Hydraulic shutoff valves installed at the reser-
voir in each engine-driven pump supply line
can be closed from the cockpit in the event of
fire or when maintenance is to be performed.
An accumulator precharged with dry air or
nitrogen dampens pressure surges and helps
maintain system pressure. An indicator on the
copilot’s instrument panel displays system
pressure. An optional amber annunciator light
warns of low pressure.
There are two filters in the system—one in the
pressure line, and one in the return line.
Systems with reservoirs pressurized by an as-
pirator incorporate an air filter.
A pressure regulator on airplanes SNs 24-296 and
prior, and 25-180 and prior maintains system
pressure at 1,500 psi. On subsequent airplanes,
pressure is regulated by the engine-driven pumps.
A system relief valve set to relieve at 1,700 psi
prevents system damage by relieving excessive
pressure into the return line.
The system accumulator and reservoir are
l o c a t e d i n t h e t a i l c o n e . A c c u m u l a t o r a i r
precharge, indicated by a gage on the accu-
mulator, should be 850 psi when hydraulic
pressure is zero. A second, bladder-type ac-
cumulator precharged to 600 psi and connected
to the hydraulic system through a one-way
check valve, is located in the tail compartment
on airplanes with thrust reversers.
HYDRAULIC SYSTEM
OPERATION
The engine-driven pumps are supplied with
fluid from the reservoir through hydraulic
shutoff valves (Figures 13-1 and 13-2). The
valves are DC motor-driven and controlled
by the guarded FIRE switches on airplanes
w i t h g l a r e s h i e l d w a r n i n g l i g h t s , a n d b y
guarded FIRE WALL SHUTOFF switches on
all other airplanes.
After starting the first engine, the HYDRAULIC
PRESSURE indicator should be checked to
verify engine-driven pump operation.
When the second engine is started, there is no
change in pressure indication, but output is
doubled. There is no positive indication that
the second pump is operating properly; there-
fore, after landing, shut down the engine
started first and actuate a hydraulic subsystem.
If pressure drops slightly then returns to nor-
mal, the second pump is functioning properly.
If an engine-driven pump fails in flight, the
other engine-driven pump is capable of meet-
ing system demands.
The electrically driven auxiliary hydraulic
pump, controlled by a switch on the instrument
panel and a pressure switch (Figures 13-1 and
13-2), draws fluid from the bottom of the
reservoir. With the HYD PUMP switch in the
on (HYD PUMP) position, the auxiliary pump
automatically cycles between 1,200 and 1,400
psi (approximately). It may be used to pres-
surize the system in the event of a main hy-
d r a u l i c s y s t e m f a i l u r e , o r f o r o p e r a t i n g
subsystems when the engines are not running.
Output is .5 gpm; therefore, a slower-than-
normal flap extension should be anticipated
subsequent to a hydraulic system failure.
Loss of fluid due to a system leak is the most
probable cause of complete loss of hydraulic
pressure. If installed, the amber LO HYD
PRESS light will illuminate as pressure de-
creases below 1,200 psi. Do not operate the
auxiliary pump until emergency landing gear
extension procedures have been accomplished,
as directed by the approved AFM. Otherwise,
the auxiliary pump may discharge the .4 gal-
lon of reserve fluid through the same leak.
Circuit protection is provided by a circuit
breaker on SNs 23-003 through 24-129 and by
a current limiter on all subsequent airplanes.
13-2
FOR TRAINING PURPOSES ONLY
LEARJET 20 SERIES PILOT TRAINING MANUAL
FlightSafety
international