CESSNA 550 SII RECURRENT TRAINING 1

Douglas S. Carmody, SafePilot Publishing, LLC. and Executive Flight Training, LLC. are not liable for the accuracy, effectiveness or safe use of this Online Course and do not warrant that this aircraft publication contains current information and/or revisions. Aircraft manuals and publications required for any reason other than training, study or research purposes should be obtained from the original equipment manufacturer. Reference herein to any specific commercial products by trade name, trademark, manufacturer, or otherwise, is not meant to imply or suggest any endorsement by, or affiliation with that manufacturer or supplier. All trade names, trademarks and manufacturer names are the property of their respective owners. No part of this Course may be copied without the expressed written permission of Douglas S. Carmody. All rights reserved.

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CITATION SII LIMITATIONS 3

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WEIGHT LIMITATIONS Maximum Ramp Weight . . . . . . . . . . . . . . . 15,300 Pounds Maximum Takeoff Weight . . . . . . . . . . . . . . 15.100 Pounds Maximum Landing Weight . . . . . . . . . . . . . 14,400 Pounds Maximum Zero Fuel Weight . . . . . . . . . . . . . 11,200 Pounds This airplane is certificated in the transport category of FAR Part 25.



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ENGINE OPERATING LIMITS OPERATING CONDITIONS THRUST SETTING

OPERATING LIMITS

TIME LIMIT (MINUTES)

OIL PRESSURE PSIG (Note 2)

ITT Temp °C

N2% RPM

N1% FAN RPM

5

710

97

106 Note 4

70 – 85 (Note 3)

10 - 121°C

MAX CONTINUOUS

CONTINUOUS

690

97

106 Note 4

70 – 85

0 - 121

MAX CRUISE

CONTINUOUS

690

97

106 Note 4

70-85

0-121

IDLE

CONTINUOUS

580

49 Note 5

40 (Min)

-40 - 121

STARTING

-

Note 1

-

-

-40 (Min)

TRANSIENT

-

710

97

TAKEOFF

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OIL TEMPERATURE

(Note 3)

0 - 121

NOTE 1

The maximum start limit is 700°C for 2 seconds;

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The maximum transient oil pressure can be 95 PSIG for 90

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Normal oil pressures are 70 to 85 PSIG above 60% TURBINE

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Refer to the appropriate thrust setting charts for % FAN RPM setting.

RPM. Oil pressures below 70 PSIG are undesirable and should be tolerated only for the completion of the flight, preferably at reduced power setting. Oil pressures below 35 PSIG are unsafe and require

seconds.

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With ignition ON, idle turbine rpm is 49 ±0.5%. A minimum decrease of 0.5% will be noted with ignition OFF.

that either the engine be shut down or a landing be made as soon as possible, using the minimum power required to sustain flight.



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STARTER CYCLE LIMITATIONS

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BATTERY CYCLE LIMITATIONS Battery limitation - Three engine starts per hour. Three generator assisted cross starts are equal to one battery start. If an external power unit is used for start, no battery cycle is counted. Use of an external power source with voltage in excess of 28 volts DC or current in excess of 1000 amps may damage the starter.

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PROLONGED GROUND OPERATION Continuous ground operation of the engine at takeoff thrust is limited to 5 minutes with ambient temperatures not exceeding 39°C above ISA.

Continuous ground operation of the starter-generator above 225 amperes is prohibited. Limit ground operation of pitot/ static heat to two minutes to prevent damage to the pitot static tubes and the angle-of-attack probe 9

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GROUND OPERATION To assure accurate fan speed thrust indications, inspect the fan for damage prior to each flight.

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WINDSHIELD ICE PROTECTION FLUID Use TT-I-735 isopropyl alcohol for windshield anti-ice system.

HYDRAULIC FLUID Use Skydrol 500A, B, B-4, C, or LD-4; or Hyjet, Hyjet W, III, IV, IVA or IVA Plus only.

APPROVED OILS EXXON TURBO OIL 2380

AERO SHELL TURBINE OIL 560

CASTROL 5000

AERO SHELL TURBINE OIL 500

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FUEL LIMITATIONS Anti-icing additive must be added to all approved fuels unless premixed. Boost Pumps - ON; when low fuel lights illuminate or at 169 pounds or less indicated fuel. The following fuels are approved for use: COMMERCIAL JET FUEL JET A, JET A-1, JET B, JP-4, JP-5 and JP-8

AVIATION GASOLINE Permitted for a maximum of 50 hours or 3500 gallons between overhauls.

Fuel remaining in the fuel tanks when the fuel quantity indicator reads zero is not usable 12

FUEL LIMITATIONS Maximum Fuel imbalance for Normal Operations . . . 200 Pounds Emergency Asymmetrical Fuel Differential . . . . . . . . . 600 Pounds Useable Fuel Capacity . . . . . . . . . . . . . . . . . . . . . . . .5818 Pounds NOTE Flight characteristics requirements were not demonstrated with unbalanced fuel above 200 pounds. A lateral fuel imbalance of 600 pounds has been demonstrated for emergency return to the airport only. 13

The Cessna Citation S/II is approved for ight into known icing conditions. The engine anti-ice system must be on during all ground and ight operations when icing conditions exist or are anticipated. NOTE Icing conditions exist on the ground or in ight when operating in visible moisture and the indicated OAT is +10°C and below. The Fluid Minimum Quantity chart must be used to determine the minimum uid required for takeoff.

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ICE AND RAIN PROTECTION LIMITATIONS

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NOTE This chart is a copy of one in the AFM and is not kept current. When guring minimum uid quantity, use the AFM Fluid Minimum Quantity chart prior 15 to operating in icing conditions.

