basler bt-67 systems training manual

BASLER BT-67 PILOT GROUND SCHOOL TRAINING

MANUAL

ii

THIS MANUAL IS FOR TRAINING PURPOSES ONLY

Pilot Ground School Training Manual This manual is for training and reference use only, it is not a part of the FAA approved Pilot’s Operating Handbook, nor is it a Basler Turbo Conversions Approved Manual. For ease of viewing many diagrams of aircraft gauges and radio control heads are drawn with the instrument colours inverted, so that they are now white faced with black labelling, rather than the actual black facing with white labelling. However all coloured instrument markings remain in the actual colours. Every effort has been made to ensure the completeness and accuracy of information contained within this training manual. However, should any conflicts arise between this manual and Basler manuals, only approved publications (AFM, pilot operating handbooks, maintenance manuals, etc) will be the final authority. All training manual data, procedures, and techniques are superseded by current official data.

THIS MANUAL IS FOR TRAINING PURPOSES ONLY

iii

TABLE OF CONTENTS

Basler BT-67

General Airplane Section 1 Annunciator & Warning Systems Section 2 Engines Section 3 Propeller Systems Section 4 Fuel System Section 5 Electrical System Section 6 Avionics Section 7 Autoflight Section 8 Pitot-Static System Section 9 Hydraulic Systems Section 10 Landing Gear System Section 11 Environmental Systems Section 12 Ice & Rain Protection Section 13 Performance Section 14 Spectrem AEM Survey System Section 15 Spectrem AEM Performance Section 16

© 2002 African Flying Boat Company - All Rights Reserved

Issued 2002

© African Flying Boat Company 1-1 Aug-2002

GENERAL AIRPLANE

Table of Contents Basler BT-67 Description..................... 1-1 Basler BT-67 Conversion Features ...... 1-2 Specifications ...................................... 1-5 PT6A-67R Engine Limitations .............. 1-6 Operational Limitations ........................ 1-7 Cabin Entry and Exits .......................... 1-9 Flight Deck ........................................ 1-10 Cockpit Lighting................................. 1-11

Cabin Lighting ....................................1-11 Exterior Lighting .................................1-11 BT-67 Instruments..............................1-12 Instrument Panels ..............................1-14 Circuit Breakers..................................1-16 Parking and Securing .........................1-19 Towing ...............................................1-20 Pre-flight Inspection............................1-21

Basler BT-67 Description The Basler BT-67 is a non-pressurised, twin-engine, turboprop airplane (Figure 1-1) designed and equipped for day or night flight in IFR conditions. The BT-67 combines the proven ruggedness and dependability of the Douglas DC3 series of aircraft, with the added reliability of modernised systems and the Pratt and Whitney PT6A-67 turboprop engine.

Figure 1-1 Basler BT-67

The Basler BT-67 STC does not cover any modifications to the certificated DC3 interiors. Basic configurations, dimensions, weights and specifications are summarised in Table 1-1. Refer to the current airplane AFM for detailed, up-to-date information.


Basler BT-67 Conversion Features

Figure 1-3 Basler BT-67 Conversion Features POWER PLANT: • Pratt & Whitney Canada PT6A-67R PROPELLER: • Hartzell 5 Blade Metal Propellers. AIRFRAME: • Fully engineered, designed and strengthened airframe after total inspection and

rejuvenation of normal wear items. Includes a structural reinforced conversion package. • Fuselage stretched 40 inches (forward of wing) and cockpit bulkhead moved forward 60

inches resulting in 35% more cabin volume and optimal centre of gravity parameters. • Redesigned outer wing leading edge and BT-67 wing tip. • All metal control surfaces. • Centre and outer wings reinforced to reduce loads on lower wing attach angles, and to

support the increased maximum gross weight. • Complete new instrument panel and control pedestal. SYSTEMS: • Electrical system replaced by complete new system designed To FAR Part 25. • Hydraulic system upgraded. Improved gear retraction time. • Fuel system engineered to FAR Part 25, including new filler caps. • Brakes, B.F. Goodrich H2-445 expander tube brake assembly installed. • Anti-ice and deice equipment - all new including hot windshield • Low infra-red signature due to an augmented exhaust system which disperses over the

upper wing surface.


FLIGHT DECK: • Flight deck is virtually 100% new, right down to the control wheel. • Flight Management and Systems Management is efficient with low crew workload due to

state-of-the-art avionics & systems design that eliminate many of the tedious tracking and routine chores that burden pilots of other lesser equipped aircraft.

• The visibility is excellent and all instruments & controls are easy to see and reach. • The cockpit environment is very comfortable. Improvements compared to the basic DC-3 in

noise, vibration, heating, ventilation and lighting are dramatic. Possibly the most distinctive feature of the Basler BT-67 are the new nacelles and cowlings. The PT6A engine is considerably longer and slimmer than the original radial engines, and has a very different airflow pattern, necessitating a redesign of the original DC3 cowlings. To increase the military effectiveness of the BT-67, augmented exhaust ducting has been included within the engine cowlings. Rather than having the exhausts discharging directly into the airstream, the exhaust pipes are enclosed in shrouds within the nacelle. Considerable cooling of the exhaust gases takes place prior to venting the exhaust over the wing. Tests have shown the infrared signature of the exhaust gases to be effectively eliminated.

Figure 1-2 Nacelles and Exhaust Shrouds

With the change in centre of gravity position and reduction in aircraft weight with the replacement of the original radial engines with the PT6A-67R turboprop engines, the Basler BT-67 fuselage has been extended with the insertion of a 40” fuselage plug forward of the wing leading edge. For added safety the plane of rotation of the propellers is now behind the cockpit bulkhead. Reinforcements have been made to the centre and outer wings to reduce loads on lower wing attach angles, and to support the increased maximum gross weight. These reinforcements consist of internal straps, stringer angles and bars on the upper and lower wing surfaces, and externally mounted plates on the lower wing surface.

Figure 1-3 Wing Strengthening Plates Ailerons, rudder and elevators are cable operated by conventional dual control wheels. Optional Basler metalised control surfaces may be installed in place of the standard DC3 fabric or ceconite covered control surfaces. External control locks should be installed to prevent potential wind damage to control surfaces whenever the aircraft is parked.


Figure 1-4 Basler BT-67 Three View

Figure 1-5 Basler BT-67 Turning Circle


Specifications Table 1-1 Table 1-2 Basler BT-67 Douglas DC3C Minimum Crew .......................................................... 2 ..................................................... 2 Engines......................................... 2x P&W PT6A-67R ............................... 2x P&W R1830 Maximum Horsepower................................. 1281 SHP ..........................................1200 HP Propellers................... 2x 5-blade Hartzell HC-B5MA-3 .........2x 3-blade Hamilton Standard Fuselage Length............................................ 67 ft 9 ½”........................................... 64 ft 5½” Wingspan ......................................................... 95 ft 8” ................................................ 95 ft Wing Area ......................................................989 sq ft ..........................................985 sq ft Fuel Types ............................................................. Jet ............................................. Avgas Maximum Certificated Weights Maximum Takeoff Weight ............................ 28,750 lbs ....................................... 26,900 lbs Maximum Landing Weight ........................... 28,750 lbs ....................................... 26,900 lbs Maximum Zero Fuel Weight......................... 26,200 lbs ....................................... 26,200 lbs Basic Operating Weight ............................... 15,750 lbs ....................................... 17,815 lbs Cabin Dimensions Cabin Width............................................................ 7 ft .................................................. 7 ft Cabin Length.................................................... 42 ft 2” .......................................... 33 ft 10” Cabin Height ...................................................... 6 ft 6” .............................................. 6 ft 6” Cabin Volume..............................................1,225 cu ft ..........................................905 cu ft Specific Loadings Wing Loading (lbs per sq foot) ............................ 29.07 ...............................................27.31 Power Loading (lbs per sHP) .............................. 11.22 ...............................................11.21 Fuel Capacities

Tank Usable Capacity. Total Capacity Main - Left 189.7 USg 718 ltr 1,271 lbs 195.0 USg 738 ltr 1,306 lbs Main - Right 189.7 USg 718 ltr 1,271 lbs 195.0 USg 738 ltr 1,306 lbs Aux – Inboard L 186.6 USg 706 ltr 1,250 lbs 189.0 USg 715 ltr 1,266 lbs Aux – Inboard R 186.6 USg 706 ltr 1,250 lbs 189.0 USg 715 ltr 1,266 lbs

Total 752.6 USg 2,849 ltr 5,042 lbs 768.0 USg 2,907 ltr 5,146 lbs Aux – Outboard L 384.0 USg 1,454 ltr 2,573 lbs 385.0 USg 1,457 ltr 2,580 lbs Aux – Outboard R 384.0 USg 1,454 ltr 2,573 lbs 385.0 USg 1,457 ltr 2,580 lbs Total (ALL) 1,520.6 USg 5,756 ltr 10,188 lbs 1,538.0 USg 5,822 ltr 10,332 lbs Fuel and Oil Specifications Fuel..................................................................................... Jet-A, -A1, -A2, -B, JP-4, -5, -8 Hydraulic Fluid...................................................................................................MIL-H-5606 Landing Gear Brakes.................................................................Goodrich H-2-445 Expander Tube Brakes Main wheels tyre size .......................................................................45 x 17:00 x 16, 10-ply Tail wheel tyre size ................................................................................ 22 x 9:00 x 6, 8-ply Operating Speeds KCAS KIAS VMO Maximum Operating Speed.................174................170 VA Maximum Manoeuvring .......................111................107 VFE¼ Maximum Flap Extension ¼................135................133 VFE ½ Maximum Flap Extension ½..................99..................97 VFE ¾ Maximum Flap Extension ¾..................97..................95 VLE Maximum Landing Gear Extended ......144................141 VMCA Minimum Control - Air............................73..................67 VMCG Minimum Control - Ground ....................64..................56


PT6A-67R Engine Limitations

OPERATING CONDITIONS FOR BASLER BT-67 INSTALLATION

POWER SETTING

(2) sHP

TORQUE % (1)

ITT ºC

NG % (7)

NP RPM (7)

OIL PRESS psi (3)

OIL TEMP ºC (6)

Takeoff 1281 32.8ºC 100 825 104 1700 90 to 135 10 to 110

Maximum Continuous

1220 48.3ºC 95.0 840 104 1700 90 to 135 10 to 105

Minimum Idle 700 72 60 Min -40 to 110

Starting 1000

(4) 200 Max -40 Min

Transient

129 (5)

870 (5) 104 1870

(4) 40 to 200 -40 to 110

Maximum Reverse 900 765 1650 90 to 135 10 to 99

1. Torque limit applies within a range of 1000 - 1700 NP RPM, below 1000 NP RPM,

torque is limited to 50.5%. 2. Engine Inlet condition limits for engine operation: Altitude: Sea level to 25,000 ft. 3. Normal Oil Pressure is 90 - 135 psi at NG speeds above 72%. With engine torque

below 75%, minimum oil pressure is 85 psi at normal oil temperature (60º - 70ºC). Oil pressures under 90 psi are undesirable. Under emergency conditions, to complete a flight, a lower oil pressure of 60 psi is permissible at reduced power levels not exceeding 50% torque. Oil pressures below 60 psi are unsafe, and require an engine be shut down, or landing be made as soon as possible using minimum power required to sustain flight.

4. Time limited to 5 seconds. 5. Time limited to 20 seconds. 6. Oil Temperature limits are -40º to 110ºC with limited operating periods of 10 minutes at

105º - 110ºC. 7. 100% NG corresponds to 37,468 RPM. 100% N2 corresponds to 29,894 RPM.

Figure 1-6 PT6A-67R ITT Temperature Limitations


Figure 1-7 PT6A-67R Torque Limitations

Operational Limitations GROUND LIMITATIONS Stabilised engine ground operation is prohibited between the following propeller RPM’s: • 400 to 900 NP RPM • 1050 to 1200 NP RPM • or above 1600 NP RPM unless headed into wind. TAKEOFF LIMITATIONS Check engine power set prior to reaching 60 knots For downwind takeoffs, the aircraft is to be rolling prior to exceeding 1600 NP RPM. MINIMUM FLIGHT CREW A minimum flight crew of pilot and co-pilot is required for all operations. TYPE OF AIRPLANE OPERATION This airplane is certified for day, night and instrument conditions when required equipment is installed and operative. Flight into known icing conditions is prohibited. TAKEOFF AND LANDING Takeoffs and landings are limited to pressure altitudes between -1,000ft and +10,000ft, and temperatures between -40ºC minimum to ISA +35ºC maximum.


MAXIMUM OPERATING ALTITUDE AND ENROUTE TEMPERATURE LIMIT The maximum operating pressure altitude is 25,000ft. Enroute temperature limits are between -40ºC minimum to ISA +35ºC maximum. ENGINE ICE PROTECTION Engine inlet and propeller icing protection must be on whenever the outside air temperature is at or below +5ºC and visible moisture is present. CROSSWINDS The maximum demonstrated crosswind at 50ft-height is 24 knots. This is not limiting. REVERSE THRUST LIMITATIONS Reverse thrust is limited to ground operations only. Prior to lowering the tail wheel to the ground, caution should be exercised as directional control may be reduced during reverse thrust operations. Positioning of the power levers below the flight idle stop while the airplane is in flight is prohibited. Such positioning may lead to loss of airplane control, or may result in an engine overspeed condition with subsequent loss of engine power. FUEL LOADING LIMITATIONS Fuel must be loaded in the following manner: Fill left and right main tanks first. Additional fuel may then be loaded in either the inboard auxiliary tanks, or if installed, the optional outboard auxiliary tanks located in the outer wing. Fuel loaded in the outer wing auxiliary tanks must be loaded symmetrically. STARTER LIMITATIONS Use of the starter is limited to cycles of 30-seconds ON. A period of one minute OFF must be follow each start cycle, with a maximum of three 30-second ON start cycles allowed consecutively. After the third start cycle has been attempted, a minimum of 30-minutes OFF must occur before a fourth start cycle is attempted. GENERATOR LIMITATIONS The maximum continuous generator load on the ground is 250 amps per generator. In flight the maximum load is 300 amps per generator. ALTERNATE STATIC SYSTEM Use of the alternate static system during takeoff is prohibited. FUEL ANTI-ICING ADDITIVES Anti-icing additive is required. Lack of anti-icing additive may cause fuel filter icing and subsequent engine flameout. ENGINE AIRSTART ENVELOPE For airborne engine restarting, use of the starter is optional only with: • propeller unfeathered, and • airspeed greater then 160 knots, and • aircraft below 15,000ft Minimum recommended NG at FUEL ON is 10 - 12%. In an emergency, regardless of whether the propeller is FEATHERED or not, if 8 - 10% NG is observed, a WINDMILLING AIR START may be attempted. Observe ITT carefully, as hot starts can occur.


Cabin Entry and Exits

Figure 1-8 Basler BT-67 Doors and Exits The standard Douglas C47 doors consist of a forward hinged passenger door, and a rearward hinged cargo door. The Basler BT-67 also has an optional upper cargo door to allow the loading of standard LD3 containers (Figure 1-9). The cargo door is the first door closed, and the last opened. It is held in a locked position by locking pins manually actuated by handles at the top and bottom of the door. The forward passenger door closes against the rear cargo door, and is opened and closed with a standard door handle. A second handle actuates locking pins at the top and bottom of the door. Always ensure all locking pins are engaged before flight. A micro-switch connected to each door will activate a DOORS UNSAFE annunciator in the cockpit in the event that either door is not closed and the aircraft’s battery is ON. The passenger door incorporates a jettison mechanism for emergency use. A guarded jettison handle is mounted on the door frame, forward of the door. Pulling this handle down will pull both passenger door hinge release pins free, allowing the passenger door to be removed.

Figure 1-9 Basler BT-67 Cargo Doors

The Basler BT-67 has either two or three overwing emergency exit hatches in the cabin, depending on the configuration of the original DC3 airframe. These hatches are opened by rotating either the interior or exterior handles, and swinging up the hatch. These exits are non jettisonable.


The cockpit contains four emergency exits; two sliding pilots’ windows, a roof hatch, and a cockpit escape hatch. The roof hatch is held in position by lugs at the rear, and two latches at the front. Releasing the two latches allows the hatch to be jettisoned. The cockpit escape hatch is located on the left side of the cockpit, just behind the pilot’s seat. The hatch is opened and closed with a standard door handle. A second handle actuates locking pins at the top and bottom of the door. As there is no annunciator warning that the escape hatch is not secured, always ensure the locking pins are engaged before flight.

Flight Deck The efficient, comfortable flight deck is arranged for convenient use by two-pilot crews (Figure 1-10). Pilot and co-pilot sit side-by-side in individual seats, separated by the control pedestal. Both seats are adjustable forward and aft, as well as vertically. The inboard armrests may be swung horizontally to provide easy access to the seats. Seat belts and inertial shoulder harnesses are provided for each seat. Conventional dual controls allow the airplane to be flown by either pilot. A non-adjustable fold-away observer’s seat is provided in the cockpit aisle.

Figure 1-10 Typical Basler BT-67 Cockpit

Most aircraft circuit breakers are located on the circuit breaker panels located on the upper left and right sides of the cockpit. The breakers are arranged logically with regards to the differing electrical buses. The only circuit breakers not located there are the main power relay and bus reset circuit breakers, which are mounted next to the junction box at the cockpit door. The pilot’s overhead panel contains all anti-icing and de-icing controls, as well as lighting and engine starting controls. Propeller system control and warning system tests are mounted on the overhead panel. The engine fire protection system panel and annunciator panel are mounted in the centre of the overhead panel. All fuel system controls and gauges are laid out in schematic format on the fuel control panel. The co-pilot’s overhead panel contains the heating system and electrical system controls. The instrument panel contains the flight instruments, engine instruments, and the avionics panel. Engine instruments are mounted in a double column next to the avionics panel. Power controls are mounted on the control pedestal, and are arranged in the conventional sequence. Also included on the pedestal are the trimmers, windshield wiper controls and brakes and tail wheel lock lever.


Cockpit Lighting Variable backlighting is provided for the cockpit instruments and for the cockpit overhead and circuit breaker panels. Each lighting group is controlled by its own rheostat switch placarded BRT – OFF, mounted on either the pilot’s overhead panel or the co-pilot’s overhead panel (Figure 1-11).

Figure 1-11 Cockpit Lighting Controls

Both pilot and co-pilot have a moveable map reading lamp mounted above their respective seats. Separate switches are provided for these lights on each pilot’s overhead panel. Internal lighting for the compass is provided by a placarded switch on the pilot’s overhead panel. A push button switch placarded COMPANIONWAY LIGHT is mounted into the cockpit roof above the flight observer’s seat. This switch controls the cockpit illumination lamp mounted in the roof. As this cockpit roof light is hot-wired to the aircraft battery, through a circuit breaker, care should be made to ensure that the light is not inadvertently left on while the aircraft is left parked.

Cabin Lighting Lamps mounted along the centre of the cabin roof provide lighting in the cabin. The lamps are controlled by two switches wired in parallel, one by the main cargo door, and one on the pilot’s overhead panel (Figure 1-12). Changing the switch setting of either light switch will either turn on or off the cabin lights. There is no variable intensity selection available for the cabin lights.

Figure 1-12 Cabin and Exterior Lighting Controls

As the cabin lights are hot-wired to the aircraft battery, through a circuit breaker, care should be made to ensure that the lights are not inadvertently left on while the aircraft is left parked.

Exterior Lighting Switches for the anti-collision rotating beacons, position navigation lights, wing inspection ice lights and landing lights are located on the pilot’s overhead panel. They are appropriately placarded.


BT-67 Instruments

Figure 1-13 BT-67 Airspeed Indicator

Figure 1-14 Engine Gauges

Figure 1-15 Fuel Quantity Gauges

Figure 1-16 Fuel Totaliser


Figure 1-17 Hydraulic System Gauges

Figure 1-18 Wing Flap Position Gauge

Figure 1-19 Electrical System Gauges

Figure 1-20 Propeller De-icing Gauges

Figure 1-21 De-ice System Pneumatic Pressure Gauge


Instrument Panels

Figure 1-22 Pilot’s Overhead Panel

Figure 1-23 Overhead Panel


Figure 1-24 Fire Protection Panel

Figure 1-25 Annunciator Lights Panel

Figure 1-26 Fuel Control Panel


Figure 1-27 Co-pilot’s Overhead Panel

BT-67 Circuit Breakers

Figure 1-28 Left Distribution Bus Circuit Breaker Panel


Figure 1-29 Right Distribution Bus Circuit Breaker Panel

Figure 1-30 Emergency Bus Circuit Breaker Panel

Figure 1-31 Essential Bus Circuit Breaker Panel


Figure 1-32 Typical Avionics Circuit Breaker Panel

Figure 1-33 Bus Circuit Breaker Panels


Parking and Securing When the brakes have received severe use in landing, permit them to cool before setting the park brake. Park brakes can only be set from the pilot’s pedals. With the pilot’s pedals depressed, pull out the park brake handle (Figure 1-34). When set the handle will remain out. The hydraulic pressure gauge must indicate at least 300 psi when the aircraft is parked to ensure sufficient pressure is applied to the brakes. To release, depress the pilot’s brake pedals.

Figure 1-34 Park Brake Handle

The tail wheel must be in the centre position before the tail wheel lock lever will engage the tail wheel in the lock position. In all cases, lock the tail wheel when the aircraft is parked, moored, during any jacking of the aircraft, before a flight, and during landing. The tail wheel will be unlocked only while taxiing on the ground and during ground handling which involves turns. Turning without first unlocking the tail wheel will result in shearing the locking pin or damaging the wheel assembly. The tail wheel is locked in the trail position by pushing to the right and forward the lever-type control located on the lower section of the control pedestal (Figure 1-35). To unlock the tail wheel, pull the control aft to the limit of travel and to the left.

Figure 1-35 Tail-wheel Lock Lever

Do not operate flaps while the aileron locks are installed. Aileron gust locks must be installed only on the designated aileron. Installation of the aileron locks on the incorrect aileron will result in structural damage to the wing if flaps are subsequently lowered. When the aircraft is parked outside a hangar, install the external control locks on the rudder, ailerons and elevators. These locks are to be installed whenever the aircraft is left unattended, regardless of weather conditions. Pitot covers, engine plugs and propeller tie downs should also be installed whenever the aircraft is left unattended.


Towing To prevent retraction of the landing gear, install the ground safety pins at all times when the aircraft is parked, and especially prior to any maintenance of the aircraft. The aircraft can be either towed from the front, by attaching tow lines to the main wheel axles as shown in Figures 1-36 and 1-37, or from the tail by attaching a tow bar to the rear axle. The tow bar hooks over the towing lugs which protrude from both ends of the axle. When towing the aircraft, there should always be a crewmember in the pilot’s seat who is qualified to operate the brakes and tail wheel lock, and one crewmember should be at each wing tip and at the tail to assist in the towing operations. When towing from either the front or the rear, make certain that the tail wheel is in the UNLOCK position. Do not attempt to move the aircraft by pushing or pulling on the control surfaces, stabilisers, tabs or flaps, but only by applying force to the wheel axles or to other approved locations.

Figure 1-36 Forward Towing Diagram

Precautions during towing: • Towing from the rear of the aircraft should be employed whenever possible. • Do not attempt to tow while the parking brakes are engaged. • Avoid towing the aircraft over rough ground, whenever possible. • In starting, when towing from the front, the acute angle between the tail wheel and the line

of towing force is not to exceed 60º.


Figure 1-37 Towing Attachments • When towing from the tail, the acute angle between the tail wheel and the centreline of the

aircraft is not to exceed 60º at any time; and at starting, this angle is not to exceed 45º. • Use caution when pushing the aircraft forward by means of the tail wheel tow bar. • Do not tow from the tail up any grade that is more than 11% on hard ground. Do not tow

from the tail on soft, boggy, or rough ground. • Tow the aircraft from the front when it is loaded or when it is moved over soft ground. • Governed by local terrain conditions, tow at low speed, do not exceed 8kph. Avoid any

sudden stops.

