Pilot’s Operating Handbook and Flight Training Supplement Cessna Model 162

This POH consists of ten sections that cover all operational aspects of a standard equipped airplane. Section 10 contains the supplements which provide amended operating limitations, operating procedures, performance data and other necessary information for airplanes conducting special operations for both standard and optional equipment installed in the airplane. Supplements are individual documents, and may be issued or revised without regard to revision dates which apply to the POH itself. These supplements contain a Log of Effective Pages, which should be used to determine the status of each supplement. 162PHUS-03 vi CESSNA INTRODUCTION MODEL 162 GARMIN G300 U.S. ORIGINAL ISSUE AND REVISIONS This Pilot’s Operating Handbook is comprised of the original issue and any subsequent revisions. To make sure that information in this manual is current, the revisions must be incorporated as they are issued. As revisions are issued, they will be noted in the Log of Effective Pages. The part number of this manual has also been designed to further aid the owner/operator in determining the revision level of any POH. Refer to the example below for a breakdown: It is the responsibility of the owner to maintain this POH in a current status when it is being used for operational purposes. Owners should contact a Cessna Service Station whenever the revision status of their POH is in question. Revisions are distributed to owners of U.S. Registered aircraft according to FAA records at the time of revision issuance, and to Internationally Registered aircraft according to Cessna Owner Advisory records at the time of issuance. Revisions should be read carefully upon receipt and incorporated in this POH. 162 PHUS -00 Revision Level (Original Issue) Manual (Pilot’s Operating Handbook, U.S.) (Serials 16200001 and On) Airplane Model - (162) 162PHUS-00 vii INTRODUCTION CESSNA MODEL 162 GARMIN G300 U.S. REVISION FILING INSTRUCTIONS REGULAR REVISIONS Pages to be removed or inserted in the Pilots’ Operating Handbook are determined by the Log of Effective Pages located in this section. This log contains the page number and revision level for each page within the POH. As revisions to the POH occur, the revision level on effected pages is updated. When two pages display the same page number, the page with the latest revision level shall be inserted into the POH. The revision level on the Log Of Effective Pages shall also agree with the revision level of the page in question. TEMPORARY REVISIONS Under limited circumstances, temporary revisions to the POH may be issued. These temporary revisions are to be filed in the applicable section in accordance with filing instructions appearing on the first page of the temporary revision. Temporary Revisions will remain current until they have either been incorporated into the next POH revision or another temporary revision has been issued that supersedes that temporary revision. Each temporary revision is issued with a current List of Temporary Revisions that is to be inserted opposite the first page of the Log of Effective Pages in the front of the POH and will supersede any previously issued List of Temporary Revisions. This list is used to track the status of temporary revisions issued against this POH and is to be removed and discarded at the next revision to the POH. Removal of temporary revisions from the POH is accomplished per the removal instructions on each temporary revision. 162PHUS-00 viii CESSNA INTRODUCTION MODEL 162 GARMIN G300 U.S. IDENTIFYING REVISED MATERIAL A bar will extend the full length of deleted, new, or revised text added on new or previously existing pages. This bar will be located adjacent to the applicable text in the margin on the left side of the page. A bar in the footer will indicate a revision to the header/footer, a new page, format or spelling/grammar changes and/or that information has slipped to or from that page. A bar located adjacent to the figure number in the margin on the left side of the page will be used to indicate that the figure number only has changed. An asterisk located at the end of a figure number will be used to indicate that an illustration has been revised or is all new material (Ex: Figure 3-4*). All revised pages will carry the revision number opposite the page number on the applicable page. A list of revisions is located at the beginning of the Log Of Effective Pages. ix 162PHUS-00 INTRODUCTION CESSNA MODEL 162 GARMIN G300 U.S. WARNINGS, CAUTIONS AND NOTES Throughout the text, warnings, cautions and notes pertaining to airplane handling and operations are utilized. These adjuncts to the text are used to highlight or emphasize important points. WARNING OPERATING PROCEDURES, TECHNIQUES, ETC., WHICH CAN RESULT IN PERSONAL INJURY OR LOSS OF LIFE IF NOT CAREFULLY FOLLOWED. CAUTION OPERATION PROCEDURES, TECHNIQUES, ETC., WHICH CAN RESULT IN DAMAGE TO EQUIPMENT IF NOT CAREFULLY FOLLOWED. NOTE An operating procedure, technique, etc., which is considered essential to emphasize. 162PHUS-00 x CESSNA INTRODUCTION MODEL 162 GARMIN G300 U.S. LOG OF EFFECTIVE PAGES Use this page to determine the currency and applicability of your POH. Pages affected by the current revision are indicated by an asterisk (*) preceding the pages listed under the Page Number column. (Continued Next Page) Revision Level Date of Issue Original Issue 22 July 2009 Revision 2 26 April 2010 Revision 1 2 November 2009 Revision 3 28 September 2010 Page Number Page Status Revision Number * Title Revised 3 * i/ii Revised 3 iii Revised 1 * iv thru vi Revised 3 vii thru x Original 0 * xi thru xv/xvi Revised 3 1-1/1-2 Original 0 1-3 thru 1-4 Revised 1 * 1-5 Revised 3 1-6 thru 1-7 Original 0 * 1-8 Revised 3 * 2-1 thru 2-26 Revised 3 * 2-27 thru 2-55/2-56 Added 3 162PHUS-03 xi INTRODUCTION CESSNA MODEL 162 GARMIN G300 U.S. LOG OF EFFECTIVE PAGES (Continued) (Continued Next Page) Page Number Page Status Revision Number 3-1/3-2 Revised 1 3-3 Original 0 3-4 Revised 2 3-5 Revised 1 * 3-6 Revised 3 3-7 Original 0 * 3-8 thru 3-9 Revised 3 3-10 Revised 1 * 3-11 thru 3-13 Revised 3 3-14 Original 0 * 3-15 thru 3-18 Revised 3 3-19/3-20 Revised 1 4-1/4-2 thru 4-4 Original 0 * 4-5 thru 4-7 Revised 3 4-8 Original 0 4-9 thru 4-10 Revised 1 4-11 thru 4-14 Original 0 4-15 thru 4-16 Revised 1 4-17 thru 4-22 Original 0 4-23 thru 4-26 Revised 1 4-27 thru 4-30 Original 0 5-1/5-2 Original 0 5-3 thru 5-9 Revised 1 5-10 thru 5-11 Original 0 5-12 thru 5-16 Revised 1 5-17 thru 5-18 Original 0 162PHUS-03 xii CESSNA INTRODUCTION MODEL 162 GARMIN G300 U.S. LOG OF EFFECTIVE PAGES (Continued) (Continued Next Page) Page Number Page Status Revision Number 6-1 thru 6-3/6-4 Original 0 6-5 Revised 2 6-6 Original 0 6-7 thru 6-8 Revised 1 6-9 thru 6-12 Original 0 * 6-13 Revised 3 6-14 Original 0 * 6-15 Revised 3 6-16 Original 0 6-17 thru 6-18 Revised 1 6-19 Original 0 6-20 Revised 1 6-21 thru 6-22 Original 0 6-23 Revised 1 * 6-24 Revised 3 6-25 thru 6-26 Revised 1 6-27 thru 6-30 Original 0 * 6-31 Revised 3 6-32 thru 6-33 Revised 1 6-34 Original 0 * 6-35 Revised 3 6-36 thru 6-39 Original 0 6-40 Revised 1 * 7-1 thru 7-3 Revised 3 7-4 Original 0 * 7-5 Revised 3 7-6 thru 7-7 Original 0 7-8 thru 7-9 Revised 1 * 7-10 Revised 3 7-11 thru 7-12 Original 0 * 7-13 thru 7-20 Revised 3 162PHUS-03 xiii INTRODUCTION CESSNA MODEL 162 GARMIN G300 U.S. LOG OF EFFECTIVE PAGES (Continued) Page Number Page Status Revision Number 7-21 Revised 1 7-22 thru 7-24 Original 0 7-25 thru 7-26 Revised 1 7-27 thru 7-29 Original 0 7-30 Revised 1 * 7-31 Revised 3 7-32 thru 7-41 Revised 1 7-42 thru 7-43 Original 0 7-44 Revised 1 * 7-45 thru 7-46 Revised 3 8-1 Original 0 8-2 thru 8-4 Revised 1 * 8-5 Revised 3 8-6 Revised 1 * 8-7 thru 8-8 Revised 3 8-9 Revised 1 * 8-10 Revised 3 8-11 thru 8-12 Original 0 8-13 Revised 1 8-14 thru 8-18 Original 0 8-19 thru 8-23 Revised 1 * 8-24 Revised 3 8-25/8-26 Revised 1 * 9-1/9-2 Revised 3 9-3 Revised 1 9-4 Original 0 * 9-5 thru 9-7 Revised 3 9-8 Revised 1 10-1/10-2 Original 0 162PHUS-03 xiv CESSNA INTRODUCTION MODEL 162 GARMIN G300 U.S. TABLE OF CONTENTS SECTION GENERAL . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .1 AIRPLANE AND SYSTEM DESCRIPTION . . . . . . . . .2 OPERATING LIMITATIONS . . . . . . . . . . . . . . . . . . . . .3 WEIGHT AND BALANCE/EQUIPMENT LIST . . . . . . .4 PERFORMANCE . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5 EMERGENCY PROCEDURES . . . . . . . . . . . . . . . . . .6 NORMAL PROCEDURES . . . . . . . . . . . . . . . . . . . . . .7 AIRPLANE HANDLING, SERVICE AND MAINTENANCE . . . . . . . . . . . . . . . . . . . . . . . . .8 PLACARDS AND MARKINGS . . . . . . . . . . . . . . . . . . .9 SUPPLEMENTARY INFORMATION . . . . . . . . . . . . .10 162PHUS-03 xv/xvi CESSNA SECTION 1 MODEL 162 GENERAL GARMIN G300 U.S. GENERAL TABLE OF CONTENTS Page Three View - Normal Ground Attitude . . . . . . . . . . . . . . . . . . . . . . . . 1-3 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-5 Descriptive Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-5 Engine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-5 Propeller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-5 Fuel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-6 Fuel Capacity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-6 Oil . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-7 Oil Specification. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-7 Oil Capacity. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-7 Maximum Certificated Weights . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-8 Maximum Weight In Baggage Compartment . . . . . . . . . . . . . . . . 1-8 Standard Airplane Weights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-8 Cabin And Entry Dimensions . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-8 Baggage Space And Entry Dimensions . . . . . . . . . . . . . . . . . . . . 