TWIN COMANCHE PILOT’S OPERATING MANUAL

.-- **- * - *. . . . . . * . . 1 -N. Placards . . II::;::: . . . . . . . . . . . . . . . . ..~ . . . . . . . . . . . . . . 1:::: 5 0. Instrument Markings . . . . . . . . . 5 . . . *. .. *‘- -*---*‘-** . . * . * * . . 5 P. LandingGearDownLigh;.::::::::::::::::::::::::::::::::‘..” 6” . . . . . SECTION III Procedures . . . . . . . . . . . . . . . . . A. Standard Procedures . . * . . * . . . * *----**.-* . . . - * -* *. * 1. StoppingEngines . . . . . . 7 2. Warning .::::::::: . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7 . . * - . * * . . . . . . * . - * 3. Spins .::::: . . . . . .._................ ~...:::::::::::: 7 . . - . * . - . . 7 * *. . . . . *. * * * * * - - * * - B. Systems Operation and Checks . . . . . . : . . . . . 7 1. Voltage Regulating and Alternator System for’A&raft’skrial Numbers j9:1’ ’ . . ’ ’ . thru 39-145 _ . 9 ~*~~1*~~--* . . . . . . . . . 2. Alternator Paralleling System for Aircraft Serial Numbers 39-146 and up ’ : : 1 : : 1 1 3. Circuit Breakers . . . . . .’ 9 4. FuelManagement 9.;...:: . . . . . . . . . . . . . . . . . . . . . . . . . . . 10 5. Manual Extension of: i&ding Gear . . . : : : : : : : : . ’ * ’ ’ * ’ ’ * * . . ’ . ’ * ’ . * lo . . . . . . . . * - * * * * * *. * . 11 C. Emergency Procedures . . 1. Detecting a Dead Engine . 1 ; 1 : 1 : 1 1 : 1 : 1 . . . . . . - . . . . . . . . . . . . . . . . 13 2. Feathering Procedure . . . . . . . . . . . . . . . . . . . . 3. Unfeathering Procedure . . 1 . . . . 13. . 4. Engine Failure During Take-Off . : 1 1 : 1 1 . . : : 1 : : 1 1 ’ ’ ’ . * * ’ ’ * . ’ . * ’ ’ . . ’ :3” .-*--l-*m * - . * * * * . - 5* EngineFailureDuringaimb ..::::::::::‘.‘..-....‘.......... I4 6. Single Engine Landing . . * . . . . . **-***-*-- . . . 147. Single Engine Go-Around’ : 1 ; 1 : ; : 1 : ; 1 ; : 1 - * * * * * * * * * . . + - * . * - . . * l4 -**--.-.-- ----*--*** . 14 FAA APPROVED November 28, 1969 RElSSUED: June 1, 1970 REVISED: November 30,197l REPORT: 1605 PART I P -- .-. - MODE1 .+ /’ SECTION L’ Supplements ................. A. Oxygen System installation .... 1. Limitations ........... 2. Procedures ........... 3. Placards ............. 4. Emergency ........... . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..................... 19 ..................... 21 ..................... 21 ..................... 21 ..................... 21 ..................... 21 B. Piper AutoControl III (Model AK161) Installation ....................... 23 1. Limitations ........................................... 23 ’ Operating Instructions -. .................................... 23 3. Normal Operation ....................................... 23 4. Emergency Procedures .................................... 24 C. Brittain Industries, Inc. Model TT-5 Tip Tanks Installation .................. 25 1. Limitations ........................................... 25 ’ Operating Procedures - Normal -. .............................. i 25 3. Performance .......................................... 26 4. Emergency ........................................... 26 D. Piper AltiMatic IILB Installation .................................. 27 1. Limitations ........................................... 27 1 Operating Instructions -. .................................... 27 3. Normal Flight Operation ................................... 28 4. With Glide Slope Coupler Installed (Flight) ........................ 28 5. Emergency Procedures .................................... 28 6. Placards ............................................. 29 E. Anti-Collision (Strobe) Light Installation ............................. 3 1 1. Operating Limitations ..................................... 3 1 F. Anti-Icing Equipment Installation ................................. 33 1. Placards ............................................. 33 2. Operating Instructions .................................... 33 G. Piper AltiMatic IIIB- 1 ........................................ 3 5 1. Limitations ........................................... 35 2. Procedures ........................................... 35 3. Performance .......................................... 40 FAA APPROVED November 28.1969 REISSUED: June 1.1970 REVISED: November 30,197 1 REVISIONS The information compiled in this manual will be kept current by revisions distributed to the airplane owners through their local dealersor distributors. There are two types of revisions used to keep the Pilot’s Operating Manual current: Temporary Revisions and Permanent Revisions. The material compiled in the revisions will consist of information necessaryto update the present information or add information to cover added equipment. I. Temporary Revision This revision will be distributed at any time it is necessary to forward information to the owners and operators of the airplane. The revision will usually consist of one or two pages which may be inserted in the appropriate section of the manual. This revision will include deletions and/or additions of material pertinent to different paragraphsof the manual. II. Permanent Revision This revision will be distributed periodically and will supersedeall previous temporary revisions. These revisions will be complete page replacement and shall be inserted in the manual in accordancewith the instructions given below. 1. Replace the obsolete pageswith revised pagesof the same page number. 2. Insert pageswith page numbers followed by a small letter in direct sequencewith the same common numbered page. III. Identification of RevisedMaterial Revised text and iliustrations shall be indicated by a black vertical line along the left hand margin of the page, opposite revised, added or deleted material. A line opposite the page number or section title and printing date, will indicate that the text or illustration was unchanged but that material was relocated to a different page or that an entire page was added. Symbols will indicate only current revisions with changesand additions to or deletions of existing text and illustrations. Changes in capitalization, spelling, punctuation or the physical location of the material on the page will not be identified by symbols. i ’ ___--.- REVISIONS ISSUED Current Permanent and Temporary Revisions to the PA-39 Pilot’s Operating Manual issued June 1970 are as follows: 761 453 (PR710317) 761 453 (PR710827) Permanent Revision Dated March 17, 1971 Permanent Revision in F/M Added propeller synchrophaser II-IO & added fuel pump revision II-l 1. Dated August 27, 1971 761 453 (PR7111 IS) 761 453 (PR720220) 761 453 (PR7203 17) 761 453 (PR720428) 761 453 (PR730330) 761 453 (PR730516) 761 453 (PR730622) 761 453 (PR750321) 761 453 (PR751226) Permanent Revision Permanent Revision Permanent Revision Misc. Revision to P/O/M; Added nose wheel alternate to Weight & Balance. Misc. Revision to P/O/M, F/M, W/B. Permanent Revision to F/M (Rev. 10) Permanent Revision to W/B Permanent Revision to P/O/M, added Rev. 11 to A F/M. Permanent Revision to P/O/M Dated Nov. 15: 1971 Dated Feb. 20, 1972 Dated March 17, 1972 Dated April 28, 1972 Dated March 30, 1973 Dated May 16, 1973 Dated June 22, 1973 Dated March 2 1, 1975 Dated December 26, 1975 TWIN COMANCHE C/R SECTION 1 LOG OF REVISIONS Revision Revised Pages Description and Revision FAA Appi oved Date This manual, which was completely revised June 1, 1970, applies to Piper Model PA-39-1 and up. 1. 31. I-ii I-19 PageI-32 added PagesI-33 & 34 added II-2 1 II-24 I-26 I-28 I-29 II-5 II-6 II-24 II-29 Under Section V, added Item F. Nov. 2, 1970 Added Item F. Added information for Supplement F. Added Supplement F. Heated Glass Panel Installation added. Appropriate was appropriated. March 17, 1971 Revised Supplement D - Item 2. - d. Added to Supplement D Item 6. - f. C. G. Chart Clarified. Basic weight was basis weight. Added Radair Exhaust GasTemperature Indicator and Delta Mixture Control Indicator. Added Narco ATS-A Transponder and King Digital ADF KR-85. 3. 4. 5. 6. 7. II-29 II-29 II-I 0 II-I 1 I-3 H-9 II-IO I-9 Added Bendix ADF-T 12D. April 19, 1971 Added Marker Beacon Installation MBT-R-I 2. June 1, 1971 Added Propeller-Synchrophaser. August 13, 1971 Added Piper Dwg. Nos. to Fuel Pumps. Added additional Hartzell Propellers. Sept. 3, 1971 Added additional Hartzell Propellers. Added additional Hartzell Propellers to Spinners. Revised instructions for Voltage Regulating Nov. 1, 1971 System item e. and g. FAA APPROVED November 28,1969 REISSUED: June 1,197O REVISED: November 1,197l REPORT: 1605 PART I PAGE 1’ MODEL: PA-39 TWIN COMANCHE C/R LOG OF REVISIONS (cant) Revision , Revised Pages Description and Revision FAA Approved Date 8. Revised Section III Item B. . i ii Added Supplement G. AltiMatic IIIB-1. Nov. 30, 1971 9 Under B. Systems Operation and Checks - Items 1 and 2 combined to 1. a & 1. b, added Ser. Nos. 39-1 thru 39-145. Revised item 2 to Alternator Paralleling information for Ser. Nos. 39-146 and up. 10 Added information from page 9. 11 Added information from page 10. 13 Emergency Proceduresfor Feathering and Unfeathering revised to include Alternator Paralleling. 19 Added Supplement G. AltiMatic IIIB-1. Pages35-36-3 7 38-3940 added 9. 6 10. 11 Added Warning Placard information. - Added item (b) to Cruising under Crossfeed Operation and Single Engine Operation. Mar. 30,1973 May 16, 1973 11. 6 Revised Warning Placard information. March 21: 1975 c,’ r*; .s ,* . I %. Everly rsec .+t- REPORT: 1605 PART I PAGE 2 MODEL: PA-39 FAA APPROVED November 28,1969 REISSUED: June 1,197O REVISED: March 21., 1975 TWIN COMANCHE C/R SECTION I LOG OF REVISIONS Revision RevisedPages Description and Revision FAA Approved Date This manual, which was completely revised June I, 1970, applies to Piper Model PA-39-l and up. 1. 2. I-ii I-19 PageI-32 added PagesI-33 & 34 added II-2 1 II-24 I-26 I-28 I-29 II-5 II-6 II-24 II-29 Under Section V, added Item F. Nov. 2, 1970 Added Item F. Added information for Supplement F. Added Supplement F. Heated Glass Panel Installation added. App>opriate was appropriated. March 17, 1971 RevisedSupplement D - Item 2. - d. Added to Supplement D Item 6. - f. C. G. Chart Clarified. Basic weight was basis weight. Added Radair Exhaust GasTemperature Indicator and Delta Mixture Control Indicator, Added Narco ATS-A Transponder and King Digital ADF KR-85. 3. 4. 5. 6. 7. II-29 II-29 II-I 0 II-1 1 I-3 II-9 II-IO I-9 Added Bendix ADF-T 12D. April 19, 1971 Added Marker Beacon Installation MBT-R-12. June 1, 1971 Added Propeller Synchrophaser. August 13, 1971 Added Piper Dwg. Nos. to Fuel Pumps. . Added additional Hartzell Propellers. Sept. 3, 1971 Added additional Hartzell Propellers. Added additional Hartzell Propellers to Spinners. Revisedinstructions for Voltage Regulating Nov. 1, 1971 System item e. and g. FM APPROVED November 28,1%9 REISSUED: June 1.1970 REVISED: November 1,197l REPORT: 1605 PART I Py- --- MODE! ,’ #--- I I i,!3,‘:‘ ,-’ lr~h~COMANCHE C/R LOG OF REVISIONS (cant) Revision Revised Pages Description and Revision FAA Approved Date 8. . i Revised Section III Item B. Nov. 30. 197 1 ii Added Supplement G. AltiMatic IIIB-1. 9 Under B. Systems Operation and Checks - Items 1 and 2 combined to 1. a & 1. b, added Ser. Nos. 39-l thru 39-145. Revised item 2 to Alternator Paralleling information for Ser. Nos. 39-146 and up. 10 Added information from page 9. 11 Added information from page 10. 13 Emergency Procedures for Feathering and Unfeathering revised to include Alternator Paralleling. 19 Added Supplement G. AltiMatic IIIB-I. Pages 35-36-37 38-39-40 added 9. 6 10. 