Pre-flight Inspection BEFORE ENTERING AIRCRAFT 1. Remove left and right aileron control locks. COCKPIT 2. Verify all circuit breakers are properly set. 3. Select battery switch to the ON position. Confirm voltage is sufficient for the start if a

battery start is required, or above 20 volts for a ground power start. 4. If using a ground power cart, select ground power switch to ON. When using a power

cart, the main battery switch shall remain selected to the ON position. This will provide protection during the start sequence if the power cart fails or drops offline.

5. Push STALL WARN test button to check the stall warning system 6. Push the AUDIBLE WARN button to check the audible warning system. 7. Select the SMOKE DET switch to FWD and then AFT to check the cockpit and cabin

smoke detection systems. 8. Select annunciator light test switch to TEST and confirm all annunciator lights and the fire

T-handles are lit. Select to Dim or BRT as appropriate 9. Operate the fire protection test switch to FIRE TEST. Note that all four squib lights and

the fire T-handles illuminate, and the fire aural warning sounds. Push fire warning mute button to verify function.

10. Operate the test switch to FAULT TEST and check to see that the fault annunciators are on.

11. Confirm that both DISCH FIREX switches are in the centre OFF position. Confirm that both fire T-handles are IN, the FUEL SHUT OFF annunciators are not lit, and the FUEL LO PRESS annunciators are illuminated. Pull both T-handles OUT and confirm that both FUEL SHUT OFF annunciators are now lit. Select both standby fuel pumps to AUTO and confirm that both STBY PUMP ON annunciators are lit, and that the fuel pressure indications remain at ZERO. Push both fire T-handles IN and confirm that both fuel pressures rise into the green range, the FUEL LO PRESS annunciators are extinguished, and the FUEL SHUT OFF annunciators are both not lit. Return the standby fuel pumps to OFF.

12. If fuel has been added the fuel totaliser shall be reset. 13. Cycle the elevator trim wheel and return to the neutral position. 14. Lower wing flaps to the full down position. 15. If required, return battery and ground power switches to OFF. LEFT WING 16. Check left fuel and oil quantity and security of caps.


17. Inspect left wing flap turnbuckle rods for condition. Check flap sections for general condition and security.

18. Check left aileron for general condition and attachment. Check for full travel and freedom of movement. Check fixed trim tab, if installed, for security. Check static wicks for condition.

19. Check condition of left wing tip and nav. light assembly. 20. Inspect landing light lens and retaining wires for condition. Broken landing light lens

retaining wires are a grounding item. 21. Inspect left wheel well for general condition. Inspect left engine fire bottle pressure gauge

readings. Check left engine fuel filter for water. 22. Inspect left landing gear struts for proper inflation (2” of strut extension minimum). Inspect

left wheel and tyre assembly for general condition and wear. 23. Remove left main landing gear safety pin. 24. Inspect left engine cowling for general condition, oil leaks and proper security. Check inlet

to oil cooler for obstructions. 25. Check propeller assembly for apparent blade damage and condition of anti-ice boots.

Remove propeller retainer strap and engine inlet cover. 26. Check main and auxiliary fuel tank sumps for water. NOSE 27. Check condition of any antennae on lower centre section and forward fuselage. 28. Check battery compartment area for evidence of leaking battery acid. 29. Check pitot tubes for general condition. RIGHT WING 30. Conduct inspection items 16 - 26 in reverse order as appropriate to the right side of the

aircraft. RIGHT REAR FUSELAGE 31. Check right side of aft fuselage for general condition. 32. Check tail wheel assembly for general condition, tyre wear, and strut inflation. 33. Check rotating beacon for damage. 34. Check right horizontal stabiliser for general condition. 35. Remove right and left elevator control locks. Check right and left elevators for full travel

and freedom of movement. Check for general condition and static wicks. Check adjustable and fixed trim tabs for security.

36. Remove rudder lock. Check rudder travel and trim tab. Check rudder stop cables by checking the “spring back” at full left and right travel. Check for general condition and static wicks.

37. Check tail cone and nav. light for general condition. LEFT REAR FUSELAGE 38. Check left side of aft fuselage for general condition. 39. Check cabin/ cargo doors for general condition.


ANNUNCIATON & WARNING SYSTEMS

Table of Contents Annunciator and Warning Systems ...... 2-1 Annunciator Lamp Test........................ 2-2 Annunciator Panel ............................... 2-3 Fire Detection System ......................... 2-4 Fire Detection System Tests ................ 2-5 De-icing & Ignition Annunciators .......... 2-6 Propeller Annunciators ........................ 2-7

Fuel System Annunciators................... 2-8 Electrical System Annunciators ........... 2-8 GPS Annunciators............................... 2-9 Landing Gear Position Annunciators.... 2-9 Stall Warning......................................2-10 Overspeed Warning ...........................2-10 Gear Warnings...................................2-10

Annunciator and Warning Systems The annunciator and warning systems (Figure 2-1) consists of the following components: • Annunciator Panel • Fire Protection Control Panel • Engine Firewall T-Handles • De-icing system annunciators • Propeller annunciators • Ignition system annunciators • Fuel system annunciators • Electrical system annunciators • GPS annunciators • Landing gear position annunciators • Stall warning • Overspeed warning • Gear warning The annunciator panel on the centre overhead panel provides general warning of system faults and malfunctions. Control for the engine fire protection system is provided on the overhead Fire Protection control panel. Indications of engine fire, and arming of the engine fire extinguisher bottles are controlled by two firewall T-handles located on the main instrument panel. De-icing system annunciators are located on the pilot’s overhead panel. Propeller annunciators are located on the overhead panel. Fuel system annunciators are located on the fuel panel, and electrical annunciators on the co-pilot’s panel. GPS and landing gear position annunciators are on the main instrument panel.


Figure 2-1 Cockpit Annunciators

An aerodynamic angle of attack vane set to trigger an aural warning is mounted on the fuselage, underneath the pilot’s window. A test button is mounted on the overhead panel. An airspeed overspeed detection unit providing an aural warning is mounted into the pitot system, with a test button on the overhead panel. Audible warnings are provided to detect when the gear is not locked down, and the power levers are retarded to idle, or the wing flaps are selected to a landing position.

Annunciator Lamp Test The annunciator lamp test switch is labelled WARN LTS, and is mounted on the overhead panel (Figure 2-2). This three-position switch also selects the light. The intensity of the lights on the annunciator panel can also be set to DIM for night flight.

Figure 2-2 Annunciator Lamp Test Switch


Annunciator Panel The annunciator panel is mounted in the centre of the overhead panel (Figure 2-3). It contains warning (red), cautionary (amber) and advisory (white) annunciator lights. A press-to-test switch is mounted on the overhead panel, to the left of the annunciator panel. In addition to colour coding the annunciators incorporate labelling to facilitate interpretation. If covered by the annunciator system, an aircraft system fault generates a signal which illuminates the appropriate warning light.

Figure 2-3 Annunciator Panel

LABEL COLOUR CAUSE FOR ILLUMINATION

L GEN FAIL Red Left generator offline

PILOT PITOT HEAT Amber Lower pitot heat turned off, or heating element failed

CO-PILOT PITOT HEAT Amber Upper pitot heat turned off, or heating element failed

R GEN FAIL Red Right generator offline

L OIL LO PRESS Red Left engine oil pressure below 60 psi

L HEAT DUCT O’HT Amber Left wheel well duct temperature over 232ºC

R HEAT DUCT O’HT Amber Right wheel well duct temperature over 232ºC

R OIL LO PRESS Red Right engine oil pressure below 60 psi

L FUEL LO PRESS Red Fuel pressure below 5.5 psi after L firewall shutoff

L FUEL FLTR CLOGGING Amber Left fuel filter bypass valve open

R FUEL FLTR CLOGGING Amber Right fuel filter bypass valve open

R FUEL LO PRESS Red Fuel pressure below 5.5 psi after R firewall shutoff

L ENG CHIP DET Amber Metallic contamination in left engine oil is detected

L MAIN OVER PRESS Amber Fuel transfer stopped by 1.5 psi L tank overpressure

R MAIN OVER PRESS Amber Fuel transfer stopped by 1.5 psi R tank overpressure

R ENG CHIP DET Amber Metallic contamination in right engine oil is detected

L INLET LIP DE-ICE Amber Left inlet lip de-ice temp controller detecting fail

(Blank)

(Blank)

R INLET LIP DE-ICE Amber Right inlet lip de-ice temp controller detecting fail

L IGN ON White Left engine ignitor powered

INV FAIL Amber The selected inverter is inoperative

DOORS UNSAFE Amber Cabin door is open or not secure

R IGN ON White Right engine ignitor powered


Fire Detection System The left and right engine fire detection systems are completely independent. The detection system monitors the resistance of wire elements, or loops, in their respective engine compartment area (Figure 2-4). A rise in compartment temperature will lower the resistance to a predetermined trip point, and activate both visual and aural warnings.

Figure 2-4 Engine Fire Loop

The visual warnings consist of a red ENG FIRE – PULL firewall T-handle for each engine located on the main instrument panel. Once illuminated the firewall T-handle light will remain illuminated until the temperature in the associated engine compartment cools below the sensing loop trip point. The aural warning consists of a speaker-amplifier located in the right overhead panel. The aural and visual warnings activate simultaneously and the aural warning may be manually cancelled by pressing the FIRE WARNING SILENCE button on the fire protection control panel. Each engine has two elements located in the accessory section, and one element in the forward power section. These three elements are connected to form one continuous loop and are set to trip at the appropriate elevated temperature. The system will still operate effectively with a break in the loop, provided there is no short circuit to ground at the break point. A circuit monitor continually senses the loop to determine faults caused by a breakdown or a short circuit in the system. An amber FAULT annunciator light, located in the overhead fire protection control panel will illuminate whenever a fault is detected.


Fire Detection System Tests The fire protection system test is conducted using a three-position toggle switch located in the overhead Fire Protection control panel (Figure 2-5). The switch is spring loaded to the centre OFF position. Selecting the FIRE TEST position will cause the engine firewall T-handles to illuminate, sound the fire warning horn, and illuminate the SQUIB TEST lights.

Figure 2-5 Fire Protection Panel

PANEL LABEL COLOUR CAUSE FOR ILLUMINATION

SQUIB TEST FIREX 1 Green Electrical continuity in firex squib 1 during test

SQUIB TEST FIREX 2 Green Electrical continuity in firex squib 2 during test

LOW PRESS 1 Amber Firex bottle 1 discharged, or pressure too low

LOW PRESS 2 Amber Firex bottle 2 discharged, or pressure too low

LOW PRESS FAULT Amber Breakdown or short circuit in fire detection loop

Selecting the FAULT TEST position causes the fire detection FAULT lights to illuminate. The fire bottle LOW PRESS. light may be tested by selecting the WARNING LIGHT TEST switch located on the overhead Control Panel to the TEST position.


De-icing System And Ignition System Annunciators The de-icing system and ignition system annunciators are located on the pilot’s overhead panel (Figure 2-6).

Figure 2-6 Pilot’s Overhead Panel

PANEL LABEL COLOUR CAUSE FOR ILLUMINATION

WSHLD HEAT W/S Blue Power applied to the left windshield heater

SURFACE DE-ICE WING Blue Wing surface de-ice switch on

INLET LIP LH L/H Blue Left inlet lip de-ice boot powered

INLET LIP RH R/H Blue Right inlet lip de-ice boot powered

INERTIAL SEP LH L/H Blue Left inertial separator is in icing position

INERTIAL SEP RH R/H Blue Right inertial separator is in icing position

STARTER ENGAGED

PUSH TO ABORT Amber Left engine starter powered

STARTER ENGAGED

PUSH TO ABORT Amber Right engine starter powered


Propeller Annunciators The propeller annunciators are located on the overhead panel (Figure 2-7).

Figure 2-7 Propeller Annunciators


PROP SYNC ON Blue Prop synch button pressed in

ARMED Amber Power applied to autofeather system

READY Green Propeller autofeather system activated

BETA Amber Propeller beta switch contact made

REVERSE Green Propeller reverse switch contact made


Fuel System Annunciators The fuel system annunciators are located on the fuel panel (Figure 2-8).



FUEL SHUT OFF Red Firewall shutoff valve closed

STBY PUMP ON Amber Standby fuel pump operating at at least 7.5 psi

XFEED VALVE OPEN Amber Crossfeed valve open

XFER PUMP ON Amber Transfer fuel pump operating at at least 1.5 psi

XFER XFEED OPEN Amber Transfer crossfeed valve open

Electrical System Annunciators The electrical system annunciators are located on the co-pilot’s overhead panel (Figure 2-9).

Figure 2-9 Electrical Annunciators


BUS TIE CLOSED Amber Manual closing of Bus Tie RCCB

MASTER OFF Red Power to electrical system, but battery relay open

GROUND POWER Green Ground power relay closed


GPS Annunciators The GPS annunciators are located on the main instrument panel (Figure 2-10).

Figure 2-10 GPS Annunciators


WPT Green GPS Destination Waypoint message active

APR Amber GPS Approach mode active

PTK Blue GPS Parallel Track offset engaged

WRN Red GPS Warning message active

NAV Green HSI information from Nav #1

GPS/LRN Blue HSI GPS mode engaged and power to GPS relay

Landing Gear Position Annunciators The landing gear position annunciators are located on the main instrument panel (Figure 2-11).

Figure 2-11 Landing Gear Annunciators


WARNING HORN SILENCE Amber Gear warning horn sounding or cancelled

GEAR UP Red Landing gear not fully down, or handle not neutral

GEAR DOWN Green Landing gear down micro made, and handle neutral


Stall Warning System The stall warning system consists of a an angle of attack vane located below the pilot’s side window (Figure 2-12), and a computer located in the main electrical junction box. The vane is electrically heated, and is controlled by the pilot’s pitot heat switch, but is protected by a circuit breaker on the ESSENTIAL bus. The angle of attack vane is moved by the aircraft’s airflow to a position relative to the aircraft’s angle of attack. When the vane reaches a position denoting a high angle of attack condition the stall warning tone is generated through the warning horn. The stall warning horn can be tested by the pilots by pressing the STALL WARN test switch on the left overhead panel.

Figure 2-12 Stall Vane

Overspeed Warning The overspeed warning system functions through of an airspeed switch, connected to the pilot’s pitot system, and mounted in the nose forward of the main instrument panel. The overspeed switch functions by allowing pitot pressure to impact onto a diaphragm, making an electrical contact at a setting of 170 knots +6 –0, and causing the overspeed warning tones to be generated through the warning horn. The overspeed warning horn can be tested by the pilots by pressing the AUDIBLE WARN test switch on the left overhead panel.

Gear Warnings The landing gear warning horn will sound if either of these two conditions exist: • Power levers are fully retarded and the main landing gear is not down. • Wing flaps are lowered beyond 3/8 travel and the main landing gear is not down. Micro-switches on each power lever will activate the landing gear warning horn if either power lever is reduced to IDLE and the main gear is not down and locked. The warning horn may be silenced by pressing the push button located on the overhead Fire Protection control panel. Another switch will activate the warning horn if the wing flaps are extended beyond 3/8 of total travel and the main gear is not down and locked. This switch is mounted on the wing flap push/pull rod. The warning horn cannot be silenced if this switch causes it to sound.


ENGINES

Table of Contents PT6A-67R Powerplant ......................... 3-1 Engine Data ........................................ 3-2 Free-turbine Reverse-flow Operation ... 3-2 Engine Description and Operation ....... 3-3 Compressor Bleed Valve ..................... 3-5 Igniters ................................................ 3-5 Exhaust ............................................... 3-5 Accessory Section............................... 3-6 Lubrication System.............................. 3-6 Magnetic Chip Detector ....................... 3-8

Engine Fuel System ............................ 3-8 Fuel Manifold Purge System................ 3-9 FCU Operation.................................... 3-9 Fuel Additives ....................................3-10 Control Pedestal.................................3-10 Engine Power Control.........................3-11 Condition Levers ................................3-11 Engine Gauges ..................................3-11 Engine Limitations..............................3-12 Starter Operating Limits......................3-12

PT6A-67R Powerplant The Pratt and Whitney PT6A-67R engines (Figure 3-1 and 3-2) are equipped with composite five-blade, full-feathering, reversing, constant-speed propellers mounted on the output shaft of the engine reduction gearbox. Engine oil supply and single-action, engine-driven governors control propeller pitch and speed. When the engines are shut down the propellers automatically feather, and will unfeather when engines are started, as engine oil is pumped into the propeller dome.

Figure 3-1 PT6A-67R Engine, Three-quarter View, Right Side


Figure 3-2 PT6A-67R Engine, Internal Arrangement

The engine utilises two independent turbine sections: one driving the compressor in the gas generator section, and the second (two-stage power turbine) driving the propeller shaft through a reduction gearbox. The engine is self-sufficient since its gas generator driven oil system provides lubrication for appropriate areas of the engine, pressure for the torque meter and power for the propeller pitch control.

Engine Data

PT6A-67R ENGINE SPECIFICATIONS AND LEADING PARTICULARS

OPERATING CONDITION ESHP SHP

JET THRUST

lb

NP RPM

SPECIFIC FUEL CONSUMPTION

lb/ESHP/hr

Takeoff 1358 1281 (1) 192 1700 0.533

Max. Cont 1294 1220 (2) 187 1700 0.542

Max. Climb 1083 1020 (3) 157 1425 0.575

Max. Cruise 1083 1020 (3) 157 1425 0.575

(1) Available to 32.8ºC (2) Available to 48.3ºC (3) Available to 35.0ºC

Type of Combustion Chamber..................................................................Annular Compression Ratio........................................................................................ 12:1 Propeller Shaft Rotation ....................................................................... Clockwise 100% NG ................................................................................................... 37,468 100% NF ................................................................................................... 29,894 Engine Diameter (at room temperature) (approx.)............................19” (483 mm) Engine Length (at room temperature) (approx.) ............................. 74” (1880 mm) Dry Weight .....................................................................................515lb (234 kg) Max. Oil Consumption per 10 Hour........................................0.3lb/hr (0.136kg/hr)

Free-turbine Reverse-flow Operation The free-turbine design of the PT6A engine refers to the fact that the turbine sections rotate freely, having no physical connection between them (Figure 3-3). The compressor turbine drives the engine compressor and engine accessories. Dual power turbines drive the power section and propeller through the planetary reduction gearbox. The compressor and power turbines are mounted on separate shafts and are driven in opposite directions by the gas flow across them.


Figure 3-3 Free-turbine Reverse-flow Turboprop Engine

Reverse flow refers to the direction of airflow through the engine. Inlet air enters the compressor at the aft end of the engine, moves forward through the compressor section and the turbines, and is exhausted at the front of the engine.

Engine Description and Operation Inlet air enters the engine through an annular plenum chamber (Figure 3-4), formed by the compressor inlet case, where it is directed to the compressor. The compressor consists of four axial stages combined with a single centrifugal stage and assembled as an integral unit. The engine is also equipped with a water wash ring at the compressor air inlet case. A row of stator vanes, located between each stage of the compressor, diffuse the air, raise static air pressure and direct the air to the next stage of compression. The compressed air passes through diffuser tubes which turn the air through 90º in direction, and converts the velocity to static pressure. The diffused air then passes through straightening vanes to the annulus surrounding the combustion chamber liner assembly. The combustion chamber liner consists of two annular sections bolted together at the front, dome-shaped end. The outer wrapper incorporates an integral large exit duct. The liner assembly has perforations of various sizes that allow entry of compressor delivery air. The flow of air changes direction 180º as it enters and mixes with fuel. The fuel/air mixture is ignited and the resultant expanding gases are directed to the turbines. The location of the inner liner eliminates the need for a long shaft between the compressor and compressor turbine, reducing the overall length and weight of the engine. Fuel is injected into the combustion liner through 14 simplex nozzles arranged in two sets of seven for ease of starting. Each nozzle is supplied by a dual manifold consisting of primary and secondary transfer tubes and adapters. The fuel/air mixture is ignited by two spark igniters which protrude into the liner. The resultant gases expand from the liner, reverse direction in the exit duct zone and pass through the compressor turbine inlet guide vanes to the single-stage compressor turbine. The guide vanes ensure that the expanding gases strike the turbine blades at the correct angle, with minimum loss of energy. The still expanding gases are then directed forward to the power turbine section. The two-stage power turbine consists of the first stage inlet guide vane and turbine, and the second-stage inlet guide vane and turbine. The turbine drives the propeller shaft via a reduction gearbox. The compressor and power turbines are located in the approximate centre of the engine, with their respective shafts extending in opposite directions. This feature provides for simplified installation and maintenance inspection procedures. The exhaust gas from the power turbine is collected and ducted in the bifurcated exhaust duct assembly and directed to atmosphere via twin opposed exhaust stubs. Inter turbine temperature, T5, is monitored by an integral bus bar, probe and harness assembly installed between the compressor and power turbines, with the probes projecting into the gas path. A terminal block mounted on the gas generator case provides a connection point to the cockpit instrumentation.


Figure 3-4 PT6-67R Engine Cross Section


The engine oil supply is contained in an integral oil tank, which forms the rear section of the compressor inlet case. The tank has a capacity of 2.5 US gallons, 9.5 litres, and is provided with manual filler access and a direct oil level sight gauge. Fuel supplied to the engine from an external source is further pressurised by an engine-driven fuel pump, and its flow to the fuel manifold is controlled by the FCU. The power turbine drives a propeller through a two-stage planetary reduction gearbox, located at the front of the engine. The gearbox contains an integral torque meter device, which is instrumented to provide an accurate indication of engine power. A magnetic chip detector is installed at the bottom of the gearbox.

Compressor Bleed Valve At low N1 RPM, the compressor axial stages produce more compressed air than the centrifugal stage can use. A pneumatic piston compressor bleed valve compensates for excess air flow at low RPM by bleeding axial stage air, P2.5, to reduce back pressure on the axial stages. This pressure relief helps prevent axial stage compressor stall. At low N1 speeds the compressor bleed valve is open. As power is increased the valve begins to progressively close. Above approximately 90% N1 the bleed valve is closed. If the compressor bleed valve were to stick closed at low N1 speeds, compressor stall could result from an attempt to accelerate the engine to higher power. If the valve were to stick open at high N1 speeds, power output would be considerably reduced. With the valve open, at a given N1 RPM, ITT will increase slightly and torque will decrease.

Igniters Two spark-type igniters in the combustion chamber provide ignition during the engine start sequence. Although the engine is equipped with two igniters, only one is needed for the engine start. The system is designed such that should one igniter malfunction, the other ignitor will continue to operate. The igniters are activated during the start sequence by the IGNITION switch being set to START, and the starter being engaged (Figure 3-5).

Figure 3-5 Ignition Controls

The system consists of an airframe-mounted ignition exciter, two individual high-tension cable assemblies, and two spark igniters. It is energised from the aircraft’s 28-volt supply, and will operate in the 9-volt to 30-volt range. The igniter system can produce up to 3,000 volts.

Exhaust Serving primarily to reduce the Basler BT-67’s infrared exhaust signature, the PT6A’s exhaust elbows are enclosed within the streamlined engine cowling, rather than exposed to the airflow. Within the cowlings, the exhaust augmenter tube and heat shield assemblies overlap, but are not attached to, the exhaust elbows (Figure 3-6). This assembly removes hot gases from inside the cowling area, then directs and disperses them above and behind the cowling. The augmenter tube also mixes outside ambient air with the exhaust gases, and with venturi action, aids in pulling hot gases out of the engine.


Figure 3-6 PT6A Engine Separated From Exhaust Ducting

Accessory Section All engine-driven accessories, with the exception of the propeller governor, overspeed governor and NF tach-generator, are mounted on the accessory gearbox at the rear of the engine. These components are driven by the compressor by means of a coupling shaft which extends the drive through a tube at the centre of the oil tank. One lubricating oil pressure pump and two scavenge oil pumps are mounted inside the accessory gearbox. Two additional scavenge pumps are externally mounted. The starter/generator, high pressure fuel pump and FCU, N1 tach generator and hydraulic pump are mounted on pads on the rear of the accessory drive case. Each mounting pad has its own specific gear ratio. The rear location of accessories provides for a clean engine and simplifies maintenance procedures.

