear  motor,  pull  the  alternate  extension  handle  up  and turn   it   clockwise.      When   this   handle   is   pulled,   the landing gear motor is disconnected from the system and the alternate drive system is locked to the gearbox and motor.    With  the  alternate  drive  locked  in,  the  landing gear    may    be    manually    extended    by    pumping    the alternate extension handle until the three GEAR DOWN position    indicator    lamps    are    illuminated.        Refer    to Chapter 9 for additional information on emergency gear extension procedures. i. Tires.  The aircraft is equipped with dual 22 x 6. 75  x  10,  8  ply  rated,  tubeless,  rim  inflation  tires  on  the main gear.  The nose gear is equipped with a single 22 x 6.  75 x 10, 8 ply rated, tubeless, rim inflated tire. j. Steerable   Nose   Wheel.      The   aircraft   can   be maneuvered on the ground by the steerable nose wheel system.  Direct linkage from the rudder pedals (fig.  2-9) to the nose wheel steering linkage allows the nose wheel to be turned 12°  to the left of center or 14°  to the right. When  rudder  pedal  steering  is  augmented  by  the  main wheel  braking  action,  the  nose  wheel  can  be  deflected up to 48°  either side of center.  Shock loads which would normally    be    transmitted    to    the    rudder    pedals    are absorbed by a spring mechanism in the steering linkage. Retraction of the landing gear automatically centers the nose  wheel  and  disengages  the  steering  linkage  from the rudder pedals. k. Wheel   Brake   System.      The   main   wheels   are equipped with multiple-disc hydraulic brakes actuated by master  cylinders  attached  to  the  rudder  pedals  at  the pilot's  and  copilot's  position.    Braking  is  permitted  from either set of rudder pedals.  Brake fluid is supplied to the system   from   the   reservoir   in   the   nose   compartment. The    toe    brake    sections    of    the    rudder    pedals    are connected   to   the   master   cylinders   which   actuate   the system  for  the  or  responding  wheels.    No  emergency brake system is provided. WARNING Repeated and excessive application of brakes,     without     allowing     sufficient cooling time between applications, will cause  loss  of  braking  efficiency,  and may  cause  brake  or  wheel  failure,  tire blowout,     or     destruction     of     wheel assembly by fire. 2-14