Ayres S2R-T Turbo Thrush (2005)

THRUSH AIRCRAFT INC – MODEL S2R-T34 TURBO THRUSH AIRCRAFT MAINTENANCE MANUAL

THRUSH AIRCRAFT INC. TURBO THRUSH

AIRCRAFT MAINTENANCE MANUAL

SINGLE COCKPIT AND DUAL COCKPIT

Model S2R – T34

Serial Numbers T34 – 273 & Up Manual Number: T34-2

Issued January 26, 2005

Note: All serial numbers with the DC suffix indicate the dual cockpit configuration.

Manufacturer’s Serial Number:_____________________________________________ Registration Number:____________________________________________________ Thrush Aircraft Inc. P. O. Box 3149 300 Old Pretoria Road Albany, GA 31706 Telephone: 229-883-1440 Fax: 229-436-4856

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LOG OF PAGES

Page Date Page Date i ……………………….. 1/26/05 ii ……………………….. 1/26/05

SECTION (CONT’D) SERVICING

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39 ……………………….. 1/26/05 SECTION I GENERAL INFORMATION 40 ……………………….. 1/26/05

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10 ……………………….. 1/26/05 50 ……………………….. 1/26/05 11 ……………………….. 1/26/05 51 ……………………….. 1/26/05

52 ……………………….. 1/26/05 SECTION II SERVICING 53 ……………………….. 1/26/05

1 ……………………….. 1/26/05 2 ……………………….. 1/26/05

SECTION III HYDRAULICS

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SECTION IV POWERPLANT &

PROPELLER 8 ……………………….. 1/26/05 1 ……………………….. 1/26/05 9 ……………………….. 1/26/05 2 ……………………….. 1/26/05

10 ……………………….. 1/26/05 3 ……………………….. 1/26/05 11 ……………………….. 1/26/05 4 ……………………….. 1/26/05 12 ……………………….. 1/26/05 5 ……………………….. 1/26/05 13 ……………………….. 1/26/05 6 ……………………….. 1/26/05 14 ……………………….. 1/26/05 7 ……………………….. 1/26/05 15 ……………………….. 1/26/05 8 ……………………….. 1/26/05 16 ……………………….. 1/26/05 9 ……………………….. 1/26/05 17 ……………………….. 1/26/05 10 ……………………….. 1/26/05 18 ……………………….. 1/26/05 11 ……………………….. 1/26/05 19 ……………………….. 1/26/05 12 ……………………….. 1/26/05 20 ……………………….. 1/26/05 13 ……………………….. 1/26/05 21 ……………………….. 1/26/05 14 ……………………….. 1/26/05 22 ……………………….. 1/26/05 15 ……………………….. 1/26/05 23 ……………………….. 1/26/05 16 ……………………….. 1/26/05 24 ……………………….. 1/26/05 17 ……………………….. 1/26/05 25 ……………………….. 1/26/05 18 ……………………….. 1/26/05 26 ……………………….. 1/26/05 19 ……………………….. 1/26/05 27 ……………………….. 1/26/05 20 ……………………….. 1/26/05 28 ……………………….. 1/26/05 21 ……………………….. 1/26/05 29 ……………………….. 1/26/05 22 ……………………….. 1/26/05 30 ……………………….. 1/26/05 23 ……………………….. 1/26/05 31 ……………………….. 1/26/05 24 ……………………….. 1/26/05 32 ……………………….. 1/26/05 25 ……………………….. 1/26/05 33 ……………………….. 1/26/05

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SECTION IV SECTION V (CONT’D) POWER PLANT & FUEL SYSTEM PROPELLER 16 ……………………………. 1/26/05

26 ……………………………. 1/26/05 17 ……………………………. 1/26/05 27 ………………………………. 1/26/05 18 ……………………………. 1/26/05 28 ………………………………… 1/26/05 19 ……………………………. 1/26/05 29 ……………………………. 1/26/05 20 ……………………………. 1/26/05 30 ……………………………. 1/26/05 SECTION VI 31 ……………………………. 1/26/05 LANDING GEAR 32 ……………………………. 1/26/05 1 ……………………………. 1/26/05 33 ……………………………. 1/26/05 2 ……………………………. 1/26/05 34 …………………………… 1/26/05 3 …………………………… 1/26/05 35 …………………………… 1/26/05 4 …………………………… 1/26/05 36 …………………………… 1/26/05 5 …………………………… 1/26/05 37 …………………………… 1/26/05 6 …………………………… 1/26/05 38 …………………………… 1/26/05 7 …………………………… 1/26/05 39 …………………………… 1/26/05 8 …………………………… 1/26/05 40 …………………………… 1/26/05 9 …………………………… 1/26/05 41 ……………………………. 1/26/05 10 ……………………………. 1/26/05 42 ……………………………. 1/26/05 11 ……………………………. 1/26/05 43 ……………………………. 1/26/05 12 ……………………………. 1/26/05 44 ……………………………. 1/26/05 13 ……………………………. 1/26/05 45 ……………………………. 1/26/05 14 ……………………………. 1/26/05 46 ……………………………. 1/26/05 15 ……………………………. 1/26/05 47 ……………………………. 1/26/05 16 ……………………………. 1/26/05 48 ……………………………. 1/26/05 17 ……………………………. 1/26/05 49 ……………………………. 1/26/05 18 ……………………………. 1/26/05 50 ……………………………. 1/26/05 19 ……………………………. 1/26/05 51 ……………………………. 1/26/05 20 ……………………………. 1/26/05 52 ……………………………. 1/26/05 21 ……………………………. 1/26/05 53 …………………………… 1/26/05 22 ……………………………. 1/26/05 54 …………………………… 1/26/05 23 ……………………………. 1/26/05 55 ……………………………. 1/26/05 24 ……………………………. 1/26/05 56 ……………………………. 1/26/05 25 ……………………………. 1/26/05 57 ……………………………. 1/26/05 SECTION VII 58 ……………………………. 1/26/05 FLIGHT CONTROLS 59 ……………………………. 1/26/05 1 ……………………………. 1/26/05 60 ……………………………. 1/26/05 2 ……………………………. 1/26/05 SECTION V 3 ……………………………. 1/26/05 FUEL SYSTEM 4 ……………………………. 1/26/05