Lubrication System The PT6A engine lubrication system functions primarily to cool and lubricate engine bearings and bushings (Figure 3-7). It also provides oil to the propeller governor and propeller reversing control system. The main oil tank houses a gear-type engine-driven pressure pump, an oil pressure regulator, a cold pressure relief valve and an oil filter. The engine oil tank, an integral part of the compressor inlet case, is located in front of the accessory gearbox. As oil is pumped from the tank, it passes through pressure and temperature sensing bulbs mounted on the rear accessory case. At N1 RPM above 72% normal oil pressure is between 90 and 135 psi. Oil is then delivered through an external oil transfer tube to the nose case. Gear-type scavenge pumps return the oil through external oil transfer tubes and through an external oil cooler below the engine. The oil cooler is thermostatically controlled to maintain the desired oil temperature. When scavenge oil temperature reaches 71ºC, a thermostatically controlled bypass valve opens to route oil through the cooler. The externally mounted oil-to-fuel heat exchanger uses hot engine oil to heat the fuel before it enters the engine fuel system. The total oil system capacity is 3.9 US gallons, 14.75 litres, including the 2.5 US gallon, 9.5 litre, oil tank. Maximum acceptable oil consumption is one quart every 10 hours; however normal oil consumption may be as little as one quart per 50 hours. Most PT6A engines seek an oil level of one to two quarts low, and filling above this level may result in oil being vented overboard. Although the pre-flight checklist calls for checking the oil level, the best time to check oil quantity is shortly after engine shutdown, since oil levels are most accurately indicated at this time. Oil level checks during pre-flight may require motoring the engine to obtain an accurate level indication. The oil tank is provided with a filler neck and integral quantity dipstick housing. The cap and dipstick are secured to the filler neck, which passes through the gearbox housing and accessory diaphragm into the tank. Dipstick markings indicate the number of US quarts of oil less than full.


Figure 3-7 PT6-67R Lubrication System


Magnetic Chip Detector A magnetic chip detector is installed at the bottom of each engine nose gearbox to indicate the presence of ferrous particles in the lubrication system. The detector activates amber ENG CHIP DET annunciators to alert the pilots to possible oil contamination. Illumination of the CHIP DET annunciator is not in itself cause for an engine to be shut down. Engine parameters should be monitored for abnormal indications.

Engine Fuel System The engine fuel system for the PT6A-67R consists of the following basic components; oil-to-fuel heat exchanger, fuel pump, fuel control unit, flow divider and purge valve, and dual fuel manifold with 14 simplex nozzles (Figure 3-8). The system also includes the fuel topping governor, covered separately in Section 4 Propeller Systems.

Figure 3-8 Simplified Engine Fuel System Diagram

The oil-to-fuel heat exchanger regulates fuel temperature at the fuel pump inlet to prevent icing at the pump filter. An engine oil line within the heat exchanger is located next to the fuel line. Heat transfer occurs through conduction between these two lines, before fuel is delivered to the FCU. The heat exchanger melts ice particles and prevents the fuel from thickening in extremely cold temperatures. The heat exchangers operate automatically whenever the engines are running. After fuel passes through the oil-to-fuel heat exchanger, it flows into the engine-driven fuel pump. The pump is a gear-type pump, located on the accessory gearbox. Its primary purpose is to supply sufficient pressure to the fuel nozzles to ensure an adequate spray pattern during all engine operations. The flow rates and pressures will vary with changes in N1 RPM. Fuel enters the pump chamber through a spring-loaded inlet screen. Should this screen become blocked, the increase in differential pressure overcomes the spring, and allows unfiltered fuel to flow into the pump chamber. The pump gears increase fuel pressure, and deliver it to the FCU via a filter in the pump outlet. A bypass valve in the pump body enables unfiltered high-pressure fuel to flow to the FCU in the event of the outlet filter becoming blocked. The FCU meters proper fuel amounts for all modes of engine operation. Flow rates are calibrated for starting, acceleration and maximum power. The FCU compares N1 with power lever setting, and regulates fuel to the engine fuel nozzles. The FCU also senses compressor


discharge pressure, P3, and compares it to N1 to establish acceleration or deceleration fuel flow limits. A pre-set minimum flow orifice guarantees sufficient fuel flow at all operating altitudes to sustain engine operation at minimum power. Fuel cut-off is accomplished by the pump unloading valve, internal to the FCU. The valve is controlled by the condition lever, and is either open or closed; it has no intermediate position. When the valve is open, fuel flows to the minimum pressurising valve, which blocks fuel flow during start until fuel pressure is sufficient for a proper spray pattern in the combustion chamber. As fuel pump output increases the minimum pressurising valve opens delivering fuel to the flow divider. The flow divider schedules the metered fuel from the FCU to the primary and secondary manifolds as a function of primary manifold pressure. During engine start, metered fuel is delivered initially to the primary manifold only. As the engine accelerates through approximately 36% N1, fuel pressure increases sufficiently to open the valve to the secondary manifold. All fourteen nozzles then deliver atomised fuel to the combustion chamber. The progressive sequence of primary and secondary fuel nozzle operation provides for cool starts. Increased acceleration in N1 may be noted when the secondary nozzles are activated.

Fuel Manifold Purge System The fuel manifold purge system (Figure 3-9) is designed to eliminate residual fuel which remains in the flow divider and fuel manifolds during engine shutdown. The system consists of an accumulator purge tank connected to the P3 air line, and a P3 discharge line from the accumulator to the flow divider and purge valve. During normal engine operation, P3 air constantly pressurises the purge tank. As long as the engine is running, fuel pressure keeps the flow divider purge valve closed. As the fuel pressure drops to zero during engine shutdown, P3 air escapes into the flow divider, pushing the residual fuel into the combustion chamber where it is burnt. As a result, the pilots may notice a one or two second delay in initial engine shutdown after the fuel condition levers are moved to STOP.

Figure 3-9 Fuel Purge System

FCU Operation The FCU is mounted on the rear of the fuel pump. A coupling between the fuel pump and the FCU transmits a speed signal to the governing section of the FCU. The FCU determines the amount of fuel scheduled to the combustion chamber by controlling N1 speed. Compressor discharge pressure, P3, is compared to N1 to establish acceleration fuel flow limits. This fuel limiting function prevents overtemperature conditions during engine start and acceleration. The FCU receives input from the condition lever, the power lever, the N1 flyweight governor and a pneumatic bellows (Figure 3-10). A brief overview of FCU operation is set out below. The power lever positions a 3D cam in the FCU that, through a cam follower and lever, determines fuel flow corresponding to the selected N1 speed. The power lever selects speeds between idle and maximum. Engine speed is controlled by the NG governor, which contains two flyweights mounted on a ballhead driven by the engine. The flyweight governor controls the on-speed condition by positioning the 3D cam in response to speed variations of the gas generator. As N1 speed changes, resulting flyweight action changes the 3D cam setting, which changes fuel flow valve setting to maintain the selected N1 speed.


Figure 3-10 Simplified Fuel Control System

The cam follower and arm transmit 3D cam movement to the fuel metering valve. As the 3D cam moves upward the fuel flow to the engine is increased, and N1 speed increases. Similarly downward motion of the 3D cam will result in decreased N1. Therefore, the NG governor continuously adjusts N1 speed to response to variation in the gas generator speed. In an overspeed condition, increasing pressure by the governor flyweights moves the 3D cam downward, resulting in decreased fuel flow through the fuel metering valve. Balance occurs when N1 speed is reduced to the selected speed, and the cam is stationary at the new speed condition. In an underspeed condition, decreasing pressure on the governor flyweights moves the 3D cam upward, resulting in increased fuel flow through the fuel metering valve until the system reaches equilibrium. Compressor discharge air pressure, P3 air, also affects the fuel metering valve position during acceleration or deceleration. Increase in P3 causes the fuel metering valve to increase fuel flow in response to increased P3 pressure until N1 speed is stabilised. A decrease in P3 pressure causes the fuel metering valve to decrease fuel flow until N1 speed is stabilised at the lower selected value. In the event of power turbine overspeed, a decrease in P3 air pressure at the fuel metering valve allows the FCU to reduce fuel flow to the gas generator.

Fuel Additives Anti-icing additive is necessary for the BT-67. Lack of anti-icing additive may cause fuel filter icing and subsequent engine flameout. It is important to add the correct amounts of additive. Higher concentrations do not insure lower fuel freezing temperatures, and too great a concentration can damage the fuel system. The minimum additive concentration is 0.06% by volume, and the maximum concentration 0.15% by volume.

Control Pedestal The control pedestal extends between the pilot and co-pilot (Figure 3-11). The three sets of powerplant control levers, from left to right, are power levers, propeller RPM and feather levers, and fuel condition levers.


Engine Power Control

Engine power is controlled by power levers which set N1 speed, and propeller levers which adjust propeller speed. Power levers control engine power from idle to takeoff power by operation of the NG governor in the FCU. The levers control power in three regions: flight, beta and reverse. The bottom of the flight range is idle. When the power levers are moved past the idle detent and pulled back into the beta range, they control propeller blade angle only. The beta range is normally used for taxi. When the levers are moved further aft into the reverse range, they control propeller blade angle and engine power to provide reverse thrust.

Figure 3-11 Control Pedestal The propeller levers are conventional in setting the

required RPM for flight. The normal governing range is 1000 to 1700 RPM. Although the BT-67 is equipped with an automatic propeller feathering system, the propeller can be manually feathered by pulling the propeller lever back past the detent into the red and yellow striped section of the quadrant. To unfeather, advance the lever back into the governing range. Regardless of propeller lever position, the propellers will move back towards feather when oil pressure is lost as the engines are shut down.

Condition Levers The condition levers have two positions, RUN and STOP. In the STOP position all fuel flow to the engine is cut off. No intermediate setting of the condition levers is possible.

Engine Gauges

Figure 3-12 Engine Gauges


Engine Limitations Airplane and engine limits are described in the Limitations section of the AFM. These limitations have been approved by the FAA, and must be observed when operating the Basler BT-67. The following engine operating limits chart provides important limitations for all operating conditions.

OPERATING CONDITIONS FOR BASLER BT-67 INSTALLATION

POWER SETTING sHP

TORQUE % (1)

ITT ºC

NG % (7)

NP RPM (7)

OIL PRESS psi (3)

OIL TEMP ºC (6)

Takeoff 1281 32.8ºC 100 825 104 1700 90 to 135 10 to 110

Maximum Continuous

1220 48.3ºC 95.0 840 104 1700 90 to 135 10 to 105

Minimum Idle 700 72 60 Min -40 to 110

Starting 1000 (4) 200 Max -40 Min

Transient 129 (5)

870 (5) 104 1870

(4) 40 to 200 -40 to 110

Maximum Reverse 900 765 1650 90 to 135 10 to 99

1. Torque limit applies within a range of 1000 - 1700 NP RPM, below 1000 NP RPM,

torque is limited to 50.5%. 2. Engine Inlet condition limits for engine operation: Altitude: Sea level to 25,000 ft. 3. Normal Oil Pressure is 90 - 135 psi at NG speeds above 72%. With engine torque

below 75%, minimum oil pressure is 85 psi at normal oil temperature (60º - 70ºC). Oil pressures under 90 psi are undesirable. Under emergency conditions, to complete a flight, a lower oil pressure of 60 psi is permissible at reduced power levels not exceeding 50% torque. Oil pressures below 60 psi are unsafe, and require an engine be shut down, or landing be made as soon as possible using minimum power required to sustain flight.

4. Time limited to 5 seconds. 5. Time limited to 20 seconds. 6. Oil Temperature limits are -40º to 110ºC with limited operating periods of 10 minutes at

105º - 110ºC. 7. 100% NG corresponds to 37,468 RPM. 100% N2 corresponds to 29,894 RPM.

During engine start, temperature is the most critical limit. The ITT starting limit of 1000ºC is limited to 5 seconds. During any start, if the indicator needle approaches this limit, the start should be aborted before the needle approaches the red triangle. For this reason it is helpful during starts to keep the condition lever out of the RUN detent so that the lever can be quickly pulled back to STOP.

Starter Operating Limits Engine starters are time-limited during the starting cycle to prevent the possibility of starter damage due to overheating. Use of the starter is limited to cycles of 30 seconds ON, followed by a period of one minute OFF, with a maximum of three 30 second ON start cycles allowed consecutively. After the third start cycle has been attempted, a minimum of 30 minutes OFF must occur before a fourth start cycle may be attempted.


PROPELLER SYSTEM

Table of Contents Propeller System ................................. 4-1 Blade Angle......................................... 4-2 Primary Governor ................................ 4-2 Low Pitch Stop .................................... 4-3 Beta and Reverse Control.................... 4-3 Beta and Reverse Control Operation ... 4-4 Overspeed Governor ........................... 4-5

Fuel Topping Governor........................ 4-6 Power Levers ...................................... 4-6 Propeller Control Levers...................... 4-7 Propeller Feathering............................ 4-7 Autofeather System............................. 4-7 Propeller Synchrophaser ..................... 4-9

Propeller System Each engine is equipped with a composite, five-bladed, counter-weighted, full-feathering, variable-pitch, constant speed, reversing Hartzell HC-B5MA-3/M11276 115 inch propeller mounted on the output shaft of the reduction gearbox (Figure 4-1). Since the engines are free turbines, with no mechanical connection between compressor and power turbines, the propeller can rotate freely on the power shaft when the engine is shut down. Propeller bungees and engine intake plugs are provided to prevent windmilling at zero oil pressure when the airplane is parked.

Figure 4-1 Hartzell Propeller Propeller pitch and speed are controlled by engine oil pressure supplied to the propeller dome through engine-driven propeller governors. A governor oil pump boosts oil pressure delivered by the engine oil system to a pressure high enough to control movement of the propeller blades. When oil pressure is present in the propeller dome, propeller pitch (blade angle) is controlled normally by the propeller governor or by the beta valve, depending upon the


propeller’s mode of operation. As oil pressure increases, the propeller moves toward low pitch (high RPM). Loss of oil pressure will cause centrifugal counterweights and feathering springs to move propeller blades toward high pitch (low RPM) and, eventually, into the feathered position. As oil pressure decreases during engine shutdown, the propeller automatically moves toward feather. The minimum low pitch propeller position is determined by a mechanically-actuated hydraulic stop, referred to as the primary low pitch stop. The power levers control beta and reverse blade angles by adjusting the low pitch stop position in beta and reverse ranges. Two governors (a primary governor and an overspeed governor) control propeller RPM. The primary governor controls the propeller through its normal governing range. The propeller control lever selects propeller RPM by adjusting the primary governor condition. Should the primary governor malfunction, the overspeed governor prevents propeller speed from exceeding 1802 RPM. The fuel topping governor acts as a backup governor limiting propeller speed to 106% of that selected by the propeller lever. In the reverse range, the fuel topping governor is reset, limiting propeller RPM to approximately 95% of the primary governor setting. The fuel topping governor limits propeller RPM by reducing fuel flow to the engine. The propeller RPM is displayed in the cockpit on a gauge that receives its input from the crankcase.

Blade Angle Blade angle is the angle between the chord of the propeller and the propeller’s plane of rotation. Because of the normal twist of the propeller, blade angle is different near the hub than it is near the tip. Blade angle for the BT-67 is measured at the chord, 42 inches from the propeller’s hub. This position is referred to as the “42 inch station”. All blade angles specified in this section are approximate values (Figure 4-2).

Figure 4-2 Blade Angles

Primary Governor

The primary governor modulates oil pressure in the propeller dome to change blade angle to maintain a constant propeller speed. As oil pressure in the dome changes, propeller blade angles change to maintain the propeller speed the pilot has selected. The primary governor can maintain any selected propeller speed from approximately 1000 RPM to 1700 RPM.

Figure 4-3A Over Speed Figure 4-3B Under Speed For example, suppose an airplane is in normal cruising flight with the propellers set at 1450 RPM. If the pilot begins a descent without changing power, the airspeed will increase. This decreases the angle of attack of the propeller blades, causing less drag on the blades, thus causing the RPM to increase. The governor will sense this overspeed condition and increase blade angle to a higher pitch. The higher pitch increases the blade’s angle of attack, slowing it back to an onspeed 1450 RPM. Likewise if an airplane changes from cruise to climb attitude without a power change, the propeller RPM tends to decrease. The governor responds to this underspeed condition by decreasing blade angle to a lower pitch, and the RPM returns to its original value. Thus the governor is able to keep the variable pitch propeller at a constant speed.


Figure 4-3C On Speed Power changes, as well as airspeed changes, cause the propeller to momentarily experience overspeed and underspeed conditions, but again the governor reacts to maintain the onspeed condition. Due to the smooth action of the governor, the pilot will notice few, if any, of these minor adjustments. There are times, however, when the primary governor is incapable of maintaining the selected RPM. For example, imagine an airplane approaching to land with its propellers set at 1700 RPM. As power and airspeed are both reduced, underspeed conditions exist which cause the governor to decrease blade angle as it attempts to restore the onspeed condition. If blade angle were allowed to decrease to its full reverse limit, aircraft control would be dramatically reduced as

the propeller blades moved into a high drag discing position. To prevent this undesirable situation, a device is provided to stop the governor from selecting blade angles that are too low for safety. As blade angle is decreased by the governor, eventually the low pitch stop is reached. Blade angle then becomes fixed, preventing its continued movement toward a lower pitch. At the low pitch stop, the governor is prevented from restoring the onspeed condition, and propeller RPM decreases below the selected governor RPM setting. Once the low pitch stop is reached, blade angle cannot decrease further until the pilot selects beta or reverse.

Low Pitch Stop With a non-reversing propeller the low pitch stop is simply at the low pitch limit of travel, determined by the propeller’s construction. But with a reversing propeller, extreme travel in the low pitch direction is past 0º , into reverse or negative blade angles. Consequently, the BT-67’s propeller system has been designed to allow the low pitch stop to be repositioned when reversing is desired. The low pitch stop is created by mechanical linkage which senses blade angle. At the low pitch stop the linkage causes a valve to close, stopping the flow of oil into the propeller dome. Since more oil causes low pitch and reversing, blocking off oil flow creates a low pitch stop. The low pitch stop valve, commonly referred to as the “beta” valve, is spring-loaded to provide redundancy in the event of mechanical loss of beta valve control. Low pitch stop operation is determined by a mechanically monitored hydraulic stop. The propeller servo piston is connected by four spring-loaded sliding rods to the slip ring mounted behind the propeller. A carbon brush block riding on the slip ring transfers the movement of the slip ring through the propeller reversing lever to the beta valve of the governor. The initial forward motion of the beta valve blocks off the flow of oil to the propeller. Further motion forward dumps the oil from the propeller into the reduction gearbox sump. A mechanical stop limits the forward motion of the beta valve. Rearward motion of the beta valve does not affect normal propeller control. When the propeller is rotating at a speed lower than that selected on the governor, the governor pump provides oil pressure to the servo piston and decreases the pitch of the propeller blades until the feedback of motion from the slip ring pulls the beta valve into a position blocking the supply of oil to the propeller, thus preventing further pitch changes.

Beta and Reverse Control The position of the low pitch stop is controlled from the cockpit by the power lever. Whenever the power lever is at IDLE or above, this stop is set at 14.5º . Bringing the power lever aft of idle progressively repositions the low pitch stop to lower blade angles. The geometry of the power lever linkage through the cam box is such that power lever movement from idle to full forward thrust has no effect on the beta valve’s position. When the


the power lever is moved from idle to the reverse range, it repositions the beta valve to direct governor pressure to the propeller piston, decreasing blade angle through zero into a negative range. The travel of the propeller servo piston is fed back to the beta valve to null its position and, in effect, to provide infinite negative blade angles all the way to maximum reverse. The opposite will occur when the power lever is moved from full reverse to any forward position up to idle, thus providing the pilot with manual blade angle control for ground handling. The region of propeller travel between idle and ground fine is referred to as “beta for taxi”, or simply beta. In this range N1 remains at 72%. To enter the beta range, the pistol-trigger catch on the power lever must be lifted, and the lever brought aft past the idle stop. On entering the beta range the BETA annunciator on the overhead panel will illuminate (Figure 4-4). The aft stop of the beta range is called ground fine. With aft movement of the power lever, blade angle moves progressively from idle to ground fine.

Figure 4-4 Beta/Reverse Annunciators

The region between ground fine and maximum reverse is referred to as “beta plus power”, or simply the reverse range. In this range, N1 progressively increases to a maximum of 85% while the blade angle decreases. To enter the reverse range the power levers are moved aft of the beta range. Once the reverse range has been entered the REVERSE annunciator on the overhead panel will be illuminated. With further aft movement of the power lever, the blade angle will progressively decrease from ground fine to maximum reverse.

Beta and Reverse Control Operation

Figure 4-5 Beta Range and Reverse Diagram


When the propeller blade angle reaches approximately 18º the four flanges extending from the dome make contact with four beta nuts (Figure 4-5). As propeller pitch angle continues to decrease each flange on the propeller dome pushes each beta nut and attached rod forward. As the rod moves forward it pulls the feedback ring forward. In turn a beta valve inside the governor is pulled into the oil cut-off position. The linkage is set to cut off oil supply to the dome when blade angle reaches 14.5º . This provides the governor with a hydraulic low pitch stop of 14.5º for flight operations. If the low pitch stop were fixed at 14.5º the propeller could not enter the beta and reverse range; however, the low pitch stop can be reset to allow the aircraft to enter into the beta and reverse ranges while the aircraft is on the ground. When the power levers are lifted back over the idle detent into the beta range, they are pulling back on the top of the reverse lever. As the reverse lever moves back, the beta valve is pushed back, re-establishing oil flow to the propeller dome. This allows propeller blade angle to go below the low pitch stop. As the propeller blades go below the low pitch stop, the propeller dome and feedback ring continue forward, eventually pulling the beta valve back into the oil cut-off position.

Overspeed Governor Since the PT6’s propeller is driven by a free turbine (independent of the engine’s compressor), overspeed can rapidly occur if the primary governor fails. The overspeed governor provides protection against excessive propeller speed in the event of primary governor malfunction. The hydraulic overspeed governor (Figure 4-6) is located on the left side of the propeller reduction gearbox. The operating point of the overspeed governor is 106% of the primary governor’s maximum speed, or 1802 RPM. If a propeller’s speed reached 1802 RPM, the overspeed governor would begin increasing blade angle to a higher pitch to prevent the RPM from continuing to rise. From a pilot’s point of view, an NP gauge stabilised at 1802 RPM would indicate a failure of the primary governor and proper operation of the overspeed governor.

Figure 4-6 Overspeed Governor Operation

For pre-takeoff check purposes, the set point of the overspeed governor is rescheduled using the propeller overspeed governor test switch on the overhead panel. During testing, propeller speed should not exceed approximately 1583-1647 RPM.


Fuel Topping Governor The primary propeller governor contains a fuel topping governor which prevents power turbine overspeed if a propeller malfunctions. An overspeed could occur, for example, if a propeller blade were to stick in a fixed position during normal primary governor operation. In addition, during reverse thrust operation, the fuel topping governor is set below the speed selected by the primary governor to permit indirect control of propeller speed by the FCU servo system. The speed at which fuel topping governor operation occurs is determined by the speed selected with the propeller levers, and by the position of the reset lever. In the flight range the reset lever is set to regulate power turbine, N2, speed at approximately 106% of the propeller lever setting. In the ground range the lever is reset to 96% of the propeller lever position. If propeller speed exceeds levels sensed by the fuel topping governor, fuel flow to the N1 section will be reduced, and engine power will decrease. When this occurs, NP normally remains constant, but it may decrease if propeller blades are frozen in a fixed position.