TAKEOFF AND LANDING OPERATIONAL LIMITS Maximum Altitude Limit. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14,000 Feet Maximum Tailwind Component . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10 Knots Maximum Water/Slush on Runway . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.5 Inches Maximum Ambient Temperature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ISA +39°C Minimum Ambient Temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -54°C Maximum Tire Ground Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .165 Knots

The autopilot and yaw damper must be OFF for takeoff and landing.

Crosswind Component Limitations Without thrust reversers (demonstrated).................... 29 knots (not limiting) With thrust reversers (demonstrated)......................... 25 knots (not limiting) 16

STALL LIMITATIONS

• Aerobatic maneuvers and spins are prohibited. • No intentional stalls permitted above 25,000 feet or at engine speed between 61.0 and 65.0% N1. 17

TAKEOFF AND LANDING OPERATIONAL LIMITS

If taking off after completing a landing or an aborted takeoff that required applying brakes above 50 knots, refer to minimum turnaround time chart.

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TAKEOFF AND LANDING OPERATIONAL LIMITS Do not takeoff with ice, snow or slush adhering to the following critical areas: Wing Leading Edge and Upper Wing Surface Flight Control Surfaces including all hinge gaps Upper surface of nose forward of the windshield Horizontal Stabilizer

Top of Fuselage

Vertical Stabilizer

Windshield

Engine Inlets

All Static Ports

Top of Engine Pylons

Angle of Attack Vanes 19

TAKEOFF AND LANDING OPERATIONAL LIMITS A visual and tactile (hand on surface) check of the wing leading edge and wing upper surface must be performed to ensure the wing is free from frost, ice, snow, or slush when the outside air temperature is less than 10°C (50°F) or if it cannot be determined that the wing fuel temperature is above 0°C (32°F) and any of the following conditions exist: There is visible moisture present (rain, drizzle, sleet, snow, fog, etc.); or Water is present on the wing upper surface; or The difference between dew point and OAT is 3°C (5°F) or less or The atmospheric conditions have been conducive to frost formation.

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APPROACH AND LANDING IN ICING CONDITIONS When any residual ice is present or can be expected during approach and landing, VREF and VAPP must be increased. VREF and VAPP, the landing distance, and the maximum landing weight permitted by climb requirements or brake energy must be determined per by utilizing Approach and Landing Performance Tables.

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ENROUTE OPERATIONAL LIMITS Maximum Operating Altitude . . . . . . . 43,000 Feet Maximum Generator Load . . . . . . . . . 300 Amperes

CABIN PRESSURIZATION LIMITATIONS Normal Cabin Pressurization Limitations Differential . . . . . . . 0.0 to 8.8 PSI, +0.1 or -0.1 PSI

MANEUVERS No acrobatic maneuvers, including spins, are approved. No intentional stalls permitted above 25,000 feet. 22

ANGLE-OF-ATTACK SYSTEM The angle-of-attack and stall warning system must be operable with a satisfactory preflight test completed. The angle-of-attack system may be used as a reference system but does not replace the airspeed indicator as a primary instrument. The angle-of-attack system can be used as a reference for approach speed (1.3 VS1) at all airplane weights and center-of-gravity locations at zero, takeoff, takeoff/ approach and landing flap positions.

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AUTOPILOT One pilot must remain seated with the seat belt fastened during all autopilot operations. The SPZ-500 autopilot and flight director is a complete automatic flight control system. The TEST EACH FLT button actives a test of the torque monitor and must be checked prior to each flight.

Minimum use height: 1000 Feet AGL - Enroute 300 Feet AGL - Non-precision Approach 180 Feet AGL - Category I ILS Approach 24

THRUST REVERSERS Reverse thrust power must be reduced to the idle reverse detent position at 60 KIAS on landing roll. Deploying the T/R’s at higher than normal landing speeds may cause “porpoising” and lead to nose wheel damage. Maximum allowable thrust reverser deployed time is 15 minutes in any one hour period. Engine static ground operation is limited to idle power (if thrust reversers are deployed). Use of thrust reversers is prohibited during touch and go landings. The thrust reversers must be verified to be operational by the Before Takeoff test. 25

FUEL SYSTEM The fuel system consists of a single fuel tank feeding the right engine and a single tank feeding the left engine. No fuel management is required in normal operation of the airplane. If necessary to balance the fuel load due to asymmetric fueling, both engines may be operated from one tank or, for single-engine operation, the operating engine may be fed from either tank. When selecting crossfeed, allow sufficient time for the in-transit light to illuminate, prior to reselecting OFF, or the opposite tank. If the airplane is parked on a slope, be sure fuel is not being lost through the fuel vents. 26

LOW FUEL LEVEL WARNING SYSTEM The low fuel level warning system provides a visual warning to the pilot when a minimum of 169 pounds of usable fuel remains in either fuel tank. The system consists of an electromagnetic float switch in each fuel tank and left and right FUEL LOW LEVEL lights. These lights are tested by the annunciator panel test switch and dimmed by the same control as the annunciator panel. A minimum usable fuel quantity between 169-219 pounds will cause an amber FUEL LOW LEVEL light to illuminate, indicating left or right tank low fuel level.