1 ……………………………. 1/26/05 5 ……………………………. 1/26/05 2 ……………………………. 1/26/05 6 ……………………………. 1/26/05 3 ……………………………. 1/26/05 7 ……………………………. 1/26/05 4 ……………………………. 1/26/05 8 ……………………………. 1/26/05 5 ……………………………. 1/26/05 9 ……………………………. 1/26/05 6 ……………………………. 1/26/05 10 ……………………………. 1/26/05 7 ……………………………. 1/26/05 11 ……………………………. 1/26/05 8 ……………………………. 1/26/05 12 ……………………………. 1/26/05 9 ……………………………. 1/26/05 13 ……………………………. 1/26/05 10 ……………………………. 1/26/05 14 ……………………………. 1/26/05 11 ……………………………. 1/26/05 15 ……………………………. 1/26/05 12 …………………………… 1/26/05 16 ……………………………. 1/26/05 13 …………………………… 1/26/05 17 ……………………………. 1/26/05 14 …………………………… 1/26/05 18 …………………………… 1/26/05 15 ……………………………. 1/26/05 19 ……………………………. 1/26/05 20 ……………………………. 1/26/05

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Page Date Page Date

SECTION VII (CONT’D) FLIGHT CONTROLS

SECTION IX (CONT’D) DISPERSAL SYSTEMS

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SECTION X ELECTRICAL

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11 ……………………….. 1/26/05 12 ……………………….. 1/26/05 SECTION VIII

INSTRUMENTS 13 ……………………….. 1/26/05 1 ……………………….. 1/26/05 14 ……………………….. 1/26/052 ……………………….. 1/26/05 15 ……………………….. 1/26/05 3 ……………………….. 1/26/05 16 ……………………….. 1/26/05 4 ……………………….. 1/26/05 17 ……………………….. 1/26/05 5 ……………………….. 1/26/05 18 ……………………….. 1/26/05 6 ……………………….. 1/26/05 19 ……………………….. 1/26/05 7 ……………………….. 1/26/05 20 ……………………….. 1/26/05 8 ……………………….. 1/26/05 21 ……………………….. 1/26/05 9 ……………………….. 1/26/05 22 ……………………….. 1/26/05

10 ……………………….. 1/26/05 23 ……………………….. 1/26/05 11 ……………………….. 1/26/05 24 ……………………….. 1/26/05 12 ……………………….. 1/26/05 25 ……………………….. 1/26/05 13 ……………………….. 1/26/05 26 ……………………….. 1/26/05 14 ……………………….. 1/26/05 27 ……………………….. 1/26/05 15 ……………………….. 1/26/05 28 ……………………….. 1/26/05 16 ……………………….. 1/26/05 29 ……………………….. 1/26/05 17 ……………………….. 1/26/05 30 ……………………….. 1/26/05 18 ……………………….. 1/26/05 31 ……………………….. 1/26/05 19 ……………………….. 1/26/05 32 ……………………….. 1/26/05 20 ……………………….. 1/26/05 33 ……………………….. 1/26/05 21 ……………………….. 1/26/05 34 ……………………….. 1/26/05

35 ……………………….. 1/26/05

SECTION IX DISPERSAL SYSTEMS 36 ……………………….. 1/26/05

1 ……………………….. 1/26/05 37 ……………………….. 1/26/05 2 ………………………... 1/26/05 38 ……………………….. 1/26/05

39 ……………………….. 1/26/05

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Page Date SECTION X (CONT’D)

ELECTRICAL 40 ……………………………. 1/26/05 41 …………………………… 1/26/05 42 …………………………… 1/26/05 43 ……………………………. 1/26/05 44 ……………………………. 1/26/05 45 …………………………….. 1/26/05 46 ……………………………... 1/26/05 47 ……………………………. 1/26/05 48 ……………………………. 1/26/05 49 ……………………………. 1/26/05 50 ……………………………. 1/26/05 51 ……………………………. 1/26/05 52 ……………………………. 1/26/05 53 ……………………………. 1/26/05 54 ……………………………. 1/26/05 55 ……………………………. 1/26/05 56 ……………………………. 1/26/05 57 ……………………………. 1/26/05 SECTION XI AIRWORTHINESS LIMITATIONS 1 ……………………………. 1/26/05 2 ……………………………. 1/26/05

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LOG OF REVISIONS

Rev. No.