11 Added Waming Placard information. . Added item (b) to Cruising under Crossfeed Operation and Single Engine Operation. Mar. 30,1973 May 16, 1973 11. 6 Revised Warning Placard information. March 21, 1975 p;; >{* m Everly’ L REPORT: 1605 PART I PAGE 2 FM APPROVED November 28,1%9 REISSUED: June 1,197O REVISED: March 21,1975 TWIN COMANCHE C/R SECTION 11 LIMITATIONS The following limitations must be observed in the operation of this airplane: A. ENGINES Lycoming LIO-320-BIA (Right) Lycoming 10-320-B 1A (Left) ENGINE LIMITS For all operation 2700 RPM, 160 HP (See Maneuvers) B. C. I I I I D. E. FUEL 1OO/130 Octane Aviation Gasoline(Minimum) PROPELLERS Hartzell HC-E2YL-2B or HC-E2YL-2BF or HC-E2YL-2BS or HC-E2YL-2BSF (Left) Constant Speed Full Feathering; Blades 7663-4 or F7663-4 Pitch Settings at 30 in. station Station: High 76’ - 77 ‘, Low 12’ Diameter: Not over 72 inches Not under 70 inches (No further reduction permitted) Hartzell HC-E2YG2BL or HC-E2YL-2BLF or HC-E’LYL-2BLS or HC-E2YL-2BLSF (Right) Constant Speed Full Feathering; Blades J7663-4 or FJ76634 Pitch Settings at 30 in. station Station: High 76” - 77 ‘, Low 12 ’ Diameter: Not over 72 inches Not under 70 inches (No further reduction permitted) COWL FLAPS Cowl flaps are provided to allow manual control of engine temperatures. The cowl flaps should be open during ground operations and in climbs. In no caseshould the cylinder head temperatures be allowed to exceed 500” F. and the oil temperatures allowed to exceed 245” F. INSTRUMENT MARKINGS (POWER PLAN?) One One I One One OIL TEMPERATURE Green Arc (Normal Operating Range) Yellow Arc (Caution) Red Line (Maximum) 120” to 245°F 60” to 120’F 245” F FAA APPROVED November 28,1969 REISSUED: June 1,197O REVISED: September 3,197l REPORT: 1605 PART I PAGE 3 MODEL: !A-39 y TWIN COMANCHE C/R F. G: OIL PRESSURE Green Arc (Normal Operating Range) Yellow Arc (Caution) Red Line (Minimum) Red Line (Maximum) 60 PSI to 90 PSI 25 PSI to 60 PSI and 90 PSI to 100 PSI ‘8 PSI 100 PSI TACHOMETER Green Arc (Normal Operating Range) Red Line (Maximum) 500 RPM to 2700 RPM 2700 RPM FUEL FLOW Green Arc (Normal Operating Range) Red Line (Maximum at SeaLevel) OGPHto16GPH 16 GPH (7 PSI) CYLINDER HEAD TEMPERATURE Green Arc (Normal Range) Red Line (Maximum) 200” to 500” F 500” F AIRSPEED LIMITATIONS AND INDICATOR MARKINGS (Calibrated Airspeed) Normal Category NEVER EXCEED SPEED MAXIMUM STRUCTURAL CRUISING SPEED DESIGN MANEUVERING SPEEDS - Minimum Weight (2825 lbs.) 0 Maximum LWeight(3600 lbs.) 230 MPH 194 MPH 1.45MPH 162 MPH FLAPS EXTENDED SPEEDS 70 MPH to 125 MPH MAXIMUM GEAR EXTENDED SPEED 150 MPH MINIMUM CONTROL SPEED (Single Engine) 80 MPH STALL SPEED Gear and Flaps Up Gear and Flaps Down 76 MPH 70 MPH AIRSPEED INDICATOR MARKINGS Green Arc (Normal Operating Range) Yellow Arc (Caution Range - Smooth Air) White Arc (Flaps Extended Range) Radial Red Line (Never Exceed - Smooth Air) Radial Red Line (Minimum Control Speed- Single Engine) Radial Blue Line (Best R/C Speed Single Engine) 76 MPH to 194 MPH 194 MPH to 230 MPH 70 MPH to 125 MPH 230 MPH 80 MPH 105 MPH FLIGHT LOAD FACTORS Positive Load Factor (Maximum) Negative Load Factor (Maximum) (No Inverted Maneuvers Approved) 3.8 G REPORT: 1605 PART I PAGE 4 ; MODEL: PA-39 , FAA APPROVED November 28,1969 REISSUED: June 1.1970 H. SlAXIMUM WEIGHT TWIN COMANCHE C/R IT IS THE RESPONSIBILITY OF THE AIRPLANE OWNER AND THE PILOT TO ASSURE THAT THE AIRPLANE IS PROPERLY LOADED. MAXIMUM ALLOWABLE GROSS WEIGHT 3600 POUNDS. SEE “WEIGHT AND BALANCE SECTION” FOR PROPER LOADING INSTRUCTIONS. 1. C. G. RANGE Weight Forward Limit Pounds Inches Aft of Datum Aft Limit Inches Aft of Datum I 3600 86.5 92 3200 83.0 92 2825 82.0 92 1. Straight line variation between the points given. 2. Datum is 79 inches ahead of the wing leading edge at spanwise Station 97.0 (First leading edge skin lap outboard of enginenacelle). J. MANEUVERS All intentional acrobatic maneuvers]including spins) are prohibited. Avoid abrupt maneuvers.When performing power on stalls do not exceed2 100 RPM. K. WING FLAP SETTINGS Take-Off 0” to 15” Landing 27” The flaps are electrically operated and the deflection is displayed on a flap position indicator. Take-off range indicated by White Arc on flap indicator. L. UNUSABLE FUEL The unusable fuel in this aircraft has been determined as 3 gallons in each inboard tank inn flight attitudes. - M. USABLE FUEL Inboard tanks - 27 gallons each Auxiliary tanks (outboard) for use in level flight only - 15 gallons each. N. PLACARDS On instrument panel in full view of the pilot: THIS AIRPLANE MUST BE OPERATED AS A NORMAL CATEGORY AIRPLANE IN COMPLIANCE WITH THE AIRPLANE FLIGHT MANUAL. ACROBATIC MANEUVERS (INCLUDLN_C;SPINS)PROHIBITED. FAA APPROVED November 28,1969 REISSUED: June 1. 1970 REPORT: 1605 PART I PAGE 5 MODEL: PA-39 ‘WIN COMANCH OIL PRESSC Green A Yellow , Red Lin Red Lin TACHOMET Green A Red Lir FUEL FLOW Green P Red Lir CYLINDER Green 1 Red Lir AIRSPEED : NEVER EX(’ MAXIMUM DESIGN Mb * &&.&.& 0 Maxim1 FLAPS EXT MAXIMUM ; MINIMUM c STALL SPE Gear ai, Gear ai AIRSPEED Green I Yelloi White, Radial Radial Radial’ FLIGHT Li Positive Loi NegativeL? :No Inverte .i ,RT: 16d EL: PA-39 ‘. On the baggagecompartment door: EMERGENCY EXIT HOLD KNOB UP TURN LATCH CLOCKWISE On landing gear operating motor accessdoor: EMERGENCY GEAR EXTENSION. REMOVE COVER, EXTENSION INSTRUCTIONS ON REVERSE SIDE. Orfinstrument panel: STALL WARNING The stall warning system is inoperative when the master switch is off. At the fuel strainer compartment: FUEL STRAINERS DRAIN ‘ONLY TANK INDICATED BY FUEL SELECTOR. ALLOW SUFFICIENT DRAIN TIME. On fuse panel accessdoor: FUSE PANEL ACCESS DOOR On right rear window moulding in baggagearea: MAXIMUM BAGGAGE AND/OR PASSENGER WEIGHT 250 LBS. IN BAGGAGE AREA INCLUDING SEATS. SEE WEIGHT AND BALANCE. On the instrument panel; MINIMUM SINGLE ENGINE CONTROL SPEED MANEUVERING SPEED MAXIMUM GEAR DOWN SPEED On the instrument panel: 80 MPH 162 MPH 150 MPH WARNING - UNCOORDINATED MANEUVERS, INCLUDING SIDE SLIPS OF 30 SECONDS OR MORE, FOR ANY REASON, AND FAST TAXI TURNS JUST PRIOR TO TAKE-OFF CAN CAUSE LOSS OF POWER IF FUEL TANKS IN USE ARE LESS THAN l/4 FULL INSTRUMENT MARKINGS Wing Flap Setting Take-Off (White Arc) 0” to 15” Landing (Full) 27 ’ P. LANDING GEAR DOWN LIGHT The green gear down light on the instrument panel indicates the landing gear is down and lock1 GEAR INDICATORLIGHTS ARE DIMMED WHILE THE INSTRUMENT LIGHTS ARE ON. 1 REPORT: 1605 PART 1 PAGE 6 MODEL: PA-39 1 FAA APPROVED November 28,Ig REISSUED: June 1, I4 REVISED: March 21,1$ BW 98-05 PIPER, THE NEW AIRWORTHINESS DIRECTIVE SMALL AIRCRAFT & ROTORCRAFT 98-04-27 * THE NEW PIPER AIRCRAF’T CORPORATION: Amendment 39- 10339; Docket _ No. 97-CE-61-AD. Applicability: Models PA-23, PA-23-160, PA-23-235, PA-23-250, PA-E23-250, PA-30, PA-39, PA-40, PA-31, PA-31-300, PA-31-325, PA-31-350, PA-34-200, PA-34-200T, PA-34-220T, PA-42, P&42-720, PA-42- 1000 airplanes (all serial numbers), certificated in any category. NOTE 1: This AD applies to each airplane identified in the preceding applicability provision, regardless of whether it has been modified, altered, or repaired in the area subject to the requirements of this AD. For airplanes that have been modified, altered, or repaired so that the performance of the requirements of this AD is affected, the owner/operator must request approval for an alternative method of compliance in accordance with paragraph (d) of this AD. The request should include an assessment of the effect of the modification, alteration, or repair on the unsafe condition addressed by this AD; and, if the unsafe condition has not been eliminated, the request should include specific proposed actions to address it. Compliance: Required as indicated, unless already accomplished. To minimize the potential hazards associated with operating the airplane in severe icing conditions by providing more clearly defined procedures and limitations associated with such conditions, accomplish the following: (a) Within 30 days after the effective date of this AD, accomplish the requirements of paragraphs (a)(l) and (a)(2) of this AD. NOTE 2: Operators should initiate action to notify and ensure that flight crewmembers are apprised of this change. (1) Revise the FAA-approved Airplane ,Flight Manual (AFM) by incorporating the following into the Limitations Section of the AFM. This may be accomplished by inserting a copy of this AD in the AFM. “WARNING Severe icing may result from environmental conditions outside of those- for which the airplane is certificated. Flight in freezing rain, freezing drizzle, or mixed icing conditions (supercooled liquid water and ice crystals) may result in ice build-up on protected surfaces exceeding the capability of the ice protection system, or may result in ice forming aft of the protected surfaces. This ice may not be shed using the ice protection systems, and may seriously degrade the performance and controllability of the airplane. . During flight, severe icing conditions ‘that exceed those for which the airplane is certificated shall be determined by the following visual cues, If one or more of these visual cues exists, immediately request priority handling from Air Traffic Control to facilitate a route or an altitude change to exit the icing conditions. Unusually extensive ice accumulation on the airframe and windshield in areas not normally observed to collect ice. Accumulation of ice on the upper surface of the wing, aft of the protected area. .- Accumulation of ice on the engine nacelles and propeller spinners farther aft than normally observed. . Since the autopilot, when installed and operating, may mask tactile cues that indicate adverse changes in handling characteristics, use of the autopilot is prohibited when any of the visual cues specified above exist, or when unusual lateral trim requirements or autopilot trim warnings are encountered while the airplane is in icing conditions, l All wing icing inspection lights must be operative prior to flight into known or forecast icing conditions at night. JNOTE: This supersedesany relief provided by the Master Minimum Equipmei~k List(MMEL).)" 2 98-04-27 .I “THE FOLLOWING WEqTeR CONjXTIONS MAY BE ~C@@VE TO SEVERE IN-FLIGHT ICtiG: . Visible rain at temperatures below 0 degrees Celsius ambient air temperature. . Droplets that splash or splat&- on impact at temperatures below 0 degrees Celsius ambient air temperature. PROCEDURES FOR EXITING THE SEVE,j?J3 ICtiG ENWROtiENT: I, ..,. These procedures are applicable to all flight phtises from takeoff to landing. Monitor the ambient air temperature. While severe icing may form at temperatures as cold as -18 degrees Celsius, increased vigilance is warranted at temperatures around free+ng with visible moisture present. If the visual cues specified in the Limitations Section of the AFM for identifying severe icing conditions are observed, accomplish the following: Immediately request priority handling from Air Traffic Control to facilitate a route or an Altitude change to exit the severe icing conditions in ‘order to avoid extended exposure to flight conditions more severe than those for which the airplane has been certificated. . Avoid abrupt and excessive maneuvering that may exacerbate control difficulties. t’ . Do not engage the autopilot. . If the autopilot is engaged, hold the control wheel firmly and disengage the autopilot. If an unusu& roll response ;le angle-of-attacti.’ or uncommanded roll control movement is observed, reduce . Do not extend flaps when holding in icing conditions. Operation with flaps extended can result in a reduced wing angle-of-attack, with the possibility of ice forming on the upper surface further aft on the wing than normal, possibly aft of the protected area. . If the flaps are extended, do not retract them until the airframe is clear of ice. . Report these weather conditions to Air Traffic Control.” (b) Incorporating the AFM revisions, as required by this AD, may be performed by the owner/operator holding-at least a private pilot certificate as authorized by section 43.7 of the Federal Aviation Regulations (14 CFR, 43.7), and must be entered into the aircraft records showing compliance with this AD in accordance with section 43.9 of the Federal Aviation Regulations (14 CFR 43.9). (c) Special flight permits may be issued in accordance with sections 2 1.197 and 2 1.199 of the Federal Aviation Regulatioxis (14 CFR 21.197 and 21.199) to operate the airplane to a location where the requirements of this AD can be accomplished. (d) An alternative method of compliance or adjustment of the compliance time that provides an equivalent level of safety may be approved by the Manager, Small Airplane Directorate, FAA, 1201 Walnut, suite’.900,; Kansas City, Missouri 64106. The request shall be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and then send it to the Manager, Small Airplane Dire&orate. NOTE 3: Information concerning the existence of approved alternative methods of compliance with this AD, if any, mayibe obtained from the Small Airplane Directorate. (e) All persons affected by this directive may examine information related to this AD at the FAA, Central Region, Office of the Regional Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri 64 106. (r) This amend&nt (39-10339) becomes effective on March 13, 1998. YFOR FURTHER INFORMATION CONTACT: Ns, John R* Dow, Sr,, Aerospace Engineer, Small Airplane Directorate, Aircraft Certification Servisls, 1201 w&mt, aujte 900, Kansas City, Missouri 64106; telephone (8 16) 426-6932; f~imiiet (816) 426-2169, . 4 I I II TWIN COMANCHE C/R SECTION III PROCEDURES A. STANDARD PROCEDURES 1. STOPPING ENGINES When operating under high ambient temperature conditions engine shut down by mixture alone may not be positive. Shutting down the engine under these conditions should be as follows: i: Pull the mixture controls to idle cut-off. Depressbutton left side of quadrant. C. Pull back the throttles and hold until engine stops. 2. WARNING This airplane is certified as a normal category airplane and must be operated in compliance with the Airplane Flight Manual. Acrobatic Maneuvers(including spins) are prohibited. Stalls and slow flight should be performed only in accordancewith the Airplane’ Flight Manual. Avoid abrupt maneuvers. Maneuvers at speedsand weights in excessof maneuvering speeds and loadings listed under Limitations Section of this Airplane Flight Manual may subject the airplane to load factors beyond which it is certificated. Maintain at least 5000 feet of terrain clearancewhen practicing stalls. 3. SPINS c All spins are prohibited, however in the event an unintentional spin.is encountered recovery can be accomplished by immediately using the following procedures: a. Retard both throttles to the idle position. b. Apply full rudder in the opposite direction to the spin. C. Push control wheel full forward. While it is not necessary for recovery, the use of ailerons against the turn (i.e. right aileron if spin is to the left) will expedite recovery. d. Maintain controls in these positions until the spin stops. Then neutralize rudder and ailerons. e. Recover from dive with smooth back pressureon the control wheel. No abrupt control movement should be used during recovery from the dive, as the maneuvering speedand positive limit maneuvering load factor may be exceeded. , FAA APPROVED November 28,1969 ’ F ISSUED: June 1,197O 3 REPORT: 1605 PART I PAGE 7 MODEL: PA-39 TWIN COMANCHE C/R B. SYSTEMS OPERATION AND CHECKS I. VOLTAGE REGULATING AND ALTERNATOR SYSTEM (FOR AlRCRAFT SER. NOS. 39-l thru 39-145) a. Alternator System Press-to-test switches, in conjunction with the ammeter, are used to determine the output of each alternator. These switches are located directly below the ammeter. In the normal position the ammeter indicates battery charge or discharge current. Depressing an alternator press-to-test switch causes the ammeter to indicate the output current for the corresponding alternator, viz. the left switch checks the left alternator. A preflight check of the alternators should be made during engine run-up. With both engines operating at approximately 2000 RPM, depress the alternator press-to-test switches individually and check alternator outputs. These should be approximately equal. b. Voltage Regulating System In event of failure of the voltage regulating system an auxiliary regulating system may be switched into the circuit. Abnormal operation may be indicated by zero output on both alternator test positions and a discharge indication for battery. To energize the auxiliary regulating system the following procedure shall be followed: (1) Reduce aircraft electrical load to minimum for continued safe flight. (2) Switch “VOLTAGE REGULATOR SELECTOR” to “AUXILIARY” position. ‘(3) Reset tripped breakers but do not reset “MAIN” Voltage Regulator Breaker. (4) Return to normal required electrical load. If the electrical system still fails to maintain correct output while using the AUX. VOLTAGE REGULATOR system, an alternator failure has probably occurred. To isolate the faulty component the following procedure should be followed: (1) Reduce aircraft electrical load to minimum for continued safe flight. (2) Turn aircraft MASTER SWITCH “OFF.” (3) Place both alternator output circuit breaker switches “OFF.” (4) Reset both MAIN and AUX. Voltage Regulator Circuit Breakers, if tripped. Return Voltage Regulator Selector to “MAIN.” (5) Turn aircraft MASTER SWITCH “ON.” Reset Voltage Regulator Circuit Breaker, if tripped. (6) Close one alternator output circuit breaker switch. Observe if electrical system is operating normally by checking for alternator output current on the ammeter. If not operating properly, open the alternator output circuit breaker; turn aircraft MASTER SWITCH “OFF” for approximately six seconds to reset the overvoltage relay. (7) Turn aircraft MASTER &ITCH “ON.” ResetVoltage Regulator.Circuit Breaker, if tripped. Close other alternator output circuit breaker switch and observe if electrical system is operating normally by checking ammeter indication as above. (8) Check that aircraft electrical load does not exceed the output capability of the operating alternator causing the battery to discharge. CAUTION Use of the voltage regulator selector switch and alternator circuit breakers should be limited to the above conditions. FAA APPROVED November 28,1969 REISSUED: June 1.1970 REVISED: November 30,197l REPORT: 1605 PART I PAGE 9 MODEL: PA-39 TWIN COMANCHE C/R 7 7 L. ALTERNATOR PARALLELING SYSTEM (FOR AIRCRAFT SER. NOS. 39-146 AND UP) On aircraft equipped with the subject system, each alternator is controlled independentI>-b>. its own voltage regulator. These regulators are interconnected electronically so as to provide paralleled outputs from their associatedalternators under normal operating engine RPM ranges. An ammeter that can be switched into either alternator output lead is provided for system monitoring, along with alternator “INOP” warning lights that illuminate when their associated alternator is not producing a voltage. In the event of an alternator “INOP” indication the following steps should be taken: i: Reduce electrical load to minimum for continued safe flight. Turn OFF one section of the Master Switch (L or R, as appropriate) to open the corresponding alternator field circuit. Reset all circuit breakers which may have tripped. C. Turn ON the section of the Master Switch which had been previously turned OFF and if “INOP” light goes out - reinstate electrical load. d. If, after turning ON the section of the Master Switch, the “INOP” light remains lit and/or the alternator circuit breaker switch has tripped, turn the same section of the Master Switch OFF and continue flight with reduced electrical load. e. In the event that both alternator “INOP” lights come on simultaneously, repeat the above procedure for each alternator individually. f. If both alternators fail to return to normal operation, turn ON Master Switch and turn OFF both alternator circuit breaker switches. Terminate flight as soon as possible. CAUTION The alternator circuit breaker switches should not be opened manually wheneverthe alternators are functioning normally. ..-: 3. CIRCUIT BREAKERS AlI circuit breakers are grouped in the lower right comer of instrument panel. To reset the circuit breakers push in on the reset button. 4. FUEL MANAGEMENT a. Normal Operation (1) Take-off and landing (a) Fuel valve “ON” main tanks. (b) Electric fuel pumps “ON”. (2) Cruising (a) Fuel valves “ON” (Main or Auxiliary). (b) Electric fuel pumps “OFF”. b. Crossfeed Operation and Single Engine Operation : A crossfeed is provided to increase the range during single engine emergency operating conditions. Fuel system operation is as follows: ( 1) Cruising (a) When using fuel from tanks on the same side as the operating engine, the following will apply: (1) Fuel Valve “ON” (main or auxiliary) on Operating engine side. (2) Fuel Valve “OFF” on Inoperative engine side. ,- (3) Electric fuel pumps “OFF” (except in case of engine driven pump failure, electric fuel pump on operating engine side must be used). REPORT: 1605 PART I PAGE 10 MODEL: PA-39 . FAA APPROVED November 28,1969 REISSUED: June 1,197O REVISED: November 30,197 1 1 I. I . 1 I 3 rc \ I 1 li9.I TWINCOMANCHEC/R (b) When using fuel from tanks on the opposite side of the operating engine, the following will apply: (1) Fuel Valve “ON” (main or auxiliary) on Inoperative engine. (2) Electric fuel pumps “OFF” (except in case of engine driven pump failure, electric fuel pump on operating engine side must be used). (3) “CROSSFEED ON” on Operative engine side. WARNING DO NOT PUT BOTH FUEL SELECTOR VALVES IN THE CROSSFEED POSITION AT THE SAME TIME. (2) Landing (a) Fuel Valve “ON” main tank on operating engine side. (b) Fuel Valve “OFF” on inoperative engine side. (c) Electric fuel pump “ON” on operating engine side. 5. MANUAL EXTENSION OF LANDING GEAR Check the following before extending the gear manually: t. Circuit breakers- check Master switch - ON ii. Alternators - check Instrument lights - OFF (daytime) To extend the gear, remove the plate covering the emergency disengagecontrol and proceed in these steps as listed on the underside of the cover plate: it. Reduce power - airspeed not to exceed 100 MPH. Place Landing Gear Selector Switch in “GEAR DOWN LOCKED” position. ii: Disengagemotor. Raise motor releasearm and push forward thru full travel. Remove gear extension handle from stowage. If left socket is not in clear position, place handle in right socket. Engageslot and twist clockwise to secure handle. Extend handle and rotate forward until left socket is in clear position. Remove handle and place in left socket and secure. Extend handle. Rotate handle FULL forward to extend landing gear and to engage emergency safety lock. Handle locked in full forward position indicates landing gear is down and emergency safet;lock engaged.Gear “DOWN LOCKED” indicator light should be “ON". NOTE Reducing