LOW FUEL LEVEL This message indicates the fuel quantity is at or below 169 pounds usable in the left and/or right tanks as determined by a float switch. 27

FUEL SYSTEM QUICK REVIEW

3. PUMPS: • Electric boost ........................... 17 to 26 psi

1. TOTAL CAPACITY: • 862 U.S. gallons (approximately 5,818 pounds).

• Primary ejector........................... 8 to 22 psi

2. BOOST PUMP SWITCHES:

• Transfer ejector pumps (3)

• ON—Boost pump receives power continuously. • OFF—Automatic boost pump is activated for start and crossfeed only. • NORM—Automatic boost pump is activated for start, crossfeed, and low fuel pressure. • Transfer rate during crossfeed is approximately 900 pounds per hour. • Low fuel pressure light illuminates at a decreasing pressure of 5 psi; it goes out at an increasing pressure of 7 psi. • Low fuel level light illuminates at 185 pounds of usable fuel remaining; input is from the oat switch.

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OXYGEN SYSTEM Oxygen for flight crew and passengers is supplied from a 22 or 64-cubic foot oxygen cylinder. On UN 0002-0254 the cylinder is located in the nose. After 0254, the cylinder is in the tail cone area. The oxygen cylinder pressure gage is located on the instrument panel. On UN 0002 through 0500, the oxygen is controlled by two levers instead of one rotary valve. On UN 501 and subsequent, a three position oxygen control switch (OXYGEN CONTROL VALVE) is located on the pilot's left console. The three positions are CREW ONLY / NORMAL / MANUAL DROP. In the NORMAL position, if the cabin altitude exceeds approximately 14,000 feet, the passenger masks will automatically drop. 29

OXYGEN SYSTEM Oxygen will flow to these masks when the lanyard is pulled as the mask is donned. Therapeutic / medical oxygen may be supplied to the passengers at any cabin altitude by placing the OXYGEN CONTROL VALVE selector in the MANUAL DROP position. This will cause all masks in the cabin to deploy. Oxygen flow may be shut off from passenger masks by positioning the oxygen priority valve to the CREW ONLY position.

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OXYGEN MASKS The oxygen mask is a quick donning sweep-on mask with a microphone and regulator attachment. The mask is a diluter demand with pressure breathing available by selecting the EMER position. The crewmember is assured that oxygen is being received when no restriction to breathing is present with the mask donned and in the 100% position. Selection of the EMER position will provide a steady flow of pressurized oxygen in the face cone. To qualify as a quick donning mask, the mask must be properly stowed in the retainer located just below each crew member's side window. To conserve oxygen when using the mask, the regulator may be set to normal if the cabin altitude is below 20,000 feet. When using an oxygen mask for smoke protection, 100% position should be selected. The emergency position may be used with the oxygen mask. 31

OXYGEN SYSTEM QUICK REVIEW • 22-cubic-foot bottle is standard; 64-cubic-foot bottle is optional. • The bottle pressure green arc is marked from 1,600 to 1,800 psi. This does not ensure oxygen availability to the crew. • Automatic mask drop occurs at 13,500 ± 600 feet cabin altitude only if normal DC power is available; the solenoid closes at 8,000 feet.







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FLIGHT INTO ICING Icing conditions exist any time the indicated RAT is +10°C or below, and visible moisture in any form is present. Cessna Citations, which have installed properly operating anti-ice and deice equipment, are approved to operate in maximum intermittent and maximum continuous icing conditions. The equipment has not been designed to provide protection against freezing rain or severe conditions of mixed or clear ice. During all operations, the pilot is expected to exercise good judgement and be prepared to alter the flight plan if conditions exceed the capability of the aircraft and equipment.

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FLIGHT INTO ICING Minimum airspeed for sustained flight in icing conditions (except during takeoff, approach and landing) is 160 KIAS. Prolonged flight with the flaps and/or landing gear extended is not recommended. Trace or light amounts of icing on the horizontal tail can significantly alter airfoil characteristics and affect the stability and control of the aircraft.

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ENGINE ANTI-ICE SYSTEM Bleed air flows continuously through the bullet nose of the engine. When the engine anti-ice switches (one for each engine) are positioned to LH and/or RH, bleed air flows through the applicable engine inlet, and engine stators if the throttle position is above 63% N2. If bleed air flow is insufficient to maintain the proper engine inlet temperature, the stator bleed air valve does not open or the throttle lever is below approximately 75% N2, the engine anti-ice fail light on the annunciator panel will illuminate. In flight, an initial setting of greater than 80% N2 is required to extinguish the anti-ice fail annunciator. The annunciator will remain extinguished unless N2 is reduced below 75%.

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ENGINE ANTI-ICE SYSTEM (CONTINUED) Operation of the system may be checked by observing engine ITT and fan speed when the engine anti-ice is turned on. The ITT should increase and the fan speed should decrease. If this check is made on the ground, it will require approximately two minutes to extinguish the engine anti-ice fail light with turbine speed set at approximately 70%. Maximum engine power setting values are reduced when using anti-ice.

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SURFACE DEICE Engine anti-ice switches must be in HI for the surface system to operate. • When the switch is placed to ALL, all panels receive TKS uid. • The WING ICE FAIL light illuminates if the switch is in ALL and pressure is less than 3.5 psi as supplied to each wing proportioner. • The TAIL ICE FAIL light illuminates if the switch is in ALL and pressure to any one or more of the horizontal stabilizer sections is less than 1.5 psi. • The ENG ANTI ICE light illuminates if the switch is in ALL or ENG and pressure at either the cuff or fairing panel falls below 1.5 psi. • The ICE FLUID PUMP FAIL ENG or SUR light indicates failure of the pump. If this occurs, the other pump will automatically activate to supply system requirements when pressure is below 5.5 psi. • Do not operate the TKS if in clear air and the OAT is below –20° C or it will form ice. fl

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SURFACE DEICE The surface anti-ice uid reservoir is located in the right-hand nose compartment over the nose wheel well. This tank contains 7 or 8.5 gallons of anti-ice uid which is approximately 80% ethylene glycol. The tank is serviced through a ller neck located at the upper forward end of the right nose baggage compartment.