FAA Acceptance

Date Section Pages Description of Revisions FAA Accepted

New FEB 22 2005 All All New Book

Effective: 1/26/05 vi


INTRODUCTION This publication provides information for the Thrush Aircraft, Inc. Model S2R-T34 Turbo Thrush Aircraft. Installations or equipment will vary from model to model due to the wide range of optional equipment. The information contained within this manual is based on data available at the time of publication and will be kept current by changes or service publications. This manual contains information on aircraft systems and operating procedures required for safe and effective maintenance. It shall not be used as a substitute for sound judgment. In this manual: *** WARNING *** -- Indicates a strong possibility of severe personal injury or loss of life if instructions are not followed. ** CAUTION ** -- Indicates a possibility of personal injury or equipment damage if instructions are not followed. * NOTE * -- Gives helpful information. CAUTION: Detailed descriptions of standard workshop procedures, safety principles and service operations are NOT included in this manual. Please note that this manual DOES contain warnings and cautions against some specific service methods which could cause PERSONAL INJURY or could damage an aircraft or MAKE IT UNSAFE. Please understand that these warnings cannot cover all conceivable ways in which service, whether or not recommended by Thrush Aircraft Inc., might be done or of the possible hazardous consequences of each conceivable way, nor could Thrush Aircraft Inc. investigate all such ways. Anyone using service procedures or tools, whether or not recommended by Thrush Aircraft Inc. must satisfy himself thoroughly that either personal safety nor aircraft safety will be jeopardized. All information contained in this manual is based on the latest product information available at the time of printing. We reserve the right to make changes at any time without notice.

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Section 1 GENERAL INFORMATION

TABLE OF CONTENTS GENERAL INFORMATION ..............................................................................................................................2 GENERAL DESCRIPTION...............................................................................................................................2 PRINCIPAL DIMENSIONS...............................................................................................................................2

GENERAL .....................................................................................................................................................2 WING.............................................................................................................................................................3 HORIZONTAL STABILIZER AND ELEVATORS ..........................................................................................3 VERTICAL STABILIZER AND RUDDER ......................................................................................................3 AREAS ..........................................................................................................................................................4 SUPPLIER FURNISHED COMPONENT MANUALS....................................................................................4

AIRCRAFT STRUCTURES..............................................................................................................................4 FUSELAGE ...................................................................................................................................................4 WING.............................................................................................................................................................5 EMPENNAGE ...............................................................................................................................................5 COCKPIT.......................................................................................................................................................5

AIRCRAFT SYSTEMS .....................................................................................................................................6 HYDRAULIC SYSTEMS ...............................................................................................................................6 POWER PLANT & PROPELLER ..................................................................................................................6 FUEL SYSTEM .............................................................................................................................................6 LANDING GEAR, WHEELS & BRAKES.......................................................................................................8 FLIGHT CONTROLS.....................................................................................................................................8 INSTRUMENTS ............................................................................................................................................8 ELECTRICAL SYSTEM ................................................................................................................................8 AIRCRAFT WEIGHT & BALANCE................................................................................................................9 FIGURE 1-1.................................................................................................................................................10 FIGURE 1-2.................................................................................................................................................11

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GENERAL INFORMATION

SECTION ONE GENERAL DESCRIPTION

The Thrush Aircraft Inc Turbo Thrush is designed especially for agricultural flying. It is a monoplane

featuring a full cantilever low wing and all metal construction. The design and construction of the

airframe components assure all structural integrity, flight safety, and minimum maintenance

requirements. The Turbo Thrush is designed for the highest crash load factors in the industry.

Safety and reliability of operation and maximum pilot crash protection are proven and effective

features of the design. The high strength overturn structure is a proven design. The fuselage and

overturn structure, constructed throughout of chrome-moly steel tubing, is immensely strong in the

cockpit area.

CONTACT INFORMATION For further information related to this manual, please contact our Product Support Manager at

(229) 883-1440.

PRINCIPAL DIMENSIONS

GENERAL

Wing Span Extended Tip 47.5 feet

Overall Length 32.71 feet

Height To Top Of Canopy 11.42 feet

Main Gear Tread 9.00 feet

Main Gear To Tail Wheel 19.17 feet

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WING

Type Full Cantilever

Airfoil Section Naca 4412

Dihedral 3.50 Degrees

C. G. Range (See Airplane Flight Manual for pertinent data)

-Fwd Limit 26.5 Inches Aft Of Datum

-Aft Limit 30.0 Inches Aft Of Datum

(Datum Is The Leading Edge Of The Wing.)

Aileron Travel

-Up 21 Degrees ±1 Degree

-Down 17 Degrees ±1 Degree

Flap Travel Down 15 Degrees ±1 Degree

HORIZONTAL STABILIZER AND ELEVATORS

Span 204 Inches (17')

Elevator Travel



Trim Tab Travel



VERTICAL STABILIZER AND RUDDER

Rudder Travel 24 Degrees ±1 Degree

INTENTIONALLY LEFT BLANK INTENTIONALLY LEFT BLANK

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AREAS

Wing 336.43 Square Feet

Aileron (Each) 23.40 Square Feet

Flaps (Each) 15.30 Square Feet

Stabilizer 39.30 Square Feet

Elevators 20.40 Square Feet

Elevator Tabs (Each) 1.30 Square Feet

Fin 9.00 Square Feet

Rudder 11.40 Square Feet

SUPPLIER FURNISHED COMPONENT MANUALS

MANUAL PART #

PT6A-34AG Maintenance Manual Vol. I & II 3021242

Parts Manual 3021244

Propeller Owner’s Manual 139

AIRCRAFT STRUCTURES

FUSELAGE The fuselage comprises a welded tubular steel frame, fiberglass hopper, and detachable skins. An

overturn structure forms an integral part of the fuselage frame. The frame structure, fittings,

bushings, brackets, and so forth are fabricated from 4130 chrome-moly seamless steel tubing.