1-8 Specific Loadings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1-8 1-1/1-2 162PHUS-00 CESSNA SECTION 1 MODEL 162 GENERAL GARMIN G300 U.S. THREE VIEW - NORMAL GROUND ATTITUDE Figure 1-1* (Sheet 1 of 2) 1-3 162PHUS-01 SECTION 1 CESSNA GENERAL MODEL 162 GARMIN G300 U.S. THREE VIEW - NORMAL GROUND ATTITUDE NOTE • Wing span shown with standard strobe lights installed. • Wheel base length is 62.40 inches (1.58 m). • Propeller ground clearance is 8.50 inches (215.90 mm). • Wing area is 120.0 square feet (11.15 sq. m). Figure 1-1* (Sheet 2) 162PHUS-01 1-4 CESSNA SECTION 1 MODEL 162 GENERAL GARMIN G300 U.S. INTRODUCTION This POH contains 10 sections, and includes the material required to be furnished to the pilot by American Society for Testing and Materials International (ASTM) standards F2245 for Light Sport Aircraft (LSA). It also contains supplemental data supplied by Cessna Aircraft Company. Section 1 provides basic data and information of general interest. DESCRIPTIVE DATA ENGINE Number of Engines: 1 Engine Manufacturer: Teledyne Continental Motors Engine Model Number: O-200-D Engine Type: Normally aspirated, direct drive, air-cooled, horizontally opposed, carburetor equipped, four cylinder engine with 201.0 cu. in. displacement. Horsepower Rating and Engine Speed: 100 rated BHP at 2750 RPM PROPELLER Propeller Manufacturer: McCauley Propeller Systems Propeller Model Number: Standard Composite Propeller - 1L100/LSA6754 Optional Aluminum Propeller - 1A162/TCD6754 Number of Blades: 2 Propeller Diameter: 67 inches (1.70 m) Propeller Type: Fixed Pitch (Continued Next Page) 1-5 162PHUS-03 SECTION 1 CESSNA GENERAL MODEL 162 GARMIN G300 U.S. DESCRIPTIVE DATA (Continued) FUEL WARNING USE OF UNAPPROVED FUELS MAY RESULT IN DAMAGE TO THE ENGINE AND FUEL SYSTEM COMPONENTS, RESULTING IN POSSIBLE ENGINE FAILURE. Approved Fuel Grades (and Colors): 100LL Grade Aviation Fuel (Blue) 100 Grade Aviation Fuel (Green) NOTE Isopropyl alcohol or Diethylene Glycol Monomethyl Ether (DiEGME) may be added to the fuel supply in accordance to TCM Service Information Letter (SIL99-2B). Refer to Section 8 for additional information. FUEL CAPACITY Total Capacity . . . . . . . . . . . . . . . . . 25.46 U.S. GALLONS (96.34 l) Total Usable . . . . . . . . . . . . . . . . . . . 24.00 U.S. GALLONS (90.82 l) Total Capacity Each Tank . . . . . . . . 12.73 U.S. GALLONS (48.17 l) Total Usable Each Tank . . . . . . . . . . 12.00 U.S. GALLONS (45.41 l) NOTE • To ensure maximum fuel capacity and minimize crossfeeding when refueling, always park the airplane in a wings level, normal ground attitude. Refer to Figure 1- 1 for normal ground attitude dimensions. • The fuel filler assembly is equipped with indicator tabs for 3/4, 1/2 and 1/4 fuel quantities. • Maximum full capacity is indicated when fuel reaches the upper hole of the indicator tab. This fuel level allows for proper thermal expansion. Filling the fuel tank above the upper hole eliminates expansion space resulting in fuel venting overboard through the fuel vent. (Continued Next Page) 162PHUS-00 1-6 CESSNA SECTION 1 MODEL 162 GENERAL GARMIN G300 U.S. DESCRIPTIVE DATA (Continued) OIL OIL SPECIFICATION SAE J1966 Aviation Grade Non-Dispersant Mineral Oil: Used when the airplane was delivered from the factory and should be used to replenish the supply during the first 25 hours. This oil should be drained and the filter changed after the first 25 hours of operation. Refill the engine with SAE J1966 Aviation Grade Non-Dispersant Mineral Oil and continue to use until a total of 50 hours has accumulated or oil consumption has stabilized. SAE J1899 Aviation Grade Ashless Dispersant Oil: Oil conforming to Teledyne Continental Motors (TCM) Service Information Letter SIL99- 2B, and all revisions and supplements thereto, must be used after first 50 hours or oil consumption has stabilized. RECOMMENDED VISCOSITY FOR TEMPERATURE RANGE Multiviscosity or straight grade oil may be used throughout the year for engine lubrication. Refer to the following table for temperature versus viscosity ranges. NOTE When operating temperatures overlap, use the lighter grade of oil. OIL CAPACITY Sump. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.0 U.S. QUARTS (4.73 l) Total. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.5 U.S. QUARTS (5.20 l) Minimum Operating Quantity . . . . . . . . . . 3.5 U.S. QUARTS (3.31 l) (Continued Next Page) Temperature SAE J1966 Non-Dispersant Mineral Oil SAE Grade SAE J1899 Ashless Dispersant Oil SAE Grade Above 4°C (40°F) 50 50 or 15W-50 or 25W-60 Below 4°C (40°F) 30 30 or 15W-50 or 25W-60 All Temperatures M20W-50 15W-50, 20W- 50 or 25W-60 1-7 162PHUS-00 SECTION 1 CESSNA GENERAL MODEL 162 GARMIN G300 U.S. DESCRIPTIVE DATA (Continued) MAXIMUM CERTIFICATED WEIGHTS Ramp Weight:. . . . . . . . . . . . . . . . . . . . . . . . 1324 POUNDS (600.6 kg) Takeoff Weight . . . . . . . . . . . . . . . . . . . . . . . 1320 POUNDS (598.8 kg) Landing Weight . . . . . . . . . . . . . . . . . . . . . . 1320 POUNDS (598.8 kg) MAXIMUM WEIGHT IN BAGGAGE COMPARTMENT Baggage Area (Station 155 to 190) . . . . . . . 50 POUNDS (22.68 kg) STANDARD AIRPLANE WEIGHTS Standard Empty Weight . . . . . . . . . . . . . . . . . 834 POUNDS (378.3 kg) Maximum Useful Load . . . . . . . . . . . . . . . . . . 490 POUNDS (222.3 kg) CABIN AND ENTRY DIMENSIONS Detailed dimensions of the cabin interior and entry door openings are illustrated in Section 4. BAGGAGE SPACE AND ENTRY DIMENSIONS Dimensions of the baggage area are illustrated in detail in Section 4. SPECIFIC LOADINGS Wing Loading . . . . . . . . . . . . . . . . . . . . . 11.0 lbs/sq. ft. (53.7 kg/sq. m) Power Loading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.2 lbs/HP 162PHUS-03 1-8 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. AIRPLANE AND SYSTEMS DESCRIPTION TABLE OF CONTENTS Page Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3 Airframe . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-3 Flight Controls and Trim System . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-4 Electric Trim System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-9 Instrument Panel, Flight and System Instruments . . . . . . . . . . . . . . 2-10 Instrument Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-12 Ground Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-14 Wing Flap System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-15 Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-16 Baggage Area . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-17 Seats, Seat Belt and Shoulder Harness . . . . . . . . . . . . . . . . . . . . . 2-18 Cabin Doors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-21 Engine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-24 Engine Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-24 Engine Instruments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-25 Tachometer (RPM) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-25 Oil Pressure (OIL PSI) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-26 Oil Temperature (OIL °F) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-27 Exhaust Gas Temperature (EGT °F) (if installed) . . . . . . . . . . . . . 2-27 Carburetor Temperature (CARB °F) . . . . . . . . . . . . . . . . . . . . . . 2-28 New Engine Break-In And Operation . . . . . . . . . . . . . . . . . . . . . . 2-28 Engine Lubrication System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-29 Ignition And Starter System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-29 Air Induction System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-30 Exhaust System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-30 Cooling System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-30 Propeller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-31 Fuel System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-31 Fuel Distribution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-31 Fuel Indicating System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-33 Fuel Venting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-34 Reduced Tank Capacity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-35 Fuel Drain Valves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-36 (Continued Next Page) 162PHUS-03 2-1 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. TABLE OF CONTENTS (Continued) Page Brake System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-36 Electrical System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-38 G300 Crew Alerting System (CAS) . . . . . . . . . . . . . . . . . . . . . . . 2-40 Master Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-40 Electrical System Monitoring . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-40 Main Battery Current (Amps). . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-40 System Voltage (Volts) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-41 Circuit Breakers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-41 12V Power Outlet . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-43 External Power Receptacle (if installed) . . . . . . . . . . . . . . . . . . . 2-44 Lighting Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-45 Exterior Lighting. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-45 Interior Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-45 Cabin Heating and Ventilating System . . . . . . . . . . . . . . . . . . . . . . 2-46 Pitot-Static System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-48 Stall Warning System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-48 Standard Avionics. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-49 Garmin Display Units (GDU). . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-49 Air Data, Attitude and Heading Reference System (ADAHRS) and Magnetometer (GRS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-51 Garmin GTX 327 Transponder . . . . . . . . . . . . . . . . . . . . . . . . . . 2-52 Garmin SL 40 VHF Comm Transceiver . . . . . . . . . . . . . . . . . . . 2-52 XM Weather and Radio Data Link (GDL) (if installed) . . . . . . . . 2-52 Avionics Support Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-53 Avionics Cooling Fans . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-53 Antennas. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-53 Microphone And Headset Installations . . . . . . . . . . . . . . . . . . . . 2-54 Auxiliary Audio Input Jack (if installed) . . . . . . . . . . . . . . . . . . . . 2-54 Cabin Features. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-55/2-56 Emergency Locator Transmitter (ELT) . . . . . . . . . . . . . . . . 2-55/2-56 Cabin Fire Extinguisher . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-55/2-56 162PHUS-03 2-2 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. INTRODUCTION This section provides description and operation of the airplane and its systems. Some equipment described herein is optional and may not be installed in the airplane. Refer to Section 10, Supplements, for details of other optional systems and equipment. AIRFRAME The airplane is an all metal, two-place, high wing, single-engine airplane equipped with tricycle landing gear and is designed for sport flying and introductory training purposes. The construction of the fuselage is a conventional formed sheet metal bulkhead, stringer, and skin semi-monocoque construction. Wing struts and main landing gear strut assemblies attach to an I-beam structure under the baggage area behind the seats. Nose gear and engine mount assembly attach to the firewall separating the cabin from engine compartment. The aft fuselage empennage (tail assembly) consists of horizontal and vertical stabilizers, the rudder, a left elevator, and a right elevator with elevator trim tab. Attaching on either side the fuselage above the cabin, the wing contains fuel tanks at the wing root between the forward and aft spars. Aft of the rear wing spar are mounted the flaps (inboard) and ailerons (outboard). In addition to the fuselage carry-through spars, the wing is attached by forward-spar-to-fuselage struts. The wing, empennage, and flight control surfaces are also made of sheet metal spars, ribs, and skin semi-monocoque construction with balance weights incorporated into the rudder, elevators, and ailerons. 