All surface anti-ice uids meeting British Deicing Fluid Speci cation DTD 406B (NATO Symbol S-745) are approved. Some companies meeting the speci cation are listed below with their uid designation:

• Canyon Industries AL-5

• Aeroshell Compound 07

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• BP Aero Deicing 2

The ENGINE ANTI ICE switch positions include OFF, HI, or LOW. The HI or LOW selection activates bleed-air heating for the engine. It also determines the amount of TKS uid ow to the inboard wing leading-edge panels if they are selected by the SURFACE ANTI ICE switch. The SURFACE ANTI ICE switch directs TKS uid to either the four inboard wing panels only or to the entire wing leading edge and the horizontal stabilizer leading-edge panels. The switch positions are ENG, OFF/RESET, and ALL. Selecting ENG directs uid only to the four inboard wing panels. Selecting the ALL position directs uid to all the wing and tail panels. fl

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SURFACE (TKS) CONTROLS

Anti-icing for the wings and horizontal stabilizers is provided by a ow of TKS uid, monoethylene glycol in solution with deionized water, through porous leadingedge panels. Each wing contains six panels. The empennage has two on each horizontal stabilizer. Either an 8.5-gallon or a 7-gallon reservoir, in the nose, stores the TKS uid. An ICE FLD LOW light on the annunciator panel, activated by a oat switch, illuminates when the uid level is suf cient for less than 20 minutes of continuous operation at high ow through all panels, or approximately 54 minutes at high ow through the inboard wing panels only. fl

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SURFACE DEICE

SURFACE DEICE Two variable-speed pumps provide uid pressure from the reservoir to the wing and empennage panels. The pump speed is automatically adjusted to ensure proper ow to the panels selected. Fluid is routed from the pumps to proportioning units, which meter ow to each panel. Whenever the system is on, a green ICE FLD SYS ON light illuminates. Fluid pressure output from the pump is monitored by a pressure switch which illuminates a corresponding ICE FLUID PUMP FAIL-ENG/SUR light on the annunciator panel. A pressure switch at each inboard wing panel illuminates the corresponding ENG ANTI-ICE LH/RH light if the pressure falls too low. A pressure switch in the line to either outer wing proportioner illuminates the WING ICE FAIL light if the pressure is below acceptable limits. A pressure switch in the line between the tail proportioner and each horizontal stabilizer panel monitors each individual panel and illuminates the TAIL ICE FAIL light if the pressure falls below limits. fl

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WARNING The surface uid anti-ice system is not a deice system and will not remove signi cant accumulations of ice. The system must be turned on immediately upon detecting ice. If more than one-eighth inch of ice is accumulated prior to turning the system on, leave the icing environment.

CAUTION Both ENGINE ANTI ICE switches must be positioned to HI for the surface protection to function. If they are not, various combinations of annunciator panel lights will indicate improper system operation.

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WATER/SLUSH OPERATION

The airplane should not be operated when standing water or slush depths exceed 0.5 inch. If the 0.5 inch depth is inadvertently exceeded, compressor surges may result.

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WINDSHIELD DEICE SYSTEMS The windshield bleed air system provides windshield anti-ice under all normal operating conditions. This system also provides external windshield rain removal. The system supplies engine bleed air through an electrically actuated pressure regulating shutoff valve in the tailcone of the airplane and manually positioned valves which regulate air to each windshield. The manual valves are located at each bleed air nozzle and are left in the OFF position for all normal operation. Ensure that the rain removal knob is pushed IN for windshield anti-icing.

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WINDSHIELD ANTI-ICE SYSTEM

When windshield anti-icing is required, the W/S bleed valves are turned ON and the W/S bleed switch is turned to LO if the indicated OAT is above -18°C or to HI if the indicated OAT is -18°C or below.

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WINDSHIELD DEICE SYSTEMS (CONTINUED) • W/S BLEED switch controls the temperature, but not the volume. • The manual valves control volume. • W/S AIR O'HEAT light comes on at 146°C.

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WINDSHIELD DEICE SYSTEMS (CONTINUED) Self-test of the temperature monitor system is normally accomplished during the preflight warning systems check by turning the windshield bleed air switch to either the HI or LO position and selecting the W/S temperature position on the rotary test switch. Proper system function is verified by illumination of the W/S AIR O'HEAT annunciator light. Self-tests may also be accomplished in flight, if desired. If the windshield bleed air anti-ice system fails, a backup alcohol anti-ice system is provided for the left windshield only. Sufficient alcohol is provided for 10 minutes of operation; so depart the icing environment without delay. 47

RAIN REMOVAL The windshield bleed air system provides rain removal during flight and ground operations. This system also serves as the windshield anti-ice system when used as described in the windshield anti-ice paragraph of this section. When rain removal is desired, the W/S BLEED switch should be positioned to LOW and the rain removal push-pull knob pulled out. A check should be made to ensure the WINDSHIELD BLEED AIR rotary controls are in the MAX position. The engine ignition should be turned ON when flying in heavy rain. 48

PITOT-STATIC AND ANGLE-OF-ATTACK ANTI-ICE Electric heating elements are incorporated in the pilot's and copilot's pitot tubes, static ports and the angle-ofattack probe. The PITOT & STATIC Switch turns on all of these elements. Correct operation may be checked during preflight by turning the switch ON for approximately 30 seconds, then OFF; and touching each element during the external inspection. Ground operation of the pitot-static heat should be limited to less than 2 minutes to avoid damage to the angle-of-attack probe. Failures of pitot heating elements and of the angle-of-attack probe element are annunciated by P/S HTR OFF and AOA HTR FAIL lights, respectively, in the annunciator panel. Since no annunciation is provided for static port heater failures, a thorough preflight check of the static port heaters is required. 49