Power Levers The power levers (Figure 4-7) are located on the power lever quadrant on the control pedestal. They are mechanically interconnected through a cam box to the FCU, reverse lever, beta valve and follow-up mechanism, and the propeller governor. The power lever quadrant permits movement of the power lever in the forward thrust (alpha) range from idle to maximum thrust, and in the beta/ reverse range from idle to maximum reverse. Mechanical stops in the power lever quadrant at the IDLE positions prevent inadvertent movement of the levers into the beta/ reverse range. To select beta or reverse, the pilot must lift up catches on the power levers to allow the power levers to be brought back past the stops.

Figure 4-7 Control Pedestal

In the forward thrust (alpha) range the power levers establish NG by selecting a gas generator governor speed which results in a fuel flow that will maintain the selected N1 RPM. In the beta range, the power levers control the beta valve to reduce propeller blade angle, thus reducing residual propeller thrust. N1 RPM is not affected in the beta range. In the reverse range, the power lever: • selects a blade angle proportionate to the aft travel of the lever • selects a fuel flow that will sustain the selected reverse power • resets the fuel topping governor NP from its normal 106% to 95% Therefore, RPM in reverse is a function of the primary propeller governor acting through the FCU to limit fuel flow and control propeller RPM in relation to power lever position.


Propeller Control Levers Propeller RPM, within the primary governor range of 1000 to 1700 RPM, is set by the position of the propeller control levers. The full forward position sets the primary governor at 1700 RPM. In the full aft position, forward of the feathering detent, the primary governor is set at 1000 RPM. Intermediate propeller RPM positions can be set by moving the propeller levers to select desired RPM as indicated on the NP gauges. A detent at the low RPM position prevents inadvertent movement of the propeller lever into feather. The feather position is indicated by red and yellow stripes at the bottom of the propeller lever slots in the power quadrant.

Propeller Feathering The propeller can be manually feathered by moving the propeller lever full aft, past the detent, into feather. This action locks the governor’s pilot valve in the full up position, opens the feather valve, and all oil quickly drains from the propeller pitch mechanism. As oil is dumped from propeller servo chambers, the counterweights and springs drive the propeller blades to the feather position. Since the propeller shaft and the N1 shaft are not connected, the propeller can be feathered with the engine running; however to avoid excessive torque loads on the propeller gearbox, the engine should be at idle power when propellers are manually feathered. If an engine fails with the Autofeather system inoperative, the propeller will maintain onspeed RPM unless it is feathered manually.

Autofeather System The autofeather system (Figures 4-8 to 4-10) provides a means of dumping oil from the propeller dome automatically in the event of an engine failure. This is accomplished by opening a solenoid valve in the propeller over-speed governor which relieves propeller governor oil pressure. Loss of governor oil pressure allows the feathering spring, within the propeller, to rotate the propeller blades to the feather position. The system is controlled by a three position toggle switch in the overhead panel. Dual indicator lights located on each side of the control switch will indicate ARMED when the control switch is moved to the ARMED or TEST position, and READY when the respective system is activated.

Figure 4-8 Autofeather System Armed Through Test


When the power levers are advanced beyond approximately 92% NG the autofeather system will activate, and the READY lights will illuminate after approximately a five second delay. The Autofeather system will be inoperative when either engines’ power lever is retarded below 92% NG.

Figure 4-9 Autofeather System Armed and Ready

Engine power is monitored by a torque switch on each engine. If the engine power should fall below approximately 25%, with the power lever still advanced, this sensor will close. Closing of the torque-sensing switch will activate the solenoid valve in the over-speed governor and feather the propeller. The torque switch will simultaneously deactivate the remaining autofeather system, and its respective READY light will go out. In order to unfeather the propeller, the autofeather control switch must be selected to OFF. This will allow the solenoid valve on the overspeed governor to close.

Figure 4-10 Autofeather System Armed, Left Engine Failure

The TEST position of the control switch is used to bypass the switches in the quadrant normally activated by the power levers. This allows ground testing of both systems at a reduced torque setting of approximately 40%. After the autofeather system has been tested,


the propeller manual feathering system should be checked. Be sure to verify that engine power is at idle, then bring both power levers into the feather detent. Propellers may be allowed to completely feather without engine damage; however ground operations while propellers are feathered should be kept to a minimum. The autofeather system is required to be armed and operable for flight during takeoff, approach and landing, and should be turned off in cruise.

Propeller Synchrophaser The synchrophaser system is an electronic system that automatically matches the speeds of the two propellers to synchronise RPM. It is not a designated master-slave system, but rather matches the RPM of the slower propeller to the RPM of the faster propeller. The synchrophaser has a limited range of authority, the maximum increase possible is approximately 25 RPM. In no case will the RPM fall below that selected by the propeller control lever. Normal governor operation is unchanged, but the synchrophaser continuously monitors propeller RPM and resets either governor as required. Propeller RPM is sensed by a magnetic pickup mounted adjacent to each propeller spinner bulkhead. The magnetic pickup transmits electrical pulses from each propeller to a single control box. The control box converts any pulse rate difference between the propellers into correct commands, and transmits the commands to coils mounted close to the flyweights of each primary governor. As coil voltages vary, the governor speed settings are biased until propeller RPM’s match exactly.

Figure 4-11 Propeller Synch Control

A synchrophaser is activated by a push-button toggle switch, mounted on the overhead panel (Figure 4-11). When the synchrophaser is off, propeller governors operate at the manual speed settings selected by the pilot. Propeller RPM can be reset when the synchrophaser is on. If the synchrophaser is on, and is not able to adjust the propeller RPM’s to match, the system has reached the end of its operating range. Increasing the RPM of the slower propeller, or decreasing the speed of the faster propeller, will bring the speeds within the synchrophaser range.


NOTES:


FUEL SYSTEM

Table of Contents Fuel System ........................................ 5-1 Fuel Tank System................................ 5-2 Fuel Capacities.................................... 5-2 Fuel System Operation ........................ 5-2 Fuel Transfer....................................... 5-3 Fuel Crossfeed .................................... 5-4 Firewall Shutoff Valves ........................ 5-4

Fuel Filters .......................................... 5-5 Fuel Control Panel............................... 5-5 Fuel Quantity Indication....................... 5-6 Fuel Drains.......................................... 5-6 Refuelling Procedures ......................... 5-6 Fuel Handling Procedures ................... 5-7 Fuel Anti-Icing Additive........................ 5-7

Fuel System The Basler BT-67 fuel system (Figure 5-1) is designed to be as simple and easy to use as possible in both normal and emergency situations. Optional auxiliary fuel tanks may be installed in each outer wing. Separate, overwing refuelling is incorporated for each fuel tank.

Figure 5-1 BT-67 Fuel System Diagram


Fuel Tank System The Basler BT-67 fuel tank system (Figure 5-2) consists of two integral fuel tanks in each wing centre section. The left and right main tanks are located forward, with the left and right inboard auxiliary tanks located aft. Fuel is drawn from a sump located at the inboard end of each tank.

Figure 5-2 BT-67 Fuel Tank Locations

Optional outboard auxiliary tanks may be installed, and are located in the left and right outer wings. This optional system consists of two tanks manifolded together in each outer wing. Inboard and outboard tanks are vented to the atmosphere through underwing vents.

Fuel Capacities

Tank Usable Capacity. Total Capacity Main - Left 189.7 USg 718 ltr 1,271 lbs 195.0 USg 738 ltr 1,306 lbs Main - Right 189.7 USg 718 ltr 1,271 lbs 195.0 USg 738 ltr 1,306 lbs Aux – Inboard L 186.6 USg 706 ltr 1,250 lbs 189.0 USg 715 ltr 1,266 lbs Aux – Inboard R 186.6 USg 706 ltr 1,250 lbs 189.0 USg 715 ltr 1,266 lbs Total 752.6 USg 2,849 ltr 5,042 lbs 768.0 USg 2,907 ltr 5,146 lbs Aux – Outboard L 384.0 USg 1,454 ltr 2,573 lbs 385.0 USg 1,457 ltr 2,580 lbs Aux – Outboard R 384.0 USg 1,454 ltr 2,573 lbs 385.0 USg 1,457 ltr 2,580 lbs Total (ALL) 1,520.6 USg 5,756 ltr 10,188 lbs 1,538.0 USg 5,822 ltr 10,332 lbs

Fuel System Operation Fuel is supplied to the engines from the two main tanks only. Fuel in the auxiliary tanks must be first transferred to the main tanks before it can be utilised. Fuel is supplied from each main tank to its respective engine by means of two submerged pumps in each tank. One is designated the main pump, and the other the standby pump, but are otherwise identical units. The main pump has a two-position selector switch (ON/OFF) and operates from the time engine start is initiated until the engine is shutdown.

Figure 5-3A Fuel System At Takeoff


Figure 5-3B Fuel System During Cruise Figure 5-3C Main Fuel Pump Failure The standby pump has a three-position switch (ON/OFF/AUTO). When selected to the ON position, the standby pump operates and functions identically to the main pump. The ON position is utilised during takeoff in case of main pump failure. When selected to the AUTO position, the standby pump operates automatically when fuel pressure falls below 7.5psi. A low pressure sensing switch inline with the main pump senses low pressure and switches on the standby pump. When this occurs, a pressure switch inline with activate the amber STBY PUMP ON annunciator (Figure 5-3C). This annunciator will be illuminated any time the standby pump is operating in excess of 7.5psi. The AUTO position is selected during all phases of flight, other than takeoff (Figures 5-3A and 5-3B).

Fuel Transfer The left and right inboard and outboard auxiliary fuel is transferred to the respective main tanks by means of two electric inline pumps, one on each side. Fuel transfer is initiated manually by the pilots by first selecting the auxiliary tank from which fuel is to be transferred, and then selecting the transfer pump switch to the ON position and holding until the amber XFER PUMP ON annunciator illuminates (Figures 5-3D and 5-3E). The annunciator verifies pump operation. Fuel will automatically stop transferring when any of the following conditions occur: 1. The selected auxiliary tank becomes empty.

The loss of pump pressure, sensed by the pressure switch, will automatically switch off the transfer pump and reset the selector switch to the OFF position.

2. The main tank is full. A fuel level sensor in the main tank will automatically switch off the transfer pump and reset the selector switch to the OFF position.

3. Excessive pressure in the main tank. An overpressure switch in the main tank will activate and switch off the transfer pump and reset the selector switch to the OFF position. When this occurs, an amber MAIN OVERPRESS annunciator illuminates on the annunciator panel. The remainder of the fuel transferred should be controlled manually.

The transfer crossfeed valve allows auxiliary fuel from either side to be transferred to the opposite main tank in the event of a transfer pump failure. The main tank crossfeed valve would then be opened to supply fuel to the engine on the side of the failed transfer pump. Although fuel may be transferred at the pilots’ discretion it is recommended that the main fuel tank level be maintained at above ¼ full. It is also desirable, when the optional outboard auxiliary tanks are installed, to transfer outboard fuel first. This will result in the least amount of unavailable fuel in the event of an auxiliary tank fuel transfer valve failure.


Figure 5-3D Fuel Transfer Figure 5-3E Fuel Cross Transfer

Fuel Crossfeed A main tank crossfeed valve is provided to allow fuel from a single main tank to be used by one or both engines. The valve is normally closed, and is opened only as required to balance fuel load or during single engine operation. When crossfeeding from a specific main tank, the main and standby fuel pumps for that tank should be selected to ON. Both main and standby pumps for the opposite main tank must be selected to the OFF position (Figure 5-3F). This assures fuel is supplied from the desired main tank only.

Figure 5-3F Fuel Crossfeed

Firewall Shutoff Valves An electrically driven firewall shutoff valve for each engine is located in its respective wheel well (Figure 5-4). They are actuated by pulling the respective red ENG FIRE – PULL firewall T-handle located on the main instrument panel. The red FUEL SHUT OFF annunciator will illuminate on the fuel control panel whenever the fuel shutoff valve is closed.

Figure 5-4 Firewall Shutoff Valve and Fuel Filter


A mechanical indicator is located on the underside of the valve to allow visual confirmation of the valve position during the pre-flight inspection.

Fuel Filters Fuel filters are installed in the left and right wheel wells (Figure 5-4). The filters are equipped with bypass valves which open in the event of filter restriction. When a bypass valve opens, an amber FUEL FLTR CLOGGING annunciator will illuminate on the overhead annunciator panel. A manual drain line is installed on the fuel bowl of the filter. When fuel pressure is applied to the system by the fuel pumps, manually pushing the drain line up will drain fuel from the filter. A purge valve purges air from the fuel lines and FCU during engine start. The purge valve operates automatically as part of the engine starting sequence.

Fuel Control Panel All fuel controls for the BT-67, with the exception of the firewall T-handles, are located on the fuel control panel above the pilots (Figure 5-5).


Two and three position switches control the settings of the six installed fuel pumps. The main and standby pump switches are manually latched into position. The transfer pumps however will only latch into the ON position when the 1.5 psi pressure switch has closed. This ensures that the transfer pumps will stop as the auxiliary tank empties. Amber annunciator lights indicate the operation of the standby and transfer pumps. Separate left and right switches control the action of the auxiliary fuel valves for each side. For aircraft with the optional outboard auxiliary tanks these switches are three position switches, controlling both inboard and outboard auxiliary fuel valves. Only one valve can be opened at any time. For aircraft with only the inboard auxiliary fuel tanks the switches are two position switches. Both the crossfeed and transfer crossfeed valves are controlled by separate push button annunciators. When either valve is open its amber annunciator will be illuminated. Red annunciator lights indicate whenever the firewall shutoff valves are closed.


Fuel Quantity Indication Fuel quantity is measured using a variable capacitance system. One sensor probe is mounted in each main and each inboard fuel tank, with two probes in each outboard auxiliary tank. Quantities are measured in pounds, using 6.7lbs per US gallon. When the fuel quantity indicator reads zero in level flight, any fuel remaining in that tank cannot be used safely. A digital fuel flow/ totaliser is installed on the main instrument panel. It is used to monitor each engine’s fuel flow and total fuel burned. When the correct total fuel level is entered into the totaliser, it is able to display the total fuel remaining in either pounds or time remaining. This system is not to be used for determining the total fuel on the aircraft. The primary means of determining fuel quantity is the fuel quantity gauges.

Figure 5-6 Fuel Totaliser

Fuel Drains During the pre-flight inspection, the fuel sumps on the tanks and filters should be drained to check for fuel contamination. There are four tank drains on the aircraft centre section, a fuel filter bowl drain in each wheel well, and a sump drain for each optional outboard auxiliary tank under each wing (if installed). Since jet fuels and water are of similar densities, water does not settle out of jet fuels as easily as from aviation gasolines. For this reason, the airplane must sit perfectly still, with no fuel being added, for approximately three hours prior to draining the sumps if water is to be removed. Although water ingestion is not as critical for turbine engines as it is for reciprocating engines, water should still be removed periodically to prevent the formation of fungus and contamination-induced inaccuracies in the fuel gauging system.

Refuelling Procedures Because of the hazards involved in handling fuel, accomplish all fuelling operations, including both fuelling and defuelling, outside.

Figure 5-7 BT-67 Fuel Cap Locations

When refuelling, service the main tanks first, auxiliary tanks second. When a fuel tank is to be filled, attach the free end of the nozzle bonding wire to the uninsulated metallic part of the aircraft to assure an electric bond between the tank and the delivery hose. Do not attach the bonding wire adjacent to the filler neck of the tank. Connect this bond before removing the filler cap of the tank to be filled, and do not disconnect until the cap has been replaced.


Fuel Handling Procedures All hydrocarbon fuels contain some dissolved, and some suspended water. The quantity of water contained in the fuel depends on its type and temperature. Jet fuels, with their higher aromatic content, tend to absorb more water than aviation gasolines. Along with water, jet fuels will suspend rust, lint and other foreign materials longer. Eventually suspended contaminants will settle to the bottom of the tank. The settling time for jet fuels is five times that of avgas; therefore jet fuels require good fuel handling practices to ensure servicing with clean fuel. If recommended ground procedures are carefully followed, solid contaminants will settle, and free water can be reduced to 30 parts per million (ppm), a value considered acceptable by the major airlines. Dissolved water has been found to be the major potential fuel contaminant. Its effects are multiplied in aircraft that operate primarily in humid regions and in warm climates. Since most suspended matter, including water, can be removed from the fuel by allowing sufficient settling time and by proper filtration, fuel contamination is usually not a major problem. Dissolved water cannot be filtered from the fuel by the micronic-type filters used in the fuel filters; however, water in the fuel can be released by lowering the fuel temperature, as happens in flight. These supercooled water droplets need only a piece of solid contaminant or an impact shock to convert them into ice crystals. Tests indicate that released, supercooled water droplets will not settle out in flight. Droplets are pumped freely through the system. If they become ice crystals in the tank, they will not settle, since the specific gravity of ice is approximately equal to that of kerosene. Although severe fuel system icing can occur at fuel temperatures from -18ºC to -29ºC, water droplets can freeze at any temperature below 0ºC. Water in jet fuels also creates an environment favourable for the growth of a microbial “sludge” in settlement areas of the fuel tanks. Sludge and other fuel contaminants can cause corrosion of metal parts in the fuel system and clogging of the fuel filters.

Fuel Anti-Icing Additive Even if the fuel does not contain water, or if water has been drained, the possibility of fuel icing exists at very low temperatures. The oil-to-fuel heat exchanger prevents fuel icing during most normal operating conditions; however it is essential that the required fuel anti-icing additives are blended into the fuel; either while refuelling, or by the refinery. Fuel anti-icing additive is necessary at all times for the BT-67.


NOTES:


ELECTRICAL SYSTEM

Table of Contents Electrical System................................. 6-1 DC Power Distribution ......................... 6-2 Significant Load Summary ................... 6-3 Distribution Bus Connected Loads ....... 6-4 Bus Tie System ................................... 6-5 Load Shedding .................................... 6-5 Batteries.............................................. 6-5 Starter/Generators............................... 6-5

DC Generation .................................... 6-6 Voltage Regulation.............................. 6-6 Electrical System Protection ................ 6-6 Generator Tests .................................. 6-6 External Power.................................... 6-7 Avionics Power.................................... 6-7 Circuit Breakers................................... 6-7 Electrical Schematics .......................... 6-8

Electrical System The aircraft electrical system is a 28-volt DC (nominal) system, with the negative lead of each power source grounded to the main aircraft structure. Direct Current electrical power is provided by two 12-volt 88-ampere-hour lead-acid batteries connected in series, and two 28-volt, 300-ampere-hour starter/generators, connected in parallel. This system is capable of supplying power to all subsystems necessary for normal airplane operation. The electrical power control switches are located on the co-pilot’s overhead panel (Figure 6-1).

Figure 6-1 Electrical Switches

The two aircraft batteries are mounted in battery trays, under the floor below the cockpit. Connected to the battery are both the cabin lights and the battery bus. Operation of the cabin lights is not dependant upon the battery switch position. The battery switch closes a battery relay, connecting the battery to the rest of the electrical system. Individual generator control units regulate output to supply constant voltage to the buses, compensating for variations in engine speed and electrical loads. The load on each generator is indicated by left and right loadmeters on the co-pilot’s overhead panel (Figure 6-2). A normal system charge of 28.25 (+0.25) volts maintains the batteries at full charge.


Figure 6-2 Volt and Ammeters

DC Power Distribution The Basler BT-67 utilises a multi bus electrical system. The buses are the battery bus, left and right distribution buses, essential bus, emergency bus, avionics bus, and emergency avionics bus (Figure 6-3). Electrical loads are divided among the buses as noted in Tables 6-1 and 6-2. Equipment is arranged so that items with duplicate functions (such as LH and RH landing lights) are connected to different buses.

Figure 6-3 Electrical System Schematic


All buses are supplied power under normal operating conditions (ground or flight). However, under most normal operating conditions (and always when under normal flight conditions), the electrical system is in a “split-bus” configuration, with the left and right distribution buses not tied together in common. Thus the left and right engine driven generators are not operated in parallel with both generators supplying power to the same buses. There are exceptions to the “split-bus” convention, and during these specific conditions, all seven buses will be ganged together in common, and will receive power from the same source(s). These conditions will occur when any of the following conditions exist: • The aircraft is on the ground with the battery switch ON, and neither, or only one, generator

is online. • Ground power is ON, regardless of battery switch position. • During any engine start cycle on either engine, on the ground or in-flight. • Whenever the bus tie switch is CLOSED. The first two conditions are controlled through an airspeed sensing switch which operates an associated ground/flight relay. The third condition is controlled by either engine start switch. Most of the electrical power is supplied by the left generator in normal operation. The left generator charges the battery and supplies power to the left distribution bus, the essential bus, and the avionics buses. The right generator supplies power to the right distribution bus. In the event that the left generator fails, the right generator then charges the battery and automatically supplies power to the essential, emergency and avionics buses in addition to the right distribution bus. The left distribution bus is shed. In the event that the right generator fails, the right distribution bus is shed.

Significant Load Summary Table 6-1 Left Distribution Bus Normal Load LIGHTS - L LANDING - PWR 21.5 Amps LIGHTS - WING INSP 7.1 Amps DE-ICE - R PITOT HEAT 5.2 Amps DE-ICE - LH PROP - PWR 26 to 30 Amps DE-ICE - LH INLET - PWR 45 to 50 Amps DE-ICE - SURF 8.0 Amps Right Distribution Bus LIGHTS - R LANDING - PWR 21.5 Amps DE-ICE - RH PROP - PWR 26 to 30 Amps DE-ICE - RH INLET - PWR 45 to 50 Amps DE-ICE - WINDSHIELD - RIGHT 25 to 27 Amps DE-ICE - WINDSHIELD - LEFT 25 to 27 Amps HEAT BLOWER 8.0 Amps Notes: These loads represent normal loads for each identified item of equipment when that item is receiving power, and thus serve as factors toward the calculation of the possible peak loads the 100-Amp Remote Control Circuit Breaker and the 150-Amp circuit breaker associated with either Distribution Bus could see as a consequence. These loads are those which must be considered in establishing individual and combined Distribution bus loading during Bus-Tie closed operation. Other loads on the Distribution Buses are not significant. Optional equipment connected to either Distribution Bus, if any, that draws any significant load and which is not listed above, must be included in possible total bus loading calculations.


Distribution Bus Connected Loads Table 6-2 Left Distribution Bus CB Panel Lights LH Landing Light Control LH Prop De-ice Power RH Auto-feather Surface De-ice

L Inlet Lip De-ice Control LH Landing Light Power Prop Governor Test RH Pitot Heat

L Inlet Lip De-ice Heater LH Prop De-ice Control Prop Sync Wing Inspection Light

Right Distribution Bus Cockpit Heat Blower Overhead Panel Lights RH Landing Light Control RH Prop De-ice Control

Cockpit Heat Control R Inlet Lip De-ice Control RH Landing Light Power LH Windshield Heat

LH Auto-feather R Inlet Lip De-ice Heater RH Prop De-ice RH Windshield Heat

Essential Bus Avionics Panel Lights Fuel Flow Indication Outboard Fuel Tank Qty LH Generator Cont Unit LH NG RPM Indicator Outside Air Temperature RH Directional Gyro Ind RH Fuel Purge Valve RH Main Fuel Pump RH Oil Temp/Press Ind Stall Warning

Engine Inst Panel Lights RH Fuel Xfer Pump LH Fuel Pressure Ind LH Inertial Sep Actuator LH Oil Temp/Press Ind Position Lights RH Fuel Pressure Ind RH Generator Cont Unit RH Map Light RH Stby Fuel Pump Cont Stall Warn Vane Heater

Flap Position Indicator RH Fuel Xfer Pump Cont LH Fuel Purge Valve LH Main Fuel Pump LH Stby Fuel Pump Cont RH Attitude Gyro Ind RH Flight Inst Panel Lts RH Inertial Sep Actuator RH NG RPM Indicator RH Turn & Bank Indicator One Inverter

Emergency Bus Airspeed Warning Fuel Xfeed Shutoff Valve LH Attitude Gyro Ind LH Fire Detection LH Fuel Xfer Pump LH Ignition LH Map Light LH Fire Squib A LH Turn & Bank Indicator RH Fire Detection RH Ignition RH NP Indication RH Standby Fuel Pump Warning Lights Power

Annunciator Panel Fuel Xfer Xfeed SO Valve LH Directional Gyro Ind LH Flight Inst Panel Lts LH Fuel Xfer Control LH ITT Indication LH NP Indication LH Fire Squib B LH Xfer Shutoff Valve RH Fuel Shutoff Valve RH ITT Indication RH Fire Squib A RH Xfer Shutoff Valve

Anti-collision Lights Landing Gear Warning LH Engine Torque Ind LH Fuel Shutoff Valve LH Hydraulic SO Valve LH Main/Aux Fuel Tank Qty LH Pitot Heat LH Standby Fuel Pump RH Engine Torque Ind RH Hydraulic SO Valve RH Main/Aux Fuel Tank Qty RH Fire Squib B Tone Generator

Avionics Bus See individual aircraft. Emergency Avionics Bus See individual aircraft.