162PHUS-03 2-3 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. FLIGHT CONTROLS AND TRIM SYSTEM The airplane's flight control system consists of conventional aileron, rudder, and elevator control surfaces, refer to Figure 2-3. The control surfaces are manually operated through cables and mechanical linkage using a control stick for the ailerons and elevator, and rudder/brake pedals for the rudder. Dual under panel control sticks are used for aileron and elevator control, refer to Figure 2-1. The under panel control sticks mimic the control movement of a floor mounted control stick while providing ease of entry to the cabin. The control stick location is designed so that the pilot's hand naturally falls on the control stick with the outboard arm on the door armrest. The control lock provided with the airplane pins the left control stick through a bracket attached to the instrument panel to immobilize the ailerons and elevator. Rudder is held centered by the springs used for increasing rudder pedal force. CONTROL STICK Figure 2-1* 162PHUS-03 2-4 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. FLIGHT CONTROLS AND TRIM SYSTEM (Continued) Rudder pedals, with toe brakes, provide rudder control through forward and aft individually adjustable pilot and copilot pedal assemblies. Rudder pedals should be adjusted before each flight, so that it is possible to have full brake pedal deflection when the same side rudder is fully deflected. Rotating the rudder pedal adjuster knob clockwise will move the rudder pedals forward allowing more leg room for the pilot or copilot. Rotating the knob counter-clockwise will move the pedals aft bringing the pedals closer to the pilot. RUDDER PEDALS Figure 2-2* 162PHUS-03 2-5 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. FLIGHT CONTROLS AND TRIM SYSTEM Figure 2-3* (Sheet 1 of 2) 162PHUS-03 2-6 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. FLIGHT CONTROLS AND TRIM SYSTEM Figure 2-3* (Sheet 2) 162PHUS-03 2-7 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. FLIGHT CONTROLS AND TRIM SYSTEM (Continued) During the preflight walk-around inspection it is possible to move the ailerons into an over-centered position by moving the aileron down from the centered or faired with the flaps UP flaps position. The down- deflected aileron may appear to be stuck or difficult to move up. This is normal characteristic of the Cessna 162 ailerons system that only occurs when the ailerons are moved down from outside the airplane. It is recommended that the aileron movement be checked by moving the ailerons up from the flaps UP faired position then returned to this position. However if an aileron is externally moved down and becomes over-centered, simply re-center the cockpit control stick and continue the walk-around inspection. CAUTION AILERON OVER CENTERING ONLY OCCURS WHEN AILERONS ARE MOVED FROM OUTSIDE THE AIRPLANE. CONTROL STICK STIFFNESS OR STICKING IN ANY DIRECTION OF ANY CONTROL SURFACE IS NOT NORMAL AND SHOULD BE INSPECTED BY MAINTENANCE PERSONNEL PRIOR TO FLIGHT. 162PHUS-03 2-8 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. FLIGHT CONTROLS AND TRIM SYSTEM (Continued) ELECTRIC TRIM SYSTEM An elevator trim tab is located on the right elevator. Trim tab operation is by direct linkage to a elevator mounted electric servo motor controlled by a control stick mounted electric trim switch. Sliding the trim switch up to the TRIM DN position will bring the nose of the airplane down while sliding the trim switch down to the TRIM UP position will trim the nose of the airplane up. Elevator trim tab position is displayed on the Engine Indicating System (EIS) area of the Garmin G300 display. ELECTRIC TRIM SYSTEM Figure 2-4* 162PHUS-03 2-9 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. INSTRUMENT PANEL, FLIGHT AND SYSTEM INSTRUMENTS The Model 162 Skycatcher instrument panel is oriented around the Garmin G300 Flight and Engine Instrumentation Display(s). Directly in front of the pilot is the G300 Primary Flight Display (PFD) which has all flight, navigation, and engine instruments depicted in an electronic format on a liquid crystal display. A vast amount of additional flight and navigational information is accessible through the G300 system’s various selectable pages. Left of the PFD on the upper panel is the optional AUTOPILOT power switch (if installed) and below it is the Ignition/Magneto key switch. NOTE Refer to the Garmin Pilot's Guide for details and operating instructions of the G300 Avionics System. The electrical system switch/circuit breaker panel, located to the right of the PFD, has switches in the top row with two rows of circuit breakers underneath. On the right end of the switch row is the PANEL LIGHT control knob. Under the electrical system panel is the radio stack with the Garmin SL40 Communication radio over the Garmin GTX 327 Transponder. An optional Garmin Multifunction Display (MFD) (if installed) is located to the right of the electrical system switch/circuit breaker panel. The MFD (when installed) takes over the moving map and engine instrument functions from the PFD as well as other display page functions. On the far right side of the panel is the remote ELT switch and optional intercom control panel. (Continued Next Page) 162PHUS-03 2-10 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. INSTRUMENT PANEL, FLIGHT AND SYSTEM INSTRUMENTS (Continued) Engine controls are located below the radio stack and MFD display (if installed) along the bottom center of the instrument panel. The controls are (left to right): Carb Heat, Throttle Control, Mixture Control, and Cabin Heat Control. The Parking Brake T-Handle, Fuel Shut-off Control Knob and optional Fuel Primer Control Knob (if installed) are located beneath the engine controls along the top of the vertical pedestal. Headset Jacks, Auxiliary 12 Volt Cabin Power Outlet and Audio Input Jack are positioned further down on the vertical pedestal. Behind the instrument panel are the Autopilot Controller and Servo Units for the Optional Autopilot System (if installed), Garmin ADAHRS (Air Data, Attitude, and Heading Reference System) Unit, and the Secondary Battery. The ADAHRS unit is a multifunctional processing unit which takes pitot-static pressures from the pitot tube on the left wing and the static source button aft of the left cowl for altitude, airspeed, and vertical speed information, tailcone mounted magnetometer data for heading, and engine data for display on the Engine Indication System (EIS) portion of the displays. Fuel indicators mounted in each wing root utilize floating balls for visual indication of fuel quantity. Two fuel placards are used to differentiated between level-flight (LVL FLT) and on-ground (GRND) fuel quantities. The MIN/TO reference mark on the GRND side indicates the ¼ fuel requirement for take-off. In-tank fuel tabs are equipped with indicator holes for full, ¾, ½, and the bottom of the tab indicates ¼ (minimum takeoff). See Fuel System description section for more information. The optional magnetic compass (if installed) is mounted to the airplane structure above the center of the windshield. 162PHUS-03 2-11 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. INSTRUMENT PANEL Figure 2-5* 162PHUS-03 2-12 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. INSTRUMENT PANEL 1. AUTOPILOT Power Switch (If installed) 2. Autopilot Controller Cover Plate - Autopilot controller located behind instrument panel (If Installed). 3. Primary Flight Display (PFD) 4. MASTER Switch (ALT and BAT) 5. Switch Control Panel 6. Circuit Breaker Panel 7. PANEL LIGHTS Control Knob 8. Multifunction Display (MFD) (If installed) 9. ELT Remote Switch/Annunciator Panel (If installed) 10. PM3000 Intercom System Control Panel (If installed) 11. CABIN HEAT Control Knob 12. MIXTURE Control Knob 13. FUEL SHUTOFF Valve Control Knob 14. Headset Jacks 15. 12V POWER OUTLET (7.5A) 16. AUX AUDIO INPUT Jack (If installed) 17. FUEL PRIMER Control Knob (If installed) 18. PARKING BRAKE Lock T-Handle 19. THROTTLE Control Knob (With Friction Lock) 20. CARB HEAT Control Knob 21. Garmin GTX 327 Transponder (If installed) 22. Garmin SL 40 VHF Comm Transceiver (If installed) 23. MAGNETOS/START Switch 162PHUS-03 2-13 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. GROUND CONTROL Effective ground control while taxiing is accomplished through a free- castoring nosewheel by using the rudder pedals and differential braking; left rudder pedal to steer left and right rudder pedal to steer right. By applying either left or right brake, the degree of turn may be increased up to 59° each side of center. Moving the airplane by hand is most easily accomplished by attaching a towbar to the nosewheel. If a towbar is not available, or pushing is required, use the wing struts as push points. Do not use the vertical or horizontal surfaces to move the airplane. If the airplane is to be towed by vehicle, never turn the nosewheel more than 57.5° either side of center or structural damage to the nose gear could result. The minimum turning radius of the airplane, using differential braking during taxi, is approximately 38 feet. To obtain a minimum radius turn during ground handling, the airplane may be rotated around either main landing gear by pressing down on a tailcone bulkhead just forward of the horizontal stabilizer to raise the nosewheel off the ground. Care should be exercised to ensure that pressure is exerted only on the bulkhead area and not on skin between the bulkheads. Pressing down on the horizontal stabilizer to raise the nosewheel off the ground is not recommended. 162PHUS-03 2-14 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. WING FLAP SYSTEM The single slot type wing flaps, refer to Figure 2-6, are manually operated down by a flap control handle located between the seats and returned to the faired UP position by air load and return spring assist. Flap control handle can be repositioned to the following detent positions only after depressing the release button in the end of the flap control handle. Flaps UP . . . . . . . . . . . . . . . . . . . . . . . Fully Stowed (Flush Position) Flaps 10º . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .First Notch Flaps 25º . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Second Notch Flaps FULL . . . . . . . . . . . . . . . . . . . . . . Last Notch (Full up Position) Figure 2-6* 162PHUS-03 2-15 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. LANDING GEAR The tricycle type landing gear on the Model 162 features tapered spring steel main gear strut assemblies and a spring steel nose gear with a urethane spring for shock absorption, refer to Figure 2-7. The nose wheel is free-castoring with directional control done via differential braking. This permits very tight turning and maneuvering in confined spaces. Full rudder deflection can assist directional control when taxiing. Figure 2-7* 162PHUS-03 2-16 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. BAGGAGE AREA The baggage area is accessed by tilting the seats forward and can accommodate up to 50 pounds of cargo. Tie-down rings are supplied to secure cargo in the baggage area. A aft bulkhead close-out net is installed separating the baggage area from the aft tailcone. The net is easily removed for maintenance access, but must be installed for flight to prevent any loose objects or baggage from migrating into the tailcone area causing control and/or center of gravity issues. Figure 2-8* 162PHUS-03 2-17 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. SEATS, SEAT BELT AND SHOULDER HARNESS SEATS The seating arrangement consists of two forward pivoting crew seats for the pilot and copilot, which allow easy access to the baggage area. Both seat positions are equipped with manually adjustable seat and shoulder belts, refer to Figure 2-9. SEAT BELTS The seat belts are attached to airplane structure on the floorboard. The buckle portion of the seat belt is located inboard of each seat, while the link portion is located outboard of each seat location, refer to Figure 2- 9. To use the seat belts, lengthen the link portion of the lap belt as needed by grasping the sides of the link and pulling outward against the belt. Insert and lock the belt link into the buckle. Positive locking has occurred when a distinctive “snap” sound is heard. Tighten belt as needed to achieve proper fit. Proper installation of the lap belt can be verified by ensuring that the lap belt is snug and low on the waist as worn normally during flight. To release the seat belts, grasp the top of the buckle opposite the link and pull outward. CAUTION IF IT BECOMES NECESSARY TO REMOVE SEAT BEAT OR SHOULDER HARNESS DURING FLIGHT USE CARE TO MAKE SURE THAT SEAT BELT OR SHOULDER HARNESS DOES NOT FALL UNDER THE SEAT WHERE IT MAY NOT BE RETRIEVABLE WHILE SEATED IN THE AIRPLANE. (Continued Next Page) 162PHUS-03 2-18 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. SEATS, SEAT BELT AND SHOULDER HARNESS (Continued) SHOULDER HARNESSES Each shoulder harness, refer to Figure 2-9, is attached to the airplane structure at the aft bulkhead. To use a seat shoulder harness fasten and adjust the seat lap belt first. Lengthen the shoulder harness as required by pulling on the connecting link on the end of the harness and the narrow release strap. Snap the connecting link firmly onto the retaining stud on the seat belt link. Then adjust shoulder harness to proper length. A properly adjusted harness will permit the occupant to lean forward enough to sit completely erect, but prevent excessive forward movement and contact with objects during sudden deceleration. Also, the pilot will want the freedom to reach all controls easily. Removing the shoulder harness is accomplished by pulling upward on the narrow release strap, and removing the harness connecting link from the stud on the seat belt link. In an emergency, the shoulder harness may be removed by releasing the seat belt first, and allowing the harness, still attached to the link portion of the seat belt, to drop to the side of the seat. CAUTION MAKE SURE THAT THE SEAT BELTS ARE PROPERLY STOWED PRIOR TO CLOSING CABIN DOOR. THE DOOR SEAL AND/OR DOOR STRUCTURE CAN BE DAMAGED IF DOOR COMES IN CONTACT WITH SEAT BELT OR SHOULDER HARNESS BELT LINK WHILE CLOSING. 162PHUS-03 2-19 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. SEAT BELT AND SHOULDER HARNESS Figure 2-9* 162PHUS-03 2-20 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. CABIN DOORS Entry to and exit from the airplane is accomplished through either of two upward opening cabin doors, one on each side of the cabin, refer to Section 4 for cabin and cabin door dimensions. The cabin doors incorporate a gas strut to assist in opening and holding the door in the full up position, refer to Figure 2-10. Each door assembly is equipped with a recessed exterior door handle/latch, interior door latch, fabric- cover door strap/handle, key operated door lock and zippered cabin door storage pocket. WARNING MAKE SURE BOTH EXTERIOR CABIN DOOR LOCK ASSEMBLIES ARE UNLOCKED PRIOR TO FLIGHT. CABIN DOORS CANNOT BE OPENED FROM INSIDE THE AIRPLANE IF EXTERIOR CABIN DOOR LOCK IS ENGAGED. To open the cabin doors from outside the airplane, insert key and unlock the cabin door lock by rotating the lock assembly counterclockwise to the unlocked position. Rotate the door handle downward to the OPEN position as indicated on the placard and gently pull outward on the door. Allow the gas strut to raise door to the fully open position and verify gas strut will support the door assembly and that the door does not come in contact with the wing lower skin. CAUTION PULLING OUT HORIZONTALLY ON THE EXTERNAL DOOR HANDLE/LATCH ASSEMBLY WILL DAMAGE THE DOOR HANDLE MECHANISM AND OUTER DOOR SKIN. To close cabin doors from outside the airplane, lower the cabin door partially down to verify the exterior door handle/latch assembly is in the OPEN position prior to pushing door closed. Make sure that both the forward and aft latch mechanisms are fully seated prior to rotating the exterior door handle/latch assembly to the closed or full up position. Rotate cabin door lock clockwise to secure cabin door. Make sure door is locked by carefully trying to rotate the exterior door handle/latch assembly. (Continued Next Page) 162PHUS-03 2-21 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. CABIN DOORS (Continued) CAUTION THE CABIN DOOR ASSEMBLY SHOULD NOT BE SLAMMED SHUT. THE DOORS DO NOT USE A PRE- CATCH MECHANISM AND SHOULD NOT BE SLAMMED CLOSED AS THIS CAN DAMAGE THE DOOR AND/OR DOOR LATCH ASSEMBLY. To close the cabin doors from inside the airplane, first secure seat belts and make sure door openings are clear of obstructions. Then from a seated position, reach out and grab the cabin door strap and carefully pull the door to the fully closed position. Make sure that both the forward and aft latch mechanisms are fully seated prior to sliding the interior door latch assembly to the LATCHED full forward position. On later models the interior door latch assembly will not move to the LATCHED or full forward position if the forward latch mechanism is not fully engaged. Lightly push on the front and rear lower areas of the door to make sure the door is properly latched and secured for flight. To open the cabin doors from inside the airplane, first hold the cabin door strap and carefully pull in on the strap while sliding the interior door latch assembly to the OPEN full aft position. Carefully push out on the inside of door while holding onto the cabin door strap and slowly allow door to raise to the full up position. CAUTION DURING HIGH WIND CONDITIONS EXTREME CARE SHOULD BE TAKEN TO PREVENT THE CABIN DOOR FROM BLOWING OPEN UPON OPENING, WHICH COULD DAMAGE THE DOOR AND/OR LOWER WING SKIN. If cabin door should accidently come open during flight, leave cabin door open and do not attempt to close, refer to Section 6, Emergency Procedures, DOOR OPEN IN FLIGHT. WARNING INTENTIONAL FLIGHT WITH DOOR(S) OPEN IS PROHIBITED. 162PHUS-03 2-22 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. CABIN DOORS Figure 2-10* 162PHUS-03 2-23 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. ENGINE The airplane is powered by a normally aspirated, direct drive, air- cooled, horizontally opposed, carbureted four cylinder engine with a wet sump lubrication system. The engine is a Teledyne Continental O- 200-D rated at 100 horsepower at 2750 RPM. Major accessories include a starter, gear-driven alternator, dual magnetos and a full flow oil filter mounted on the rear of the engine accessory case. ENGINE CONTROLS Engine power is set using the throttle control. The throttle control is a smooth black knob located at the center of the instrument panel. The throttle control is configured so that the throttle is open in the forward position and closed in the full aft position. A friction lock, located at the base of the throttle control knob, is operated by rotating the lock clockwise to increase friction or counterclockwise to decrease friction. Engine fuel mixture is controlled by the mixture control. The mixture control is a red knob, with raised points around the circumference, located immediately to the right of the throttle control and is equipped with a lock button in the end of the knob. The rich position is full forward, and the idle cutoff position is full aft. For small adjustments, the control may be moved forward by rotating the knob clockwise, and aft by rotating the knob counterclockwise. For rapid or large adjustments, the knob may be moved forward or aft by depressing the lock button in the end of the mixture control knob, and then positioning the control as desired. (Continued Next Page) 162PHUS-03 2-24 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. ENGINE (Continued) ENGINE INSTRUMENTS The G300 Engine Indication System (EIS) provides the pilot graphical indicators and numerical values for Tachometer (RPM), Oil Pressure (OIL PSI), Oil Temperature (OIL ºF), Carburetor Temperature (CARB ºF), Battery Current (AMPS), and Elevator Trim Position (TRIM) along the top of the appropriate G300 Display. Additional indicators and numerical values for Exhaust Gas Temperature (EGT ºF), Electrical Bus Voltage (VOLTS), and Engine Hours (ENG HRS) are displayed on the G300 ENGINE Page when selected. In normal operation, EIS information is displayed on the Primary Flight Display (PFD) in the single display installation and on the optional Multifunction Display (MFD) in the dual display configuration (if installed). During engine start or when the AVN MASTER is OFF, the EIS is only displayed on the PFD even if the optional MFD is installed. TACHOMETER (RPM) A speed sensor, mounted on the right magneto, provides a digital signal to the ADAHRS which processes and outputs the RPM data to the EIS. Engine speed (RPM) is shown by the tachometer indicator found on all EIS pages. The tachometer indicator uses a circular scale with moving pointer and a digital value. The pointer moves through a range from 0 to 3000 RPM. The numerical RPM value is displayed in increments of 10 RPM in white numerals below the pointer. The normal engine speed operating limit (top of green arc) is 2750 RPM. When engine speed is more than 2750 RPM, the pointer, digital value, and label (RPM) turn red to show engine speed is more than the limit. The digital value and label (RPM) will flash. A red-X is displayed when the instrument is invalid or out of the data range. (Continued Next Page) 162PHUS-03 2-25 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. ENGINE (Continued) ENGINE