ANGLE-OF-ATTACK/STALL WARNING SYSTEM The aerodynamic stalls of this airplane are characterized by a pre stall rolling tendency concurrent with mild buffeting as the airplane is decelerated towards a fully stalled condition. This rolling tendency can be readily controlled with prompt use of aileron control. The airplane is equipped with a stick shaker. 50

ANGLE-OF-ATTACK/STALL WARNING SYSTEM (CONTINUED) The stall warning consists of a stall strip on the leading edge of each wing. The stall strips create turbulent airflow at high angles of attack, causing elevator buffet to warn of approaching stall conditions. Buffet occurs prior to the actual stall at approximately VSI + 10 knots in the clean configuration and VSO + 5 knots in the landing configuration. 51

ANGLE-OF-ATTACK/STALL WARNING SYSTEM (CONTINUED) The AOA indicator is calibrated from 0.1 to 1.0 with red, yellow, and white arcs. The 0.1 mark represents a very low angle of attack; 1.0 indicates the aircraft has exceeded the critical angle of attack and has stalled. The area from 0.1 to 0.57 represents the normal operating range, except for approach and landing. The white arc from 0.57 to 0.63 covers the approach and landing range, with the middle of the arc (0.6) indicating the optimum landing approach airspeed (VAPP or VREF). The yellow range of 0.63 to 0.85 represents a caution area, indicating approach of the critical angle of attack. The red arc from 0.85 to 1.0 is a warning zone that indicates the beginning of low-speed buffet followed by a full stall. 52

ENGINES Engine thrust is supplied by two aft fuselagemounted JT15D-4B turbofan engines manufactured by Pratt & Whitney Aircraft of Canada. These lightweight engines are high bypass, twin-spool turbofans developing 2,500 pounds of thrust in standard day, sea level conditions. Bypass ratio is 2.7-1

JT15D-4B

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ENGINES SYSTEM QUICK REVIEW GENERAL

IGNITION:

OIL

• Pratt and Whitney JT15D-4B

• Two exciter boxes and two plugs. • Step modulator on FCU is energized whenever ignition circuit is energized. • IGN Switch: • NORM—Start and engine antiice. • ON—Takeoffs, landings, heavy rain, heavy turbulence, stalls, emergency descents; power from DC buses in the cockpit. • For start, the power source is a circuit breaker on the hot battery bus.

• Overall capacity is 2.08 U.S. gallons.

• 2,500-pound thrust • Bypass ratio 2.7:1 • N1 rpm at 106% = 16,540 RPM • N2 rpm at 97% = 31,450 RP

• Maximum allowable oil consumption is one quart every four hours mea-sured over a ten-hour period. • Check oil level approximately ten minutes after shutdown. • Do not mix approved brands (not more than two quarts of another approved brand may be added in a 400-hour period).

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IGNITION SYSTEM • Comprised of two exciter boxes and two plugs. • The step modulator on the FCU is energized whenever the ignition circuit is energized. SWITCH POSITIONS: • NORM—Start and engine anti-ice. • ON—Takeoffs, landings, heavy rain, heavy turbulence, stalls, emergency descents. Power for the system is from DC buses in the cockpit. 55

FIRE DETECTION

The aircraft is equipped with engine fire detection and extinguishing system. The systems include heat detection devices which illuminate fire warning lights in the cockpit and controls to activate one or both fire extinguisher bottles. There is a test function for the fire detection system. Two portable fire extinguishers are stowed inside the airplane.

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FIRE DETECTION

The engine re protection system is composed of two sensing loops, two control units (one for each engine) located in the tail cone, one ENG FIRE warning switchlight for each engine, two re extinguisher bottles which are activated from the cockpit, and a re detection circuit test. Detection and extinguishing system electrical power is supplied from the main DC system.

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FIRE DETECTION Within each engine nacelle are two heatsensing cables, or loops, one mounted around the lower engine accessory section and one surrounding the engine combustion section. The loops are connected to control units that monitor their electrical resistance. As the loop is heated, its electrical resistance decreases until, at a temperature of 500° F, a circuit is completed to the control unit to illuminate the applicable red ENG FIRE switchlight.

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HYDRAULIC SYSTEM The landing gear, flaps, speedbrakes and thrust reversers are hydraulically operated. The hydraulic system is an open center system constantly pressurized to 60 PSI by a hydraulic pump on each engine. The system is only pressurized to 1500 PSI during system operation. This is indicated by the HYD PRESS ON light on the annunciator panel. When a cycle of actuated system is complete, the light extinguishes.

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• Flaps beyond approach aps regardless of throttle position; the horn cannot be silenced.

• Flaps are held in each selected position by trapped uid.

1. QUANTITIES:

• The rotary test switch is in the LDG GEAR position.

• Flap actuators are interconnected to prevent split aps.

5. SPEEDBRAKES:

• Normal DC power is required to position control valve.

• Total system...... 4.4 gal • Reservoir capacity .......... 0.65 gal

2. ENGINE-DRIVEN PUMP: • Flow rate is 3.25 gpm maximum. 3. OPEN-CENTER SYSTEM: • Bypass valve open (normal) .................. 60 psi • Bypass valve closed (system operation) .......... 1,500 psi

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4. LANDING GEAR WARNING HORN: • Either throttle less than 70% N2 rpm and the speed below 150 KIAS; the horn can be silenced.