Bus Tie System A manually operated bus tie selector switch is provided. For dual generator operation, the bus tie should be in the OPEN position, maintaining separation of the left and right distribution buses. During single generator operation, the inoperative generator’s distribution bus can be powered by closing the bus tie, subject to the following limitations: • The operating generator load limit must not exceed 300 Amps (250 Amps for ground

operations). • The inoperative generator distribution load must not exceed 100 Amps (bus tie circuit

breaker rating). • The sum of BOTH the left and right distribution bus loads must not exceed 150 Amps (each

distribution bus circuit breaker rating).

Load Shedding In the event both left and right generators fail, both left and right generator buses will be shed from the power load automatically. This will leave the battery to power the essential, emergency and avionics buses. Further load shedding can be accomplished manually. A two-position emergency power selector switch on the co-pilot’s overhead panel, when selected to the EMER position, will shed the essential and main avionics buses. This will leave only the emergency and emergency avionics buses online. The emergency power switch should be selected to EMER immediately after the failure of both generators. A fully charged battery will supply all emergency and emergency avionics bus loads for approximately one hour. This may be further extended by turning off any items not required for safe flight.

Batteries Two 12-volt 88-ampere-hour lead-acid batteries, connected in series, are located in the bottom of the fuselage forward of the wing leading edge. They are accessible through trap doors on the lower surface of the fuselage. Each battery is mounted on a platform which may be lowered until the battery is clear of the aircraft. Special terminals on the batteries plug into receptacles in the fuselage, so that when the batteries are raised into position they are automatically connected into the system.

Starter/Generators The starter/generators are dual-purpose, engine driven units. The same unit functions as a starter during engine starting, and as a generator when supplying electrical power. A series starter winding is used during starter operation, and a shunt field winding is used during generator operation. Regulated generator output is 28.25 (+0.25) volts, and 300 Amps maximum continuous load. Starter power to each individual starter/generator is provided from the battery bus through a generator bus relay, and a starter/generator relay. With one engine running, and its generator online, the operating generator can be used to assist the battery in starting the opposite engine. This is known as a generator cross-start. Normally the first engine is started on battery power alone, and the second engine is started using a generator-assisted cross-start. To prevent damage to the starter motors, starter operation time limitations must be observed. Use of the starter is limited to cycles of 30-seconds ON. A period of one minute OFF must be follow each start cycle, with a maximum of three 30-second ON start cycles allowed consecutively. After the third start cycle has been attempted, a minimum of 30-minutes OFF must occur before a fourth start cycle is attempted.


DC Generation The generator system consists of the starter/generator units, generator control switches, generator control units (GCU’s), line contactors and loadmeters. Generator switches, marked ON, OFF and RESET are mounted on the co-pilot’s overhead panel. The generating system is self-exciting and does not require electrical power from the aircraft electrical system for operation. The system uses generator residual voltage for initial generator build-up. Two GCU’s regulate generator output and provide constant bus voltage during variations in engine speed and electrical load requirements. Each generator’s load is indicated on the loadmeters on the co-pilot’s overhead panel.

Voltage Regulation Each generator is controlled by a three-position switch on the co-pilot’s overhead panel (Figure 6-4). When the generator control switch is held to RESET, the generator residual voltage is applied through the GCU to the generator shunt field, causing the generator output voltage to rise. As generator output approaches the 28-volt regulator setting, the voltage regulator circuit begins controlling the generator shunt field to maintain a constant output voltage. The voltage regulator circuit varies shunt field excitation, as required, to maintain a constant 28-volt generator output Figure 6-4 RH Gen. for all rated conditions of generator speed, load and temperature. Controls Selecting the generator control switch to ON applies generator voltage to the GCU. The GCU compares generator output voltage to aircraft bus voltage. If the generator output voltage has risen to within 0.5 volts of the aircraft bus, the GCU will close the starter/generator relay and connect the generator to the aircraft bus.

Electrical System Protection The GCU provides overvoltage protection to prevent excessive generator voltage from being applied to aircraft electrical equipment. If either generator output exceeds the maximum allowable 32-volts, the overexcitation circuits of the GCU will detect which generator is producing excessive voltage, and will disconnect that generator from the electrical system. Ground fault transformers sense current flow from the generators on both the ground and electrical system sides. If a current imbalance is sensed, signifying a short to aircraft, the GCU will turn the generator off.

Generator Tests Each generator’s protection circuitry is tested by a spring-loaded three-position switch on the co-pilot’s overhead panel (Figure 6-5). To test the overexcitation circuit in the GCU, the test switch is held to the OVERVOLT position. Correct functioning will be indicated by the selected generator being disconnected. The ground fault circuitry in the GCU is tested through the GRD FAULT position of the test switch. If the circuitry is functioning correctly the selected generator will be disconnected.


External Power External power may be supplied to the aircraft electrical system through an external power receptacle (Figure 6-5) located on the lower fuselage forward of the wing leading edge. A selector switch mounted on the co-pilot’s overhead panel is used to control the ground power relay. When the relay is closed, power is supplied from the external source to all aircraft buses. The green GROUND POWER annunciator illuminates when ground power is supplying power to the aircraft buses.

Figure 6-5 GPU External Receptacle

The aircraft batteries may be charged from a ground power source by selecting both the ground power and battery switches to ON. The external power source should be capable of delivering adequate power for aircraft starts. Using an inadequate ground power source can cause voltage drop, which may cause the starter to intermittently drop offline, resulting in relay chatter and possible welded contacts. Prior to attempting an external power start, aircraft electrical loads should be reduced to the minimum level practical. In the event that either generator is brought online while ground power is selected, ground power will automatically be disconnected, so as to avoid paralleling the ground power unit with the aircraft generator. To restore ground power, it will be necessary to turn the generator OFF, and then select the ground power switch to OFF to reset the latching circuit. Then select the ground power switch to ON. Observe the following precautions when using an external power source: • Ascertain that the external power source has a minimum capacity of 1000 Amps

(intermittent) and 300 Amps (continuous) output at 28.0 to 28.4 volts. Never attempt to start the aircraft’s engines with external power unless the aircraft battery indicates a charge of at least 20 volts.

• Use only an external power source that has a negative ground. • Be sure that the external power source is turned off before connecting it to the aircraft. All

avionics, the external power switch and the generator switches should be off. Generators should remain off until auxiliary power is disconnected.

• External power voltage must be regulated to 28.0 to 28.4 volts before it is plugged in to the external power receptacle. Voltages higher than 30 volts over extended periods of time may damage the battery.

Avionics Power The avionics electrical system is covered separately in Section 7 Avionics Systems.

Circuit Breakers Both AC and DC power are distributed to aircraft systems through circuit breaker panels, which protect most components in the airplane. Each circuit breaker is stamped with its amperage rating. The main circuit breaker panels are mounted beside the pilot’s head, the avionics circuit breaker panel beside the co-pilot’s head, and the bus and control circuit breakers are mounted beside the junction box at the cockpit door (Figure 1-28 to 1-33).


The avionics bus and emergency avionics bus are each equipped with a manual bypass circuit breaker switch. These red guarded switches are mounted on the electrical power junction box behind the pilot’s seat. They are used to manually supply power to their respective buses in the event avionics power is lost. Detailed procedures for tripped circuit breakers, and other electrical system malfunctions are contained in the AFM Emergency Procedures section.

Electrical Schematics The following diagrams (Figures 6-6 to 6-9) illustrate various conditions of the Basler BT-67 electrical system.


Figure 6-6 Electrical System Energised By GPU


Figure 6-7 Electrical System In Flight


Figure 6-8 Electrical System With Left Generator Failed


NOTES:


AVIONICS

Table of Contents Introduction ......................................... 7-1 Bendix/King Avionics Suite .................. 7-1 Avionics Power Distribution ................. 7-2 Intercom.............................................. 7-3 VHF Communications.......................... 7-4 HF Communications ............................ 7-4 Transponders ...................................... 7-5

VOR/ILS Navigation ............................ 7-6 DME Navigation .................................. 7-6 ADF Navigation................................... 7-7 Marker Beacons .................................. 7-8 GPS Navigation................................... 7-8 Radar Altimeters ................................. 7-9 Weather Radar...................................7-10

Introduction The purpose of this section is not to describe the theory of operation, nor to provide full operating instructions, but merely to provide an overview of the key features of the installed avionics. For full information refer to the applicable manufacturers’ handbooks.

Bendix/King Avionics Suite The avionics suite installed in the Basler BT-67 will differ according to customer requirements. To follow is a description of a typical Bendix/King installation, as fitted to the BT-67 operated by Spectrem Air.

Figure 7-1 Main Instrument Panel


Figure 7-2 Antenna Locations

Avionics Power Distribution The avionics of the BT-67 are distributed over two buses; the Avionics Bus and the Emergency Avionics Bus (Figure 7-3).

Figure 7-3 Simplified Avionics Electrical Schematic

The Number 1 avionics and all intercoms are located on the Emergency Avionics Bus, whilst all other secondary avionics and the HF radio, are located on the Avionics Bus. A complete listing of equipment powered by each bus can be found from the circuit breaker panel itself (Figure 7-4). Both avionics buses are triple-fed buses, receiving power from both the battery bus and from the generators. This ensures that the avionics can be powered at all times there is electrical power available. Under normal circumstances power to both avionics buses is applied simultaneously through the Avionics Master Switch via the Avionics Power Relay and Emergency Avionics Power Relay. In the event that the switch or either relay fails, two guarded emergency override switches are provided on the junction box circuit breaker panel behind the pilot’s seat, one for each avionics bus. Activation of either of these switches will bypass the respective power relay to provide power to that bus. In the event that both generators fail in flight, the Avionics Bus can be manually shed, leaving


only the avionics on the Emergency Avionics Bus powered. This is accomplished by selecting the EMER POWER switch on the co-pilot’s overhead panel to EMER. With this reduced electrical loading (the Emergency Bus will also still be powered) a fully charged battery should provide power for approximately one hour. This may be extended by turning off any items not required for safe flight.

Figure 7-4 Avionics Circuit Breaker Panel

As with all avionics, the radio equipment of the BT-67 should be turned ON only after engine start, and turned OFF prior to shutdown. This can be done either through the Avionics Master switch, or individually for each radio. Ensuring the avionics are not powered at these times will prevent voltage transients from damaging the solid-state circuitry of the avionics.

Intercom Intercom functions are provided by the installation of dual Bendix/King KMA 24H Audio Controls (Figure 7-5). In the dual installation both pilots can talk on different transmitters at the same time; however the captain will automatically have priority if he keys the mike while another crew member is using the same transmitter.

Figure 7-5 KMA 24H Audio Control

Each KMA 24H has a built in five-station intercom with two dedicated amplifiers. Intercom operation may be either continuous “hot mike”, voice activated, or keyed activation using separate press-to-talk buttons. Selection of the desired method of microphone activation is accomplished with the VOX sensitivity control. Turning the left outer concentric knob fully clockwise gives hot mike operation. Fully counter-clockwise gives keyed microphone operation. Separate intercom key switches must be installed in order to use this function. Any intermediate setting will give varying sensitivities to the voice-activated intercom. The inner concentric knob is the volume control for the intercom, which does not affect the levels of the other audio inputs. The rotary switch on the right of the KMA 24H selects the desired transmitter for the cockpit


microphones. The emergency EMG position connects the pilot’s mike and headphones directly to the COMM1 as a means of fail-safe communication in the event of a failure within the KMA 24H. The numbered positions on the rotary selector correspond to the various communications radios, in this case 1 is VHF COMM1, 2 is VHF COMM2, 3 is HF, and 4 is unused. In aircraft equipped with a cabin speaker, the PA selection allows the pilots to make cabin announcements. The EXT position is for use with the optional external ramp hailer. When either pilot keys the microphone to transmit, all other intercom microphone inputs are muted, which ensures that the keyed microphone is the single source of transmitted audio. All receiver inputs are also muted during transmission. Receiver selection is through the two rows of pushbuttons in the centre of the unit. The top row of buttons control the audio selection for the cockpit speaker, and the bottom row selects audio for the headphones. The selections are independent, and any audio input can be selected for speaker or headphones, or both. The volume of the audio input is set with the volume controls of the individual radios.

VHF Communications The KX 165 Nav/Comm (Figure 7-6) simultaneously displays two frequencies, one in use, and the other on standby. Toggling between these frequencies is achieved by pressing the ↔↔ button. Only the standby frequency can be adjusted using the frequency selector knobs. Selection of frequencies at 25 kHz spacing is possible with the smaller frequency knob. Pushed in it selects frequencies in steps of 50 kHz, pulled out frequencies are selected in 25 kHz steps.

Figure 7-6 KX 165 Nav/Comm

Whenever the unit is transmitting the T annunciator will be illuminated between the in-use and standby frequencies. To aid in receiving a weak station, pulling out the volume knob overrides the automatic squelch setting.

HF Communications As opposed to most of the other avionics, the HF system is not duplicated. It consists of a single, panel mounted KCU 951 Control Head (Figure 7-7), and remotely mounted power amplifier/antenna coupler and receiver/exciter.

Figure 7-7 KCU 951 Control/Display


• The MODE button is used to select emission mode between LSB, USB and AM. • Toggling between the pilot selected frequency and up to 99 pre-programmed channels is

effected by pressing the FREQ/CHAN button. • The CLARIFIER knob adjusts the received frequency to clarify speech in SSB operating

mode. • As signal strength varies greatly during HF transmissions, SQUELCH may need to be

continually adjusted during flight. • The STO channel store button, and the PGM programme buttons allow frequencies to be

programmed into the stored channels and also edited. To allow for proper tuning, allow the HF to warm up for at least two minutes before use. Once a frequency has been selected, key the microphone to tune the antenna. When the tuning sequence is complete the TX annunciator will stop flashing and the frequency display will reappear. Ensure no personnel are near the HF antenna when transmitting. Serious RF burns can result from direct contact with the HF antenna or antenna terminal when transmitting.

Transponders For IFR redundancy, dual KT 79 Transponders are fitted (Figure 7-8). The operation and functioning of both transponders is identical, with the exception of the source for their altitude information. The number 1 transponder receives altitude information from the pilot’s KEA 346 Encoding Altimeter (Figure 7-9). The number 2 transponder receives its information from the Blind Encoder.

Figure 7-8 KT 79 Transponder

The inner adjustment knob is used to select the required squawk code. The outer knob selects one of the four operating modes. • In SBY standby mode, the KT 79 is energised, but inhibited from replying to any

interrogation. • When in the ON mode the KT 79 will reply to all mode A and C interrogations, however

altitude information is suppressed. • In the ALT mode, replies with altitude information are made to all valid mode A and C

interrogations. • The TST test mode temporarily inhibits all replies. Internal tests are performed, and all

segments of the display panel are illuminated.

Figure 7-9 KEA 346 Encoding Altimeter

The R annunciator is illuminated in the centre of the screen when the unit is replying to an


interrogation in either mode A or C. In ALT mode the transponder displays the altitude reported in the left window. Switching between the two transponders is achieved through a two-position switch mounted on the main instrument panel between the two transponders.

VOR/ILS Navigation Control for the VOR and ILS navigation is through the dual KX 165 Nav/Comm’s (Figure 7-6). As with the VHF comm’s portion of the unit, both an in-use and a standby frequency can be simultaneously displayed, and swapped back and forth. When the inner frequency selector knob is pulled OUT, the VOR radial FROM the station in use is displayed in the right window. The standby frequency will go into non-displayed storage, while the frequency of the in use station can now be directly adjusted. If the station is too weak, or is an ILS frequency, the digital flag “- - -” will be displayed. Pulling OUT the volume knob allows both voice and Morse ident of the transmitting and station to be heard on the intercom. When the knob is pushed IN the ident tone is muted.

Figure 7-10 KI 525A Horizontal Situation Indicator and KA 51B Slaving Control

Both VOR and ILS navigation information is displayed on KI 525A HSI’s (Figure 7-10). Each HSI only displays the information from its respective Nav radio, no facility is installed to transfer between indicators. Both HSI’s are normally operated slaved to the aircraft’s compass system. However in the event of a failure in the compass system each HSI can be operated as a free compass, and manually corrected through individual KA 51B Slaving Controls.

DME Navigation DME information is displayed on separate indicators for both the pilot and co-pilot. The pilot display is the master indicator (Figure 7-11) with the mode selector switch, while the co-pilot’s is a slaved repeater unit. The indicators are capable of displaying DME distance, groundspeed and time to station simultaneously.

Figure 7-11 KDI 572 DME Master Indicator

The mode selector on the master indicator allows the DME unit to be channelled by either NAV1 or NAV2. Selecting the HLD position allows the DME to remain channelled to the


previously selected frequency, and is annunciated by either 1H or H2, depending on whether NAV1 or NAV2 was previously used. The mode selector also allows the DME to be turned off.

ADF Navigation The KR 87 ADF (Figure 7-12) has two basic modes of operation, ANT (antenna) and ADF. In the ANT mode, the bearing needle in the RMI indicator will not point to the station, but provides improved audio reception. The ANT mode is selected with the ADF button pressed out, and is annunciated ANT on the left of the ADF display. The ADF mode is used for navigation purposes, allowing the bearing needle to point to the station. The ADF mode is selected with the ADF button pressed in, and is annunciated ADF on the display.

Figure 7-12 KR 87 ADF

This unit incorporates a beat frequency oscillator circuit, which allows NDB's to be idented which are not modulating the carrier wave with audio. The BFO circuit, when activated by pressing the BFO button in, generates a 1020 Hz tone which will be heard each time the NDB transmitter is turned on. This allows the Morse coding to be identified in the normal manner. Two frequencies are able to be simultaneously displayed on the KR 87 ADF. The frequency on the left is always the frequency in use, however the right display is shared by several different functions. Normally displayed will be the standby frequency, however pressing in the FLT/ET pushbutton changes the function of the right window to the timer. When FLT is annunciated on the far right of the display window, the right window is displaying flight time, or more accurately the time since the unit was turned ON. Momentarily pressing the FLT/ET button again will bring up the elapsed timer. This timer can be reset to zero by pressing the SET/RST button. The elapsed timer also has a countdown timer function, which can be initiated by holding down the SET/RST button for approximately three seconds, or until the ET annunciator begins to flash. The countdown time can then be set by rotating the concentric frequency control knobs. In order to start the countdown timer, push the SET/RST button again. On reaching zero, the counter will commence counting up from the originally set time, and the time will flash for fifteen seconds to alert the pilot. Pressing the FREQ button initially changes the right window to the standby frequency mode. Subsequent pushes of the FREQ button transfers the in-use and standby frequencies back and forth.

Figure 7-13 KI 227 ADF Indicator

ADF station information is displayed on KI 227 indicators (Figure 7-13). The information from


each ADF is displayed only on the respective RMI, no facility is installed to transfer between indicators. Gyro heading information from the number 1 compass system is displayed on the co-pilot’s RMI, whilst the number 2 compass systems is displayed on the pilot’s indicator. In this way both pilot’s have heading information displayed from two independent sources. As these RMI’s are unslaved, each indicator’s compass rose can be manually set to the correct heading by rotating the SYNC knob.

Marker Beacons A single Marker Beacon Receiver is mounted on the main instrument panel. A three-position control switch to the right of the indicator selects the mode of operation; H high sensitivity, L low sensitivity, and T test. In the H mode the receiver is at its most sensitive and should detect the outer marker at approximately one mile during the approach. Selecting the L mode will set the receiver to its lower sensitivity to detect actual station passage.

GPS Navigation The GPS unit selected for the BT-67 is the IFR certified Trimble TNL 2000T (Figure 7-14).

Figure 7-14 Trimble TNL 2000T GPS

This unit features a removable Jeppesen subscription database card holding location and data on worldwide airports and navigation facilities. Never insert or remove the datacard from the GPS when the power is ON. The system automatically resets when the card is removed, and there is a risk of data corruption and other system errors. The TNL 2000T is turned on through the power switch located at the top centre of the unit. The GPS annunciator indicates when the unit is in 3-D mode, when the annunciator is flashing, operation is in 2-D mode. When the unit has sufficient integrity monitoring, RAIM, the IFR annunciator will be illuminated. Information is displayed on the TNL 2000T on a two-line LED screen. Several different sets of information are displayed in each operating mode. The LED screen features automatic dimming for varying cockpit light intensities. The pilot controls the operation of the TNL 2000T with the lower Mode keys, the right Function keys, and the Selector knobs. Annunciator lights alert the pilot to any system messages. • The NAV key selects the navigation mode, displaying the GPS navigation information. • Pressing the WPT key selects the waypoint mode. Complete information is available on

database waypoints of airports, VORs, NDBs, intersections, and for user-defined waypoints. User waypoints can be manually defined or modified in waypoint mode.

• The FPL key selects the flight plan mode. Flight plans can be created, edited, reversed or erased. The TNL 2000T can store up to 20 flight plans.

• Pressing the CALC key selects the calculator mode. Many common flight computer functions are available, including TSD, TAS, winds, fuel management and VNAV calculations. Autosaving the present position is also a function of the calculator mode.


• The AUX key selects the auxiliary mode, for access to the less commonly used functions of checklists, system status, sensor status, system set-up and installation.

• The APT/VOR button cycles between the 20 nearest waypoints in each waypoint category. • The D key selects the direct-to navigation function. • Pressing the MSG key will display any messages present. • The ENT key is used to enter, select or change the information displayed on the LED

screen. • For NAV functions, the inner selector knob changes the information on the top line of the

LED, the outer selector knob changes the information on the bottom line. • There are usually several different lines of information available in each mode of operation.

Turning either knob clockwise scrolls forward through the information, while turning counter clockwise scrolls backwards.

• The selector knobs also control the change of information on these lines and are used for data entry. The inner knob is used to select a data item such as a letter of the alphabet. The outer knob is used to move from one character field to the next.

• The red WRN annunciator will flash to indicate that an accuracy or problem message is

waiting. • The ADV annunciator indicates that an advisory message is waiting. Pressing the MSG

button will cycle through any messages. Once the messages have been read the annunciators will stop flashing, but will remain lit until the condition has been rectified.

• The PTK annunciator will be illuminated to indicate that the parallel track mode has been selected.

Figure 7-15 GPS Annunciators

The GPS annunciator lights and navigation display function switch are mounted above the flight instruments on the main instrument panel. The four-section annunciator light will illuminate to display the annunciator messages displayed on the GPS. The NAV/GPS/Loran switch allows the navigation information displayed on the pilot’s HSI to be toggled between the Nav1 and the GPS. In GPS mode the crosstrack indication is not affected by the selected heading of the indicator needle.

Radar Altimeters The Spectrem AEM aircraft is fitted with two independent radar altimeter systems. Both systems display through separate KNI 415 Radar Altimeter Indicators (Figure 7-16) mounted on the main instrument panel. The KNI 415 gives accurate altitude indications from –20 feet to +2,000 feet.

