For flight planning, use only TOTAL pound indication on the fuel quantity gage. DO NOT use separate tank indications for this purpose. Steady illumination of any of the amber fuel system valve lights on the fuel control panel, except FUEL JETTISON valve lights (if installed), indicates the corresponding valves are not in the position selected. Momentary illumination while corresponding valve is in transit after switching modes indicates proper operation. NORMAL FUSELAGE FUEL TRANSFER - Model 25B CAUTION DO NOT energize fuselage fuel transfer system when wing and tip tanks are full. NOTE Plan fuel requirements to ensure return to takeoff point or alternate if fuselage tank fails to transfer. FUS TANK XFER-FILL Switch - XFER when tip tank fuel quantity indicates 760 pounds or less in each tank. When XFER is selected, the STANDBY PUMPS will be deenergized, the cross flow valve will open, the fuselage tank valve will open, and the transfer pump will be energized. CAUTION If the crossflow valve fails to open, fuselage fuel cannot be transferred forward to the wing tanks using the transfer pump. However, fuselage fuel will gravity flow into the right wing tank until heads are equal. A right wing heavy condition will occur. 2. FUS TANK XFER-FILL Switch - OFF after EMPTY light on fuel control panel illuminates. , NORMAL FUSELAGE FUEL TRANSFER - Model 25C CAUTION DO NOT energize fuselage fuel transfer system when wing and tip tanks are full. NORMAL TRANSFER PROCEDURE: 1. FUS TANK XFER-FILL Switch - XFER when tip tank fuel quan- tity indicates zero fuel, each tank. When XFER is selected, the STANDBY PUMPS will be deenergized, the crossflow valve will open, the fuselage tank valve will open, and the transfer pump will be energized. FUEL CROSSFLOW (WING-TO-WING TRANSFER) 1f lateral unbalance occurs: CAUTION DO NOT crossflow with jet pump inoperative. Engine starvation may occur due to fuel being pumped through open crossflow valve into opposite wing. 1. CROSSFLOW Switch - OPEN. 2. STANDBY PUMP Switch (heavy wing) - ON. 3. STANDBY PUMP Switch (light wing) - OFF. 4. CROSSFLOW Switch - CLOSED after fuel balance is obtained. 5. STANDBY PUMP Switches - OFF. 6. If preceding procedure fails to balance fuel, unequal power set- ting may be used to balance the wings. STALL WARNING SYSTEM OPERATIONAL CHECK (Cont) AIRCRAFT EQUIPPED WITH 8° FLAP SYSTEM APPROACH AND CLEAN CONFIGURATION (Cont) I. O,On Aircraft not modified per AMK 81-9. Lift angle-of-attack vane until lower edge of vane is at lower line labeled 9 . Shaker should actuate and applicable STALL warning light should flash. 0 On Aircraft modified per AMK 81-9. Lift angle-of-attack vane until lower edge of vane is at lower line labeled e . Nudger and Shaker should actuate and applicable STALL warning light should flash. J. Lift angle-of-attack vane until upper edge of vane is at upper line labeled 0 . Pusher should actuate. 0 NOTE Repeat the above procedure using opposite vane and ° switch. A steady stall warning light indicates a mal function or that the system is off. The shaker and flashing light will continue during pusher operation. WARNING On Aircraft modified per AMK 81-9, the action of the nudger verifies operation of the pitch torquer prior to pusher actuation. If, during ground test, the shaker is not accompanied by the nudger - Do not dispatch. If, during flight, the shaker is not accompanied by the nudger - Do not decelerate further. Each engine is equipped with an ignition exciter and two igniter plugs; either plug is capable of starting the engine. Operational time limits are: 2-minutes ON, 3-minutes OFF; then 2-minutes ON, 23-minutes OFF: or 5-minutes ON, 25-minutes OFF. When flight conditions dictate, the ignition system may be operated continuously as required. If extended duty ignition exciters are not installed and the system time limits are exceeded, the ignition exciters must be replaced. Regardless of which ignition exciters are installed, if the system time limits are exceeded, the igniter plugs must be removed, inspected and tested per G. E. Accessory Overhaul Manual. 1. Attach Anti-ice Additive blender tube to refuel nozzle (see figure 2-1), ensuring blender tube discharges into fuel stream. 2. Start fuel flow then start additive flow by fully depressing and slipping ring over trigger or pushing button hold down. Refueling rate shall not be less than 30 GPM (113.6 LPM) nor greater than 60 GPM (227.1 LPM). During topping off, refueling rate may be less than 30 GPM (113.6 LPM). 