As a corrosion preventative, hot linseed oil is pumped throughout the entire welded structure. On

an average, 12 gallons are pumped into the frame and 11 to 11 1/2 gallons drain out, leaving a

residual coating on all members. The exterior of the frame is sandblasted, etched, and primed,

which is followed by two coats of polyurethane paint that is resistant to chemical reaction. The

fuselage is covered with heat treated Alclad panels attached with camloc fasteners. Side skins can

be removed using only a screwdriver, thus exposing the fuselage frame for thorough cleaning and

inspection. All skins are supported clear of the fuselage tubing to prevent accumulation of corrosive

chemicals. The seams and lap joints of the skin panel support structure are sealed with a special

compound to eliminate chemical action between the mating surfaces. Each skin panel is etched,

primed, and painted before assembly to insure complete coverage. All lower fuselage skins around

the hopper opening and aft to the tail post are made of stainless steel. The skin fasteners in the

high corrosion areas are also stainless steel.

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WING

The wing is a constant chord of 90 inches, all metal, and full cantilever design. The massive main

spar is a tension field beam structure constructed from Alclad webs and high strength heat-treated

steel caps. All wing skins, ribs, and leading edges are constructed from Alclad heat-treated

material. The leading edge structure is made especially strong to minimize denting and is riveted

with universal rivets for strength. The fuel tanks, which are located in the inboard section of the

wing, are an integral part of the structure. Close pitch riveting of the seams, substantial

reinforcement, and flexible sealants minimize chances of rupture in crash conditions. Drain holes

are provided in adjacent bays to prevent accumulation of fuel in the event of a leak. The ailerons

and flaps are all metal construction and are hinged on ball bearings. The flaps are electrically

operated by push rods and are completely sealed against chemical entry. Flap hinges are stainless

steel.

EMPENNAGE

The horizontal stabilizer, elevator, rudder and vertical fin are an all-metal structure. All skins, ribs

and leading edges are constructed from alclad material. The movable surfaces are hinged on

sealed bearings that can be easily replaced. The rudder and the elevator have aerodynamic

balances that are protected by overhangs on the fixed surfaces.

COCKPIT

For some models, there are two choices of the enclosed cockpit canopies for the Turbo Thrush (1)

the SINGLE cockpit canopy or (2) the DUAL cockpit canopy. The overturn structure is exceptionally

strong and welded to "hard points" in the fuselage frame. The forward bracing supports the

windshield support channels and is welded to a lateral tube that is curved to provide more head

clearance. The fiberglass canopy shell has extra thickness on the top portion and is well attached

to the extra large steel tube structure so that it will serve as a skid in case of overturn. The large

canopy doors permit easy entrance to one or both cockpits. The doors should not be removed for

flight, as the aircraft performance will be lowered. The cockpit seat belts are anchored to the seat

structure, and the shoulder harnesses are secured to a steel channel at the bottom of the seat

structure. The seats adjust vertically. The rudder pedals adjust fore and aft. The windshield is a

three-piece construction. The center section is tempered safety plate glass for better resistance to

scratching and is enclosed in a stainless steel frame. The windshield side panels are Plexiglas and

are curved to provide streamlining.

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AIRCRAFT SYSTEMS

HYDRAULIC SYSTEMS

The hydraulic system consists of two master cylinders and hydraulic brake lines connecting the

master cylinders to the wheel brake cylinders. Applying toe pressure on the rudder pedals actuates

the master cylinders, which are located just aft of the pilot’s rudder pedals. A small reservoir is

incorporated within each master cylinder to supply the system with brake fluid.

POWER PLANT & PROPELLER (Refer to manuals listed in Chart on Page 1-4 in this Section.)

The Turbo Thrush is powered by the PT6, a lightweight free turbine engine incorporating a reverse

flow combustion path, is designed for aircraft propulsion use. It utilizes two counter rotating turbine

sections. One drives the compressor, and the other drives the propeller through a reduction

gearbox. The latter turbine is "free" or independent of the compressor turbine. More recent and

higher-powered models incorporate a two-stage free turbine. The PT6 has been produced in

several models and has been adapted to a multitude of uses.

The propeller has three blades mounted on a hollow hub in the front end of which is a servo-piston

that moves forward under servo-oil pressure or rearward under feather returns spring pressure.

There are three links from the servo-piston. One goes to each blade root, and these links transmit

forward motion of the servo-piston to the blade roots and pivot the blades in the decrease pitch

direction. When servo-piston pressure is relieved, the servo-piston moves rearward under feather

return spring pressure and pivots the blades in the increase pitch direction. This action is assisted

by centrifugal force of the counterweight on each blade root.

FUEL SYSTEM

A 230-gallon fuel supply is available for the Turbo Thrush. In each wing, fuel is contained inside

integral wing tanks (wet wing fuel tanks) just outboard of the center section subwings. The left wing

and right wing fuel tanks are interconnected through a 5 U.S. gallon header tank that is located in

the fuselage. The fuel supply lines, to the engine, are routed from the header tanks outlet finger

screen through a fuel shutoff (on/off) valve to an electric driven fuel boost pump. The electric driven

fuel boost pump discharge is then routed through a 25-micron main fuel filter to an engine driven

fuel boost pump. The electric driven fuel boost pump serves two purposes, first as a backup

system to provide continuous fuel pressure to the engines high pressure fuel pump in case the

engine driven fuel boost pump fails and secondly to provide boosted fuel pressure to the engines