Figure 7-16 KNI 415 Radar Altimeter Indicator

7-10

A self-test button is used to test the radar altimeter R/T and indicator. The Decision Height (DH) lamp will also be lit if the DH bug is set above 50 feet during the test. The DH knob controls the DH bug, which indicates height during an approach. The DH lamp illuminates when the DH is reached. The lamp can be turned off by pushing it in, and can be turned on again by depressing it a second time. Once turned off, the DH lamp will be automatically armed climbing again through the DH.

Weather Radar No weather radar is currently installed on the BT-67 aircraft operated by Spectrem Air.


AUTO-FLIGHT

Table of Contents Introduction ......................................... 8-1

Introduction The Basler BT-67 operated by Spectrem Air is not fitted with the optional Basler supplied autopilot.


NOTES:


PITOT-STATIC SYSTEM

Table of Contents Pitot-Static System .............................. 9-1 Alternate Static System ....................... 9-2

Airspeed Switches............................... 9-2

Pitot-Static System The pitot-static system provides a source of ram air and static air for operation of the flight instruments. A pair of heated pitot-static tubes are located on the fuselage under the cockpit. Tubing from the upper pitot tube is connected to the co-pilot’s airspeed indicator, and tubing from the lower pitot tube is connected to the pilot’s airspeed indicator (Figure 9-1). The pilot’s pitot pressure source is completely independent of the co-pilot’s pitot source.

Figure 9-1 Pitot Static System Schematic

The normal static system provides sources of static air from both pitot-static tubes to both pilot and co-pilot flight instruments. The static ports are incorporated into the sides of each pitot-static tube, and are open to the atmosphere, providing the source for normal static pressure. Pitot-static lines are interconnected to provide a redundant static air source to both sides of the airplane. The pilot’s static source is completely independent of the co-pilot’s static air source.


The pitot-static tubes can be heated electrically for flight in icing conditions. As a precautionary measure, it is customary to have the pitot heat on during flight in visible moisture at OAT +5º C and below. It is not advisable to operate the pitot heat system on the ground, except for testing or for short intervals to remove ice or snow from the tubes. Operating the pitot heat on the ground for extended periods can damage the internal heating elements in the pitot-static tubes.

Alternate Static System If the normal static source fails, alternate static air lines can be selected as the static source for the pilot’s and co-pilot’s flight instruments. If, for example, ice accumulations obstruct the static air ports, the alternate source should be selected (Figure 9-2). The alternate line obtains air from the alternate static source located in the left auxiliary fuel bay.

Figure 9-2 Alternate Static Selector

Airspeed and altimeter indications change when the alternate static air source is selected. Refer to the Airspeed Calibration – Alternate System and Altimeter Correction – Alternate System graphs in the Performance section of the AFM for correct indications when using the alternate static air source. When the alternate static air source is not needed, ensure that the static air source valve is in the NORMAL position.

Airspeed Switches There are three airspeed switches in the pitot-static system. They are located in the nose, forward of the main instrument panel. Two airspeed switches are connected to the pilot’s pitot system, and one to the co-pilot’s system. The pilot’s forward switch senses high airspeed and controls the overspeed warning system. It is operated by pitot pressure acting on a diaphragm. The pitot pressure is countered by spring pressure on the other side of the diaphragm. The pilot’s rear switch senses low airspeed and controls the activation of the inlet lip de-ice system and the distribution bus system. The switch is connected to both the pitot and static air sources. The co-pilot’s switch also senses low airspeed. This switch controls the activation of the inlet lip de-ice system and the stall warning system. The switch is operated by pitot pressure acting on a diaphragm, measured against a spring.


HYDRAULIC SYSTEMS

Table of Contents Hydraulic Power Supply..................... 10-1 Hydraulic Fluid Reservoir................... 10-2 Hydraulic System Accumulator .......... 10-2 Hydraulic Hand Pump........................ 10-3

Wing Flaps.........................................10-4 Landing Gear and Brakes...................10-5 Windshield Wipers..............................10-5

Hydraulic Power Supply A pressure accumulator hydraulic power system (Figure 10-1) operates the landing gear, wheel brakes, wing flaps and the windshield wipers. An engine-driven hydraulic pump is installed on the accessory drive of each engine. Approximately 221/2 litres of hydraulic fluid are required to fill the system.

Figure 10-1 Hydraulic System Schematic


The engine-driven variable-displacement hydraulic pumps maintain a system operating pressure of 875 +25 psi. The main system relief valve functions to protect the hydraulic system from excessive fluid pressure, opening at 1000 +50 psi to return fluid to the reservoir. Hydraulic system pressure is shown on the hydraulic pressure gauge located on the main instrument panel (Figure 1-17).

Hydraulic Fluid Reservoir Hydraulic fluid is gravity-fed from the hydraulic reservoir, located behind the bulkhead behind the co-pilot’s seat (Figure 10-2), to the two engine-driven variable-displacement hydraulic pumps, which supply fluid pressure for the hydraulic system. The fluid capacity of the reservoir is 10 quarts. Seven quarts are available to the engine-driven hydraulic pumps, while the remaining 3 quarts in the reservoir sump are available only to the hydraulic hand pump for emergency operation.

Figure 10-2 Hydraulic Fluid Reservoir

A sight gauge is provided on the reservoir to allow an accurate reading of hydraulic fluid levels when hydraulic systems are not being utilised. Due to the varying fluid demands of the hydraulic actuators, hydraulic fluid may or may not be visible during flight. Consequently, the reservoir should not be replenished during flight unless a known loss of hydraulic fluid occurs. The 3 quarts of reserve fluid are not visible in the sight gauge.

Hydraulic Pressure Accumulator The hydraulic pressure accumulator is attached to a bracket adjacent to the hydraulic fluid reservoir. The lower chamber of the accumulator is charged with nitrogen to an initial pressure of 250 psi. The nitrogen charge forces the diaphragm against the wall of the upper chamber. When the system pressure builds above 250 psi, hydraulic pressure partially overcomes the nitrogen pressure and forces the diaphragm down as fluid flows into the upper chamber. When hydraulic system pressure reaches 875 psi, the upper chamber contains approximately 5 quarts of hydraulic fluid.


The highly compressed nitrogen in the lower chamber forces fluid out of the upper chamber to replenish the system when: • the landing gear is lowered. The weight of the landing gear and the air-load against it cause

the gear to fall rapidly. Without the accumulator to replace fluid immediately, the system would be depleted faster than the engine-driven pumps could replenish it, causing a vacuum in the system.

• the engine-driven pumps fail or a leak develops. The hand pump can be used to operate any hydraulically-driven service or to pressurise the accumulator.

The accumulator also acts as a shock absorber to dampen pressure surges in the hydraulic system.

Hydraulic Hand Pump The hydraulic hand pump, located at the bottom of the hydraulic control panel (Figure 10-3), provides pressure to hydraulically operated services when there is insufficient pressure in the system, or when the hydraulic fluid supply (except the hand pump reserve) has been lost. It is a double-acting piston-type pump, supplying pressure with every stroke of the pump handle.

Figure 10-3 Main Hydraulic System Controls

Hydraulic fluid is supplied to the hand pump from the bottom of the hydraulic fluid reservoir. The reservoir is designed to prevent the engine-driven pumps from using the emergency supply of fluid (three quarts) reserved for the hand pump. The hydraulic hand pump has two operating modes; either supplying pressure directly to hydraulic services, or pressurising the hydraulic accumulator. A hand pump-to-pressure accumulator valve on the hydraulic control panel is used to route hand pump pressure directly to a hydraulically driven service, or to the pressure accumulator. In flight, the shutoff valve must be closed, so that hydraulic services are directly actuated by operation of the hand pump. The hand pump may also be used to increase pressure in the hydraulic pressure accumulator for ground operation of the hydraulic system when it is not practical to run the engines. Opening the shutoff valve provides a bypass around the check valve in the pressure manifold when it is necessary to increase pressure in the accumulator by operating the hand pump.


Wing Flaps The wing flaps are of the split trailing edge type, and are composed of four sections interconnected with push/pull rods to function as one unit. The flap sections are located on the underside of the centre section and outer wings, beginning at the inboard end of each aileron. The range of travel is 0º to 45º , and is shown on the flap position indicator on the main instrument panel (Figure 1-17). The wing flaps are controlled by the three-position wing flap control valve located on the hydraulic control panel (Figure 10-3). The flaps are lowered or raised by moving the flap control valve handle from the NEUTRAL position to the DOWN or UP position, respectively (Figure 10-4). Upon reaching the desired flap setting, the selector valve handle is returned to the NEUTRAL position.

Normal Down Flow: The DOWN position of the operating valve handle directs the fluid pressure in the down line. Valve piston remains in normal position, allowing down flow unless down line pressure exceeds approximately 485 psi. Wing flaps move down.

Normal Return Flow: The by-pass line is an emergency measure for almost instantaneous return of relief valve piston to its normal position. When by-pass line flows the pressure drops in the down line allows the spring to raise relief valve piston. Normal return flow of fluid results.

System Flow Blocked:

The DOWN position of the operating valve handle directs fluid pressure into the down line. Wing flaps move down. Back pressure in the down line caused by wind resistance on the wing flaps force down the relief valve piston blocking the entire system flow. Downward movement of the wing flaps is halted.

Relief Flow: The NEUTRAL position of the operating valve handle blocks flow in the wing flap system. Strong wind resistance on the wing flaps causes fluid back pressure to flow through the relief port. Wind raises the wing flaps until the back pressure is relieved.

Figure 10-4 Wing Flap Hydraulic System – Normal Flow


A relief valve prevents the lowering of the wing flaps to the full DOWN position at an indicated air speed greater then 97 KIAS. If excessive airloads are encountered while the wing flaps are extended 0 to 1/2, the relief valve will allow the flaps to raise until excessive loads are relieved. If wing flaps are extended beyond the 1/2 position, mechanical leverage prevents flap retraction when excessive airloads are encountered.

Bypass Emergency Flow:

The operating valve handle in the UP position directs fluid pressure into the up line. If the pressure in the down line has not been relieved, the relief valve piston will not return to its normal position. Since the relief valve is closed, the entire down line return flow routes through the bypass line to the reservoir. Wing flaps start up.

Normal Return Flow: The UP position of the operating valve handle directs the fluid pressure into the up line. Wing flaps move up. The orifice check valve restricts the return flow of fluid from the actuating cylinder adding a time lag which prevents rapid raising of wing the flaps, and consequent sudden loss of lift.

Figure 10-5 Wing Flap Hydraulic System – Bypass Flow A flap/landing gear aural warning transmitter is installed on the wing flap actuating rod. Whenever the wing flaps are extended beyond the 3/8 position, and the landing gear is not fully extended, the landing gear warning horn will activate and is non-cancellable.

Landing Gear and Brakes The landing gear and brake systems are covered separately in Section 11 Landing Gear Systems.

Windshield Wipers The windshield wipers are covered separately in Section 13 Ice and Rain Protection - Windshield Wipers.


NOTES:


LANDING GEAR SYSTEMS

Table of Contents Landing Gear Components................ 11-1 Retraction.......................................... 11-3 Extension .......................................... 11-3 Safety Latch Mechanism.................... 11-3 Out Of Sequence Rectification........... 11-5 Hand Pump Alternate Operation ........ 11-5

Ground Safety Pins ............................11-5 Tail Wheel Lock..................................11-6 Tyres..................................................11-6 Brakes ...............................................11-7 Brake Operation .................................11-7

Landing Gear Components The landing gear consists of the main retractable landing gear assemblies (Figure 11-1) and a non-retractable, full-swivelling tail wheel. The tail wheel incorporates a centre locking pin, which is controlled by the pilot to aid in directional control during takeoff, landing and taxiing. The upper end of each main landing gear shock truss, and the rear brace strut are attached to the aircraft structure. The lower section of each shock truss incorporates an oleo-pneumatic shock strut. Airplane weight is borne by the air charge in the shock struts. At touchdown, the lower portion of each strut is forced into the upper cylinder. This moves fluid through an orifice, further compressing the air charge and thus absorbing landing shock. Orifice action also reduces bounce during landing. Both shock struts for each main wheel are connected by an airline running along the oleo truss, which balances forces on each shock strut. During retraction, the actuating cylinder pulls on the upper section of the shock Figure 11-1 Main Landing Gear


truss, causing the truss to pivot at its centre knuckle, and retract the gear into the gear well. Similarly, during extension, the upper section of the shock truss is forced down by the action of the actuating cylinder, aided by gravity. To assist in locking the gear in the extended position, the shock truss knuckle is moved into an over-centre position by the hydraulic actuator. A micro-switch mounted on the rear wall of the gear well is contacted by the main landing gear reaching the fully down position. To ensure that the main landing gear cannot be accidentally retracted, a mechanical safety latch system is installed. This will be more fully explained under the SAFETY LATCH MECHANISM of this section. Landing gear system controls are located on the hydraulic panel, between the two pilots’ seats (Figure 11-2).

Figure 11-2 Landing Gear Hydraulic Controls

Main landing gear position indication (Figure 11-3) consists of a red GEAR UP light, a green GEAR down light, and a landing gear hydraulic pressure gauge. The GEAR UP light will illuminate anytime the gear is not fully down, and/or the landing gear handle is not in the NEUTRAL position. The GEAR DOWN light will illuminate only when the gear is fully down and the landing gear handle is in the NEUTRAL position.

Figure 11-3 Landing Gear Indicators

The landing gear hydraulic pressure gauge only indicates pressure applied to the down side of the landing gear hydraulic actuator.


Retraction Retraction of the main landing gear is accomplished as follows: • Select the mechanical safety latch control handle to the LATCH RAISED position. This

raises the safety latches of each main gear hydraulic actuating cylinder. • Select the landing gear control valve to the UP position. This directs system hydraulic

pressure to the lower end of each actuating cylinder and retracts the landing gear. The GEAR DOWN light will go out as the micro-switch is released, and the GEAR UP light will illuminate.

• When the gear has fully retracted, select the landing gear control valve to the NEUTRAL position. This locks the hydraulic pressure in each landing gear actuating cylinder, holding the gear in the up position, and repositions the mechanical safety latch control handle to the SPRING LOCK position. The landing gear hydraulic pressure indicator will read zero and the GEAR UP light will be illuminated.

Extension Extension of the main landing gear is accomplished as follows: • Select the landing gear control valve handle to the DOWN position. This relieves the

hydraulic pressure on the lower side of each main actuating cylinder, directs hydraulic system pressure to the upper side of each cylinder, and forces the landing gear to extend.

• Check the landing gear hydraulic pressure gauge. When the main landing gear has fully extended the landing gear pressure indication should be equal to the hydraulic system pressure.

• Select the landing gear control valve handle to the NEUTRAL position. The GEAR UP light will go out, and the GEAR DOWN light will illuminate when both the gear down micro-switch and the handle neutral micro-switch are both contacted.

• Clip the safety latch control valve handle to the floor in the POSITIVE LOCK position. Note: if the aircraft is left unattended for extended periods of time, move the landing gear control valve handle to the DOWN position. This will prevent damage to the landing gear hydraulic system due to thermal expansion.

Safety Latch Mechanism A spring-loaded mechanical safety latch is installed in each nacelle on the forward side of the front spar. These automatically latch when the landing gear is fully extended, by engaging a slot in the lower end of the actuating cylinder piston rod. The latches for both gear legs are controlled simultaneously by cables connected to a single control handle located on the cockpit floor, between the pilots’ seats (Figure 11-4).

Figure 11-4 Landing Gear Safety Latch


The control handle has three positions: • UNLOCKED, which lifts the lock (also called the spade), • SPRING LOCK, which receives and latches the landing gear actuating rod hook when the

gear is extended, • POSITIVE LOCK, which may be locked by mechanical linkage rather than spring

pressure after the gear is spring locked down. This is accomplished by positioning the latch handle against the floor and maintaining it there by means of the locking clip ring.

Figure 11-5 Landing Gear Latch Lever Operation The mechanical safety latch mechanism is interconnected with the landing gear handle to provide the following operations (Figure 11-5):


• Holds the spade assembly up to clear the actuating cylinder hook when the latch handle is positioned to LATCH RAISED, so that when the gear handle is selected to UP, the gear will be free to retract.

• When the gear lever is moved from UP to NEUTRAL, the mechanism automatically positions the mechanical safety latch to SPRING LOCK.

• When the gear selector is positioned to NEUTRAL or DOWN, the mechanism puts a stop on the gear selector so that it cannot be moved to UP while the mechanical safety latch is positioned to SPRING LOCK or POSITIVE LOCK.

Out Of Sequence Rectification It is possible that the latch handle is incorrectly moved to positive lock when the landing gear handle is positioned to NEUTRAL, with the landing gear UP. If this takes place, the situation may not be noticed, and the gear handle may be positioned to DOWN to lower the gear. The gear will move down normally, however, when the hook strikes the spade attempting to automatically lock, the spade would not move and serious structural damage may result to the latch mechanism and the front spar. Likewise there would be no GEAR DOWN annunciator, as the gear would not be locked down. Should the latch handle become jammed in the UNLOCKED position by an out of sequence operation, the dog should be pulled forward to allow the catch to spring into place. The latch handle will then return to the SPRING lock position.

Hand Pump Alternate Operation The hydraulic hand pump, located on the hydraulic panel (Figure 11-2), has a handle that extends between the two pilots’ seats. The operation of the hand pump is covered separately in Section 10 Hydraulic Systems.

Ground Safety Pins Landing gear safety pins are provided to prevent inadvertent retraction of the landing gear when the aircraft is on the ground (Figure 11-6). For increased visibility, the pins are attached to red fabric streamers. When not installed, the safety pins are hooked onto the hydraulic panel, so that both pilots can confirm the pins are not installed.

Figure 11-6 Main Landing Gear Ground Safety Pin

The ground safety pins should be installed at all times when the aircraft is parked, and especially prior to any maintenance of the aircraft.


Tail Wheel Lock The spring-operated lock for the tail wheel is a duralumin pin hinged to the fuselage structure. The pin is pulled from a slot in the tail wheel strut by a cable connected to handle on the lower section of the control pedestal (Figure 11-7), allowing the tail wheel to turn freely. A spring locks the tail wheel assembly when the handle is released.

Figure 11-7 Tail-wheel Lock Lever

In the event of excessive side loads being applied to the tail wheel, the pin will shear, allowing the tail wheel to swing free before damage occurs to the fuselage structure in this area. The lock may only be engaged when the tail wheel is in the trail position.

Tyres The main landing gear wheels are equipped with a 45 x 17:00 x 16, 10-ply rated, tubed tyre. The tail wheel is equipped with a single 22 x 9:00 x 6, 8-ply rated, tubed tyre. All tyres should be inflated to between 30 psi and 60 psi, depending on aircraft takeoff weight.

18

19

20

21

22

2324

2526

27

2829

3031

3233

30 40 50 60PSI

Gro

ss T

akeo

ff W

eig

ht

(100

0 lb

s)

Main GearTyres

Tail WheelTyre

Figure 11-8 Tyre Inflation Pressures


Brakes The main landing gear expander tube-type wheel brakes may be applied simultaneously or independently by means of brake control valves contained in a single housing and linked to the rudder-brake pedals. Application of toe pressure on the rudder-brake pedals allows hydraulic fluid under pressure to flow through the brake control valves and brake lines to the brake assemblies. In the brake assemblies, expander tubes are forced outward by hydraulic pressure to press brake pucks against the brake drums. The pressure applied to the brakes is proportional to the toe pressure applied to the rudder pedals. When the rudder-brake pedals are released, hydraulic pressure is relieved and leaf springs force the pucks inward against the expander tubes. The hydraulic fluid flows back through the brake lines, through the brake control valve and into the return line to the hydraulic reservoir. A parking brake mechanism secures the rudder-brake pedals in the depressed position. The parking brake is covered fully in Section 01 General Airplane – Parking and Securing.

Brake Operation Use extreme care when applying brakes immediately after touchdown, or at any time when there is considerable lift on the wings, to prevent skidding the tyres and causing flat spots. Heavy brake pressure can result in locking the wheel more easily immediately after touchdown, than when the same pressure is applied after the full weight of the aircraft is on the wheels. A wheel, once locked in this manner immediately after touchdown, will not become unlocked as the load is increased, as long as brake pressure is maintained. Proper braking action cannot be expected until the tyres are carrying heavy loads. Although brakes can stop the wheels from turning, stopping the aircraft is dependant on the friction of the tyres on the runway. For this purpose, it is easiest to think in terms of the coefficient of rolling friction, which is the frictional force divided by the loads on the wheel. It has been found that optimal braking occurs with approximately a 15 to 20 percent rolling skid; that is the wheel continues to rotate, but has approximately 15 to 20 percent slippage on the surface, so that the rotational speed is 80 to 85 percent of the speed which the wheel would have were it in free roll. As the amount of skid increases beyond this amount, the coefficient of friction decreases rapidly so that, with a 75 percent skid, the friction is approximately 60 percent of the optimum and, with a full skid, becomes even lower. There are two reasons for this loss in braking effectiveness with skidding. First, the immediate result of the skid is to scuff the rubber, tearing off little pieces which act almost like rollers under the tyre. Second, the heat generated by the skid friction starts to melt the rubber, and the molten rubber acts as a lubricant. NACA figures have shown that, for an incipient skid with an approximate load of 10,000 lbs per wheel, the coefficient of friction on dry concrete is as high as 0.8, whereas the coefficient is in the order of 0.5 or less with a 75 percent skid. Therefore, if one wheel is locked during application of brakes there is a very definite tendency for the aircraft to turn away from that wheel and further application of brake pressure will offer no corrective action. Since the coefficient of friction goes down when the wheel begins to skid, it is apparent that a wheel, once locked, will never free itself until brake pressure is reduced so that the braking effect on the wheel is less than the turning moment remaining with the reduced frictional force. The following procedures will apply for brake operation: • If maximum braking is required after touchdown, lift should first be decreased as much as

possible by raising the flaps before applying brakes. This procedure will improve braking action by increasing the frictional force between the tyres and the runway. However, immediately following maximum braking while landing, little or no braking action may be available because of brake fade.


• For short landing rolls, a single, smooth application of the brakes with constantly increasing pedal pressure is most desirable.

• When the brakes are used to stop the aircraft, it is recommended that a minimum of 15 minutes elapse between landings where the landing gear remains extended in he slipstream, and a minimum of 30 minutes between landings where the landing gear has been retracted, to allow sufficient time for cooling between brake applications. Additional time should be allowed for cooling if brakes are used for steering, cross-wind taxiing operation, or a series of landings are performed.

• The full landing roll should be utilised to take advantage of aerodynamic braking and to use the brakes as little and as lightly as possible.

• After the brakes have been used excessively for an emergency stop, and are in a heated condition, the aircraft should not be taxied into a crowded parking area or the parking brakes set. Peak temperatures occur in the wheel and brake assembly from 5 to 15 minutes after a maximum braking operation. To prevent brake fire and possible wheel assembly explosion, the specified procedures for cooling brakes should be followed.

• The brakes should not be dragged when taxiing, and should be used as little as possible for turning the aircraft on the ground.

• Taxiing with one engine inoperative is not recommended.


ENVIRONMENTAL SYSTEMS

Table of Contents Heating System................................. 12-1 Heating System Controls ................... 12-2

Cockpit Windows................................12-2 Supplemental Oxygen ........................12-2

Heating System The Basler BT-67 standard environmental system (Figure 12-1) utilises P3 bleed air from each engine to provide heating for the cabin. A separate, optional, air-conditioning system may also be installed to provide cabin cooling.

Figure 12-1 Heating System


The bleed air is directed through a shut-off valve to an ejector pump located in each outboard wheel well. The ejector pump mixes outside air with bleed air to obtain a specified temperature, which is controlled by a thermostatic sensor in the cockpit. The mixed air from each engine is then directed to a common plenum located below the floor in the main cabin. A centrifugal blower is attached to the plenum which mixes ambient air with the bleed air to hold the air temperature entering the cabin ducting system to a predetermined value. Vents along the main cabin floor and in the cockpit deliver the heated air to the cabin.