3. Stop additive flow then stop fuel flow. CAUTION Ensure that additive is directed into the fuel stream and that additive flow starts after fuel flow starts and stops before fuel flow stops. Do not allow concentrated additive to con tact interior of fuel tank or aircraft painted surfaces. Use not less than 20 fluid ounces (1 can) of additive per 260 gallons (984.2 liters) nor more than 20 fluid ounces (1 can) per 104 gallons (393.7 liters) of fuel. CHECKING FUEL ADDITIVE Prolonged storage of the aircraft will result in water buildup in the fuel which "leaches out" the additive. This is indicated when an excessive amount of water accumulates in the fuel sumps. Check the additive concentration using a Differential Refractometer. Minimum additive concentration shall be 0.035% by volume and the maximum concentration shall be 0.15% by volume. Identification if ice can be made by the following methods: A. Ice formation on the lower corner of the windshield or the nose of the tip tank. B. During night flights, two red ice detect lights will cause red areas (approximately 1-1/2 inches in diameter) to appear on the windshield when particles of moisture or ice form on the windshield. The light on the pilot's side is located in the defog airflow stream and the light on the copilot's side is located outside the defog airflow stream. If the windshield heat or alcohol system is operating, the copilot must moni tor the light on his side for indication of ice or moisture formation. The windshield ice detection lights will indicate moisture encounters when OAT is above freezing. At below freezing OAT, the lights will indicate _ice encounters. For temperature conversion, see figure entitled RAM AIR - OUTSIDE AIR TEMPERATURE CONVERSION in Section IV. C. The wing structure temperature indicator is marked with green, yellow, and red temperature range arcs. When the indicator pointer is in the green arc, the wing structure is above 35° F and is warm enough so that ice will not adhere. When the pointer is in the yellow arc, the wing is approaching a "too hot" condition (refer to WING ANTI-ICE, this section). When the pointer is in the red arc, the wing structure is below 35° F and indicates the wing heat system should be used or if the wing heat system is on, it indicates wing heat system failure or too low engine rpm. D. A visual inspection may be used to check for ice accumula- tions on the wing leading edges. For night operation, the optional wing inspection light on the right fuselage may be used to check for ice buildup on the wings. The wing inspection light is illuminated by setting the WING INSPECTION switch to ON. NOTE • The wing inspection light, in itself, is inadequate for ° detecting the presence of ice near the wing tips. • If the presence of ice on the wing leading edge is suspected during night operations in atmospheric conditions conducive to icing, the normal approach speeds must be increased per the WING HEAT FAILURE procedure in Section III - EMERGENCY PROCEDURES. WING ANTI-ICE The wing anti-ice system utilizes engine bleed air and should be used in concurrence with Nacelle Heat, Windshield Heat, and Radome Anti-Ice. A. Stab/Wing Heat Switch - ON. Monitor wing temperature gage and WING OV HT light. If the wing temperature gage is in the yellow arc, the wing structure is approaching a "too hot" condi tion. Corrective action can be taken by reducing RPM. How ever, power should not be retarded below 80% RPM to retain effective wing and engine anti-ice. 0 NOTE • On aircraft incorporating AAK 82-8 or AMK 83-4, in ° icing condtions below -18 C (0°F) OAT, ice buildup may occur on the outboard wing leading edge at reduced power, such as holding. Satisfactory handling qualities have been demonstrated with this ice accumulation. The ice may be dissipated by adding power to above 85% RPM (spoilers may be required for added drag). If it is necessary to land with any ice, or suspected ice, on the wing leading edge, perform the landing procedure for WING HEAT FAILURE in Section III, EMERGENCY PROCEDURES. Ice on the wing leading edge should be suspected any time icing conditions are encounted below -18°C (0°F) OAT and it cannot be confirmed that the wing leading edge is clear, such as during night operation. 