INSTRUMENTS (Continued) OIL PRESSURE (OIL PSI) The oil pressure transducer, connected to the engine forward oil pressure port, provides a signal to the engine display that is processed and shown as oil pressure. Engine oil pressure is shown on all EIS pages. The analog instrument range is 0 to 110 PSI with digital information range from 0 to 140 PSI. The instrument has a red band from 0 to 10 PSI (low warning), a yellow band from 10 to 30 PSI (low caution), a green band from 30 to 60 PSI (normal operating range), a yellow band from 60 to 100 PSI (high caution) and a red band from 100 to 110 PSI (high warning). A white pointer indicates actual oil pressure. When oil pressure is 0 to 10 PSI or 100 to 140 PSI, the pointer, digital value, and label (OIL PRES) will change to red to show that oil pressure is outside normal limits. A red-X is displayed when the instrument is invalid or out of the data range. When the engine speed (RPM) is in the green arc and the oil temperature is in the green band, the oil pressure should be in the green band. If oil pressure is below the green band or above the green band, adjust the engine speed to maintain adequate oil pressure. When engine speed is at idle or near idle, the oil pressure indication must be above the lower red band. With the engine at normal operating oil temperature, and engine speed at or close to idle, oil pressure below the green band, but above the lower red band, is acceptable. In cold weather, the oil pressure will initially be high (close to the upper red band when the engine is started). As the engine and oil warm up, the oil pressure will come down into the green band range. (Continued Next Page) 162PHUS-03 2-26 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. ENGINE (Continued) ENGINE INSTRUMENTS (Continued) OIL TEMPERATURE (OIL °F) The oil temperature sensor is installed in the engine oil filter adapter and provides a signal to the engine display that is processed and shown as oil temperature. Engine oil temperature is shown on all EIS pages. The instrument range is 0 to 250°F with digital information range from 30 to 265°F. The instrument has a yellow band from 0 to 75°F (low caution), a green band from 75 to 220°F (normal operating range), a yellow band from 220 to 240°F (high caution) and a red band from 240 to 265°F (high warning). A white pointer indicates actual oil temperature. When oil temperature is in the red band, 240 to 265°F, the pointer and OIL TEMP turn red and flash to show oil temperature is higher than the limit. A red-X is displayed when the instrument is invalid or out of the data range. EXHAUST GAS TEMPERATURE (EGT °F) (if installed) A thermocouple installed in the exhaust down pipe measures EGT and provides a signal to the engine display that is processed and shown on the EIS ENGINE page by the EGT °F vertical tape display and digital information. The tape indicator range is from 1000 to 1600°F, and the digital information range is from 1000 to 2000°F. The white pointer indicates actual EGT temperature. Red digits are shown for values less than 0°F and more than 2000°F. A red-X is displayed when the instrument is invalid or out of the data range. (Continued Next Page) 162PHUS-03 2-27 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. ENGINE (Continued) ENGINE INSTRUMENTS (Continued) CARBURETOR TEMPERATURE (CARB °F) Carburetor temperature is provided as a reference to determine the potential for carburetor ice formation. The CARB °F indication is provided on vertical tape and in digital form on the right side of the EIS display atop the PFD and on the G300 ENGINE page. The tape display range is from 20 to 80°F and the digital indication range is from -40°F to 100°F. A yellow caution range is depicted from 5°F to 40°F. A red-X is displayed when the instrument is invalid or out of the data range. NOTE Although carburetor ice is more likely to form at temperatures within the yellow band range, it can form at temperatures outside the yellow caution range. If engine roughness or unexplained RPM loss is encountered, full carburetor heat should be immediately applied. NEW ENGINE BREAK-IN AND OPERATION The initial engine run-in was accomplished at the factory and is ready for the full range of use. It is recommended that cruising be accomplished at 75% power as much as practicable until a total of 50 hours has accumulated or oil consumption has stabilized. This will ensure proper seating of the piston rings. Low cruise power settings should be used for extended descents during the first 25 hours of new engine operations. Avoid extended descents with reduced power settings. (Continued Next Page) 162PHUS-03 2-28 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. ENGINE (Continued) ENGINE LUBRICATION SYSTEM The engine utilizes a full pressure, wet sump type lubrication system with aviation grade oil as the lubricant. The capacity of the engine sump, located on the bottom of the engine, along with the engine oil filter and oil cooler is 5.5 quarts. Oil is drawn from the sump through a filter screen on the end of a pickup tube to the engine driven oil pump. Oil from the pump then passes through a full-flow oil filter, a pressure relief valve at the rear of the right oil gallery, and an oil cooler before lubricating engine parts. The oil returns to the sump by gravity. The filter adapter in the full-flow filter is equipped with a bypass valve which will cause lubricating oil to bypass the filter in the event the filter becomes plugged, or the oil temperature is extremely cold. An oil dipstick/filler tube is located at the left rear of the engine case. The oil dipstick/filler tube is accessed through a door located on the left side of the engine cowling. The engine should not be operated on less than 3.5 quarts of oil. For extended flight, fill to 5.0 quarts (dipstick indication only). For engine oil grade and specifications, refer to Section 8 of this POH. IGNITION AND STARTER SYSTEM Engine ignition is provided by two engine driven magnetos, and two spark plugs in each cylinder. The left magneto fires the lower left and lower right spark plugs, and the right magneto fires the upper left and upper right spark plugs. Normal operation is conducted with both magnetos due to the more complete burning of the fuel/air mixture with dual ignition. Ignition and starter operation is controlled by a rotary-type switch located on the left instrument panel. The MAGNETOS switch is labeled clockwise, OFF, R, L, BOTH, and START. The engine should be operated on both magnetos (BOTH position) except for magneto checks. The R and L positions are for checking purposes and emergency use only. When the MAGNETOS switch is rotated to the spring-loaded START position, with the MASTER switch in the ON position, the starter contactor is closed and the starter, now energized, will crank the engine. When the switch is released, it will automatically return to the BOTH position. (Continued Next Page) 162PHUS-03 2-29 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. ENGINE (Continued) AIR INDUCTION SYSTEM The engine air induction system receives ram air through an intake on the lower front portion of the engine cowling. The intake is covered by an air filter which removes dust and other foreign matter from the induction air. Airflow passing through the filter enters an air box, which is equipped with a cable operated air door for cold air or pre-heated air from a shroud around the exhaust for carburetor heat. The pre-heated air is also the source for alternate air if the air induction filter should become blocked. The pre-heated air is unfiltered and drawn through a scat tube in the cylinder baffles. Use of full carburetor heat at full throttle will result in a loss of approximately 200-250 RPM due to the increased intake temps. After passing through the air box, induction air enters a carburetor under the engine, and is then ducted to the engine cylinders through intake manifold tubes. EXHAUST SYSTEM Exhaust gas from each cylinder passes through a riser assembly to a common muffler, located below the engine, and then overboard through a single exhaust outlet. Outside air is supplied to a shroud constructed around the outside of the muffler to form a heating chamber. The air heated by the shroud is then supplied to the cabin. COOLING SYSTEM Ram air for engine cooling enters through two intake openings in the front of the engine cowling. The cooling air is directed from above the engine, around the cylinders and other areas of the engine by baffling, and then exits through an opening at the bottom aft edge of the engine cowling. A winterization kit is available for the airplane. Refer to Section 10, Supplement 4 for description and operating information. 162PHUS-03 2-30 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. PROPELLER The airplane is equipped with a 67 inches diameter, two bladed, fixed pitch, one-piece composite propeller incorporating a nickle leading edge cap. An optional two bladed, fixed pitch, one-piece forged aluminum alloy propeller, which is anodized to retard corrosion is also available. FUEL SYSTEM The airplane fuel system, refer to Figure 2-12, consists of two vented integral fuel tanks (one tank in each wing), fuel shutoff valve, and a fuel strainer. The engine-mounted portion of the system consists of the gravity-fed updraft float carburetor. WARNING UNUSABLE FUEL LEVELS FOR THIS AIRPLANE WERE DETERMINED IN ACCORDANCE WITH ASTM INTERNATIONAL COMPLIANCE STANDARDS. FAILURE TO OPERATE THE AIRPLANE IN COMPLIANCE WITH FUEL LIMITATIONS SPECIFIED IN SECTION 3 MAY FURTHER REDUCE THE AMOUNT OF FUEL AVAILABLE IN FLIGHT. FUEL QUANTITY DATA IN U.S. GALLONS FUEL DISTRIBUTION Fuel flows by gravity from the two wing tanks to a fuel shutoff valve, the fuel strainer, and to the carburetor. The carburetor meters fuel flow in proportion to induction air flow and distributes to the cylinder intake manifold. FUEL TANKS FUEL LEVEL (QUANTITY EACH TANK) TOTAL FUEL TOTAL UNUSABLE TOTAL USABLE ALL FLIGHT CONDITIONS Two Full (12.73) 25.46 1.46 24.0 Figure 2-11 162PHUS-03 2-31 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. FUEL DISTRIBUTION Figure 2-12* 162PHUS-03 2-32 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. FUEL SYSTEM (Continued) FUEL INDICATING SYSTEM Fuel quantity is measured by two sight gages located in each wing root just forward and above the cabin door opening. The sight gages are vented and include orifice fittings at the top and bottom of the sight tube. The fuel levels are aided with floating balls for increased readability. There are two scales displayed on each fuel sight gage. The aft scale shows fuel levels to be read in level flight. The markings should not be relied upon for accurate readings during skids, slips or unusual attitudes. The forward scale shows fuel levels as read on level ground. Minimum takeoff is indicated when the center of the ball is at the "MIN/TO" line on the forward scale