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HYDRAULICS SYSTEM QUICK REVIEW

• Held closed mechanically; they are held open by trapped uid. • Retracted normally with the switch in RETRACT; they are also retracted if either throttle is advanced above 85% N2. • Require normal DC power to remain extended; they will blow to trail immediately upon DC power failure.

• Flap extension time (seconds): • 0 to 7°.................. 5 • 7 to 20° ............... 8.5 • 20 to 35° ............. 5 • Flap retraction time (seconds): • 35 to 20° ............. 6 • 20 to 7° ............... 7.3 • 7 to 0°.................. 6

• System pressure is required for extension and retraction.

6. FLAPS: • Flaps are held retracted by trapped uid.

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HYDRAULICS SYSTEM QUICK REVIEW 7. LANDING GEAR: • Normal DC power is required for hydraulic retraction and extension. • Freefall/pneumatic system is the emergency backup for extension.

8. THRUST REVERSERS:

• If the thrust reversers do not operationally check, ight should not be attempted. • The thrust reversers must be in idle power by 60 knots. • Do not deploy thrust reversers until nosewheel contact.

• On the ground, one squat switch (either or both) allows the control valve to energize to the deploy position when commanded.

9. ANTISKID BRAKES:

• Emergency stow switch is powered from the opposite thrust reverser circuit breaker.

• Normal DC power required to operate the pump

• Illumination of the ARM or UNLOCK light in ight triggers the MASTER WARNING lights.

• Pneumatic brakes are a backup for the power brakes; no differential braking and no antiskid protection are available. • Antiskid protection drops out below 12 knots. • Touchdown protection and test mode

• Separate from the main airplane hydraulic system with the reservoir located in the nose

• Antiskid protection available only with power brakes

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HYDRAULIC SYSTEM Operating pressure (1500 psi) is indicated by the HYD PRESS ON light illuminating on the annunciator panel. When a cycle of the gear, speedbrakes or thrust reversers is complete, the light extinguishes. If the HYD PRESS ON light remains illuminated for an extended time, the speedbrake, landing gear or thrust reversers circuit breaker should be pulled per the Abnormal Checklist to relieve system pressure.

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THRUST REVERSERS The thrust reversers are of the “target door” design which form the aft portion of the engine nacelle when in the stowed position. Their support structure attaches directly to the aft engine bypass duct mounting ring. Each reverser is actuated by two hydraulic cylinders to deploy and/or stow. The reversers are locked into the stowed position by the design which incorporates an overcenter feature in the actuation linkage. The hydraulic power required for operation is provided by the standard airplane system through the thrust reverser isolation and control valves. Activation of the system is by pilot operation of the thrust reverser throttle levers mounted on the primary throttle levers. The reversers can only be deployed when the primary throttle levers are in the idle thrust position and the airplane is on the ground. The thrust reverser lever(s) should not be placed in the idle reverse detent position in flight since a single failure of either squat switch could permit deployment of the thrust reverser(s).

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THRUST REVERSERS

If the thrust reverser lever is inadvertently placed in the idle reverse detent position in flight, the airplane MASTER WARNING light will flash along with the illumination of the ARM and HYD PRESS ON annunciator lights. A MASTER WARNING light when thrust reversers are moved to deploy on the ground means that neither landing gear squat switch has activated. To ensure actuation of the squat switches and to eliminate any delay in the deployment of the thrust reversers, it is recommended that the speed brakes be extended immediately following touchdown.

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THRUST REVERSERS (CONTINUED) Three reverser indicator lights for each reverser are mounted on the panel for monitoring reverse functions: ARM light, UNLOCK light, and DEPLOY light. The amber ARM light indicates hydraulic pressure to the control valve. The amber UNLOCK light indicates the thrust reversers are not in the fully stowed position. The white DEPLOY light indicates that the thrust reversers are in the full deploy position. The DEPLOY light shall illuminate in less than 1.5 seconds after the hydraulic UNLOCK light illuminates. An erroneous sequencing or a delay in the illumination of the thrust reverser lights indicates a failure in the thrust reverser system. Either or both conditions require a maintenance check before further flight.

WARNING

Do not attempt to fly the airplane if the thrust reverser preflight check is unsuccessful.

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THRUST REVERSERS (CONTINUED) After deployment, power may be increased by moving the thrust reverser throttle levers aft for desired reverse thrust. For convenience, STOPS have been installed on the thrust reverser levers. These stops are set to an N1 reverse power setting that is based upon maximum thrust reverser capability down to a temperature of -18°C. and corresponds to 83.0% N1 at that temperature. Below -18°C, reverse thrust must be limited to a maximum of 79.3%. This will allow the pilot to keep his attention on the landing rollout instead of diverting his attention to the reverse power settings, except in abnormal ambient temperature condition (below -18°C).

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THRUST REVERSERS (CONTINUED) In the event of an inadvertent thrust reverser deployment, an automatic throttle retarding device will retard the throttle to approximately idle thrust depending on the throttle friction setting. After this device has activated, the throttle lever can be advanced resulting in corresponding reverse thrust. If the pilot attempts to override the retraction mechanism, the throttle cable system could be damaged. Subsequent reduction of the throttle lever to idle will not result in engine flameout unless mechanical damage has resulted from the deployment.

WARNING

Do not attempt to override the retraction mechanism or advance the throttle after retraction. This could result in a dangerous asymmetrical thrust condition. Do not use throttle friction or manually restrain the throttle levers during takeoff.