1 For ground operation to prevent wing overheating, limit RPM to 70% and monitor Wing Temperature indicator. To prevent overheating of the horizontal stabilizer heating elements, ensure that the STAB HEAT light is "on" and there is no additional DC ammeter increase. RADOME ANTI-ICE Anti-icing the radome prevents ice from shedding off the fuselage and entering the engines. Anti-icing the radome only is accomplished by setting the Alcohol Pump Switch to RADOME. Alcohol for radome anti-icing only will last for approximately 1-1/2 hours. An amber ALC AI warning light will come on when the alcohol is depleted, then set Alcohol Pump Switch OFF. 0 NOTE Use the WSHLD & RADOME position of the switch only ° if the bleed airflow for windshield anti-ice has failed. Automatic Operation A. In Normal/Out Defog Knob - Pull out. B. Windshield Heat Switch - AUTO. The switch will remain in this position. The green WSHLD HEAT light will come on when the control valve begins to open and the valve will go to the full open position. If the WSHLD OV HT lights come on, the control valve will close and remain closed until the windshield has cooled sufficiently, then open again. rNOTE Automatic operation of the system is recommended when constant use of the system is required. To deenergize the system during automatic operation, set the Wind shield Heat Switch to MAN and hold OFF until green wind shield heat indicator light goes out. Manual Operation A. In Normal/Out Defog Knob - Pull out. B. Windshield Heat Switch - MAN; ON and hold until desired amount of airflow is obtained. The green WSHLD HEAT light will come on when the control valve begins to open. The valve will remain in the selected position when the switch is released. If the red WSHLD OV HT lights come on, the pressure regulator is automatically closed, stopping hot airflow to the windshield. When the windshield has cooled sufficiently, the pressure regulator valve is again opened, supplying hot airflow to the windshield. NOTE -' If the WSHLD OV HT light continues to cycle during manual mode of operation, manually control the system by setting the Windshield Heat Switch to OFF and holding until red light goes out. The switch may again be held to ON when the windshield has cooled sufficiently. If the airflow should fail, the pilot's windshield may be anti-iced by: C. Alcohol Pump Switch - WSHLD & RADOME. This will anti-ice only the radome and pilot's windshield and will give an alcohol flow for approximately 43 minutes. The amber ALC AI light will come on when the alcohol is depleted, then set Alcohol Pump Switch OFF. INTERIOR WINDSHIELD DEFOG Interior windshield defog requires no management. The interior windshield is continuously defogged by normal cabin air conditioning. The procedures in this section of the manual have been developed by Learjet Inc. and approved by the FAA. This section contains those operating procedures requiring the use of special systems and/or regular systems in order to protect the occupants and the aircraft from harm during a critical situation requiring immediate response. The procedures located in this section supplement Normal Procedures when an emergency condition exists. Use of Normal Procedures should be continued when applicable. Sound judgement as well as thorough knowledge of the aircraft, its characteristics, and the flight manual procedures are essential in the handling of any emergency situation. OVERRIDING CONSIDERATIONS In all emergencies, the overriding consideration must be to: Maintain Airplane Control Analyze the Situation Take Proper Action TERMINOLOGY Many emergencies require some urgency in landing the aircraft. The degree of urgency required varies with the emergency; therefore, the terms 'land as soon as possible" and "land as soon as practical are employed. These terms are defined as follows: Land as soon as possible - A landing should be accomplished at the nearest suitable airfield considering the severity of the emergency, weather conditions, field facilities, ambient lighting, and aircraft gross weight. Land as soon as practical - Emergency conditions are less urgent, and although the mission is to be terminated, the degree of the emergency is such that an immediate landing at the nearest adequate airfield may not be necessary. SINGLE-ENGINE LANDING Single-engine approach and landing is flown essentially the same as with both engines except for the following: A. Final Approach; GEAR DOWN, Flaps - 20°. B. Final Approach Speed; VREF + 10 KIAS. C. When landing is assured; Flaps - 40°, VREF• LANDING