power and rocking gear extension handle will aid in manually extending the landing gear. DO NOT RETRACT WITH HANDLE IN SOCKET. DO NOT ,RE-ENGAGE MOTOR IN FLIGHT. After the gear has been extended manually, do not perform any unnecessaryoperation to the gear until the aircraft is placed on jacks. FM APPROVED November 28,1%9 REISSUED: June 1,197O REVISED: May l&l973 REPORT: 1605 PART I PAGE 11 MODEL: PA-39 TWIN COMA&HE C/R THIS PAGE INTENi’l’IONALLY LEFT BLANK REPORT-: 1605 PART I PAGE 12 MODEL: PA-39 ,‘,\ FM APPROVED November 28,1%9 REISSUED: June 1.1970 TWIN COMANCHE C/R c. EMERGENCY PROCEDURES 1. DETECTING A DEAD ENGINE a. Loss of Thrust b. Nose of aircraft will yaw in direction of dead engine (with coordinated controls) 1 -, FEATHERING PROCEDURE The propellers can be feathered only while the engine is rotating above 1000 RPM. LOSSof centrifugal force due to slowing RPM will actuate a stop pin that keeps the propeller from : feathering each time the engine is stopped on the ground. If an engine freezes up. it will be impossible to feather its propeller. Single engine performance will decreaseif the propeller of the inoperative engine is not feathered. I it-. C. d. e. f. g. h. i. j. k. 1. m. n. 0. P. q. r. S. t. U. Minimum Control Speed- 80 MPH Best R/C Speed Single Engine - 105 MPH Maintain Direction and Airspeed above 90 MPH Mixture Controls - forward Propeller Controls - forward Throttle Controls - forward Gear - retract Wing flaps - retracted (check visually) Electric Fuel Pumps - “ON” Magnetos - “ON” Identify inoperative engine Throttle of inoperative engine - retard to verify Propeller on inoperative engine - feather Mixture on inoperative engine - idle cut off Electric Fuel Pump of inoperative engine- “OFF” Magnetos of inoperative engine - off Cowl Flaps - close on inoperative engine, use as required on operative engine Alternator circuit breaker switch of inoperative engine - off Electrical Load - reduce to prevent battery depletion Trim - as required Fuel Management - fuel “OFF” of inoperative engine, consider crossfeeduse 3. UNFEATHERING PROCEDURE iti-. Fuel valve on inoperative engine side - “ON” Electric fuel pump - “OFF” Throttle - open l/4 inch I. i. Propeller control - forward . i I e. Mixture - rich f. Magneto switches - “ON” ! g- Starter - engagetill prop windmills h. Propeller - set to cruise RPM i. Throttle - reduced power till engine is warm j. Alternator circuit breaker switch - “ON” FAA APPROVED November 28,1969 REISSUED: June 1,197O REVISED: November 30,197l REPORT: 1605 PART I PAGE 13 MODEL: PA-39 TWIN COMANCHE C/R 4. ENGINE FAILURE DURING TAKE-OFF a. If there is adequate runway remaining for deceleration - CUT POWER IMMEDIATELY AND STOP STRAIGHT AHEAD. b. If there is inadequate runway remaining and 95 MPH is not yet atta,med, the following procedure is recommended: (1) Throttle controls - closed (2) Master switch - off (3) Fuel valves - off (4) Continue straight ahead,turning to avoid obstacles as necessary C. If engine failure occurs when airspeed is at or greater than 95 MPH the pilot must decide whether to abort the take-off or attempt a single engine take-off. His decision must be based on his judgment considering the runway remaining, density altitude, loading, obstruction. weather and his own capability. 5. ENGINE FAILURE DURING CLIMB a. Feather inoperative engine (seefeathering procedure) b. Hold single engine best rate-of-climb speedof 105 MPH C. Monitor cylinder head temperature - adjust cowl flap as required 6. SINGLE ENGINE LANDING a. Feather inoperative engine (see feathering procedure) b. Do not extend landing gear until certain of making field. ii: Do not lower wing flaps until certain of making field. Trim for landing - (rudder) Maintain additional altitude and speedduring approach, keeping in mind that landing should be made right the first time and that a go-around may require the use of fulI power on the operating engine, making control more difficult. A final approach speed of 105 miles per hour and the use of half rather than full wing flaps will place the airplane in the best configuration for a go-around should this be necessary,but it should be avoided if at all possible. Under some conditions of loading or density altitude a go-around may be impossible, and in any event the sudden application of power during single engine operation may causecontrol difficulties. 7. SINGLE ENGINE GO-AROUND If a single engine go-around cannot be avoided proceed as follows: it: Throttle - open Propeller - forward i. Landing Gear - retract Flaps - retract F: Airspeed - “One Engine Inoperative Best Rate-of-Climp Speed” 105 MPH Trim-set g. Cowl Flap - as required (operating engine) NOTE Aircraft will not climb with gear and flaps extended. REPORT: 1605 PART I PAGE 14 MODEL: PA-39 FAA APPROVED November l&l969 REISSUED. June 1 e 1970 TWIN COMANCHE C/R . V I J THIS PAGE-INTENTIONALLY LEFT BLANK FAA APPROVED November 28,1969 REISSUED: June 1,197O REPORT: 1605 PART I PAGE 15 MODEL: PA-39 TWINCOMANdHEC/R 1 THIS PAGE INTENTIONALLY LEFT BLANK REPORT: 1605 PART I PAGE 16 MODEL: PA-39 FAA APPROVED November 28,1969 REISSUED: June I,1970 I II /-‘ -I 3 TWIN COMANCHE C/R. SECTION IV PERFORMANCE The loss of altitude during a power off stall with gear and flapyretracted is 500 feet. Maintain at least 5000 feet of terrain clearance when practicing stalls. I FAA APPROVED November 28,1969 REISSUED: June 1,197O REPORT: 1605 PART I PAGE 17 MODEL: PA-39 TWIN COMANCHE C/R THIS PAGE INTENTIONALLY LEFT BLANK REPORT: 1605 PART I PAGE 18 MODEL: PA-39 FAA APPROVED November 28,1969 ‘a I REISSUED: June 1,197O I TWIN COMANCHE C/R SECTION V SUPPLEMENTS NOTE A FLIGHT MANUAL SUPPLEMENT IS REQUIRED TO BE IN THE AIRPLANE FLIGHT MANUAL ONLY IF THE EQUIPMENT WHICH IS THE SUBJECT OF THE SUPPLEMENT IS INSTALLED. A. Oxygen System B. AutoControl 111 (Model AK161) C. Brittain Tip Tanks D. AltiMatic IIIB E. AntiCollision (Strobe) Lights F. Anti-Icing Equipment G. AltiMatic IIIB-1 (Including AutoFlite II) I I I 1 r 1 1 r L FAA APPROVED November 28,1969 REISSUED: June 1, 1970 REVISED: November 30, 1971 REPORT: 1605 PART I PAGE 19 MODEL: PA-39 TWINCOMANCHEC/R THIS PAGE INTENTIONALLY LEFT BLANK REPORT: 1605 PART I PAGE 20 MODEL: PA-39 FAA APPROVED November 28,1969 REISSUED: June 1,197O TWIN COMANCHE C/R A. OXYGEN SYSTEM INSTALLATION 1. LIMITATIONS No smoking while the oxygen system is in use. ? i. PROCEDURES a. Check pressure gauge in rear of cabin for sufficient pressure for anticipated requirements for flight. Full system pressureis approximately 1850 PSIG. If necessaryrecharge cylinder. b. When oxygen is desired pull out control cable knob placing regulator in “ON” position. C. At seating positions where oxygen is to be used plug mask assembly into oxygen outlet, turn clockwise 90” and apply mask to face. d. The flow indicator located in the mask assembly oxygen line should be checked. When red indicator disappearsthe oxygen is flowing through the line normally. e. To stop flow of oxygen push in control cable knob placing regulator in “OFF” position and remove mask from face. f. Leave mask assembly connected to oxygen outlet for at least 3 minutes to completely bleed down low pressurelines. 3. PLACARDS a. At each oxygen outlet: NO SMOKING WITH OXYGEN IN USE b. At the oxygen control knob: ?i; PULL ON, OXYGEN 4. EMERGENCY In the event that during operation the red indicator appears in any of the flow indicators, the aircraft should be lowered to a safe altitude immediately. FAA APPROVED November 28,1969 REISSUED: June 1,197O REPORT:1605PARTI PAGE21 MODEL. PA-39 TWINCOMANCHEC/R II L L- t' u THIS PAGE INTENTIONALLY LEFT BLANK / , 1 I 'I REPORT: 1605 PART I PAGE 22 MODEL: PA-39 FAA APPROVED November 28,1969 REISSUED: June 1~1970 TWIN COMANCHE C,,‘R B. PIPER AUTOCONTROL III (MODEL AK 16 1) INSTALLATION 1. LIMITATIONS Disengageduring take-off and landing. Disengageabove 225 MPH CAS. 3 u. OPERATING INSTRUCTIONS To Engage: Push Console Heading Lock button (HDG) to “OFF” position. Center ROLL knob. Push ON/OFF button to “ON” position. 3. To Turn: Rotate console ROLL knob in desired direction. For Heading Lock: Set directional gyro (D.G.) with magnetic compass.Push D.G. HDG knob in, rotate to select desired heading. Push console Heading Lock button (HDG) to “IN” position. To Disengage: Push ON/OFF button to “OFF” position. NORMAL OPERATION ii. Be sure airplane is properly trimmed. (Ball Centered) Check vacuum and ascertain that the directional gyro and artifical horizon are functioning properly. L i. EngageAutoControl. (Ground Check Only) Rotate the ROLL knob fulI right and full left. Determine that the control wheel describesa corresponding right and left turn, then center knob. e. Set the directional gyro with the magnetic compass.Push D.G. HDG knob in, rotate to select desired heading. f. Push Console Heading Lock button to “IN” position. The AutoControl is now “locked-m” for directional control. iiF* Turns may be accomplished by either of the following methods: (1) Push Console Heading Lock button to “OUT” position. Rotate the ROLL knob in desired direction. (2) Push Console Heading Lock button to “IN” position. Select new heading by pushing D.G. HDG knob in and rotating. h. Maximum angle of bank should not exceed 20’. i. DisengageAutoControl by pushing the ON/OFF button to “OFF” position. With Piper Radio Coupler Installed : The Autopilot is coupled to the VOR NAV receiver in the modes indicated on the function switch. In the Heading (HDG) mode, the Autopilot is controlled by the directional gyro. FAA APPROVED November 28,1969 REISSUED: June 1,197O REPORT: 1605PARTI PAGE23 MODEL: PA-39 TWIN COMANCHE C/R I 4. EMERGENCY PROCEDURES a. In the event of a malfunction in the AutoControl, push the ON/OFF button to “OFF” position. This disengages the AutoControl from the control system. 111 b. AutoControl may be overpowered manually by exertion of 12(+ 3) pounds force on the control wheel. C. In cruise configuration AutoControl malfunction with a 3 second recovery delay resulted in a 35 degree bank and 30 foot altitude loss. I d. In approach configuration AutoControl malfunction with a 1 second recovery delay resulted in a 25 degree bank and 40 foot altitude loss. REPORT: 1605 PART I PAGE 24 MODEL: PA-39 FAA APPROVED November 28.1969 REISSUED: June 1,197O a TWINCOMANCHECiR C. BRITTAIN INDUSTRIES INC. MODEL l-T-5 TIP TANKS INSTALLATION The information in this document is FAA approved material which. together with the appropriate basic CAA-FAA approved Airplane Flight Manual is applicable and must be carried in the airplane when modified by the installation of Brittain Industries, Inc.: Model TT-5 Tip Tanks in accordancewith S.T.C. SA727WE dated June 25, 1964, revised January 23, 1970. The information in this document supersedes the basic Airplane Flight Manual only where covered in the items contained in this supplement. For limitations and procedures not contained in this supplement, consult the manual proper. 1. LIMITATIONS Same as prescribed in appropriate FAA approved Airplane Flight Manual except: a. Auxiliary wing tip tank fuel to be used in level flight only. b. When using auxiliary fuel, use wing tip tank fuel first. C. Maximum allowable gross weight 3725 Ibs. Any weight in excessof 3600 lbs. must consist of symmetrically loaded fuel in the tip tanks. d. Never exceed airspeedlimit of 230 MPH (red line). C.G. Range with symmetrical fuel quantity in each tip tank: 87.62” to 9 1.38” at 3725 lbs. z 9 C.G. Range with 0.3 gals. in each tip tank :86.72” to 9 1.87” at 36;5 Ibs. C.G. Range with tip tanks empty: 86.5” to 92” at 3600 lbs. C.G. of wing tip tank fuel is 90.5”. Straight line variation between points given. ‘. 