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THRUST REVERSERS (CONTINUED) An emergency stow switch is installed on the fire panel for each thrust reverser. They are used only for stowing the reversers when they will not stow through the primary thrust reverser controls. Each emergency stow switch receives its electrical power through the opposite thrust reverser circuit breaker. The emergency stow function can be checked on the ground by deploying the reversers normally and then actuating each emergency stow switch. When the emergency stow switch is actuated, the DEPLOY and UNLOCK light shall extinguish and the ARM and HYD PRESS ON light will remain illuminated. Return the thrust reverser lever to stowed position, then turn each emergency stow switch off. All lights shall be extinguished. 68

THRUST REVERSERS (CONTINUED) The nose wheel must be on the ground and forward pressure maintained on the control column prior to and during the deployment and actuation of the thrust reversers. Single engine reversing has been demonstrated during normal landings and is easily controllable. Also for an increased aerodynamic drag during the landing roll, the thrust reversers should remain in the idle deployed position below 60 KIAS. Care should be taken on runways with loose dirt, gravel or grit as idle reverse at low speed can cause foreign object damage.

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THRUST REVERSERS LIMITATIONS • On the ground, one squat switch (either or both) allows the control valve to energize to the deploy position when commanded. • Emergency stow switch is powered from the opposite thrust reverser circuit breaker. • Illumination of the ARM or UNLOCK light in flight triggers the MASTER WARNING lights. • If the thrust reversers do not operationally check, flight should not be attempted. • The thrust reversers must be in idle power by 60 knots. • Do not deploy thrust reversers until nosewheel contact. • Use of thrust reversers is prohibited during touch and go













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Memory Items!



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BRAKE SYSTEM The power brake system uses a multi-disc brake assembly in each main gear wheel, powered by a hydraulic system that is independent of the airplane hydraulic system. The system uses an electrically powered hydraulic pump to maintain constant pressure for brake operation. The brakes are normally used as anti-skid power brakes, but can be operated as power brakes without anti-skid protection. In the event that brake system hydraulic pressure is lost, emergency braking is available. Braking is initiated by rudder pedal-actuated master cylinders. If both the pilot and copilot attempt to apply the brakes simultaneously, the one applying the greater force on the rudder pedals has control since they are plumbed together in series. System components include a hydraulic accumulator and a reservoir pressurized by cabin air. Reservoir fluid level and accumulator air precharge are preflight inspection items.

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EMERGENCY BRAKE SYSTEM In the event the hydraulic brake system fails, a pneumatic brake system is available. The system uses air pressure from the pneumatic bottle which can also be used for emergency landing gear extension. Air bottle pressure is adequate for stopping the airplane, even if the landing gear has been pneumatically extended. Do not depress the brake pedals while applying emergency air brakes. The shuttle valve action may fail, allowing air pressure to enter the hydraulic lines and rupture the brake reservoir. The emergency brakes should be pulled only enough to obtain the desired rate of deceleration, then held until the airplane stops. Repeated applications waste air pressure. Anti-skid protection is not available during emergency braking. Do not attempt to taxi after using the emergency brakes.

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ANTISKID SYSTEM The anti-skid system permits maximum braking without wheel skid under all runway conditions. Detection of an impending skid causes the control box to command the anti-skid valve to reduce pressure being applied to the brakes. When the transducer signal returns to normal, braking pressure is restored to the brakes. Touchdown protection is a feature of the anti-skid system that prevents touching down with locked brakes. Optimum braking is obtained by deployment of speedbrakes at touch-down, then firmly applying and holding the brakes until the desired speed has been reached. Do not pump the brakes. 74

ANTI-SKID SYSTEM QUICK REVIEW • Separate from the main airplane hydraulic system with the reservoir located in the nose. • Normal DC power required to operate the pump. • Antiskid protection available only with power brakes. • Pneumatic brakes are a backup for the power brakes; no differential braking and no antiskid protection are available. • Antiskid protection drops out below 12 knots. • Touchdown protection and test mode. 75

LANDING GEAR SYSTEM The Citation SII landing gear is electrically controlled and hydraulically actuated. When retracted, the nose gear and the struts of the main gear are enclosed by mechanically actuated doors. The main gear wheels remain uncovered in the wheel wells. Gear position and warning are provided by colored indicator lights and a warning horn. Nosewheel steering is mechanically actuated through linkage from the rudder pedals. A self-contained shimmy damper is located on top of the nose gear strut. Power braking is provided with or without anti-skid. Emergency braking is also provided. 76

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ELECTRICAL SYSTEM DC power is supplied by a 300-ampere starter-generator unit on each engine and a 44 ampere-hour battery. Engine ground starts use either external power or the airplane battery for the first engine start. The second engine normally uses the generator from the operating engine to supply electrical power for the start. External power or the airplane battery may be used for starting the second engine, if desired, by turning the generators to the off position. Generator assisted cross start capability is disabled in flight. 78

EXTERNAL POWER • The external power receptacle is located below left engine nacelle. • The external power is routed to the hot battery bus. • The battery charges from the GPU regardless of battery switch position. • Set GPU to 28.5v and between 800 to 1,000 amps before connecting to the aircraft. 79

ELECTRICAL SYSTEM One generator is capable of supplying all standard electrical requirements in flight in the event of a generator failure. A protected DC power path is included which provides bus extension to the opposite circuit breaker panel. This is identified on each circuit breaker panel as RH and LH CB PANEL. The bus extensions feed DC power from one side to the bus extension on the opposite circuit breaker panel in order to allow logical grouping of corresponding LH and RH system circuit breakers.

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ELECTRICAL SYSTEM Power for the avionics system is controlled by three switches labeled AVIONICS POWER on the left meter panel. The right switch controls DC power to the system. The center switch controls DC power to two 375 volt-ampere inverters which individually provide AC power to their respective busses. Each inverter provides both 26 VAC and 115 VAC power. The number 1 inverter powers the number one - 26 VAC bus and the number one - 115 VAC bus. The number 2 inverter powers similar number two busses. If either inverter fails, the respective INVERTER FAIL light will illuminate and automatic switching will occur, causing both busses to receive power from the remaining inverter.