WITH ONE OR BOTH SPOILERS EXTENDED A. Landing Gear Switch - GEAR DOWN. B. FLAPS - UP. C. Final Approach Speed - VREF + 40 KIAS. D. Landing Distance - Increased by 40% Antiskid ON, 55% Antiskid OFF. FLAPS UP LANDING A. Final Approach Configuration - GEAR DOWN, FLAPS - UP. B. Final Approach Speed - VREF + 30 KIAS. C. Landing Distance - Increase by 35%. NOTE Use of ;drag chute or thivst reversers (if installed) ° is recommended. A. Spoiler Switch - RET. B. Landing Gear Switch - GEAR DOWN. C. Landing Gear Circuit Breaker - Pull. D. Emergency Gear Extension Lever (left side of pedestal) - Push full down to latched position. Do not attempt to retract the landing gear once Emergency Gear Extension has been selected. To do so may cause excessive air pressure to be introduced into the hydraulic return lines, thereby rupturing the reservoir. Once Emergency Gear Extension has been selected, the left and right red UNSAFE gear lights will be lighted; also the green LOCKED DOWN lights will be lighted, indicating the gear is down and locked. The red UNSAFE lights will remain lighted due to the main gear inboard doors remaining open in the emergency mode. After Gear is Down and Locked E. Emer ency Gear Extension Lever - Return to full up (latched) position only t there is no hydraulic pressure. This is accomplished by pulling aft on the small sheet metal tab on top of the handle housing and pulling the handle to the full up (latched) position. This will maintain air pressure for emergency braking if there is a leak in the air line. If the aircraft hydraulic system is operating normally, there NOTE will be no need to energize the auxiliary hydraulic pump and braking action will be normal. F. Flap Switch - Full Down. The auxiliary hydraulic pump has an operational cycle of 3 minutes ON and 20 minutes OFF. Operation at more frequent intervals may result in overheating the pump drive motor. PRESSURIZATION SYSTEM FAILURE In case of pressurization loss at altitude: A. Don oxygen masks. ~~ Refer to Section D for oxygen system operation. N'-"'~ NOTE An immediate descent to Flight Level 300 or below is required in the event of cabin decompression resulting in a cabin altitude above 15, 000 feet. B. Maintain engine RPM. C. In Normal/Out Defog Knob - Push in D. Windshield Heat - AUTO E. Air Bleed Switch - OFF No In case of cabin pressurization failure which causes cabin altitude to exceed 10, 000 feet, an aneroid switch inunediately closes "con trol pressure" to the outflow valve and the AUTOMATIC,MODE is Inoperative. F. Auto/Man Switch - MAN. Use Manual Control to pressurize to desired cabin altitude. G. If cabin pressurization cannot be maintained by this procedure, the EMERGENCY DESCENT procedure should be followed. No If cabin altitude control is regained in the manual mode and it is desired to return to automatic mode of operation: 1. Altitude Controller - Set at or below 8, 000 feet. 2. Rate Selector - Turn to full INCR. 3. Manually control cabin to altitude selected on Altitude Controller. 4. Allow cabin rate indicator to stabilize at 0. 5. Reset Rate Selector to nominal position. 9. Auto/Man Switch - AUTO. CABIN ALTITUDE EXCEEDS 8500 FEET 1. Crew Oxygen Masks - Don. Select 100% OXY In the event of cabin pressurization failure which NOTE causes cabin altitude to exceed 10,000 feet, an aneroid switch immediately closes to the outflow valve and the Automatic Mode is inoperative. 2. If aircraft is climbing, stop climbing and level off at (or descend to) the nearest appropriate altitude. 3. Pilot and Copilot OXY MIC Switches - ON. Note Communication between crew members can be accomplished by using the INPH function on the AUDIO CONTROL panel and increasing the MASTER VOL level. 4. AIR BLEED Switch - NORM. 5. Cabin Altitude - Check. EMERGENCY FUEL SYSTEM OPERATION FUEL PRESSURE WARNING LIGHT ILLUMINATED Illumination of a red FUEL PRESS warning light is an indication of loss of fuel pressure to the engine. Probable cause is jet pump failure. Affected Engine A. STANDBY PUMP Switch - ON. B. AIR IGN Switch - ON. C. CROSSFLOW Switch - CLOSE. D. 0 (25B and 25C Aircraft) Fuselage Tank XFER-FILL Switch- OFF. * (25C Aircraft) Fuselage Tank FUS VALVE Switch - CLOSE. c~ NOTE Fuselage fuel transfer deactivates the wing STANDBY ° PUMPS. The fuel in the fuselage tank will not be avail able above 25, 000 feet. Below 25, 000 feet the engine driven fuel pump will suction feed sufficient fuel to sup ply the engine