Heating System Controls When the POWER switch is selected to ON, the cabin sensor, duct sensor, bleed air shutoff valve and temperature controller are activated. The computer is in standby mode and the bleed air valves are moved to the closed position stopping the flow of air. Selection of the ON/STBY switch to ON activates the computer, allowing the system to start controlling the bleed air heating valve to allow warm air to enter the aircraft cabin. In AUTO mode, the heating system is controlled automatically by the temperature controlling computer. Cabin temperature is sensed by the cabin temp sensor located above the hydraulic quantity sight gauge. The sensed temperature is compared to the selected temperature by the temp controller, which then modulates the bleed air valves to increase or decrease the engine bleed air flow. The hi temp cutout switch, when activated at 121ºC, turns on the ambient blower which mixes ambient air with the hot air, resulting in a lower duct air temperature. The blower may also be operated manually to provide cockpit ventilating air when heat is not required. If the hi temp cutout switch fails to operate, and the duct temperature reaches 149ºC the duct temp sensor will activate the temperature computer to drive the bleed air valves to the fully closed position. No air will enter the cabin until the duct temperature has fallen to a safe value. Over-temp sensors are located in each wheel well heat duct. The sensors will activate at 232ºC and illuminate a duct over-temp HEAT DUCT O’HT warning light on the annunciator panel. The system should then be selected to the MAN position and the duct temperature manually decreased until the annunciator lights are no longer illuminated. In the event that the duct warning lights do not extinguish, it will be necessary to shut down the heating system. In the event that the temperature sensing or controlling system fails, the bleed air valves may also be controlled manually. This is accomplished by selecting the TEMP CONTROL switch to MAN, and toggling the MANUAL TEMP control switch to increase or decrease the cockpit temperature.

Cockpit Windows Both pilot’s side windows can be safely opened at any stage of flight. Due to the low pressure area around the aircraft’s nose, a slight suction is created at the open windows. This results in an increased airflow from the cabin to the cockpit with either window open. If necessary the squelch settings of the intercom may need to be reset due to the increased cockpit noise level with open side windows.

Supplemental Oxygen The Basler BT-67 operated by Spectrem Air is not fitted with either of the optional Basler supplied supplemental oxygen systems.


ICE AND RAIN PROTECTION

Table of Contents Introduction ....................................... 13-1 Ice Protection Systems ...................... 13-2 General Description and Operation.... 13-3 Precautions During Icing Conditions .. 13-3 Severe Icing Procedures.................... 13-4 Engine Inertial Separators ................. 13-5 Inertial Separator Controls ................. 13-6 Engine Inlet-Lip Heaters .................... 13-6 Inlet-Lip Heater Controls.................... 13-7 Engine Continuous Ignition ................ 13-8

Fuel System Anti-Ice ..........................13-8 Propeller De-Ice .................................13-9 Windshield Anti-Ice.............................13-9 Windshield Wipers............................13-10 Pitot-Static Heat ...............................13-10 Stall Warning Vane Heat ..................13-11 Surface De-Ice Systems...................13-11 Wing Ice Lights.................................13-12 Alternate Static Air System...............13-12

Introduction The Basler BT-67 is certified for day, night and instrument conditions when the required equipment is installed and operative. Flight into known icing conditions is prohibited. Severe icing may result from environmental conditions outside those for which the airplane is certificated. Flight in freezing rain, freezing drizzle, or mixed icing conditions (super-cooled liquid water and ice crystals) may result in ice build-up on protected surfaces exceeding the capacity 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. 1 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.

A. Unusually extensive ice accumulation on the airframe and windshield in areas not

normally observed to collect ice. B. Accumulation of ice on the upper surface of the wing aft of the protected area. C. Accumulation of ice on the engine nacelles and propeller spinners further aft than

normally observed. 2 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.


3 All wing icing inspection lights must be operative prior to flight into known or forecast icing conditions at night.

[NOTE: This supersedes any relief provided by the Master Minimum Equipment List (MMEL).]

Ice Protection Systems The Basler BT-67 has NOT been approved by the FAA for flight into known icing conditions, however the airplane is equipped with a variety of ice protection systems for operation during inadvertent icing encounters. Ice protection controls are primarily grouped together on the pilot’s overhead panel. The windshield wiper controls are located on the control pedestal.

Figure 13-1 Anti-icing and De-icing Components

Figure13-2 Cockpit De-icing Controls


General Description and Operation The Basler BT-67 is equipped with a variety of ice protection systems. Only one, the surface de-ice, is a de-icing system designed to be used AFTER ice has accumulated. All other ice protection systems are to be used as anti-icing systems to PREVENT the formation of ice on aircraft surfaces. The following is a list of ice protection systems provided for the Basler BT-67: • Inertial Separators (Ice Vanes) • Engine Inlet-lip Heaters • Propeller De-ice • Windshield Anti-ice • Pitot Heat • Stall Warning Vane Heat • Surface De-ice (Leading edge Boots) Ice protection for the engine is provided by an inertial separation system, which is driven by an electrical actuator. The leading edge of the engine air intake is protected by an electrically heated rubber boot. The propellers are protected from icing by electrically heated rubber boots on each blade. Electrical heating elements embedded in the windshield provide adequate protection against the formation of windshield ice. Windshield wipers for both the pilot and co-pilot provide increased visibility for approaches and taxi operations in rain. A heating element in the pitot-static tube prevents the pitot opening from becoming clogged with ice. The stall waning vane is electrically heated. Pneumatic de-icing boots on the wings, horizontal and vertical stabilizers remove ice after it has formed on the leading edges of these surfaces. Regulated bleed air pressure inflates the boots, and vacuum pressure deflates the boots.

Precautions During Icing Conditions During winter a careful pre-flight inspection is required before operating in cold weather, or in potential icing conditions. In addition to the normal exterior inspection, special attention should be paid to areas where frost and ice may accumulate. Even a thin layer of frost or snow can cause great harm. It is not the thickness and weight of the frost that matters, it is it’s texture. A slightly irregular surface can substantially decrease proper airflow. Never underestimate the damaging effects of frost and snow. ALL frost and accumulated snow MUST be removed from the wings, tail surfaces and propellers before the aircraft is flown. De-icing fluid should be used when needed. Fuel drains should be tested for free flow. Water in the fuel system has a tendency to condense more readily during winter months, and left unchecked large amounts of moisture may accumulate in the fuel system. Anti-icing additive is necessary for the BT-67. It is important to add the correct amounts of additive. Higher concentrations do not insure lower fuel freezing temperatures, and too great a concentration can damage the fuel system. The minimum additive concentration is 0.06% by volume, and the maximum concentration 0.15% by volume. The brakes and tyres should be checked prior to taxiing the airplane. If an anti-ice solution is needed to free the brakes, be sure that the solution does not contain oil-based lubricants. If the tyres are frozen to the ground, use undiluted defrosting fluid, or a ground heater to melt the ice, then move the airplane as soon as the tyres are free. Heat applied to the tyres should not exceed 71º Celsius. In addition to preventing unnecessary reduction gearbox wear, using the propeller tie-downs is effective as an ice preventative when the airplane is parked during cold weather. When the


propeller is properly secured, moisture is channelled down the blades, past the propeller hub, and off the lower blade. The propeller hubs should also be inspected for ice and snow accumulation. The pitot-static tubes should be covered while the airplane is parked for extended periods. Once the covers are removed make sure both tubes are free of ice or water. The Basler BT-67 is equipped with both de-icing and anti-icing equipment. However, only the surface de-ice is a true de-icing system. That is, surface de-ice is intended to eliminate ice that has already accumulated. The remaining ice protection systems are considered to be anti-ice systems, and should be used to prevent ice formation. Due to the distortion of the wing airfoil, stalling speeds should be expected to increase as ice accumulates on the airplane. For the same reason stall warning devices are not accurate and should not be relied upon. Always maintain a comfortable margin of airspeed above the normal stall airspeed when ice is on the airplane. Engine inertial separators should be operated anytime potential icing conditions are encountered. Because of ram air effect, engine icing will occur at ambient temperatures slightly above freezing. Even small pieces of ice can damage compressor blades. Engine anti-icing should be used: • before visible moisture is encountered at OAT +5º C and below. • at night when freedom from visible moisture is not assured, and the OAT is +5º C or below. Before entering icing conditions, pitot heat, prop de-ice, inlet-lip de-ice and windshield heat should also all be ON.

Severe Icing Procedures The following weather conditions may be conducive to severe in-flight icing: • visible rain at temperatures below 0º C OAT • droplets that splash or splatter on impact at temperatures below 0º C OAT. Procedures for exiting the severe icing environment: These procedures are applicable to all flight phases from takeoff to landing. Monitor the ambient air temperature. While severe icing may form at temperatures as low as -18º C, increased vigilance is warranted at temperatures around freezing 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: 1 Immediately request priority handling from ATC 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 aircraft has been certificated.

2 Avoid abrupt and excessive manoeuvring that may exacerbate control difficulties. 3 Do not engage the autopilot. 4 If the autopilot is engaged, hold the control wheel firmly and disengage the autopilot. 5 If an unusual roll response, or un-commanded roll control movement is observed, reduce

the angle of attack. 6 Do not extend the 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 de-icing boots.

7 If the flaps are extended do not retract them until the airframe is clear of ice. 8 Report these weather conditions to ATC.


Engine Inertial Separators The inertial separator system is installed in each engine to prevent foreign objects such as dust and ice from entering the engine inlet plenum. During normal flight operations when icing conditions are not present, the intake vane is retracted, and the bypass vane is extended (Figure 13-3). At temperatures above +5º C the inertial separators should be in the NORMAL position, since it is unlikely that ice will form at these temperatures.

Figure 13-3 Inertial Separators in NORMAL Position The inertial separators should be activated when visible moisture is present and the temperature is +5º C or below. Since air temperature decreases as it moves through the inlet towards the engine air intake screen, moisture can enter the engine as water or water vapour, and freeze when it reaches the engine intake screens. As the ice continues to build on the intake screens, it may break off into small pieces which enter the compressor section and may cause severe damage to the compressor blades. Therefore the inertial separators should be activated anytime the OAT reaches +5º C and moisture is present or suspected.

Figure 13-3 Inertial Separators in ICING Position When the inertial separators are in the ICING position (Figure 13-4), a sudden turn is introduced into the air inlet, creating a venturi effect. At the same time the bypass vane in the


lower cowling at the aft end of the air duct opens. As the mixture of air and ice particles or water droplets enters the inlet, it is accelerated by the venturi effect. Due to their greater mass, and therefore greater momentum, the heavier particles accelerate past the screen area and are discharged overboard through the bypass door. The airstream, however, makes the sudden turn free of ice particles and enters the engine through the inlet screen.

Inertial Separator Controls The deflector doors are extended or retracted simultaneously by electric actuators. The actuators are controlled by switches labelled INERTIAL SEP which are located on the pilot’s overhead panel (Figure 13-5).

Figure 13-5 Inertial Separator Controls

When the inertial separators are extended blue L/H and R/H annunciator lights will be illuminated. Since inlet airflow is restricted by the doors, torque will decrease proportionate to the power setting, and ITT will be increased slightly. When the inertial separators are retracted, the annunciators will be extinguished, torque will increase and ITT will decrease. The inertial separators cannot be extended to intermediate positions. They are either extended or retracted.

Engine Inlet-Lip Heaters The engine inlet lip is protected by an electrically heated rubber boot. A thermostatic sensor and a power relay control the boot temperature. The selector switches and blue INLET LIP L/H and R/H indicator lights are located on the pilot’s overhead panel (Figure 13-6). When activated the indicator light cycles in conjunction with the thermostatically controlled power relay.

Figure 13-6 Inlet Lip De-Ice Controls

In the event that the thermostatic sensor fails, the boot will continue to function, due to the presence of a standby thermostatic switch. Failure of the thermostatic sensor will be indicated by the illumination of the appropriate amber INLET LIP DE-ICE annunciator light on the main annunciator panel. The system is automatically deactivated below 57 KIAS by an airspeed switch, thus preventing boot overheating due to inadvertent use on the ground.


Inlet-Lip Heater Ground Control To allow ground testing of the inlet-lip de-ice system, a ground test switch is provided on the cockpit door upper panel (Figure 13-7). This test switch bypasses the ground/ flight relay to allow power to be supplied to the inlet-lip de-ice boots on the ground. The inlet-lip de-ice test switch is provided for ground operation ONLY. Do not operate the inlet-lip de-ice systems for more than 10 seconds with the aircraft on the ground to prevent damage to the de-ice boots.

For normal indication, the blue GRD/FLT RELAY light will be illuminated with battery or GPU power ON. Place the test switch to TEST to check that the light extinguishes. Return the test switch to NORMAL and confirm that the light illuminates. With at least 1 engine running and 1 generator ON, turn ON the left and right inlet-lip de-ice switches. Place the test switch into the TEST position, and note that the blue L/H and R/H annunciator lights illuminate. Return the test switch to the NORMAL position, and confirm that the L/H and R/H annunciates extinguish. Turn OFF both inlet lip switches. Figure 13-7 Inlet Lip De-ice Ground Test

Switch


Engine Continuous Ignition The engine continuous ignition system provides for automatic continuous ignition to prevent engine power loss due to combustion failure. Continuous ignition is activated during icing flight, or when flying into turbulence or heavy precipitation.

Figure 13-8 Ignition Controls

Control switches for the ignition are located on the pilot’s overhead panel. The system is activated by moving the switches to the CONT position (Figure 13-8). When the ignition is activated the white IGN ON annunciator light will be illuminated.

Fuel System Anti-Ice Moisture in the fuel can freeze, and fuel can thicken during flight in extremely cold temperatures. Fuel temperature in the FCU is maintained by an oil-to-fuel heat exchanger, mounted on the engine’s accessory section (Figure 13-9). An engine oil line within the heat exchanger is located next to the fuel line. Heat transfer occurs through conduction between these two lines, before fuel is delivered to the FCU. The heat exchanger melts ice particles and prevents the fuel from thickening in extremely cold temperatures. The heat exchangers operate automatically whenever the engines are running.

Figure 13-9 Oil-to-Fuel Heat Exchanger


Propeller De-Ice The propeller electric de-icer system consists of electrically heated de-ice boots, slip rings and brush block assemblies, a timer for automatic operation, and two ammeters. Two switches on the pilot’s overhead panel, labelled PROPELLER DE-ICE, control the system (Figure 13-10).

Figure 13-10 Propeller De-ice Controls

Although propeller de-ice is capable of removing ice from the propeller after it has accumulated, the system is normally used as an anti-icing system and it should be turned on before entering icing conditions. The heated boots reduce ice adhesion on the propeller blades. The ice is then removed by the centrifugal effect of the propellers and the blast of the airstream. The propeller de-ice system is controlled by two separate PROPELLER DE-ICE switches labelled ON, OFF and STBY ON. For normal operations each switch is placed in the ON position. Normal operation of the system will be indicated by a reading of 26-30 amps for approximately 1 minutes 30 seconds. In the event of a system failure, indicated by a zero reading on the ammeter, the selector switch should be placed in the STBY ON position. This selects a standby controller in the de-ice timer. Propeller de-ice must not be operated when the propellers are static, to avoid damage to the brush blocks and slip rings.

Windshield Anti-Ice The pilot’s and co-pilot’s windshields are each heated independently, however windshield heat can only be selected for both windshields simultaneously through a single WSHLD HEAT switch mounted on the pilot’s overhead panel. When power is applied to the windshields a blue W/S annunciator light will be lit (Figure 13-11).

Figure 13-11 Windshield Heat Controls

Both windshields are composed of three layers. Both inside and outside layers are thick glass. The middle layer is a polyvinyl sheet which contains the gold-filament fine wire heating grids. Windshield heating elements are connected through terminal blocks on the corner of


the glass to the control switch, and circuit breakers on the RH Bus. Windshield temperature is automatically controlled by a temperature sensing element embedded in each windshield and a temperature controller in each windshield circuit. Under normal use the blue W/S annunciator light will be seen to cycle on and off, and a small degree of optical distortion may be apparent in the windshields when heating is applied.

Windshield Wipers Separate windshield wipers are mounted on the pilot’s and co-pilot’s windshields. The wipers are driven by independent hydraulic actuators mounted below the windshield (Figure 13-12).

Figure 13-12 Wipers and Wiper Controls

The windshield wiper controls are mounted on either side of the control pedestal. The speed of the wipers can be adjusted to any setting by opening or closing the hydraulic needle-type control valves to regulate the hydraulic flow to the wiper actuating mechanisms. The wipers may be used either on the ground or in flight, as required. To protect the windshield from scratching the wipers must not be operated on a dry windshield.

Pitot-Static Heat The pitot-static tubes are protected from ice by an electric heater built into each tube. The heater elements are controlled by switches located on the pilot’s overhead panel. Individual amber PILOT and CO-PILOT PITOT HEAT annunciator lights, located on the overhead annunciator panel, will illuminate any time the heaters are switched OFF, or when there is a failure of a heating element (Figure 13-13).


Figure 13-13 Pitot Heat Controls

The pitot heat system should not be operated on the ground, except for testing or for short intervals to remove snow or ice from the tube.

Stall Warning Vane Heat The stall-warning vane is electrically heated and is controlled by the pilot’s pitot heat switch, but is protected by a separate circuit breaker on the ESSENTIAL bus.

Surface De-Ice Systems The wings, horizontal and vertical stabilisers are de-iced in flight with a system of inflatable rubber boots attached to the leading edges of these surfaces. After ice has accumulated, pneumatic pressure can be cycled through a de-ice distributor valve to inflate the boots in sequence. Ice is removed by alternately inflating and deflating the de-ice boots. The surface de-icing system is controlled by a switch labelled SURFACE DE-ICE on the pilot’s overhead panel. A blue WING annunciator indicates operation of the system (Figure 13-14), and a de-icing system pressure gauge, mounted on the co-pilot’s main instrument panel, indicates 8 psi at the peak pressure of each inflation cycle.

Figure 13-14 Surface De-Ice Controls

Placing the SURFACE DE-ICE switch into the ON position provides electrical power to the distributor valve in the wing centre section. P2 bleed air is cycled through the distributor valve to the leading edge boots in the following sequence: • centre outboard wing boot • upper and lower outboard wing boots • centre inboard wing boots • upper and lower inboard wing boots • all stabiliser boots After each boot has been inflated, the pressure air is routed to the next boot by the distributor valve, and the original boot deflates. The exhausted air is expelled overboard through a fitting at the bottom of the wing. The sequence will be completed in approximately 40 seconds. For most effective de-icing, at least 1 to 1½ inches of ice should be allowed to form before attempting ice removal. Very thin ice may crack and cling to the boots instead of shedding when the boots are inflated. Subsequent cyclings will then have a tendency to build up a shell of ice outside the contour of the leading edges of the boots, making ice removal efforts ineffective.


Wing Inspection Lights During night flight, the wing inspection lights can be used as necessary to check for wing ice accumulation. The wing inspection lights should not be used for prolonged periods on the ground.

Figure 13-15 Wing Inspection Lights Control

The wing inspection lights are on the outboard side of each nacelle. The control switches are labelled WING INSP, and are mounted on the pilot’s overhead panel (Figure 13-15).

Alternate Static Air System The alternate static air source is used any time the normal static air source is obstructed. The alternate static air switch is a guarded toggle switch on the co-pilot’s panel (Figure 13-16).

Figure 13-16 Alternate Static Selector

When the airplane has been exposed to moisture and/ or icing conditions (especially on the ground), the possibility of obstructed pitot-static tubes should be considered. Partial obstructions will result in the rate-of-climb indication being sluggish during a climb or descent, inaccurate airspeed indications, and incorrect altimeter indications. A suspected obstruction is verified by switching to the alternate system and noting a sudden, sustained change in the rate of climb. When using the alternate system, the AFM should be consulted for the corrections to the airspeed and altimeter indications. In general, whenever the alternate system is selected, the aircraft is actually lower and slower than indicated by the aircraft’s flight instruments.


PERFORMANCE

Table of Contents Introduction ....................................... 14-1 Weight and Balance Procedure ......... 14-1 Weight and Balance Example............ 14-2

Cargo Loading Considerations ...........14-5 Using Performance Charts .................14-6

Introduction The purpose of the Performance section of the AFM is to present the performance data necessary for pre-flight and flight planning. Airplane performance may vary according to pilot technique. However, these variations can be minimised by establishing standard operating procedures. All FAA-approved performance data is included in the AFM, and supersedes any information contained in this section.

Weight and Balance Procedure Complying with the weight and balance limitations is the first step in flight planning. Representative loading situations for the BT-67 have been used in this section. As with all other performance data, only the weight and balance data in the AFM will be accurate for that particular airplane. Basic Empty Weight is the weight of an empty airplane, including permanently installed equipment, fixed ballast, full hydraulic fluid and full oil. Because of fuel system design, a certain portion of fuel is not available to the engines. Only this quantity of unusable fuel is included in the Basic Empty Weight. Note that the Basic Empty Weight is the configuration from which all loading data is completed (Figure 14-1). Basic Operating Weight includes everything that is loaded onto the airplane in preparation for the flight. This consists of flight and cabin crews, catering, emergency equipment and any other removable aircraft equipment items not included in the Basic Empty Weight. Useful Load includes everything that is loaded aboard the airplane in preparation for a particular flight. Useful Load consists of flight crew, passengers, usable fuel baggage, and all other items loaded onto the airplane. Basic Empty Weight plus Useful Load equals Ramp Weight. Ramp Weight is defined as airplane weight at engine start-up, after all loading is completed.


DC3-TP67 FLIGHT MANUAL

CHAPTER 6 WEIGHT AND BALANCE MANUAL

AIRCRAFT WEIGHING FORM DATE WEIGHED 6/26/92 REGISTRATION NO. A2-ADL SERIAL NO. 33581 . PLACE WEIGHED Oshkosh WEIGHING OFFICER Lee Broderson .

WEIGHING

POINT SCALE

READING TARE

NET

WEIGHT ARM

MOMENT

LEFT MAIN 8289

8289

RIGHT MAIN 8185

8185

SUB-TOTAL (both mains)

16474 222.5 3665465

TAIL 1438

1438 583.8 839504.4

TOTAL (as weighed)

17912

4504969.4

TOTAL OF ITEMS WEIGHED BUT NOT PART OF EMPTY WEIGHT (from column 1 below)

-1467.3 240.5 - 352885.6 TOTAL OF BASIC ITEMS NOT IN AIRPLANE WHEN WEIGHED (from column 2 below)

+ + AIRPLANE EMPTY WEIGHT 16444.7 252.49 4152083.8

COLUMN 1 COLUMN 2

Items Weighed But Not Part of Basic Wt.

Weight Arm Moment Basic Items Not In Air- plane When Weighed

Weight Arm Moment

Main Fuel 1467.3 240.5 352885.6

REMARKS __________________________________________________________________________________________________________________________________________________________________________________________

Figure 14-1 Basic Empty Weight and Balance Form Exceeding maximum takeoff weight results in longer takeoff runs, reduced rate-of-climb and increased stall speeds. High ambient temperatures, high humidity and high field elevation deteriorate performance even further. Once airborne, the excess weight also results in higher fuel consumption and lower service ceilings. Loading the airplane forward of the centre-of-gravity limit requires higher than normal aft control pressures during takeoff and climb. Loading aft of the rear limit causes poor stall characteristics, greater aircraft instability in flight, and heavier than normal forward control pressures to maintain level flight. In either case, elevator trim may not be able to compensate for an extremely unbalanced condition.

Weight and Balance Example A sample weight and balance is shown using the standard format for the Basler BT-67. Representative empty weight data has been used. Consult the AFM for data specific to the actual aircraft. In this example, the aircraft will carry two pilots, 5,000 lbs of cargo, and sufficient fuel to fly to the destination with reserves. The data has been entered as follows:


• Airplane Empty Weight is obtained from the AFM, as shown in Figure 14-1. Enter Airplane Empty Weight on the Weight and Balance Loading Form (Figure 14-2).