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high pressure fuel pumps during engine starting. The aircraft’s fuel system is equipped with two

fuel filters, a ¼ inch mesh finger strainer is installed in the outlet fitting from the header tank and a

25-micron, airframe supplied, main fuel filter located on the forward L/H side of the firewall. Fuel

from the aircraft fuel system enters the engines high pressure fuel pump which has two fuel

filters, an 74-micron inlet filter and a 10-micron discharge filter (refer to the engines appropriated

maintenance manual for pertinent maintenance details for the engine supplied filters and fuel

system). The fuel tank vent system is designed to keep the fuel spillage to a minimum. The fuel

tanks are vented through tubing connected at both the inboard and outboard ends of the

individual fuel tanks to the centrally located vent system in the fuselage. Ram air enters a vent

scoop, on the fuselage, under the left wing and pressurizes the vent system to maintain positive

pressure on the fuel tanks. The vent system is provided with two quick drains, located on the

fuselage under each wing, to drain any fuel that might happened to have got in the tanks

outboard vent lines. At engine shutdown, fuel from the start control unit or the flow divider/dump

valve, located at the 6 o’clock position on the engines fuel nozzle manifold, is directed to a

residue fuel reservoir “EPA tank” mounted inboard on the L/H aft shin skin. This reservoir hold

approximately 3 engine shutdowns worth of fuel before the fuel will exit the reservoirs vent

system. (NOTE: This reservoir should be emptied after each engine shutdown.) (NOTE: It is

common and normal after an engine compressor Water Wash or Performance Recovery Wash to

have water or soap appears in the reservoirs’ drained waste fuel.) The fuel quantity gauge is

located on the lower left instrument panel. The fuel quantity indicating system consists of two

transmitters, one indicator gauge, and a L/H or R/H tank fuel quantity selector switch. A

transmitter, installed in each wing tank transmits an electrical signal to the single fuel quantity

indicator. The instrument reads both the left and right fuel tanks singularly as chosen by the

electrical control switch, adjacent to the fuel quantity indicator gauge on the instrument panel.

The two fuel tanks are serviced through filler ports located on the top of both wings. The filler

ports incorporate security chains to prevent the lost of the fuel caps. Service the aircraft from

refueling facilities that utilize proper ground handling equipment and filter systems to remove

impurities and water accumulations from the bulk fuel. If filtering facilities are not available, filter

the fuel through a quality high-grade chamois. Fuel tanks should be serviced after the last flight

of each day to reduce condensation and allow any entrapped water accumulations to settle to the

fuel system drains, to be removed, prior to the next flight.

Prior to the first flight of the day the header tank and fuel filter should be drained to check for the

presence of water or sediment in the fuel system. If there is a possibility, at any time, that any

tank may contain water, the header tank and fuel filter should be drained as necessary to ensure

no water exists in the fuel system. For fuel system servicing information, refer to Section Two.

1 – 7 Effective: 1/26/05


LANDING GEAR, WHEELS & BRAKES

The main landing gear is made using a formed chrome-moly spring steel unit. The left Main gear

and the right main gear are symmetrical. The main wheels are 11 X 10. The spring steel

construction and design of the main gear allow for absorption of landing weight and common

stresses associated with such, thus eliminating the need for shock struts. The brake system has

individual toe brakes and individual park brakes. The use of a special N-513 compound cup in each

master cylinder permits the use of MIL-H-5606, a heavy-duty aviation hydraulic fluid. The brakes

are dual caliper disc types. The tail gear is a spring steel type and uses a 500 x 5 tailwheel.

FLIGHT CONTROLS

The flight controls are of conventional design employing extensive use of ball bearings for low

friction and smoothness of operation. The aileron and elevator controls are push rod systems and

the rudder control is through cables. The elevator trim control is actuated by a lever that moves the

tab to the desired position through push rods. The wing flaps are operated electrically and

controlled by a switch located on the left side of the cockpit. The rudder controls are interconnected

by springs to the aileron system so that a wing may be lifted with the rudder alone.

INSTRUMENTS

The standard instruments are located on three separate panels: An upper panel, a left panel, and a

right panel. The left panel contains a clock, oil temperature, hour meter, fuel pressure, oil pressure,

air filter Delta “P”, and fuel quantity gauges. The right panel contains a voltmeter, ammeter, and

circuit breakers. The upper panel contains all engine-warning lights, torque pressure, ITT indicator,

and Gas Generator percent RPM, Propeller RPM and standard flight instrument package.

ELECTRICAL SYSTEM

The standard 24 volts and 105 amps electrical system consists of the starting system, the

navigation lights, the wiper/washer system, and the strobe lights. The landing lights, the working

lights, and the air conditioner system are optional. The landing and working lights may be installed

in the field, since the wiring for them is included in the standard wire bundle. The electrical system

obtains power from two 24-volt batteries and one starter/generator. An external power receptacle is

standard equipment and may be used for connecting a 24-volt ground power unit to the aircraft for

engine starting or maintenance. The ground start system utilizes the master relay so that starting is

accomplished by engaging the starter switch.

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AIRCRAFT WEIGHT & BALANCE

Refer to S2R-T34 Flight Manual for aircraft weight and balance information.