2. OPERATING PROCEDURES - NORMAL ‘. Same as prescribed in appropriate FAA approved Airplane Flight Manual except: a. Select tip tank fuel by positioning Fuel Selector Valve in “AUX” position and TIP/AUX Fuel Selector Switch in “TIP” position. b. Select auxiliary tank fuel by positioning Fuel Selector Valve in “AUX” position and TIP/AUX Fuel Selector Switch in “AUX” position. i. Aircraft fuel’gaugesregister fuel quantity of tank selected. TIP/AUX Fuel Selector Switch to be in “AUX” position when using main tank fuel. If the Tip Tanks have been run completely dry in flight, air may be trapped in the line fromeTip Tank to solenoid valve when Tip Tanks are subsequently filled. The air pocket in the line may prevent immediate feeding of fuel from Tip Tanks. To avoid this condition, purge air from lines prior to starting aircraft. (1) Turn fuel selector valve to “AUX” position. 80 85 so 95 C. G. LOCATION (INCHES AFT OF DATUM 1 FAA APPROVED November 28,1969 REPORT: 1605 PART I PAGE 25 REISSUED: June 1,197O MODEL: PA-39 TWIN COlllANCHE C/R (2) Turn on aircraft master switch and place tip tank fuel selector switch to tip tank position. Ascertain that tip tank solenoid switch. under fuel console. is operating by listening for a slight click when switch is operated. (3) Lift up appropriate fuel dram valve and allow fuel to drain. Observe for Ilow in clear plastic tube, followed by interrupted flow of no fuel for a few seconds. further followed by a bubbling flow then full flow. Total drain time should not be less than 30 seconds. (4) Procedure shall be accomplished for each tip tank separately. (5) In addition to above procedure, operate the power plant from each wing tip separately until steady fuel flow is assured during ground run-up prior to flight. 3. PERFORMANCE Same as prescribed in appropriate FAA approved Airplane Flight Manual. 4. EMERGENCY In the event of malfunction of the tip tank fuel system, place the TIP/AUX Fuel Selector Switch in the “AUX” position. NOTE Fuel Solenoid Valve failure automatically switches fuel flow from “TIP” to “AUX” tank. REPORT: 1605 PART I PAGE 26 MODEL: PA-39 FAA APPROVED November 28,1969 REISSUED: June 1,197O REVISED: March 17,197l TWIN COMANCHE C/P D. PIPER ALTIMATIC IllB INSTALLATION I. LIMITATIONS ROLL and PITCH “OFF” during take-off and landing. ROLL. PITCH, and AUTOFLITE “OFF” above 225 MPH. MINIMUM COUPLED APPRO.ACH SPEED ( 110 MPH). 3 -. OPERATING INSTRUCTIONS a. ROLL SECTION To Engage: Push console heading lock button (HDG) “OFF”. Center ROLL knob. Push ROCKER switch to “ON” position. To Turn: Rotate console ROLL knob in desired direction. (Maximum angle of bank should not exceed 30”. Maximum angle will be 70” with heading lock engaged.) Heading Lock: Set directional gyro with magnetic compass. Push directional gyro HDG knob in. rotate to select desired heading. Push console heading lock button (HDG) to “ON” position. Roil Section Ground Check: With heading lock button “OFF”, engage ROLL SECTION. Rotate ROLL knob full right and full left. Determine that the control wheel describes a corresponding right and left turn, then center knob and disengageprior to take-off. b. PITCH SECTION (ROLL SECTION must be engaged prior to engaging PITCH SECTION) : To Engage: Push altitude preselect button (ALT) to “OFF” position. Center the PITCH command disk. Push PITCH button to “ON” position. To ChangeAltitude: Rotate PITCH command disk in desired direction. Altitude Preselect: Center PITCH command disk. With airplane in level flight, rotate the altitude selector DN/UP knob until trim UP/DN indicator is level. Calibrate the altitude indicator to match altimeter by rotating the knurled altitude indicator dial. Rotate the altitude selector knob to select desired altitude. Push altitude preselect button (ALT) to “ON” position. The altitude preslect button may also be engagedwhen the aircraft is climbing or descending. Rotate the altitude selector knob until trim indicator indicates UP or DOWN as desired, then engagethe altitude preselect button. Pitch Section Ground Check: With altitude preselect button “OFF”, rotate the PITCH command disk full DOWN and ful.l UP. Determine that the control wheel describes a corresponding fore and aft movement, then center the disk and disengageprior to take-off. FAA APPROVED November 28,1969 REPORT:1605 PARTIPAGE27 REISSUED: June 1,197O MODEL: PA-39 TWIN COMANCHE C/R C. AUTOFLITE SECTION The AUTOFLITE SECTION of the AltiMatic IIIB is approved for all attitudes. including take-off and landing. The engagement of the ROLL SECTION of the AltiMatic automatically disengagesthe AUTOFLITE SECTION. To Engage: Place AUTOFLITE rocker switch in “ON” position. Correct minor heading variation by rotating AUTOFLITE “TRIM” knob in desired direction. To Turn: Push AUTOPILOT “OFF” button on control wheel. Make turn manually. Release button to re-engageAUTOFLITE on completion of turn. d. PITCH TRIM SECTION The airplane can be trimmed (1) manually with the crank or (2) by actuating the pitch trim slide switch on the pilot’s control wheel. Push switch forward for nose down trim and rearward for nose up trim. Pitch trim is automatically accomplished when the PITCH SECTION is engaged. With Pitch Trim Warning Light InstaIled: The warning light on the instrument panel will light when the pitch is out of trim for approximately 4 seconds,when the pitch section is engaged. The press-to-test feature of the indicator light assembly, when held in (with autopilot pitch function engaged)will show the operator that the lamp is good and the length of time that is required for the warning system to actuate. With Piper Radio Coupler Installed: The ALTIMATIC is coupled to the VOR NAV receiver in the modes indicated on the function switch. In the heading (HDG) mode, the ALTIMATIC is controlled by the directional gyro. 3. NORMAL FLIGHT OPERATION L. Be sure airplane is properly trimmed. (Ball Centered) EngageAUTOFLITE SECTION. . Check vacuum and ascertain that the directional gyro and artificial horizon are functioning properly. d. Engage ROLL SECTION. F: EngagePITCH SECTION. DisengageROLL and PITCH SECTIONS before landing. 4. WITH GLIDE SLOPE COUPLER INSTALLED (FLIGHT) a. Set Radio Coupler to localizer normal mode (LOC/NORM). b. “Engage” altitude hold. i. Extend landing gear and set flaps 15 ’ maximum. Adjust power to maintain desired approach speed 110 to 120 MPH. 5. EMERGENCY PROCEDURES a. In the event of a malfunction in the ROLL or PITCH SECTION, push the ROLL ON/OFF button “OFF”, or push the A/P “OFF” button on the control wheel. REPORT: 1605 PART I PAGE 28 MODEL: PA-39 FAA APPROVED November 28,1969 REISSUED: June 1.1970 REVISED: March 17,1971 TWIN COMANCHE C/R Either operation above disengages both ROLL and PITCH SECTIONS of the ALTIMATIC from the control system and will disengage the AUTOFLITE SECTION if engaged. as long as button remains depressed. After the A/P “OFF” button on the control wheel has been pushed due to a malfunction in the ROLL or PITCH SECTION. the ALTIMATIC IIIB can only be reengaged by the actuation of the ROLL ROCKER switch and PITCH button on the console. AUTOFLITE SECTION will automatically re-engage upon release of the A/P button. b. The PITCH TRIM SECTION may be overpowered manually. In the event of a malfunction in the PITCH TRIM SECTION, pull the Electric Trim circuit breaker. - The AltiMatic ROLL SECTION and AUTOFLITE SECTION may be overpowered manullly by exertion of 12 f 3 pounds of force on either control wheel. The AltiMatic PITCH SECTION may be overpowered manually by exertion of 15 + 2 pounds of force on either control wheel. d. In cruise configuration, AltiMatic malfunction with a 3 second recovery delay results in a 45 o bank and 300 foot altitude loss. e. In approach configuration, AltiMatic malfunction with a 1 second recovery delay results in a 28” bank and 220 foot altitude loss. f. In cruise configuration, AutoFlite malfunction with a 3 second recovery delay results in a 35 ’ bank and an 80 foot altitude loss. is. In approach configuration, AutoFlite malfunction with a 1 second recovery delay results in a 25” bank and 40 foot altitude loss. 6. PLACARDS a. On left control wheel: x DWN UP A v b. On left control wheel: A/P OFF b C. On instrument panel: AutoFlite Trim L R d. On instrument panel (when radio coupler is installed): A/P NAV SEL ON OFF ON e. On instrument panel (when Glide Slope Coupler is installed) G/S ENGAGED f. On instrument panel (when p$chL$n warning is installed): OUT OF TRIM .’ FAA APPROVED November 28,1969 REISSUED: June 1,197O REVISED: March l7,1971 REPORT: 1605 PART I PAGE 29 MODEL: PA-39 TWIN COMANCHE C/R z THIS PAGE INTENTIONALLY LEFT BLANK REPORT: 1605 PART I PAGE 30 MODEL: PA-39 FAA APPROVED November 28,1969 REISSUED: June 1,197O TWIN COMANCHE C’R E. ANTI-COLLISION (STROBE) LIGHT INSTALLATION 1. OPERATING LIMITATIONS a. The following shall apply when RED anti-collision (Strobe) Lights are installed: WARNING TO AVOID OPTICAL ILLUSION AND SEVERE VERTIGO, TURN ANTI-COLLISION LIGHTS OFF UPON ENTERING CLOUDS. FOG OR HAZE. b. The following shall apply when supplementary WHITE anti-collision (Strobe) lights are installed : WARNING TURN OFF STROBE LIGHTS WHEN TAXIING f-i VICINITY OF OTHER AIRCRAFT, OR DURING FLIGHT THROUGH CLOUD, FOG OR HAZE. STANDARD POSITION LIGHTS TO BE ON FOR ALL NIGHT OPERATIONS. FAA APPROVED November 28,1969 REISSUED: June 1,197O REPORT: 1605 PART I PAGE 3 1 MODEL: PA-39 I TWIN COMANCHE C/R L b L , b REPORT: 1605 PART I PAGE 32 MODEL: PA-39 THIS PAGE INTENTIONALLY LEFT BLANK FAA APPROVED November 28,1969 ADDED: November 2,197O TWIN COMANCHE C/R F. ANTI-ICING EQUIPMENT INSTALLATION The following items of equipment must be installed and operable to effectively cope with normally encountered icing conditions: I. II. III. IV. V. 1. Wiggins Airways Wing and Tail Pump Type Pneumatic Deicing Boots installed per Wiggins Airways S.T.C. No. SA848EA dated October 22, 1970, or FAA approved equivalent. Wiggins Airways Alcohol Propeller Anti-Icing Kit installed per Wiggins Airways S.T.C. No. SA8 16EA dated May 7, 1970, or FAA approved equivalent. Piper Antennas installed per P.A.C. Dwg. 25043. No special operating instructions required. ‘HeatedPitot Head installed per P.A.C. Dwg. 26732. No special operating instructions required. Piper Heated GlassPanel on Windshield installed per P.A.C. Dwg. 26711. PLACARDS a. On circuit breaker panel: b. When heated glasspanel only is installed, on instrument panel: “Warning - This aircraft is not fully equipped for flight in icing conditions.” NOTE WHEN ALL ITEMS OF EQUIPMENT LISTED ABOVE ARE INSTALLED, PLACARD NO. 1-b. IS NOT REQUIRED. 2. OPERATING INSTRUCTIONS Prior to flight in conditions’where the possibility of encountering icing exists, the HEATED PANEL assembly should be attached to the aircraft and the lead wire plug firmly inserted in socket provided. WSHLD HEAT An operational check should then be accomplished by turning the HEATED PANEL switch Q& for a period not exceeding 30 SECONDS. Proper operation is indicated by the glasssection being warm to the touch. If icing conditions are encounteredthe HEATED PANEL switch should be turned ON and remain 9 until the icing conditions cease. When icing is not prevalent the unit should be turned OFF. UNDER NO CIRCUMSTANCES SHOULD THE UNIT BE TURNED ON FOR A PERIOD EXCEEDING 30 SECONDS UNLESS: it. The aircraft is in flight, or Ice exists on the HEATED PANEL. 