• Distribute the freight amongst the cargo compartments, taking care not to exceed either the compartment capacities or the maximum floor loadings (Figure 14-3).

• Add weights and moments for the crew and crew baggage to obtain the Zero Fuel Weight. • After determining that the aircraft is loaded within the Zero Fuel Weight limits, fuel weight

and moment can be added to the loading form to obtain the Ramp Weight. To obtain the fuel moment, use the Fuel Loading Schedule table (Figure 14-4). Subtract the taxi and run-up fuel to arrive at the Takeoff Weight.

• The centre of gravity of the airplane can be determined as follows. Divide total weight by total moment to obtain the arm. If this number is within the Centre of Gravity Envelope (Figure 14-5) the aircraft is properly loaded.

• To determine the Landing Weight, compute the weight and moment of the fuel remaining at landing, and add it to the Zero Fuel Weight.

REGISTRATION

Max Zero Fuel Wt 242.35"

Max Takeoff Wt 263.10"

PASSENGERS

CREW BAGGAGEFLIGHT CREW

FREIGHT (SEE BELOW)

16,444.7

30

WEIGHT

AIRCRAFT EMPTY WEIGHT340

DATE____/____/______

MOMENT/1000ARM

4,152.0838

7,198.484

30"70"

10.2000

611.3000

259.0

252

768.02,491.65,127.4

COMPARTMENT E 436"

560.000

COMPARTMENT D 2,000 358" 716.000

lb @ 0.80 SG

2,406.42,367.2

379.4372.2

4,871.1

2,533.0

27,791.70

2.1000

5,000 296 1,478.0000

21,814.7 258.7 5,642.3838

Usable Fuel Capacities US Gallons Litres

26,200 lb Max Forw'd CoGMax Rearw'd CoG

lb @ 0.76 SG

28%

241"2,542 MAIN TANK FUEL

259.0 7,215.4838

INBOARD AUX TANK FUEL 2,500 276" 690.0000

280"

120"202"

2x Main Tanks2x Inb'd Aux Tanks

1,436.2

27,861.7

28,750 lb

TAXI AND RUN-UP FUEL BURN

TAKEOFF WEIGHT

MAIN TANK FUEL AT LANDING 1,005.0

ZERO FUEL WEIGHT

1,005 271" 271.8000

-70 -17.000

13%

OUTBOARD AUX FUEL AT LANDING 271"

LANDING WEIGHT 22,819.7 257.9 5,884.0838

2x Outb'd Aux Tanks

DC3-TP67 A2-ADL

COMPARTMENT C

COMPARTMENT A

RAMP WEIGHT

OUTBOARD AUX FUEL

ZERO FUEL WEIGHT

1,412.72,907.2

241" 241.7000

INBOARD AUX TANK FUEL AT LANDING 276"

21,814.7 259 5,642.3838

COMPARTMENT F 506"

COMPARTMENT B 1,000 202.000

TOTAL FREIGHT 5,000 295.6 1,478.000

COMPARTMENT G 560"

2,000

Figure 14-2 Weight and Balance Loading Form


FUSELAGE STATION

77

120

A

2330

COMP.

COMPART- MENT

CAPACITY (POUNDS)

FLOOR AREA

(SQ.FT.)

MAXIMUM DEMONSTRATED

FLOOR LOAD (LBS./SQ.FT.)

MAXIMUM CONCENTRATED*

FLOOR LOAD (LBS./SQ.IN.)

163

A 2330 45.0 150 66

202

B 2330

B 2330 46.5 150 66

241

C 2330 48.0 150 66

280

C 2330

D 2330 48.0 150 66

319

E 2330 43.5 150 66

358

D 2330

F 700 31.0 150 66

397

G 550 18.0 150 - - - -

436

E 2330

* This refers to concentrated loads such as heavy equipment stands or 475 * racks. Such loads must be located within six (6) inches of a cabin floor

* beam and cover a maximum surface area of 12 sq. inches.

506

F 700

538

560

G 550

583

Figure 14-3 Cargo Loading Schedule GALLONS

POUNDS

MOMENT x 1000 MAIN TANK STA. 240.5

MOMENT x 1000 INBD.AUX. STA. 276.0

MOMENT x 1000 WING AUX. STA. 270.5

25 167.5 40.3 46.2 45.3 50 335 80.6 92.5 90.6 75 502.5 120.8 138.7 135.9

100 670 161.1 184.9 181.2 125 837.5 201.4 231.1 226.5 150 1005 241.7 277.4 271.8 175 1172.5 282.0 323.6 317.2 200 1340 322.3 369.8 362.5 250 1675 402.8 462.3 453.1 300 2010 483.4 554.8 543.7 373 2500 601.2 FULL AUXS 690.0 676.2 379 2542 FULL MAINS 611.3 687.6 400 2680 724.9 450 3015 815.5 500 3350 906.2 550 3685 996.8 600 4020 1087.4 650 4355 1178.0 700 4690 1268.6 750 5025 1359.3 768 5146 FULL AUXS 1392.0

Note: Fuel Weight based on 6.7 lbs./gallon. Figure 14-4 Fuel Loading Schedule


Figure 14-5 Centre of Gravity Envelope

Cargo Loading Considerations The method of loading cargo, its placement in the airplane and the method of restraint should be determined before starting the actual loading. For loads that are evenly distributed in a given section (“A” through “G”), the Cargo Loading Schedule (Figure 14-3) should be used. For any load that cannot be located at the centroid of a compartment, or that extends over more than one compartment, it will be necessary to determine its own C.O.G. and its location in the airplane. Determine the C.O.G. arm (fuselage station) by measuring from a known location in the cabin to the C.O.G. of the load. Determine the moment for the load by multiplying the weight by the arm. This result should be divided by 1000 to be compatible with the other loading data. Once weight and balance have been considered, the performance graphs can be used to plan the flight. Refer to the AFM for specific flight planning details and examples.


Using Performance Charts Information in this section is presented for the purpose of compliance with the appropriate performance criteria and certification requirements of FAR Part 25. The maximum operating weights are limited by the following performance graphs or criteria from the Basler BT-67 AFM, and compliance therewith is mandatory. Maximum take-off weight to achieve Takeoff Climb Weight Limit graph, 5-22 Maximum take-off weight as limited by the Takeoff Brake Energy Weight Limit graph, 5-23 Takeoff Distance Required graphs 5-25, 5-27 and 5-28 to 5-32 Maximum landing weight to achieve Landing Climb Weight Limit graph, 5-35 Maximum landing weight as limited by Brake Energy Landing Weight Limit graph, 5-36 Landing Distance Over 50-Foot Obstacle graph, 5-40 Performance of the Basler BT-67 must meet the specifications under which the aircraft was certified. The BT-67 is certified as a Transport Category Airplane under FAR Part 25. The original DC3 certification is under CAR 4b. Refer to Table 14-1 for a comparison of certification requirements for aircraft performance, and what the requirements actually equate to in KIAS at ISA, sea-level conditions at MAUW.

Table 14-1 Certification Requirements Comparison

CAR 4b Airplane Airworthiness; Transport Categories

FAR Part 25 Airworthiness Standards –

Transport Category Airplanes Distance required to accelerate to V1, then come to a complete stop

Longest of: Accelerate/Go to 35 feet 115% of all-engine distance to 35ft

Takeoff Runway

Information not available 4500 ft Climb at 50ft/min at V2 Positive climb at VR, out of ground

effect V1 > VMCG

First Segment Takeoff, Gear Down 0 – 35 feet

50 ft/min at 90 kts Positive climb at 83 kts Climb at a rate of 0.035 VS1

2 at V2

Climb gradient > 2.4% at V2 Second Segment Takeoff, Gear Up 35 – 400 feet

167 ft/min at 90 kts 218 ft/min at 90 kts Climb at a rate of 0.02 VS0

2 at 5000ft standard altitude

Climb gradient > 1.2% at 1.25 VS Third Segment Enroute 400 – 1500 feet

79 ft/min at 5000ft altitude 92 ft/min at 86 kts Climb at a rate of 0.04 VS0

2

Climb gradient > 2.1% at < 1.5 VS Approach Climb Single Engine

159 ft/min 221 ft/min at 104 kts Climb at a rate of 0.07 VS0

2

Climb gradient > 3.2% at 1.3 VS Landing Climb Both Engines

278 ft/min 293 ft/min at 90 kts Distance over 50 feet at 1.3 VS0

Actual distance over 50 feet (Times 1.67 for FAR121 Ops)

Landing Distance

3,350 ft at 90 kts 3,350 ft


Figure 14-6 Takeoff Path Profile FAR Part 25

Table 14-2 Configurations

Operating Engines

Thrust Flap Setting Landing Gear

1st Segment Takeoff Climb 1 Takeoff 0º Down 2nd Segment Takeoff Climb 1 Takeoff 0º Up Final Segment and Enroute Climb

1 Maximum Continuous

0º Up

Approach Climb 1 Takeoff 0º Up Landing Climb 2 Takeoff 45º Up Landing 1 @ 45º Down @ One propeller feathered, the other engine Ground Idle after touchdown. Performance values are determined by using the performance charts found in the AFM. Be sure that you are using the correct chart for specific aircraft configurations. The following are guidelines in performance calculations: • In addition to presenting the answer for a particular set of conditions, the example on the

graph also presents the order in which the various scales on the graph should be used. For example if the first item in the example is OAT, then enter the graph at the existing OAT.

• The reference lines indicate where to begin to follow guidelines. Always project to the reference line first, then follow the guidelines to the next know item.

• The associated conditions define the specific conditions from which performance parameters have been determined. They are not intended to be used as instructions; however, performance values determined from charts can only be achieved if specified conditions exist.

• Indicated airspeeds (IAS) were obtained using the Calibrations – Normal Airspeed System graph.

• The full amount of usable fuel is available for all approved flight conditions. • Notes have been provided on various graphs and tables to approximate performance with

the inertial separators in the ICING position. The effect will vary, depending on airspeed, temperature, altitude, and ambient conditions. At lower altitudes, where operation up to the torque limit is possible, the effect of ice vane extension will be less, depending upon how much power can be recovered.


NOTES:


SPECTREM AEM SURVEY SYSTEM

Table of Contents Introduction ....................................... 15-1 Ferry Configuration............................ 15-3 Survey Configuration ......................... 15-3 Spectrem Aircraft Limitations ............. 15-3 Internally Installed Survey Equipment 15-4 External Towed Birds......................... 15-4 Bird Normal Procedures .................... 15-5 Bird Cable Wrap Procedure ............... 15-5 Bird Emergency Release ................... 15-6

EM Winch Manual Procedure .............15-6 APU ...................................................15-7 APU Cockpit Procedures....................15-7 Cabin Smoke Detection System .........15-8 Survey GPS .......................................15-8 Crew Call ...........................................15-9 Flight Following ..................................15-9 Revised AC system............................15-9

Introduction The basic Basler BT-67 aircraft has been specially modified (Figure 15-1) to allow the Spectrem AEM system to collect geophysical electromagnetic (EM) and magnetic (Mag) data. These changes have altered the appearance, performance and installed systems of the basic BT-67.

Figure 15-1 Spectrem Airborne EM System in Flight


Figure 15-2 Spectrem AEM System 3 View

Figure 15-3 Spectrem AEM Turning Circle


Ferry Configuration Visual external differences: • Nose and Tail Stinger extensions to the fuselage • Wing-tip Pods with a nose fairing blanking off the forward intake hole • A Camera Blister attached to the underside of the fuselage • An Exhaust Stack and Deflector Plate attached to the port side of the fuselage • A Window Cavity blanked off and fitted with intake holes on the port side to accommodate

the APU intake requirements • A Window Cavity blanked off and fitted with an exhaust scoop on the starboard side of the

fuselage Internal differences: • An APU installed inside the cabin on the port side of the aisle • A Winch Mechanism installed behind the APU • An Electronics Rack installed behind the winch mechanism • A Transmitter and Choke Rack installed on the forward starboard side of the aisle. • A Transformer installed behind the transmitter rack • Various Trap Doors installed in the cabin floor to facilitate access to equipment installed

below the floor

Survey Configuration Visual external differences: • Nose and Tail Stinger extensions to the fuselage • Wing-tip Pods • A Five-strand Cable Loop spanned between the wing-tip pods and the nose and tail

stingers • A Camera Blister attached to the underside of the fuselage • Space Frame Cradles, attached to the underside of the fuselage, housing the EM and Mag

birds • An Exhaust Stack and Deflector Plate attached to the port side of the fuselage • A Window Cavity blanked off and fitted with intake holes on the port side to accommodate

the APU intake requirements • A Window Cavity blanked off and fitted with an exhaust scoop on the starboard side of the

fuselage Internal differences: • An APU installed inside the cabin on the port side of the aisle • A Winch Mechanism installed behind the APU • An Electronics Rack installed behind the winch mechanism • A Transmitter and Choke Rack installed on the forward starboard side of the aisle. • A Transformer installed behind the transmitter rack • Various Trap Doors installed in the cabin floor to facilitate access to equipment installed

below the floor

Spectrem Aircraft Limitations In addition to the normal Basler BT-67 limitations the following now apply: • It is recommended that VMCA = 75 KIAS For the Spectrem AEM Survey configuration all normal aircraft limitations apply with the exception of the following: • Avoid side-slip manoeuvres • Never Exceed Speed with Birds Docked VNE = 150 KIAS • Essential flight crew and technical observers only, no passengers • Avoid flight in any icing conditions


When launching, towing, or retrieving the EM and Mag birds the following limitations apply: • Day VFR conditions only • Maximum speed during launching and retrieval of the birds is 120 KIAS • Never Exceed Speed with birds in tow is 130 KIAS • Minimum speed during launching, towing, or retrieval of the birds is 90 KIAS • Survey operations requiring the launching, towing or retrieval of the birds are not to be

conducted if the intercom between the pilot and the winch operator is unserviceable or not functioning

• Avoid retrieving birds in turbulence

Internally Installed Survey Equipment

Figure 15-4 Spectrem Operator’s Rack, APU, EM Bird Winch, EM Transmitter

External Towed Birds The Spectrem AEM system includes two different towed sensor birds for data retrieval. Mounted on the wing centre section, offset to starboard is the high drag EM bird, and towards the rear fuselage, offset to port is the smaller, low-drag Mag bird (Figure 15-5).

Figure 15-5 Birds Docked


Once fully deployed the two birds occupy very different positions (Figure 15-5). The high-drag EM bird takes a position of approximately 100 ft (31m) below the aircraft, and 300 ft (92m) behind. The low-drag Mag bird flies in a position 90 ft (28m) below the aircraft, and 60 ft (19m) behind.

Figure 15-6 Birds In Flight

Bird Normal Procedures The Spectrem operator is solely responsible for the launching and retrieving of the birds. The pilots’ responsibility is to ensure that the aircraft is flown in such a manner as to safely and expeditiously ensure bird deployment and docking. For all bird launching and retrieving the aircraft must be positioned in the smoothest air conditions possible. This normally entails climbing to at least 3,000 ft above ground, or above the condensation level. The aircraft must be flown at normal survey speed, 115 KIAS, with control inputs and turbulence movements kept to the absolute minimum. In normal flight the two bird cables appear to cross, but should not come close to touching. Occasionally turbulence or abrupt flight manoeuvres result in the birds “crossing over”. The following procedures apply: • Cease survey operations, and climb the aircraft to a safe height, and smooth air. • Slow the aircraft to 95 knots. • The Spectrem operator will manoeuvre the birds to pull the Mag bird over the EM cable. • If the Mag bird hooks to the EM cable it will be necessary to yaw the aircraft. • Once the bird cables are uncrossed return to normal survey flight. Occasionally during retrieval of the EM bird, the Spectrem operator will determine that the EM tow cable has jumped over the EM docking pins on the cradle. The simplest procedure to rectify this is as follows: • Climb the aircraft to a safe height, and smooth air. • Slow the aircraft to 95 knots. • Deploy full flaps. • Under the direction of the Spectrem operator apply left or right rudder. • Once the bird cables are uncrossed return to normal flight.

Bird Cable Wrap Procedure In the event that a cable “cross over” is unrecognised and a second cross occurs the following procedures are to be employed to attempt to successfully salvage the birds: • This is not an easy procedure and cannot be guaranteed to work. Even in a best case

scenario 20 minutes can be expected to be spent untangling the cables, some wraps have


taken up to 2 hours, and others have not been able to be successfully resolved. As such all crew members must be aware of the aircraft’s fuel state, the vicinity of alternate airfields, and any other flight planning considerations.

• If possible, the junior pilot should be left at the controls of the aircraft, while the senior pilot should assist the Spectrem operator.

• Route the aircraft towards the airport, planning this so as not to be flying over populated areas.

• Climb the aircraft to a safe height, and smooth air. • Accelerate the aircraft to 130 knots. • The Spectrem operator will man the EM winch, while the non-flying pilot will operate the

Mag winch under instruction of the operator. • Once the cable knot is close to the aircraft, use the wrap-hook to hook the Mag cable. It

may be necessary to slow the plane down, and to pitch the aircraft to catch the cable. • Pull the Mag cable into the cabin, and secure the cable and Mag bird with the emergency

rope. • Cut the Mag cable. • Normally the Mag cable and emergency rope will unwind itself from the EM cable. The EM

bird can then be docked as normal. • Ensure the Mag bird is secured before landing. • In the event that the above procedures have not been successful and the birds cannot be

secured before the aircraft reaches a critical fuel situation, it will be necessary to jettison the EM bird.

Bird Emergency Release In the event that the EM bird must be cut free from the aircraft, a guarded electrical cutter switch is mounted on the co-pilot’s circuit breaker panel (Figure 15-7). This switch electrically activates a guillotine-type cable cutter mounted on the EM winch, severing the EM bird’s cable. The cutter solenoid is hotwired to the aircraft battery via a circuit breaker, and after use must be returned to the OFF position to prevent damage to the solenoid.

Figure 15-7 EM Bird Cutter Switch

During the critical take-off phase of flight the Spectrem operator removes the EM bird safety pin, which is normally installed to prevent inadvertent activation of the EM cable cutter. In the event of an aircraft engine failure immediately after take-off, it may be necessary to remove the extra drag induced by the EM bird to enable the aircraft to be able to safely return to the airport for landing. At other times, for example an extreme bird cable wrap, it may not be possible to retrieve the birds. In these cases close liaison with the Spectrem operator may result in the need for the pilot’s to jettison the EM bird. In the event of a bird jettison, the safety of people and property on the ground must be considered, as well as the possibility of dropping the bird in a location for easy ground recovery.

EM Winch Manual Procedure In the event that the EM winch has a lose of electrical power it will be necessary to manually winch in the high-drag EM bird. The pilot’s will be required to assist the Spectrem operator as


follows: • Climb the aircraft to a safe height, and smooth air. • Slow the aircraft to 95 knots. • Deploy full flaps. • Under the direction of the Spectrem operator apply left or right rudder. • One pilot must move to the EM winch to assist the Spectrem operator by operating the

hand brake. • Once the EM bird has been safely docked resume normal flight.

APU To provide the required electrical power for the Spectrem AEM system, a Turbomach T-62T-40C auxiliary power unit is installed in the forward cabin on the port side of the aisle. The APU and the electronics rack are totally independent of the aircraft’s electrical system, and rely on their own battery for starting. Starting and operational controls for the APU are mounted on the electronics rack, and are controlled by the Spectrem equipment operator both in flight and on the ground. During normal operations the pilots control only the fuel tank selector for the APU. Cockpit controls and indicators for the APU are mounted on the overhead panel (Figure 15-8), and consist of an green APU master ON light, fuel tank selector, APU fire detection light and shutdown button, APU fire extinguisher switch, APU shutdown button, and also an APU fire audible warning.

Figure 15-8 Spectrem APU Controls

In the event of various abnormal situations being detected by the APU’s internal sensors, including APU fire, the APU will automatically shut down.

APU Cockpit Procedures Under normal conditions, the pilots will not be responsible for performing any duties with regards to starting, controlling or shutting down the APU. The pilots’ sole duty will be selecting from which main fuel tank the APU draws fuel. The APU draws fuel from either of the aircraft’s two main fuel tanks, through the selection of the FUEL SELECTOR switch. If it becomes necessary to change the fuel tank selected for the APU, the fuel selector switch must be rapidly and smoothly moved between the tank positions to ensure that the APU fuel pump does not cavitate and cause an APU flameout. In the event that an APU shutdown is required of the pilots, shutdown is initiated by pressing the PUSH TO STOP button. Once the operator confirms that the APU has stopped, the fuel selector should be placed in the OFF position. In the event that the APU FIRE warning light illuminates and the APU fire warning sounds both pilots and Spectrem equipment operators must attempt to shut down the APU and extinguish the fire. The following procedures must be followed: • The APU’s automatic shutdown protection will shutdown the APU. • Open guard and select the APU FIRE EXT switch to ON.


• Press APU STOP switch. • Select APU FUEL SELECTOR to OFF. • Verify APU FIRE warning light extinguished. • If the APU FIRE light remains illuminated, follow standard emergency procedures to

evacuate the aircraft. In the event of a fire elsewhere in the aircraft, or in preparation for a crash landing or ditching, ensure the APU is shut down.

Cabin Smoke Detection System Two smoke detector units are installed inside the fuselage, one in the main cabin, and one in the cockpit. When triggered the smoke detection warning tone is generated through the cockpit warning horn. The warning horn may be silenced by pressing the push button located on the overhead Fire Protection control panel. The two cabin smoke detectors can each be separately tested by the pilots by pressing the SMOKE DET test switch on the overhead panel to either the FWD (cockpit) or AFT (cabin) positions.

Survey GPS For navigation of the Spectrem AEM system during data gathering survey flights, the aircraft’s real-time GPS positioning and navigational commands are provided by an a Millennium 2000 AT96 GPS and an OmniStar satellite differential GPS mounted on the electronics rack, with a cockpit display and control panel mounted in front of the windshield (Figure 15-9).

Figure 15-9 Spectrem Cockpit GPS

The GPS is powered solely from the APU electrical system. During survey flight the Spectrem operator will have loaded planned survey flight lines into the computer controlling the GPS. The GPS display is able to show both a navigation data display, and a map overview display. Cycling between the navigation display and an overview of the survey lines is available by pressing the CDI MAP button. The overview is able to be zoomed in and out through pressing the ZOOM+/- buttons while in the Map mode.


While in the Navigation mode, the pilots are able to select different survey blocks, if applicable, by pressing the AREA+/- buttons. Flight line numbers can be individually increased by pressing the LINE+/- buttons, and increased by tens of lines by pressing the ZOOM+/- buttons. In the event that the GPS keypad in the cockpit fails, GPS keypad functions can be executed by the Spectrem operator from the EM rack.

Crew Call Two Crew Call buttons are installed in the cockpit, one on each windshield fairing (Figure 15-10). These buttons activate a horn installed at the Spectrem operator’s station, used to alert the operator that he is required on the intercom. The crew call system is hotwired to the aircraft’s battery through the cabin lights circuit breaker, and will operate with the aircraft’s battery master OFF.

Figure 15-10 Crew Call Buttons

Flight Following The Spectrem AEM system aircraft is fitted with an automatic flight following system that supplies position information approximately every 10 minutes to satellites worldwide. This system functions automatically whenever the aircraft’s battery master is ON.

Figure 15-11 Flight Following Emergency Switch

A guarded emergency switch is located below co-pilot’s circuit breaker panel (Figure 15-11). Activating this switch increases the update rate to every 3 minutes. This increased transmission rate is detected by the flight following system processing centre, and is relayed to the Spectrem emergency contact.

Revised AC System The Spectrem AEM system aircraft now no longer retains any functioning AC electrical equipment.


NOTES:

basler bt-67 systems training manual

Documents