1 – 9 Effective: 1/26/05


Figure 1-1

Effective: 1/26/05 1 - 10


1 – 11 Effective: 1/26/05 Figure 1-2


Section 2

SERVICING & INSPECTION

TABLE OF CONTENTS SECTION TWO ........................................................................................................................................... ....1

SERVICING ..................................................................................................................................................3 GROUND HANDLING ..................................................................................................................................3

TOWING....................................................................................................................................................3 TAXIING ....................................................................................................................................................3 PARKING ..................................................................................................................................................3 MOORING.................................................................................................................................................3 JACKING...................................................................................................................................................4 LEVELING.................................................................................................................................................4

COLD WEATHER OPERATION...................................................................................................................4 COLD WEATHER MAINTENANCE HINTS ..............................................................................................5

GROUND EMERGENCY PROCEDURES ...................................................................................................5 ENGINE FIRES .........................................................................................................................................5 ELECTRICAL FIRES.................................................................................................................................6

GROUND OPERATION OF ENGINE...........................................................................................................6 BEFORE STARTING ENGINE..................................................................................................................6 STARTING ENGINE .................................................................................................................................7 ENGINE OPERATIONAL CHECK ............................................................................................................8

SYSTEM AND COMPONENT SERVICING .................................................................................................8 HYDRAULIC SYSTEM..............................................................................................................................9 ENGINE OIL SYSTEM ..............................................................................................................................9 FUEL SYSTEM .......................................................................................................................................13 DEFUELING............................................................................................................................................15 INDUCTION SYSTEM.............................................................................................................................16 POWER PLANT INTERNAL CLEANING................................................................................................16

LANDING GEAR, WHEELS & BRAKES ....................................................................................................17 TIRES......................................................................................................................................................17 BRAKE BLEEDING .................................................................................................................................17 INSPECTION...........................................................................................................................................17 INSPECTION CHECK LIST ....................................................................................................................17 INSPECTION CHART..............................................................................................................................19

ENGINE EXTERNALS................................................................................................................................21 ENGINE OIL SYSTEM................................................................................................................................23 ENGINE FUEL SYSTEM ............................................................................................................................24 IGNITION SYSTEM ....................................................................................................................................25 PNEUMATIC SYSTEM ...............................................................................................................................25 AIRFRAME FUEL SYSTEM .......................................................................................................................26 MAIN LANDING GEAR ...............................................................................................................................26 TAIL GEAR..................................................................................................................................................27 FUSELAGE SKINS .....................................................................................................................................29 HOPPER.....................................................................................................................................................29 WINGS........................................................................................................................................................29 FUSELAGE FRAME ...................................................................................................................................31 CONTROL SYSTEMS ................................................................................................................................31 METAL EMPENNAGE ................................................................................................................................32 AILERONS AND FLAPS.............................................................................................................................33 COCKPIT ....................................................................................................................................................34 ELECTRICAL SYSTEM...............................................................................................................................35 AIRFRAME MAINTENANCE ......................................................................................................................36

Effective: 1/26/05 2 - 1

THRUSH AIRCRAFT INC - MODEL S2R-T34 TURBO THRUSH AIRCRAFT MAINTENANCE MANUAL

CORROSION CONTROL........................................................................................................................36 WINDSHIELD..........................................................................................................................................37 HOPPER REPAIR...................................................................................................................................37 FUEL TANK REPAIR ..............................................................................................................................37 BATTERY MAINTENANCE.....................................................................................................................37 LUBRICATION ........................................................................................................................................38

Effective: 1/26/05 2 - 2

THRUSH AIRCRAFT INC – S2R-T34 TURBO THRUSH AIRCRAFT MAINTENANCE MANUAL

SECTION TWO

SERVICING & INSPECTION

Standard procedure for ground handling, servicing, inspection, airframe maintenance, lubrication,

and storage are included in this Section. Adherence to these procedures on a scheduled basis can

save many hours of maintenance and aircraft down time. When a system component requires

service or maintenance other than that outlined in this Section, refer to the applicable Section of

this manual for complete information.

GROUND HANDLING

TOWING

Movement of the aircraft on the ground may be accomplished as follows:

A. Pull and guide the aircraft by means of a tow bar with the tail wheel unlocked.

B. Attach a rope harness to the main gear when there is a need to tow the aircraft forward through

snow or over soft and/or muddy ground.

TAXIING

Before attempting to taxi the aircraft, maintenance personnel should be checked out by qualified

personnel. When it is determined that the propeller area is clear, apply the power to start the taxi roll

and perform the following:

A. Push the stick full forward to unlock the tail wheel.

B. Taxi a few feet and check the brake operation.

C. While taxiing, make slight turns to determine that the tail wheel steering is operative.

D. Avoid taxiing over ground of loose stones, gravel, or other loose material that may cause foreign

object damage to the propeller or to other aircraft in the area.

E. You may taxi with the power lever in the Beta region to govern ground speed. Observe all

engines operating limits.

PARKING

Head the aircraft into the wind and set the parking brake. Do not set the parking brake during cold

wet weather because the accumulated moisture may freeze in the brakes. Do not set the parking

brake if the brakes are overheated. Install the internal control lock. Place the chocks under each

main wheel.

Effective: 1/26/05 2- 3


MOORING

Park aircraft as previously outlined. In winds up to 20 knots, secure the aircraft at the wing tie down

rings. For winds above 20 knots, tie the tail and main gear as well as the wings. Install external

control surface locks. Be sure to tie the propeller down to prevent it from wind milling with zero oil

pressure. The aircraft should be placed in a hangar when wind velocity is predicted to exceed 50

knots. When mooring aircraft, use 3/4-inch manila or nylon rope. A clove hitch or other anti-slip knot

should be employed. If a manila rope is used for tie down, allow enough slack to compensate for

contraction of the rope fiber without damaging the aircraft.