1969 FM APPROVED November 28, ADDED: November 2,197O REPORT: 1605 PART I PAGE 33 MODEL: PA-39 TWIN COMANCHE C/R CAUTION This equipment cannot be expected to cope with heavy or very prolonged moderate icing conditions. The latter can be expected to tax the equipment beyond its capacity. Pilots should always strive to avoid heavy icing conditions. If heavy icing is encountered unexpectedly or unavoidably, prompt action must be taken to get into more favorable flying weather conditions. NOTE When all items of equipment listed above are installed, the placard “Warning - This aircraft is not fully equipped for flight in icing condition.” IS NOT REQUIRED. When the heated panel is removed or any of the above listed installed equipment is inoperable (known before flight) the warning placard must be reinstalled. REPORT: 1605 PART I PAGE 34 MODEL: PA-39 FAA APPROVED November Z&l969 ADDED: November 2.1970 J z GENERALSPECIFICATIONS Periormance Weights PowerPla;lt:::::::::::::::::::::::::::::::::::::::::::::::::: ; FuelAndOil . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . BaggageArea . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ; Dimensions LandingGear::::::::::::::::::::::::::::::::::::::::::::::::: 32 TWIN COMANCHE C,‘R GENERAL SPECIFICATIONS PERFORMANCE Performance figures are for standard PA-39 airplanes equipped for cross-country transportation and flown at gross weight under-standard conditions at sea level unless otherwise stated. Performance for a specific airplane may vary from published figures depending upon equipment installed, the condition of engines, airplane and equipment, atmospheric conditions and piloting technique. Take-off Run (short field, ft) Take-off Distance Over 50-ft Barrier (short field, ft) Accelerate - Stop Distance (ft) Vmc MPH (as determined by the F.A.A.) Stalling Speed (gear and flaps down, power off, mph) Stalling Speed (gear and flaps up, power off. mph) Best Rate of Climb (ft per min) Best Rate of Climb Speed (mph) Best Angle of Climb Speed (mph) Single Engine Rate of Climb (ft per min) Best Single Engine Rate of Climb Speed (mph) Absolute Ceiling (ft) Service Ceiling (ft) Single Engine Absolute Ceiling (ft) *i Single Engine Service Ceiling (ft) Landing Roll (short field, flaps down, ft) Landing Over 50-ft Barrier (short field, flaps down, ft) ALTITUDE CRUISING SPEEDS (MPH) Normally Aspirated Power Man. Setting Press. (rpm> (in.1 Altitude (feet) Normal Intermediate Economy Longe Range 2400 26 4,200 2400 24 - . 6,400 2200 24 6,400 2200 20 11,200 Turbocharged Power Setting (rpm) Man. Press. (in.) Altitude (feet) Turbo 2400 28 20,000 Intermediate 2400 26 20,000 Economy 2200 24 20,000 Long Range 2200 22 20,000 - GENERAL SPECIFICATIONS ISSUED: June 1.1970 Revised: November 197 1 Speed (mph) 198 196 188 178 Speed Range (mph) (miles) 234 1210 22.6 222 1475 17.2 203 1605 14.4 191 1650 13.2 Normally Turbo Aspirated Charged 940 990 1530 1590 2470 2560 80 80 70 70 76 76 1460 1290 112 117 90. 90 260 165 105 105 20,000 25 >ooo+ 18,600 25 ,OOO+ 7,100 12,600 5 3800 8,800 700 7’5 1,870 1,900 Range (miles) 830 1030 1110 1200 Fuel Con. (iid 19.6 15.4 13.8 11.6 Fuel Con. (gph) PAGE 1 IF---- TWIN COMANCHE C/R Normally Turbo Aspirated Charged WEIGHTS Gross Weight (lbs) Empty Weight (Standard) (Ibs) USEFUL LOAD (Standard) (Ibs) * These weights are approximate POWER PLANT Right Engine - Lycoming LlO-320-BlA L 10-320-C1A Left Engine - Lycoming 1O-320-B 1A IO-320-ClA Rated Horsepower 160 160 Rated Speed (rpm) 2700 2700 Bore (in.) 5.125 5.125 Stroke (in.) 3.875 3.875 Displacement (cubic in.) 3 19.8 319.8 Compression Ratio 8S:l 8.511 Dry Weight (lbs) 295 308 FUEL AND OIL Fuel Capacity (U.S. gal) 90 1’0 Unusable fuel (inboard cells only) 6 6 Fuel, Aviation Grade (minimum octane) 100/130 100/130 Oil Capacity (qts) (each engine) 8 8 BAGGAGE AREA Maximum Baggage(lbs) 250 250 BaggageSpace(cubic ft) 20 20 BaggageDoor Size (in.) 19 x.21 19 x 21 DIMENSIONS Wing Span (ft) Wing Area (sq ft) Length (ft) Height (ft) Wing Loading (lbs per sq ft) Power Loading (Ibs per hp) Propeller Diameter (in.) PAGE 2 3600 3725 2270* 741h* 1330* 1309 * 36 36.8 178 178 25.2 25.2 8.2 8.2 20.2 20.9 11.3 11.7 72 72 GENERAL SPECIFICATIONS ISSUED: June 1,197O TWIN COMANCHE C/R LANDING GEAR Normally Turbo Aspirated Charged Wheel Base (ft) Wheel Tread (ft) Tire Pressure (psi) Nose Main Tire Size Nose (six-ply rating) Main (six-ply rating) , GENERALSPECIFICATIONS ISSUED: June 1,197O 1.3 7.3 9.8 9.8 42 42 42 42 6.00 x 6 6.00 x 6 6.00 x 6 6.00 x 6 PAGE 3 TWINCOMANCHEC/R PAGE 4 GENERAL SPECIFICATIONS ISSUED: June 1,197O DESCRIPTION AIRPLANEAND SYSTEMS Engine and Propeller ............................................ 1 Tile Optional Rajay Turbocharger ..................................... 1 Fuel Injection ................................................ 3 St~ctures ................................................... 3 LandingGear ................................................. 3 Control System ............................................... 4 Fuel System . ................................................. 5 Electrical System .............................................. 7 \‘dcuumSystem ................................... : ........... 9 Instrument Panel ............................................... 10 Instrument Static Pressure System ..................................... 10 Heating and Ventilating System ...................................... 12 Seats ...................................................... 14 Finish .................................................... . . 14 Baggage Area ................................................. 14 Srall Warning System ........ .;. .................................. 14 Propeller Synchrophaser ............ -.-. ............................ 15 TWIN COMANCHE C/R DESCRIPTION AIRPLANE AND SYSTEMS ENGINE AND PROPELLER The four cylinder Lycoming fuel injected engines are rated at 160 horsepower at 2700 RPM. The left engine rotates clockwise and the right engine rotates counterclockwise. These engines are equipped with geared starters, fuel injectors shielded ignition systems and special enginesare available with optional Rajay Turbochargers. Engine mounts are of steel tube dynafocal mount construction. Engine cowls are cantilever structures attached at the firewall, with side panelswhich are quickly removed by means of quick releasefasteners. The exhaust system is a cross-overtype with exhaust gasesdirected overboard at the bottom of the nacelles in the area of the cowl flaps. The cowl flaps are located on the bottom of the engine nacelles and are manually operated by push-pull controls located in the cabin to the right of the power control quadrant. Oil coolers are mounted on the left rear baffle of each engine.Air passesthrough the oil coolers before reaching the area of the cowl flaps. The propellers are Hartzell, constant-speed, controllable, full-feathering units. These are controlled entirely by use of the propeller control levers located in the center of the power control quadrant. Feathering of the propellers is accomplished by moving the controls fully aft through the low RPM detent into the feathering position. Feathering takes place in approximately six seconds.A propeller is unfeathered by moving the prop control ahead and engaging-thestarter. THE OPTIONAL RAJAY TURBOCHARGER The Rajay Turbocharger consists of a precisely balanced rotating shaft with a radial inflow turbine wheel on one end and a centrifugal compressorimpeller on the other. Normally wasted exhaust’gasesdrive. the turbine which powers the impeller that supplies air under pressure to the engine air inlet. Since engine horsepower is never used to drive the turbocharger, no additional mechanical load is imposed above those of a naturally aspirated engine. Low altitude power and engine efficiency is thus provided at high altitudes by converting exhaust gas energy to compressedair. Although true airspeedsincrease, turbocharging does not increasenormal engine speeds,loads, or brake mean effective pressurelimits. The turbocharger bearings are lubricated with engine oil at a pressureof 30-50 psi at normal operating temperatures. Pressure switches within the lubricant supply line will activate a red warning light if the turbocharger oil pressure falls below 27-30 psi. If the light signal is received, the pilot should deactivate the turbocharger. The primary control in turbocharger operation is the degree of closure of the waste-gatevalve. Boost pressureis manually regulated by gradually turning the vernier control knobs to allow matching of manifold pressurein both engines. The induction system uses the normal throttle and fuel mixture controls that are used in non-turbo operation. The air inlet box is designed to accommodate compressor discharge air without decreasing normally aspirated take-off power. If the turbo compressor fails to deliver boost pressure,a check valve door will open automatically as the engine reverts to normally aspirated operation. This function prevents a prohibitive drop in throttling pressure and the corresponding loss of power that would occur if all the engine air for normal operation were routed through the inoperative compressor. Alternate air is available automatically if the inlet duct is obstructed during normally aspirated operation. AIRPLANE AND SYSTEMS ISSUED: Jung 1,197O PAGE 1 TWIN COMAlyiCHE C/R PAGE 2 AIRPLANE AND SYSTEMS ISSUED: June 1,197O TWIN COMANCHE C/R FUEL INJECTION The Bendix RSA-5 fuel injection system is based on the principle of meaWing engine air COnSumptiOn by use of a venturi tube and using airflow to control fuel flow to the engines. Fuel distribution to the cylinders in accomplished by a fuel flow divider. Fuel pressure regulation by means of the servo valve causes a minimal drop in fuel pressure throughout the metering system. Metering pressure is maintained above vapor forming conditions while fuel inlet pressure is low enough to allow the use of a diaphragm pump. Vapor lock and associated problems of difficult starting are thus eliminated. Incorporated in the servo regulator is the airflow sensing system which contains a throttle valve and venturi. The differential pressure between the entrance and the throat of the yenturi is the measurement of air entering the engine. These pressures are applied across an air diaphragm in the regulator. A change in power changes the airflow to the engine and across the diaphragm in the regulator. Mounted on top of the engine is the ported fuel flow divider with four nozzle lines routed to the cylinders. The divider contains a spring loaded positive shut-off valve. Within each cylinder are continuous flow air bleed nozzles with provisions to eliminate the adverse effects of low manifold pressure when idling. Since fuel metering is provided by the servo regulator rather than the nozzles, more uniform cylinder head temperatures result and a longer engine life is possible. Induction air for the engine enters the opening in the nose cowl and is picked up by a large air duct at the right rear baffle. The air is directed through a filter and on to the servo regulator. An alternate air source for the induction system contains a spring loaded door at the throat of the servo regulator. This door operates automatically if primary source is obstructed or manually by the push-pull control ‘on the right side of the power control quadrant. The primary.system should always be used for take-off. ‘Q STRUCTURES