and fuel transfer can be accomplished. Replan flight accordingly. IF THE FUEL PRESS LIGHT DOES NOT EXTINGUISH E. JET PUMP Switch - OFF. IF THE FUEL PRESS LIGHT DOES NOT EXTINGUISH F. Descend to 25, 000 feet or lower. NOTE D *The engine-driven fuel pump will suction-feed sufficient ° fuel to supply the engine at altitudes of 25, 000 feet or below. ORecord time engine is operated with low fuel pressure indication. Engine operation with low fuel pressure is limited to a maximum of 10 hours between engine over hauls. FUEL TRANSFER VALVE FAILS TO CLOSE 25B and 25C Aircraft: if the left transfer valve (XF ER-FLLL valve) fails to close after completion of fuel transfer, the crossflow valve will also remain open. In this event, operation of either standby pump will transfer some fuel back into the fuselage tank. Additionally, fuel may gravity flow from both wings into the fuselage tank. • 25C Aircraft: if the right transfer valve (FUS VALVE) fails to close after completion of fuel transfer, the crossflow valve will close normally. In this event, operation of the right standby pump will transfer fuel from the right wing into the fuselage tank. Additionally, fuel may gravity flow from the right wing into the fuselage to nk. 1. Fuel crossflow operations should be conducted with caution as standby pump operation will transfer fuel back into the fuselage. 2. Periodically transfer fuel back into wings. a NOTE Since continuous operation of a standby pump may not ° be possible, a descent to 25,000 feet or below will be required if a jet pump should fail. YMCA Minimum Control Speed, Air The minimum flight speed at which the airplane is controllable with 5° of bank toward the good engine when one engine suddenly becomes inoperative and the remaining engine is operating at takeoff thrust. VMCG Minimum Control Speed, Ground The minimum speed on the ground at which control can be maintained using aerodynamic controls alone, when one engine suddenly becomes inoperative and the remaining engine is operating at takeoff thrust. V 1 Critical Engine Failure Speed The speed at which, due to engine failure or other causes, the pilot may elect to stop or continue the takeoff. If engine failure occurs at V 1, the distance to con tinue the takeoff to 35 feet will not exceed the usable takeoff distance. The dis tance to stop the airplane will not exceed the accelerate-stop distance. V1 must not be less than VMCG or greater than VR. VR Rotation Speed The speed at which rotation is initiated during takeoff to attain takeoff performance. V2 Takeoff Safety Speed The actual speed at 35 feet above the runway surface as demonstrated in flight during single-engine takeoff. V2 is maintained to 1,500 feet above the runway or until clear of obstacle to produce the maximum climb gradient. V2 must not be less than 1.2 times the stalling speed, less than 1.1 times VMCA, or less than VR plus an increment in speed attained prior to reaching 35 feet above the runway. VSO The stalling speed in the landing configuration. VS1 The stalling speed in the appropriate gear/flap configuration. VREF Landing Approach Speed The airspeed equal to 1.3 VS0 with the airplane in the landing configuration. INTRODUCTION Airplane performance is affected by many variables in addition to the usual ones, such as airplane weight, density altitude, runway length, etc. In the case of the Learjet 20 series, performance is also affected by type of engine, type of wing, design, and type of brakes installed. GENERAL Most performance data for all approved operating conditions is provided in chart form in the Performance section of the approved Airplane Flight Manual. Except for Models 23 and 24, climb, cruise, and descent data are provided in the Learjet Pilot's Manual. Airplane performance data is also provided in tabular form in the Pilot's Manual and the aircrew checklist. However, the effects of wind, runway gradient, anti-skid-off, and anti-ice on conditions are not compensated for in the tabular data in the crew checklists or the Pilot's Manual. Therefore, if any of the above are factors, the AFM charts should be used for flight planning. Weight and balance data is presented in the Weight and Balance section of the AFM. Assumed Conditions The performance data presented for each phase of operation is based on certain assumed conditions. Assumed conditions, along with the description of the corresponding charts, are given in this chapter. Standard