JACKING

Jack points are provided on each main spar and located at wing stations 120 & 193.38. When using

the jack points to lift the aircraft, all hopper loads should be removed. (Fig. 2-1) A jack point is also

provided on the tail wheel trunnion attach fitting on the lower left longeron.

LEVELING

The aircraft may be leveled by raising the tail to an approximate level flight position and by supporting

the tail on a stable jack or platform. Adjust the height of the tail wheel until the left-hand lower

longeron located under the cockpit is level.

COLD WEATHER OPERATION

Aircraft operation in cold weather creates a need for additional maintenance practices and operating

procedures that are not required in moderate temperatures. Whenever possible, shelter the aircraft in

a heated hangar to prevent frost, ice, or snow accumulation that requires added maintenance time to

remove. These weather elements, if allowed to accumulate only a fraction of an inch in thickness on

the critical airfoils and control surfaces, seriously degrade aircraft lift and flight control effectiveness.

The possibility of aircraft system failures is increased when the aircraft is parked where wind driven

snow or freezing rain can be forced into various openings of the aircraft. If the aircraft is to be moored

outside in extreme cold, the battery should be kept fully charged to prevent freezing. Make certain

that all vents, air inlets, and so forth are covered.

Locating the aircraft inside a heated hanger is the most effective method of preheating the aircraft.

The use of an external power unit is recommended to conserve the battery.

2 - 4 Effective: 1/26/05


COLD WEATHER MAINTENANCE HINTS

The information that follows is intended only for the purpose of supplementing the existing information

in this manual when operating the aircraft in cold weather. Keeping the aircraft in top maintenance

condition during cold weather cannot be over stressed.

The battery should be maintained at full charge during cold weather to prevent freezing. After adding

water to the battery in freezing temperatures, charge the battery to mix the water and electrolyte. A

frozen battery may explode when subjected to a high charge rate. Corrosive damage to the area

adjacent to an exploded battery will result if the electrolyte solution is not removed immediately.

Instructions for removing spilled electrolyte are provided in this Section. The battery should be

removed and stored in a warm place if the aircraft is to remain idle for an extended period of time.

In the fuel system, condensation is more likely to occur in cold weather due to a more rapid and

positive division of moisture content from other fuel properties. If at all possible, use fueling facilities

that filter moisture from the fuel. If fueling facilities with filters are not available, filter the fuel through a

good quality chamois. Fill the tanks with correct grade of fuel as soon as possible after landing to

reduce the possibility of condensation and ice formation in the tanks. Fuel extracted from fuel header

tank drain before starting deserves a closer examination when the aircraft is being operated in cold

weather.

Cold weather operation demands procedures that are in addition to normal Post Flight Maintenance

Procedures. Fill the fuel tanks immediately after flight. If shelter is not available, tie the aircraft down

and install covers on all vents, openings, and so forth as required.

GROUND EMERGENCY PROCEDURES

Emergency procedures must be accomplished as rapidly as possible, should an emergency arise. It

is suggested that steps pertaining to each emergency be committed to memory in order to accelerate

the procedure and minimize any possible damage.

ENGINE FIRES

The following Dry Motoring Run procedure is used to clear an engine at any time when deemed

necessary to remove internally trapped fuel and vapor or when there is evidence of a fire within the

engine. Air that passes through the engine serves to purge fuel, vapor, or fire from the combustion

sections, the gas generator turbine, the power turbine, and the exhaust system.

Effective: 1/26/05 2- 5


A. Fuel Condition Lever - Cut Off

B. Ignition Switch - Off

C. Master Switch - On

D. Fuel Shutoff Valve - On

E. Fuel Auxiliary Pump Switch - On

This will provide lubrication for the engine-driven fuel pump.

F. Engine Starter Switch – On

G. Maintain the starter operation for the desired duration. The maximum starter duration is 3

minutes.

H. Engine Starter Switch - Off

I. Fuel Auxiliary Pump Switch - Off

J. Fuel Shutoff Valve - Off

K. Master Switch - Off

L. Allow a 5-minute cooling period for the starter before going any further with the starting

operation.

ELECTRICAL FIRES

Circuit breakers will automatically trip and stop the current flow to a shorted circuit. However, as a

safety precaution in the event of an electrical fire, turn the battery switches to off. Use a fire

extinguisher approved for electrical fires to extinguish the flame.

GROUND OPERATION OF ENGINE

BEFORE STARTING ENGINE

Visually check the aircraft for general condition. Verify that all camlocs on the skin panels are

fastened. Remove all accumulations of frost, ice, or snow in cold weather from the wing, the tail, and

the control surfaces. Check that the control surfaces contain no internal accumulations of ice.

Remove the inlet and exhaust covers, if fitted. If night flight is planned, check the operation of all

lights and have a flashlight available.

2 - 6 Effective: 1/26/05

*** WARNING ***

If the fire persists as indicated by the sustained

interturbine temperature, close the fuel system

shutoff valve at this point and continue motoring.


After a complete visual inspection has been accomplished, the following checklist may be used for the

external prestart check. The aircraft should be headed into the wind and should have the wheel

chocks in place.

A. A fire extinguisher must be readily available in the event of an engine fire.

B. Check the engine oil level. Assure that the oil system has been serviced with the correct

grade of oil.

C. Verify that the internal control lock has been removed and that the controls operate freely.

D. Set the parking brake.

E. Check the fuel quantity in both tanks.

F. Set the trim tabs for takeoff.

G. Clear the area of all personnel.

STARTING ENGINE

Use the following procedure to start the PT6A engine.