Structures are of sheet aluminum construction and are designed to ultimate load factors well,,in excess of normal requirements. All components are completely zinc chromate primed and exterior surfaces are coated with acrylic lacquer. The main spars of the wings are joined with high strength butt fittings in the center of the’ fuselage, making in effect a continuous main spar. The spars are attached to the fuselage at the side of the fuselage and in the center of the structure; wings are also attached at the rear spar and at an auxiliary front spar. The wing airfoil section is a laminar flow type, with maximum thickness about 40% aft of the leading edge. This permits the main spar, located at the point of maximum thickness, to pass through the cabin under the rear seat, providing unobstructed cabin floor spaceahead of the seat. LANDING GEAR The nose gear is steerable with the rudder pedals through a 40 degree arc. During retraction of the gear, the steering mechanism is disconnected automatically to reduce rudder pedal loads in flight. The nose gear is equipped with a hydraulic shimmy dampener. Retraction of the landing gear is accomplished through the use of an electric motor and gear tram, actuating push-pull cables to each of the main gear and a tube to the nose gear. The landing gear motor is beneath the center floor panel and the selector switch on the instrument panel to the left of the power control quadrant. To guard against inadvertent movement of the landing gear selector on the ground, the handle must also be pulled aft before moving it upward. The gear selector has the shapeof a wheel to distinguish it from the electric flap control which has an airfoil shape. As an added safety feature, the warning horn is connected to the gear selector switch. The horn will then operate if the selector is moved to the UP position ARPLANE AND SYSTEMS ISSUED: June 1,19X) PAGE 3 TWIN COMANCHE C/R Landing Gear Selector Switch with the master switch on and the weight of the airplane on the landing gear. To prevent gear retraction on the ground, an anti-retraction switch is installed on the left main gear. This prevents the completion of the electric circuit to the landing gear motor until the gear strut is within 3/4 inch of full extension. . The gear indicating lights are located conveniently by the gear selector switch. The green indicating light below the selector switch shows that all gear are down and locked. The amber light above the gear selector switch is the gear up indication: it will flash if the power of one engine is reduced below 12 inches of manifold pressurewhile the gear is up and locked. The white light indicates that the landing gear is in transit. The gear up warning horn will sound when power is reduced (below approximately 12 inches of manifold pressure) on both engines and the gear is not down and locked. The pilots should become familiar with the gear warning horn to distinguish it from the stall warning horn. GEAR INDICATION LIGHTS ARE DIMMED WHILE THE INSTRUMENT LIGHTS ARE ON. The brakes are actuated by toe brake pedals mounted on the left set of the rudder pedals.Hydraulic brake cylinders above the brake pedals are accessiblein the cockpit for servicing. Parking brake valvesare incorporated in each cylinder and have two cablesattached from the parking brake “T” handle. To prevent inadvertent application of the parking brake in flight, a safety lock is incorporated in the valves, thus eliminating the possibility of pulling out the “T” handle until pressure is applied by use of the toe brakes. Toe brakes for the right side are available as optional equipment. A tow bar is provided with each airplane. When not in use it is stowed next to the main spar. It may be moved by lifting the flap covering the forward side of the spar and removing the bar from its fasteners. When towing with power equipment, caution should be used not to turn the nose gear beyond its 40 degreearc as this may causedamage to the nose gear and steering mechanism. CONTROL SYSTEM Dual flight controls are provided as standard equipment. Cables connect the movable control surfaces with the rudder pedalsand control columns. Directional and longitudinal trim is provided by an adjustable trim mechansim for the rudder and stabilator. The manual rudder trim control is located to the right of the throttle quadrant. Max-Lift electrically operated flaps are usedon the Twin Comanche C/R. The flaps are operated by an electric motor; they can be lowered and stopped in any desired position. The airfoil shapedflap control is to the right of the power control quadrant. A flap position indicator is located on the instrument panel marked to show the position of the flap relative to the wing. A range for take-off operation is also shown. PAGE4 . AIWLANEANDSYSTEMS ISSUED: June 1,19X) TWIN COMANCHE C/R J Fuel Selector f COWL FLAP I L Rudder Trim And Flap Controls Located in the inboard end of the right flap is a lock which holds the flap in the UP position so that it can be used as a step for entry or exit. A second lock is incorporated to prevent the flap from going full down in casea step load is applied and ;the up lock is not fully engaged. FUEL SYSTEM !’ The fuel is carried in four integral fuel cells located in the leading edge sections of the wings. Capacity of the two main fuel cells is 30 gallons each, of which 27 gallons is usable. “i The auxiliary fuel system consists of two 15 gallon cells (all usable) installed in the wings just outboard of the main fuel cells. Wing tip tanks are available as optional equipment on the normally aspirated Twin Comanche C/R and standard equipment on the Turbocharged Twin Comanche C/R. AUXILIARY FUEL AND TIP TANK FUEL IS TO BE USED IN LEVEL FLIGHT ONLY. The cells should be kept full of fuel during storageof the airplane to prevent accumulation of moisture and deterioration of the cells. For storage of more than ten days without fuel, the cells should be coated with light engine oil to prevent excessivedrying. During normal operation use the engine driven fuel pump to draw fuel from the cell directly adjacent to that engine. However, fuel can be drawn from any cell to both enginesthrough use of the engine driven fuel pump or the electric auxiliary fuel pump. For emergency single engine operation a crossfeedis provided to increase the range. The fuel strainers for the system are located beneath the floor panel in the center section of the fuselage. Daily draining of the strainers may be accomplished in the cockpit by opening the hinged access door located in the floor panel just aft of the fuel selector handles and pulling up on the knob located in the center of the selector valve. The general procedure for draining the fuel system is to open the strainer quick drain for several seconds with the fuel cell selector on the main cell, then to change the selector to the auxiliary cell and repeat the process. To drain tip tanks, if installed, turn the master switch on and follow the procedure for the main and auxiliary cells. Allow enough fuel flow to clear the lines as well as the strainer. Positive fuel flow shut-off can be observed through the clear plastic tube that carries the fuel ANPLANEANDSYSTEMS ISSUED: June 1,197O PAGE 5 -- TWIN COMANCHE C/R Fuel System Schematics PAGE 6 AIRPLANE AND SYSTEMS ISSUED: June 1,197O TWIN COMANCHE C/R overboard. Located inside the fuel valvesis a by-passvalve which will open at l/2 psi differential pressureif the strainer screenbecomesblocked. Fuel quantity is indicated by two electric gauges located below the instrument cluster. The instruments are connected to a transmitter unit located in each fuel cell. The gaugeswill indicate the amount of fuel available in the cells that are selected. ELECTRICAL SYSTEM Electrical power is supplied by two 12 volt 70 ampere alternators and a 35 ampere-hour battery. The battery is located in the fuselage nose section in a sealed stainless steel battery box. Refer to the Maintenance Section for servicing of the battery. There are two voltage regulator systems installed in the Twin Comanche aircraft. On aircraft up to and including serial number 39-145 the alternators are paralleled by using one voltage regulator to control field voltage of both alternators. Also incorporated in the system is an overvoltage relay. Its function is to open and remove field voltage to the unregulated alternators in the event of a failure of the voltage regulator, thus preventing an overvoltagecondition which could damagethe electrical equipment. In the event of a voltage regulating system failure, an auxiliary regulating system may be switched into use. Abnormal system operation may be indicated by zero output on both alternator test positions and a dischargeindication for the battery. NOTE Use of the Volta& regulator selector switch should be limited to the above condition unless the Service Manual is consulted. On aircraft with serial number 39-146 and up, the electrical system incorporates an alternator ’ paralleling system. The system has two voltage regulators which control the alternators. The regulators are interconnected electrically to provide parallel outputs from their respective alternators under normal operating engine RPM. Whenever the engines are operating at a hig,h-differential RPM, the alternator inoperative light for the slower engine may come on. A distinctive feature of this alternator paralleling system is the split rocker type master switch and the alternator inoperative lights. Alternator inoperative lights are located below the ammeter. The lights illuminate when the respective alternator fails to provide voltage. The lights should be checked prior to starting the engines, to seeif the bulbs are burned out, by turning on the master switch. If the bulbs do not illuminate the bulbs should be replaced. NOTE Refer to the Airplane Flight Manual Section for corrective action for an alternator failure. An ammeter, in conjunction with alternator push-to-test switches, is used to check the electrical system output. All of these items are located on the pilot’s instrument panel. The ammeter normally indicates battery charge or dischargecurrent. Depressingeither of the push-to-test switches will causethe ammeter to indicate the respectivealternator output current. AIRPLANE AND SYSTEMS ISSUED: June 1,197O REVISED: March 30,1973 PAGE 7 TWIN COMANCHE C/R Electrical switches for the various systems are located primarily in the lower left sub-panel. Rheostat II switches control the interior lighting: the instrument post light switch is on the left sub-panel just below the parking brake handle. the map light switch is just to the right of the brake handle and the switch for back lighting of the radio selector panel is centrally located at the top of the radio stack. The instrument panel II can be illuminated from the cabin ceiling by turning the switch in the. center of the overhead light panel. The alternator shunt fuses and spare fuses are located under the floor panel access door aft of the nose wheel well. The circuit breakers are located on the lower right sub-panel and automatically break the I electrical circuit if an overload occurs. It may be necessary to allow approximately two minutes for the 7 circuit breakers to cool before resetting them if an overload occurs. Corrective action should be taken in event of continual circuit breaker popping. It is possible to trip the breaker manually by pulling out the reset button. Do not open alternator switches in flight unless an emergency arises. If power loss occurs, I reduce electrical load to a minimum and terminate flight as soon as practical. PAGE 7a AIRPLANE AND SYSTEMS ISSUED: February 20.1972 - I TWIN COMANCHE C/R RIGHT LEFT VOLTAGE CENTER RIGHT BONDING BUS Gl”“I,I , FILTE- Q Q .EFT LLTERNATOR I I I I - - -.1. .