in level ground attitude which represents 3 gallons per side or 6 gallons total (1/4 tanks). . (Continued Next Page) FUEL QUANTITY SIGHT GAGE Figure 2-13 162PHUS-03 2-33 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. FUEL SYSTEM (Continued) FUEL INDICATING SYSTEM (Continued) An empty tank is shown on the fuel quantity indicator as a red line on the aft (Level Flight) scale along with the “E” indication. When an indicator shows an empty tank, approximately 0.73 gallons of unusable fuel remain in the tank. The fuel gage shows the fuel available in the tank up to the limit of the gage measurement range of approximately ¾ tank. At this level, additional fuel may be available, but no additional indication of the fuel will be visible. A visual check of each wing tank fuel level must be performed prior to each flight. Compare the visual fuel level and indicated fuel quantity to accurately estimate usable fuel. WARNING TAKEOFF IS PROHIBITED IF EITHER FUEL SIGHT GAGE INDICATES BELOW THE MIN/TO INDICATION (LESS THAN 1/4 TANK) OR BELOW THE BOTTOM OF THE IN-TANK FUEL FILLER INDICATOR TAB. FUEL VENTING Fuel system venting is essential to system operation. Complete blockage of the fuel venting system will result in decreasing fuel flow and eventual engine stoppage. The fuel venting system consists of an interconnecting vent line between the fuel tanks and a check valve equipped overboard vent in the left fuel tank assembly. The overboard vent protrudes from the bottom surface of the left wing, just inboard of the wing strut upper attachment point. The fuel filler caps are vacuum vented; the fuel filler cap vents will open and allow air to enter the fuel tanks in case the overboard vents become blocked. (Continued Next Page) 162PHUS-03 2-34 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. FUEL SYSTEM (Continued) REDUCED TANK CAPACITY The airplane may be serviced to a reduced capacity to permit heavier cabin loadings. This is accomplished by filling each tank to the bottom edge of the fuel filler indicator tab, thus giving a reduced fuel load of 3 gallons usable (minimum takeoff fuel) in each tank. Continued filling to each subsequent hole in the indicator tab will add an additional 3 gallons usable fuel per tank. Do not fuel above top hole as fuel expansion may cause overflow from fuel vent. IN-TANK FUEL FILLER INDICATOR TAB (Continued Next Page) Figure 2-14 162PHUS-03 2-35 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. FUEL SYSTEM (Continued) FUEL DRAIN VALVES The fuel system is equipped with 6 drain valves (4 in the wings and 2 in the lower cowl) to provide a means for the examination of fuel in the system for contamination and grade. The system should be examined before each flight and after each refueling, by using the sampler cup provided to drain fuel from each wing tank sump, the lower cowl fuel shutoff valve drain and the fuel strainer sump lower cowl drain. If any evidence of fuel contamination is found, it must be eliminated in accordance with the preflight inspection checklist and the discussion in Section 8. If takeoff weight limitations for the next flight permit, the fuel tanks should be filled after each flight to prevent condensation. BRAKE SYSTEM Disc brakes located on the main gear wheels are hydraulically actuated by master cylinders on each rudder pedal, refer to Figure 2-15. A parking brake valve locks brake pressure to the disc calipers when the parking brake handle on the lower instrument is pulled with the brake pedals depressed. The hydraulic brake system is connected in series running from the firewall mounted reservoir to the copilot's master cylinder then to the pilot's master cylinder then to the brake caliper through the parking brake valve. 162PHUS-03 2-36 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. BRAKE SYSTEM Figure 2-15* 162PHUS-03 2-37 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. ELECTRICAL SYSTEM The airplane is equipped with a 14-volt direct current (DC) electrical system powered by an engine driven alternator, refer to Figure 2-16. A 12-volt main storage battery is located on the right forward side of the firewall. The alternator and main battery are controlled by the MASTER switch found on the switch/circuit breaker panel. The alternator and main battery power is supplied to a relay panel located on the left forward side of the firewall. Power for all electrical circuits, except the secondary power circuit, is supplied from the relay panel to the electrical bus located in the switch/circuit breaker panel. The electrical bus is supplied with power anytime the MASTER switch is turned on. The electrical bus provides power to the avionics bus through the AVN MASTER switch. The avionics bus is supplied with power anytime the MASTER switch and AVN MASTER switch are turned on. The optional autopilot system (if installed) is powered through the avionics electrical bus via the AVN MASTER switch and is protected by the TRIM/AP circuit breaker. An autopilot power switch is located to the left of the PFD and supplies power to the autopilot system components when selected to the AUTOPILOT position. The electrical system is equipped with a secondary battery located between the firewall and the instrument panel. This secondary battery supplies power to the secondary power circuit for PANEL LIGHTS dimming as well as stable power to the G300 avionics and displays during engine start. This prevents undesired panel light dimming and loss of G300 information resulting from voltage drop during the engine start. The secondary battery is also controlled by the MASTER switch found on the switch/circuit breaker panel. Power is supplied to equipment on the secondary power circuit anytime the MASTER switch is turned on. (Continued Next Page) 162PHUS-03 2-38 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. ELECTRICAL SYSTEM (Continued) Figure 2-16* 162PHUS-03 2-39 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. ELECTRICAL SYSTEM (Continued) G300 CREW ALERTING SYSTEM (CAS) All system alerts, cautions and warnings are shown on the PFD screen. The following annunciations are supported: Refer to the Garmin G300 Pilot’s Guide for more information on CAS system annunciations. MASTER SWITCH The MASTER switch is a rocker type switch with two sides. The BAT side of the switch controls the electrical power to the airplane from both the main battery and secondary battery. The ALT side of the switch controls the alternator system. In normal operation, both sides of the switch (ALT and BAT) are operated simultaneously. However, the BAT side of the switch may be selected separately as necessary. The ALT side of the switch cannot be set to ON without the BAT side of the switch also being set to ON. ELECTRICAL SYSTEM MONITORING Main battery current, and system voltage indications are available on the PFD during single display operation and on the MFD when a second display is available. MAIN BATTERY CURRENT (AMPS) The main battery current indication is shown on all G300 EIS pages. A positive number indicates a charging battery while a negative number indicates a discharging battery. The tape display range is -35 to 35 amps with digital information range from -50 to 50 amps. A red-X is displayed when the instrument is invalid or out of the data range. A white pointer indicates actual current value. (Continued Next Page) LOW VOLTS TRIM UP or TRIM DOWN (if installed) TERRAIN SET BARO 162PHUS-03 2-40 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. ELECTRICAL SYSTEM (Continued) SYSTEM VOLTAGE (VOLTS) Electrical system voltage is shown on the G300 ENGINE page. The tape display is 8 to 18 volts with digital information range from -30 to 30 volts. The tape has a red band from 8 volts to less then 12.5 volts (low warning), a green band from 12.5 to 15.0 volts (normal operating range), a yellow band from 15.0 to 16.0 volts (high caution) and another red band from greater than 16.0 volts to 18.0 volts (high warning). A red-X is displayed when the instrument is invalid or out of the data range. A white pointer indicates actual system voltage. When the system voltage is less than 12.5 volts, a red LOW VOLTS CAS message is displayed on the bottom left side of G300 PFD attitude indicator. CIRCUIT BREAKERS Individual system circuit breakers are found on the switch/circuit breaker panel, refer to Figure 2-17. All circuit breakers on this panel are capable of being opened, or disengaged from the electrical system, by pulling straight out on the outer ring for emergency electrical load management. Using the circuit breaker as a switch is discouraged since the practice will decrease the life of the circuit breaker. The relay panel, located forward of the firewall, contains three circuit breakers associated with the alternator, main power feeder, and secondary power feeder. The alternator and main power circuit breakers are capable of being reset in flight by momentarily pushing the MAIN CB RESET switch located on the switch/circuit breaker panel. Pushing the MAIN CB RESET switch activates the reset solenoids contained in the relay panel. CAUTION EXCESSIVE ACTIVATION OF THE MAIN CB RESET SWITCH WILL DECREASE SOLENOID LIFE. Activation of the MAIN CB RESET switch is not required during normal preflight operation. (Continued Next Page) 162PHUS-03 2-41 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. ELECTRICAL SYSTEM (Continued) SWITCH/CIRCUIT BREAKER PANEL ASSEMBLY Figure 2-17 162PHUS-03 2-42 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. ELECTRICAL SYSTEM (Continued) 12V POWER OUTLET A 12 volt power outlet connector (POWER OUTLET 12V - 7.5A) is provided by an automotive style power outlet located on the center pedestal. The power outlet receives electrical power from a dedicated 7.5 amp circuit breaker located in the switch/circuit breaker panel. CAUTION • CHARGING OF LITHIUM BATTERIES MAY CAUSE THE LITHIUM BATTERIES TO EXPLODE. • TAKE CARE TO OBSERVE THE MANUFACTURER'S POWER REQUIREMENTS PRIOR TO PLUGGING ANY DEVICE INTO THE 12 VOLT CABIN POWER SYSTEM CONNECTOR. THIS SYSTEM IS LIMITED TO A MAXIMUM OF 7.5 AMPS. • USE CAUTION WITH POWER/ADAPTER CABLES IN THE CABIN TO AVOID ENTANGLING OCCUPANTS OR CABIN FURNISHINGS AND TO PREVENT DAMAGE TO CABLES SUPPLYING LIVE ELECTRIC CURRENT. • DISCONNECT POWER/ADAPTER CABLES WHEN NOT IN USE. 162PHUS-03 2-43 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. ELECTRICAL SYSTEM (Continued) EXTERNAL POWER RECEPTACLE (if installed) The External Power Receptacle, if installed, allows the use of an external power source for cold weather starting or for lengthy maintenance work on electrical equipment. The receptacle is located on the forward left side of the firewall and is accessed through a hinged door in the engine cowling. Anytime an external power source of correct voltage, 12.0 - 14.0 volts, and polarity is connected to the external power receptacle, the power will be connected to the main battery regardless of MASTER