Conditions Standard conditions which apply to all performance calculations are: • Ambient temperature and pressure altitude • Winds • Gross weight • Runway gradients • Anti-ice on or off • Anti-skid on or off • Flaps 8°, 10°, 20°, or 10° overspeed for takeoff and 40° for landing Flight§afety intarnetionel Runway Change in runway elevation per 100 feet of runway length. The Gradient values given are positive for uphill and negative for downhill gradients. Gradient The ratio of the change in height during a portion of the climb to of Climb the horizontal distance traversed in the same interval. Gross The climb gradient that the airplane can actually achieve given Climb Gradient ideal conditions. Net Climb The gross climb gradient reduced by 0.8% during the takeoff phase. Gradient This conservatism is required by FAR 25 for terrain clearance deter mination to account for variables encountered in service. Climb Segments (In Order of Occurrence) First The first segment climb begins from the point at which the airplane Segment becomes airborne and ends at the point at which the landing gear is Climb fully retracted. Refer to Table PER-1 for the applicable configura tion. Gross climb gradient with one engine inoperative and the other engine at takeoff thrust must be positive, without ground effect. This requirement is satisfied by compliance with the applicable Takeoff Weight Limits Chart. The second segment begins at the end of gear retraction and continues to height above the runway of 1,500 feet and V2 speed. It is noted that the second segment in Figure PER-1 is shown only to a height of 400 feet. However, this is a minimum requirement and for simplified flight planning, the Takeoff Flight Path charts shown in the Flight Manual present the second segment required gradients to the 1,500-foot point for obstacle clearance considerations. Table PER-1. CLIMB CONFIGURATIONS TYPE OF NO. OF ENGINES THRUST FLAP GEAR CLIMB OPERATING SETTING POSITION First Segment 1 Takeoff Takeoff Down Second Segment 1 Takeoff Takeoff Up Final Segment 1 Max cont Up Up Enroute 1 Max cont Up Up Approach 1 Takeoff 20° Up Landing 2 Takeoff Dn-40° Down PER-4 FOR TRAINING PURPOSES ONLY Revision .01 Final segment climb begins t the end of the second segment and ends at a height of at least 1,500 feet AGL. The gross climb gradient must be at least 1.2% with one engine not operating and the other engine at maximum continuous thrust. This requirement is sat tisfied by compliance with the applicable Takeoff Weight Limits chart. Airspeed for this segment is 1.25 VS. The final segment climb gradients are presented for pilot's reference and are not used in the takeoff path calculation. Enroute Enroute climb is a climb with flaps up, landing gear retracted, and Climb maximum continuous thrust on one engine. There is no minimum re quirement for enroute climb gradients. The enroute net climb gra dients are presented for pilot's reference. Velocity is presented in the Enroute Climb Speed Schedule chart. The Approach Climb is made from a missed or aborted approach. With the airplane in the appropriate configuration (flaps 20°, gear up, and takeoff thrust on one engine), the gross climb gradient must be at least 2.1 %. This requirement is satisfied by compliance with the Landing Weight Limits chart. Airspeed for this maneuver is 1.3 VS,. Landing This climb is made from an aborted landing. When the airplane is in Climb the landing climb configuration (flaps and gear down, takeoff thrust on both engines), the gross climb gradient must be at least 3.2%. This requirement is satisfied by compliance with the Landing Weight Limits chart. Landing climb airspeed is 1.3 VSO. TAKEOFF PERFORMANCE Wind Components Headwind, tailwind, and crosswind components can be calculated by using the Wind Component chart found in the General section of the AM "Performance Data" chapter. This value is entered on the Takeoff Worksheet (Figure PER-2). Maximum Allowable Takeoff Weight The maximum allowable takeoff weight at the start of takeoff roll is limited by the most restrictive of the following requirements: • Maximum certificated takeoff weight • Maximum takeoff weight to meet minimum single-engine climb gradient requirements and not exceed brake energy limits (climb or brake energy limited) • Maximum takeoff weight for runway length available • Maximum takeoff weight for obstacle clearance • Maximum landing weight for destination airport