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ELECTRICAL SYSTEM A battery in good condition should supply power to all buses for approximately 10 minutes. If only the hot battery and emergency buses are powered, battery life should be extended to 30 minutes. An emergency battery bus is provided to supply DC power to operate COMM 1, NAV 2, the copilot's HSI and DG2, copilot's attitude indicator, overhead floodlights, RH pitot-static heater, LH and RH N1 tachometers and voltmeter. Communications can be continued using a headset and selecting EMER COMM on either audio amplifier.

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MEMORY ITEM FOR ELECTRICAL SYSTEM

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PRESSURIZATION SYSTEM Two elements are required to provide cabin pressurization. One is a constant inflow of air. The other is controlling the flow of air in and out of the airplane to achieve the desired differential pressure and cabin altitude. In the Citation SII, the inflow of air to the cabin is fairly constant and the outflow of air is controlled by two outflow valves located on the aft pressure bulkhead . The cabin pressure control system includes a pressure controller, two outflow valves, two cabin altitude limit valves, and a pneumatic relay. An emergency dump valve and a regulated vacuum supply complete the cabin pressure control system. Unconditioned emergency pressurization is available from the left engine. 84

PRESSURIZATION QUICK REVIEW • OFF—All valves are closed; bleed air is still available for service air and anti-icing/deicing. • GND—The shutoff valve opens only if the airplane is on the ground; a largerthan-normal amount of bleed air is supplied from the right engine only.

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• LH—The left ow control and shutoff valve opens; the normal ow of air to the ACM is from the left engine only, at 6 pounds per minute.

• NORMAL—The left and right ow control and shutoff valves open; this is the deenergized (fail-safe) condition of the system, providing the normal ow of air from each engine to the ACM, at 6 pounds per minute per engine (12 pounds per minute total volume).

• EMER—The emergency pressurization valve opens in ight only; air- ow to the ACM is stopped; the only control of temperature is with the left throttle.

• RH—The right ow control and shutoff valve opens; the normal ow of air to the ACM is from the right engine only, at 6 pounds per minute. 85

PRESSURIZATION QUICK REVIEW BLD AIR GND LIGHT: • The ground valve is open. • The light extinguishes and the ground valve closes if the throttle is advanced too high on the right engine (primary pressure switch). Reduce power to reopen the ground valve. • The light extinguishes if the primary pressure switch fails, and the secondary pressure switch activates if power is too high on the right engine.

The ACM O'PRESS light illuminates, and the GND position cannot be reselected until the right engine power is reduced and the NORM PRESS circuit breaker is pulled and reset (ACM O'PRESS annunciator light illuminates until reset).

• Ten seconds is required to move the valve fully from one position to the other. • No freezing protection is provided for the water separator when operating in manual mode.

TEMPERATURE CONTROL: • Normal DC power is required for automatic and manual modes. 86

MEMORY ITEM FOR PRESSURIZATION SYSTEM

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AIR CONDITIONING SYSTEM The air conditioning system provides conditioned air to both cockpit and cabin areas. Engine bleed air is required to operate the system. The cabin and cockpit temperature is regulated by mixing hot bleed air with air cooled by an air cycle machine (ACM). Fans are provided to circulate cabin air. An optional vapor cycle air-conditioning system may be installed to provide additional cooling. An optional flood cooling system is available to provide a means to rapidly cool the cabin temperature.

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FLOOD COOLING The FLOOD COOLING switch is located next to the pressurization controller on center panel. It has two positions: ON and OFF. In OFF- The air ow is direct through the normal distribution system. In ON- the normal distribution system is bypassed and all air ow is directed to the ood cooling duct and vent. Inside the ductwork is a fan to increase air ow at low power settings. This is the fastest way to cool a heat soaked cabin. fl

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PNEUMATIC SYSTEM Bleed air from each engine is extracted from the engine high-pressure compressor section and routed to four different places:

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To the flow control valves for use by the air cycle machine.

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To the ground valve for use by the air cycle machine during ground operation.

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To the emergency valve for alternate pressurization.

Through check valves for distribution to the windshield anti-ice, cabin door seal, and pressurization control systems.

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FLIGHT CONTROL SYSTEM The ailerons, rudder, and elevators are manually operated by either crewmember through a conventional control column and rudder pedals. Control inputs operate the control surfaces through cables and bellcranks. The rudder pedals can be adjusted to three separate positions by depressing a spring-loaded latch on the side of the rudder pedal. A mechanical interconnection between the rudder and the ailerons provides small rudder deflections with aileron movement. A spring in the system can be manually overridden for cross controlling. 91

FLIGHT CONTROLS QUICK REVIEW RUDDER:

ELEVATORS:

AILERONS:

• Maximum travel is 22° either side of centerline.

• Maximum travel is 20° up and 15°down.

• Maximum travel:

Trim Tab: • travel is 10° either side of centerline (servo tab).

• Trim tab travel is 5° up and 17° down.

• Nosewheel de ection either side of centerline with full rudder pedal de ection: • 20° either side of centerline • Do not attempt ight if the nosewheel steering is inoperative.

• Can be electrically trimmed

• 19° up and 15° down • Trim tab on left aileron only; maximum travel is 20° up and down.

• If equipped with optional copilot electric trim, the pilot's has priority. • Trim tab(s): On both elevators

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FLIGHT CONTROL MEMORY ITEMS

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PUSH TO TEST