switch position. In order to power onboard equipment with the external power source, the MASTER switch must be turned to the ON position. Figure 2-18* 162PHUS-03 2-44 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. LIGHTING SYSTEMS EXTERIOR LIGHTING Exterior lighting consists of navigation/strobe lights located on the wing tips and a landing/taxi light located on the outboard left wing leading edge. All exterior lights are operated by switches found on the switch circuit breaker panel assembly to the right of the PFD. Exterior lights are grouped together in the LIGHTS section of the switch panel. To activate the LDG (landing/taxi light), NAV, and STROBE light(s), place the switch in the up position. Circuit breakers for the lights are found on the lower portion of the switch circuit breaker panel assembly. INTERIOR LIGHTING Instrument panel, pedestal and interior lighting is provided by an integral LED matrix light assembly mounted on the overhead cabin structure. Individual LEDs are directed at various instrument panel locations and both fuel quantity indicators for night and low-light illumination. Airplanes equipped with the BRS parachute option will also have a LED light directed at the parachute deployment handle. The PANEL LIGHTS dimming knob, located on the right side of the electrical system switch/CB panel, controls intensity of the overhead LED matrix lighting, non-stabilized magnetic compass internal lighting (if installed) and Garmin G300 displays when the G300 display setup backlight intensity is configured in the AUTO mode. Refer to the Garmin Pilot’s Guide for detailed information on manually adjusting the internal lighting on G300 displays. Rotating the PANEL LIGHTS dimming knob full clockwise provides maximum brightness while rotating the dimmer control knob counterclockwise decreases light intensity from the highest level to off. Power for the PANEL LIGHTS dimmer control is provided by the secondary power circuit and protected by the SEC PWR circuit breaker. 162PHUS-03 2-45 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. CABIN HEATING AND VENTILATING SYSTEM The cabin heating and ventilating system consists of a cabin heat system and fresh air from wing leading edge openings, refer to Figure 2-19. Cabin heat is controlled by a push-pull cable from the CABIN HEAT control knob located on the lower instrument panel. When the CABIN HEAT control knob is placed in the ON or full aft position, air is heated as it is directed around the exhaust heat shroud and ducted to a air control valve located on the firewall. With the air control valve in the on position, warm air is directed towards the pilot's and front passenger’s feet through various ducting located in the cabin. With the CABIN HEAT control knob in the OFF or full forward position, preheated air is redirected into the engine compartment through the air control valve located on the firewall. Cabin heat air volume will be reduced anytime you are operating with the CARB HEAT in the on position. This is done to allow maximum heat directed to the carburetor to prevent or alleviate carburetor ice. Always use maximum CARB HEAT when carburetor icing is suspected. Fresh air ducts direct ram air from the wing root openings to fresh air valves located overhead at the outboard corners of the windshield. These fresh air valves are equipped with directional control as well as variable flow adjustment settings. 162PHUS-03 2-46 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. CABIN HEATING AND VENTILATING SYSTEM Figure 2-19* 162PHUS-03 2-47 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. PITOT-STATIC SYSTEM The pitot-static system uses a total pressure (pitot) head mounted on the lower surface of the left wing, external static port mounted on the left side of the forward fuselage and associated plumbing to connect the G300 ADAHRS unit and the optional autopilot control unit (if installed). STALL WARNING SYSTEM Stall warning is a pneumatically operated by localized pressure differential created by low pressure airflow around the wing and higher static pressure internally. The stall warning horn will sound 5-8 knots prior to stall buffet to provide notice to the pilot of impending stall. Horn activation will not prevent a stall from occurring. When the horn sounds, the pilot should react by initiating stall avoidance procedures (decrease of angle of attack, increase airspeed, etc.). The stall warning system does not require electrical power for operation. The stall warning system opening located in the right wing should be checked during preflight to make sure it is free of debris that might interfere with airflow and correct operation. To check the system, place a clean cloth over the vent opening and apply suction; a sound from the warning horn will confirm system operation. 162PHUS-03 2-48 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. STANDARD AVIONICS The Garmin G300 Avionics System is an integrated flight control and navigation system. The system combines primary flight instruments, navigational, engine and electrical system information displayed on color Liquid Crystal Display(s) (LCD). The G300 system consists of the following pieces of equipment: GARMIN DISPLAY UNITS (GDU) The Cessna 162 can be configured with the standard single Garmin Display Unit or the optional dual display installation. SINGLE DISPLAY (Standard) The primary flight instrument information and engine system information will always be shown on the upper portion of the GDU and the navigational or other menu selected information will be displayed on the lower portion when configured in the split screen mode. Refer to the Garmin Pilot’s Guide for system information and operating instructions. The upper portion of the GDU displays roll and pitch information, heading and course navigation information, plus altitude, airspeed and vertical speed information to the pilot. Engine and electrical information are also displayed in the EIS strip as well as displaying warning/status annunciations of airplane systems. Autopilot system information will be displayed along the upper portion of the GDU if the optional autopilot system is installed. The lower portion of the GDU displays navigation information. The navigation information can be configured with the Horizontal Situation Indicator (HSI) display as default or one of the other menu driven map displays including a large scalable, moving map that corresponds to the airplane's current location. Data from other components of the system can be overlaid on this map. Lightning and weather information can all be displayed on the lower portion of the GDU display. (Continued Next Page) 162PHUS-03 2-49 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. STANDARD AVIONICS (Continued) GARMIN DISPLAY UNITS (GDU) (Continued) DUAL DISPLAYS (Optional) When the dual display option is installed, the GDU located in front of the pilot is automatically configured as a PFD and the second GDU located to the right, is automatically configured as a MFD. Refer to the Garmin Pilot’s Guide for system information and operating instructions. The top or upper portion of the PFD displays roll and pitch information, heading and course navigation information, plus altitude, airspeed and vertical speed information to the pilot as well as displaying warning/ status annunciations of airplane systems. Autopilot system information will be displayed along the upper portion of the PFD attitude display if the optional autopilot system is installed. The lower portion of the PFD will display the HSI and navigation information. All other map functions, engine and electrical information move over to the MFD. The MFD displays a large scalable, moving map that corresponds to the airplane's current location. Data from other components of the system can be overlaid on this map. Lightning and weather information can also be displayed on the MFD. The dual display MFD becomes the principle display for all of the engine, fuel, and electrical system parameters on the EIS strip along the top portion of the MFD and MFD engine pages. During a single display failure, the Garmin G300 installation will automatically revert into backup mode and display all flight information and basic engine information on the remaining display. This feature allows the pilot full access to all necessary information should either of the display screens malfunction. XM radio and WX weather systems information will continue to operate during a PFD display failure, but will become inoperative during a MFD display failure due to system components being integrated into the MFD display unit. (Continued Next Page) 162PHUS-03 2-50 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. STANDARD AVIONICS (Continued) AIR DATA, ATTITUDE AND HEADING REFERENCE SYSTEM (ADAHRS) AND MAGNETOMETER (GRS) The ADAHRS provides airplane attitude and flight characteristics information to the G300 displays integrated avionics units. The ADAHRS unit, located behind the instrument panel, contains accelerometers, tilt sensors and rate sensors that replace spinning mass gyros used in other airplanes. The magnetometer, located in the tailcone, interfaces with the ADAHRS to provide heading information. The air data portion of the ADAHARS compiles information from the airplane's pitot-static system to calculate pressure altitude, airspeed, true airspeed, vertical speed and outside air temperature. An outside air temperature probe, mounted on top of the cabin, is connected to the ADAHRS. The engine monitor portion of the ADAHARS is responsible for receiving and processing the signals from all of the engine and airframe sensors. It is connected to engine RPM, CARB TEMP and EGT sensors along with receiving information from the electrical system components. This unit transmits this information to the engine display computers. (Continued Next Page) 162PHUS-03 2-51 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. STANDARD AVIONICS (Continued) GARMIN GTX 327 TRANSPONDER Refer to Section 10, Supplement 6 for appropriate transponder operating information. GARMIN SL 40 VHF COMM TRANSCEIVER Refer to Section 10, Supplement 7 for appropriate transceiver operating information. XM WEATHER AND RADIO DATA LINK (GDL) (if installed) The optional XM weather and radio data link provides weather information and digital audio entertainment in the cockpit. The XM receiver unit is part of the optional MFD second display. XM weather and XM radio operate in the S-band frequency range to provide continuous uplink capabilities at any altitude throughout North America. A subscription to the XM satellite radio service is required for the XM weather and radio data link to be used. Refer to the Garmin Pilot’s Guide for system information and operating instructions. 