- nlwl 1. --1...1 LEFT ALTERNATOR AMMETER ALTERNATOR Electrical System Schematic Serial Number 39-146 and up AIRPLANE AND SYSTEMS ISSUED: June 1,197O REVISED: February 20.1972 PAGE 8 1 I_ I 1111 I I I FFT l-1 ’ I I I I ------ Y,;‘- - - -’ L-----JlII II I I AUX. VOLT REGULATOR AMMETER +- Pl Electrical System Schematic Up to and including Serial Number 39-145 PAGE 7b AIRPLANE AND SYSTEMS ISSUED: February 20,1972 REVISED: March 17.1972 TWIN COMANCHE C/R VACUUM SYSTEM Suction for the vacuum operated gyro instruments is supplied by two engine driven (dry type) vacuum pumps, interconnected to form a single system. Either vacuum pump has sufficient capacity to operate the gyro instruments. If suction is lost from one or the other side a check valve automatically closes and suction is supplied by the remaining system. A vacuum gauge is installed in the instrument panel to provide’a constant indication of vacuum source. Incorporated in the instrument are two red indicators (right and left systems). During normal operation the indicators are not visible, but if vacuum is lost, for example on the right side, then the right indicator will be visible. Suction is indicated on the gauge in inches of mercury; normal operating range is 4.8 to 5.1 inches. The system is controlled by two adjustable regulators, one located in each engine nacelle. After initial adjustment the regulators require very little attention. VACUUM MANIFOLD CHECK VALVES I 1 CENTRAL AIR FILTER REGULATOR a5LEFT ENGINE VACUUM PUMP 1 ARTIFICIAL HORIZON VAhJUM GAGE URPLANEANDSYSTEMS Vacuum System Schematic I ‘. REGULATOR RIGHT ENGINE VACUUM PUMP ISSUED: June 1,197O PAGE 9 TWIN COMANCHE C/R INSTRUMENT PANEL The instrument panel is designed to accommodate the customary advanced flight instruments on the left side in front of the pilot and engineinstruments on the right side. The optional instruments such as the gyro instruments of the flight group are shock mounted. The Artificial Horizon and Directional Gyro in this group are operated by an optional vacuum pump on each engine. The vacuum system standby is the optional, electrically operated Rate of Turn indicator. INSTRUMENT STATIC PRESSURE SYSTEM Static air for the airspeed indicator, altimeter and vertical speed indicator is supplied from two static ports, one located on each side of the aft fuselagesection. An optional, alternate static source is available to provide continuous operation of the pitot static instruments should the static system ports or lines become obstructed. The alternate static source is located in the cabin on the lower left side of the control quadrant. If incorrect instrument readings are suspected the alternate static source valve should be opened, venting the static system to cabin pressure.Cabin pressurewill vary, however with cabin ventilators open, cabin heater operating and various airspeeds. Use of the alternate static source may result in the following instrument indications: The altimeter reads higher than normal; indicated airspeed greater than normal; and the vertical velocity indicator momentarily shows a climb. The following table shows airspeed corrections for the standard static system and the alternate static system: IAS MPH 80 88 120 160 200 220 80 91 100 120 Standard Alternate Static System Static System CAS CAS MPH MPH 82 81 90 91 121 113 160 148 197 185 216 204 80 81 90 91 98 97 117 113 Gear and Flaps Retracted Extended PAGE 10 AIRPLANE AND SYSTEMS ISSUED: June 1,197O ,OnlN03 n31sond30 ‘OE S,O(llN03 lV3” NISI3 ‘6C ,3NVd n3nv3ns lln3NlJ ‘8V L131HO,, NV913 ‘LV ,OLllN03 N33hYO ‘9P 53H31,MS tlOlVNIJ3llV ‘Et Y--l SdVld lMD3 ‘VP U-7 tl,V 31VNY3llV ‘EP WlYl n3aantl ‘It “311MS no1v”lJv dl-ld ‘Iti L13,d”03 H,VOWdV,hVN ‘OV NOlV’3,ON, dVll ‘6E SNOl,“S lS31 NOlVNY31,V ‘LE lH9IldOlrlv ‘9E Ya13313s tlV39 DNIaNV, ‘SE “31,MS L1OlJ3-,3s ‘AVN ‘tE L1313WWV ‘EC IHO,, dVW ‘ZE 1(01V3,(1N, 3L(“lYIW ‘IE lV'3" LOlId ‘OE LI IHO, LINIONV, ‘62 7 1HOI-l ONIONI, ‘91 H31lMS SJ.H9ll 1SOd ‘LZ 3,ONVH 3WVYS ‘WtlVd ‘91 NO,S,J,03-,,lNV '31lHM NOISl,,O,-IlNV '(13(1 52HOIl AVN H311MS ’73s ‘83Ll’llOh H31,Ms 831SVW 3wa S3f)“V’f) 3NION3 ONV 13fl.i nO*vJ,aNl MaId 1 3 1 1 1 L1313WOH3Vl IHOlY 36”V9 3Y”SS3L(d alOdlNvW L(313WOHJV.l 1 3 3 1 811, 311VW,l,V SU3A133SNVYI AVN,WWOD ‘52 ‘VZ ‘EZ ‘ZZ ‘11 ‘OZ ‘61 ‘WI ‘LI ‘91 ‘SI ‘PI 0,avu dav *EI ,3NVd MO133135 Oiavu ‘21 130NOdSNVkll *I I Na3V38 )IIWMVW ‘01 (10133,3s sNlt1V38 INWO ‘6 ew,,3 JO 32VB ‘8 (1313WlLlV ‘L o&i*9 lVN01133n1a ‘9 1HI)I-I L13ldfl03 S/D ‘9 NOZltlOH OtlAD ‘V II4617 6NlNYVM 1lVlS ‘E t101V,,ON, a33dstliv ‘1 l,OlV3,aNI Nil”1 4 0 31VY *I TWIN COMANCHE C/R HEATING AND VENTILATING SYSTEM The flow of air for heating and defrosting is taken through an inlet located in the nose and regulated by controls located in the lower right side of the instrument panel. Heated air for the cabin and windshield defrosting is provided by a Janitrol heater installed in the nose section. Operation of the heater is controlled by a three position switch, labeled FAN, OFF and HEAT. The FAN position will operate the vent blower only and may be used for cabin ventilation on the ground or for windshield defogging when heat is not desired. Iz For heat the manual heater fuel valve must be on and the three position switch turned to HEAT. This will start fuel flow and ignite the burner simultaneously. With instant starting and no need for priming, heat should be felt within a few seconds. I Regulation of heat, airflow and defroster operation is controlled by levers on the heater control console. The top control regulates a thermostat and provides a wide range of temperature selections. Adjustable heat deflectors are on the cabin wall beneath the instrument panel to provide additional comfort. Cabin temperature and air circulation can be maintained by using various combinations of lever settings, to suit individual desires.To minimize the feeling of drafts, a low airflow high heat combination may be used. Windshield defrosting may be regulated by various settings of the defroster lever and in severe windshield fogging or icing conditions, it may be desirable to restrict the heater air, since this will drive more air through the defrosters. Heat may be supplied to warm the cabin before flight by turning on the master switch, the right auxiliary fuel pump, and starting the heater. The cabin heater uses gasoline from the fuel injector on the right engine. If the right fuel selector is in the off position the heater is inoperative. In case of right engine failure the heater can be operated by leaving the fuel selector on and closing the mixture control while operating the auxiliary fuel pump. Before the heater is operated under these conditions, determine that no fuel leaks are present between the tank and engine. Located in the heater is a heat limit switch, which acts as a safety device to render the heater system inoperative if a malfunction should occur causing excessively high temperatures. This control is located in the downstream end of the vent jacket, with the’reset button on the heater shroud. It is reached only through the nose section to insure that the malfunction causingthe overheat condition is corrected prior to future heater operation. To prevent activation of the overheat control during landing roll or while taxiing, turn the three position switch to FAN during the approach for landing. During ground operation the manual switch should be on FAN for severalminutes to cool the heater before turning to the OFF position. Ventilating air for the cabin interior is obtained from the two ventilators located at each side of the instrument panel. The rear seat area is furnished with air by overhead vents from the dorsal fin scoop. Each individual vent is adjustable for the desired airflow. Located in the aft section of the cabin is an exhaust vent to improve the circulation of air in the cabin interior. PAGE 12 AIRPLANE AND SYSTEMS ISSUED: June 1,197O TWIN COMANCHE C/R Heating And Ventilating System AIRPLANE AND SYSTEMS ISSUED: June 1,197O PAGE 13 TWIN COMANCHE C/R SEATS The front and rear reclining seats are adjustable fore and aft to provide comfort for pilot and passengers. Seat backs may be tipped forward to facilitate ease of entry and exit from the aircraft. They are easily removed by taking out the stops at the end of the mounting tracks and sliding the seats from their tracks. The optional family seat(s) may be removed to allow more baggage area and access to the rear fuselage panel. Release the snap fasteners that attach the seat backs to the hat shelf and turn the wing fasteners at the back of the seat cushions. Vertically adjustable pilots’ seats are offered as optional equipment. FINISH All sheet aluminum components are carefully fmished inside and outside to assure maximum service life. Both sides of all pieces are alodine treated. External surfaces are coated with durable acrylic lacquer in attractive high gloss colors. The application of primer to interior surfaces prevents corrosion of structural and non-structural parts on the inside of the airplane. BAGGAGE AREA Maximum weight in the baggage area, including baggage, passengers and family seats is 250 pounds, with up to 20 cubic feet of available space. Baggage may be placed in the aircraft through a 19 x 2 1 inch door or through the passenger entrance. Tie-down straps are available for securing baggage when the family seats are not installed. The baggage door may also be used as an emergency exit. It is opened by holding the inside door knob up while turning the latch clockwise. STALL WARNING SYSTEM An approaching stall is indicated by both a stall warning light and horn, activated by a lift detector installed on the left wing outboard of the engine nacelle. This warning horn is separate and operated by a different system from the gear warning horn previously mentioned under landing gear. PAGE 14 AIRPLANE AND SYSTEMS ISSUED: June 1,197O -- I’ I TWIN COMANCHE C/R I PROPELLER SYNCHROPHASER* I The propeller synchrophaser eliminates manual propeller adjustments to reduce “Beat” effect of the propellers. The synchrophaser automatically maintains selected RPM after the pilot manually sets the desired RPM and selects the “Prop Sync.” position on the propeller synchrophaser switch. The synchrophaser will maintain the selected RPM until the pilots readjust RPM’s. See Operating Instructions for synchrophaser operation. I I I *Optional Equipment AIRPLANE AND SYSTEMS ISSUED: June 1,197O PAGE 15 WEIGHTAND BALANCE EmptyWeight ................................................... 2 BasicWeight .................................................... Weight and Balance - Visual Plotter ....................................... i Summary of Disposable Load .......................................... 6 Equipment List .................................................. 9 A. Propellers and Propeller Accessories .................................. 9 B. Engine and Engine Accessories - Fuel and Oil Systems ....................... 11 C. Landing Gear ............................................... 15 D. Electrical Equipment .......................................... 17 E. Interior Equipment ............................................................................