162PHUS-03 2-52 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. AVIONICS SUPPORT EQUIPMENT Avionics cooling fan, antennas, and provisions for aviation-style headsets support the operation of the avionics equipment installations. AVIONICS COOLING FAN A single DC electric fan provides forced air and ambient air circulation cooling for the G300 avionics equipment. The fan is located forward of the instrument panel and removes air from between the firewall bulkhead and instrument panel, directing the warm air up at the inside of the windshield. Power is provided to the cooling fan when the MASTER (BAT) switch and the AVN MASTER switch are both ON. ANTENNAS A single combination VHF COM/GPS antenna is mounted on the top of the cabin. The dual-mode COM/GPS antenna is mounted on the left side. It connects to the VHF communication transceiver and the GPS receiver in the PFD and GDL. The transponder antenna is mounted on the bottom of the cabin and is connected to the Mode S transponder by a coaxial transmission cable. The optional XM antenna (if installed) is mounted on the top right side of the cabin. It provides a signal to the GDL-69A XM Data Link receiver. (Continued Next Page) 162PHUS-03 2-53 SECTION 2 CESSNA AIRPLANE AND SYSTEM DESCRIPTION MODEL 162 GARMIN G300 U.S. AVIONICS SUPPORT EQUIPMENT (Continued) MICROPHONE AND HEADSET INSTALLATIONS Standard equipment for the airplane includes two push-to-talk (PTT) microphone buttons on the control sticks and provisions for aviation- style headsets at each pilot station. For airplanes not equipped with the PM3000 Intercom System all radio communications are transmitted through the Garmin SL40 Comm radio. For airplanes equipped with the PS Engineering PM3000 Intercom System, all radio communications are transmitted through the PM3000 Intercom System when the system is selected to the ON position, otherwise all radio communications are transmitted through the Garmin SL40 Comm radio. For more information on system description and operation, refer to Section 10, Supplement 3. AUXILIARY AUDIO INPUT JACK (if installed) An auxiliary audio input jack (AUX AUDIO INPUT) is located on the center pedestal when the PS Engineering Intercom System is installed, refer to Section 10, Supplement 3 for additional operating information. 162PHUS-03 2-54 CESSNA SECTION 2 MODEL 162 AIRPLANE AND SYSTEM DESCRIPTION GARMIN G300 U.S. CABIN FEATURES EMERGENCY LOCATOR TRANSMITTER (ELT) Refer to Section 10, Supplements 1 or 2 for appropriate ELT operating information. CABIN FIRE EXTINGUISHER (if installed) A portable Halon 1211 (Bromochlorodifluoromethane) fire extinguisher is installed on the baggage area floorboard to be accessible in case of fire. The extinguisher is classified 5B:C by Underwriters Laboratories. The extinguisher should be checked prior to each flight to ensure that the pressure of the contents, as indicated by the gage at the top of the extinguisher, is within the green arc (approximately 125 psi) and the operating lever lock pin is securely in place. To operate the fire extinguisher: 1. Loosen retaining clamp(s) and remove extinguisher from bracket. 2. Hold extinguisher upright, pull operating ring pin, and press lever while directing the liquid at the base of the fire at the near edge. Progress toward the back of the fire by moving the nozzle rapidly with a side-to-side sweeping motion. WARNING VENTILATE THE CABIN PROMPTLY AFTER SUCCESSFULLY EXTINGUISHING THE FIRE TO REDUCE THE GASES PRODUCED BY THERMAL DECOMPOSITION. 3. The contents of the cabin fire extinguisher will empty in approximately eight seconds of continuous use. Fire extinguishers should be recharged by a qualified fire extinguisher agency after each use. After recharging, secure the extinguisher to its mounting bracket. 162PHUS-03 2-55/2-56 CESSNA SECTION 3 MODEL 162 OPERATING LIMITATIONS GARMIN G300 U.S. OPERATING LIMITATIONS TABLE OF CONTENTS Page Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-3 Airspeed Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-4 Airspeed Indicator Markings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-5 Powerplant Limitations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-6 Powerplant and Electrical Instrument Markings. . . . . . . . . . . . . . . . . 3-7 Weight Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-8 Maximum Weight in Baggage Compartment. . . . . . . . . . . . . . . . . 3-8 Center of Gravity Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-8 Service Ceiling. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-8 Maneuver Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-9 Load Factors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-9 Flight Load Factor Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-9 Kinds of Operations Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-10 Kinds of Operations Equipment List . . . . . . . . . . . . . . . . . . . . . . . . 3-11 Fuel Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-14 Flaps Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-14 System Limitations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-15 12V Power Outlet . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-15 G300 Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-15 Placards. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3-16 162PHUS-01 3-1/3-2 CESSNA SECTION 3 MODEL 162 OPERATING LIMITATIONS GARMIN G300 U.S. INTRODUCTION Section 3 includes operating limitations, instrument markings, and basic placards necessary for the safe operation of the airplane, its engine, standard systems and standard equipment. The limitations included in this section and in Section 10 have been approved. Observance of these operating limitations is required by Federal Aviation Regulations. NOTE • Refer to Section 10, Supplements, of this Pilot's Operating Handbook for amended operating limitations, operating procedures, performance data and other necessary information for airplanes equipped with specific options. • The airspeeds listed in Figure 3-1, Airspeed Limitations, are based on Airspeed Calibration data shown in Section 5. The Cessna Model No. 162 is approved under ASTM standard F2245. 162PHUS-00 3-3 SECTION 3 CESSNA OPERATING LIMITATIONS MODEL 162 GARMIN G300 U.S. AIRSPEED LIMITATIONS Airspeed limitations and their operational significance are shown in Figure 3-1. AIRSPEED LIMITATIONS SYMBOL SPEED KCAS KIAS REMARKS VS Stall Speed - Clean 44 41 Stall speed flap up. VSO Stall Speed - Landing Configuration 40 37 Stall speed flaps full VFE Maximum Flap Extended Speed: FLAPS 10° FLAPS 25° FLAPS FULL 98 84 71 100 85 70 Do not exceed this speed with flaps down. VO Maximum Operating Maneuvering Speed 1320 Pounds 1200 Pounds 1100 Pounds 88 84 81 89 85 80 Maximum speed at which the airplane may be stalled without exceeding structural limitations. VA Design Maneuvering Speed: 100 102 Do not make full or abrupt control movements above this speed. Does not provide protection from possible overstressing the airplane. VNE Never Exceed Speed 143 148 Do not exceed this speed in any operation. V NO Maximum Structural Cruising Speed 120 124 Do not exceed this speed except in smooth air, and then only with caution. Figure 3-1 162PHUS-02 3-4 CESSNA SECTION 3 MODEL 162 OPERATING LIMITATIONS GARMIN G300 U.S. AIRSPEED INDICATOR MARKINGS Airspeed indicator markings and their color code significance are shown in Figure 3-2. AIRSPEED INDICATOR MARKINGS MARKING KIAS VALUE OR RANGE SIGNIFICANCE Red Band <37 Low airspeed warning. White Band 37 - 70 Full Flap Operating Range. Lower limit is maximum weight V SO in landing configuration. Upper limit is maximum speed permissible with flaps extended. Green Band 41 - 124 Normal Operating Range. Lower limit is maximum weight V S1 at most forward C.G. with flaps retracted. Upper limit is maximum structural cruising speed. Yellow Band 124 - 148 Operations must be conducted with caution and only in smooth air. Red Line >149 Maximum speed for all operations. Flaps 25° Tick Mark 85 Maximum Flaps 25° Operation Flaps 10° Tick Mark 100 Maximum Flaps 10° Operation Vx Tick Mark 57 Best Angle of Climb Air Speed Vy Tick Mark 62 Best Rate of Climb Air Speed Figure 3-2 162PHUS-01 3-5 SECTION 3 CESSNA OPERATING LIMITATIONS MODEL 162 GARMIN G300 U.S. POWERPLANT LIMITATIONS Engine Manufacturer: Teledyne Continental Motors Engine Model Number: O-200-D Maximum Power: 100 BHP Rating Engine Operating Limits for Takeoff and Continuous Operations: Maximum Engine Speed:. . . . . . . . . . . . . . . . . . . . . . . . . .2750 RPM NOTE The static RPM range at full throttle with carburetor heat off and mixture leaned to maximum RPM is 2375 - 2475 RPM. For allowable variations in static RPM at non-standard temperatures, refer to the 162 Maintenance Manual. Maximum Oil Temperature: . . . . . . . . . . . . . . . . . . . . .240°F (116°C) Oil Pressure, Minimum: . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .10 PSI Oil Pressure, Maximum: . . . . . . . . . . . . . . . . . . . . . . . . . . . .100 PSI CAUTION ENGINE OPERATION WITH INDICATED OIL PRESSURE BELOW THE GREEN BAND RANGE WHILE IN CRUISE OR CLIMB CONFIGURATION IS CONSIDERED ABNORMAL. REFER TO SECTION 3, AMPLIFIED EMERGENCY PROCEDURES, "LOW OIL PRESSURE". Fuel Grade: Refer to Fuel Limitations Oil Grade (Specification): SAE J1966 Aviation Grade Non-Dispersant Mineral Oil or SAE J1899 Aviation Grade Ashless Dispersant Oil. Oil must comply with the latest revision and/or supplement for Teledyne Continental Motors (TCM) Service Information Letter SIL99-2B or later revision, must be used. Propeller Manufacturer: McCauley Propeller Systems Standard Composite Propeller Model Number. . . . . . . 1L100/LSA6754 Maximum Propeller Diameter . . . . . . . . . . . . 67.0 INCHES (1.70 m) Minimum Propeller Diameter . . . . . . . . . . . . . 66.5 INCHES (1.69 m) Optional Aluminum Propeller Model Number . . . . . . . .1A162/TCD6754 Maximum Propeller Diameter . . . . . . . . . . . . 67.0 INCHES (1.70 m) Minimum Propeller Diameter . . . . . . . . . . . . . 66.0 INCHES (1.68 m) 162PHUS-03 3-6 CESSNA SECTION 3 MODEL 162 OPERATING LIMITATIONS GARMIN G300 U.S. POWERPLANT AND ELECTRICAL INSTRUMENT MARKINGS Powerplant and electrical instrument markings and their color code significance are shown in Figure 3-3. Operation with indications in the red range is prohibited. Avoid operating with indicators in the yellow range. POWERPLANT AND ELECTRICAL INSTRUMENT MARKINGS INSTRUMENT RED (LOWER WARNING) YELLOW (LOWER CAUTION) GREEN (NORMAL OPERATING RANGE) YELLOW (UPPER CAUTION) RED (UPPER WARNING) Tachometer (RPM) ---- ---- 2000 to 2750 RPM 2750* to 3500 RPM Oil Temperature (OIL °F) ---- 0 to 75°F 75 to 220°F 220 to 240°F 240* to 265°F Oil Pressure (OIL PSI) 0 to 10 PSI 10 to 30 PSI 30 to 60 PSI 60 to 100 PSI 100* to 140 PSI Exhaust Gas Temperature (if installed) (EGT °F) ---- ---- 1000 to 1600°F White Advisory Carburetor Temperature (CARB °F) 5 to 40°F Battery Current (AMPS) -35 to 35 White Advisory Bus Voltage (VOLTS) -30 to 12.5 12.5 to 15 15 to 16 16 to 30 *Maximum operating limit is lower end of red. Figure 3-3 162PHUS-00 3-7 SECTION 3 CESSNA O