A. Battery and Generator Switches - On

B. Power Lever - Idle

C. Propeller Lever - Feather

D. Fuel Condition Lever - Cut Off

E. Fuel Shutoff Valve – On

F. Fuel Auxiliary Pump Switch - On

G. Fuel Inlet Pressure Indicator - Check 5 PSI Minimum

H. Engine Starter Switch - On

The minimum speed to obtain a satisfactory light is 12% Ng.

I. After approximately 5 seconds of motoring at the stabilized gas generator speed, turn the

Ignition Switch On and move the Condition Lever to the Ground (low) Idle position.

J. Observe that the engine accelerates normally to idle RPM and the maximum allowable

inter-turbine temperature-starting limit is not exceeded.

** CAUTION **

Whenever the gas generator fails to light up within 10 seconds after moving the fuel condition lever to the

ground (low) idle position: fuel condition lever – idle cutoff, ignition switch - off, starter – off. Allow a 30-

second fuel draining period that is followed by a 15-second dry motoring run before attempting another

start. If for any reason a starting attempt is discontinued, allow the engine to come to a complete stop and

then accomplish a Dry Motoring Run as described on page 2-5 under Engine Fires. That procedure is

also referred to as Dry Motoring run.

Effective: 1/26/05 2- 7


When the engine attains idle rpm:

K. Engine Starter Switch and Ignition Switch - Off

L. Oil Pressure - Check 60 PSIG Minimum

M. Fuel Auxiliary Pump - Off

N. Fuel Pressure from Engine Driven Pump - Check 5 PSI Minimum

O. Generator Charging – Check

ENGINE OPERATIONAL CHECK Refer to Section Four for specific operational checks and/or Pratt & Whitney Maintenance Manual.

Before proceeding with a ground run up, be sure that the propeller system is purged by feathering the

propeller once or twice with the power control lever in idle position.

The following procedure should be used to check the propeller overspeed governor.

A. Place the propeller lever in full increase RPM position (forward).

B. Turn prop test switch on.

C. Increase RPM with the power lever until governing occurs. This should occur at 2025±20

RPM. (In no case should any engine limitations be exceeded.)

D. Reduce power back to idle.

E. Turn prop test switch off.

SYSTEM AND COMPONENT SERVICING Servicing procedures contained in this Section are confined to those maintenance actions that occur

with routine frequency and require a reasonably short period of time to accomplish. Servicing

practices and maintenance to aircraft systems and components that require less frequent attention

are contained in the appropriate Section of this manual.

2 - 8 Effective: 1/26/05

*NOTE* If RPM is not governed at 2025 ±20 RPM with the prop test switch

on, consult Section IV of this manual for adjustment of the

overspeed-governor.

** CAUTION **

Fill hopper and hold the elevator control firmly full up during all

high power ground operations to keep aircraft from nosing over.


HYDRAULIC SYSTEM

The hydraulic system consists of two master brake cylinders and the necessary hydraulic lines

connecting the master cylinders to the wheel brake cylinders. Applying toe pressure on the rudder

pedals actuates the master cylinders, which are located just aft of the pilot’s rudder pedals. Refer to

Section Six for brake servicing procedures.

ENGINE OIL SYSTEM

The oils that are specified for the lubrication system are detailed in the Pratt and Whitney Canada

Service Bulletin 1001. All oils listed in the bulletin are approved for flight operation. It is recommended

for all turbo aircraft that the oil should be changed every 400 hours. The oil system contains 13 U.S.

quarts.

In cases where oils that are approved are not available, an operator must obtain prior approval or

recommendations for use of substitution oil from the Service Department, Pratt and Whitney Canada

Corp, 1000 Marie-Victorin, Longueuil, Quebec, Canada J4K 1A1.

A. OIL LEVEL CHECK To avoid overfilling of oil tank, and high oil consumption, an oil level check

is recommended within 30 minutes after shutdown. Ideal interval is 15 to 20 minutes. If more

than 30 minutes has passed, and the dipstick indicates that oil is needed, start the engine and

run at ground idle (low idle) for five minutes, and recheck oil level.

1. Unlock the filler cap and dipstick from the filler neck at the eleven o'clock position on

the accessory gearbox and remove the filler cap.

2. Check the oil tank contents against the markings on the dipstick. Service as

required.

Effective: 1/26/05 2- 9

** CAUTION **

Do not mix different brands, viscosity’s, or types of oil since their

chemical structures may make them incompatible. If different types of

oil become mixed, drain and flush the system. Refill with new oil.


3. If the engine is nose high or nose low, compensation must be made to avoid over

or under servicing.

B. If the oil level is too low to register on the dipstick due to possible excessive consumption or if

low or fluctuating pressures have been recorded, refer to Troubleshooting - Lubrication Problems

in the Pratt and Whitney Maintenance Manual for the action to be taken. After that has been

accomplished, proceed as follows to check the oil level.

1. Fill the oil tank to the appropriate normal level. Record the quantity of oil added to

the system.

2. Install the filler cap and dipstick. Ensure that the cap is locked securely.

3. Run the engine idle for approximately 5 minutes.

4. Check the oil level.

5. Check the oil filter per applicable Pratt & Whitney Maintenance Manual.

C. On engines which have remained stationary for a period of 12 hours or more, proceed as follows

to check the oil level.

1. Start the engine and run at idle speed for a minimum of 2 minutes.

2. Feather the propeller.

3. Shut down the engine.

4. Check the oil level.

D. Recommendations for oil change intervals are based on the performance of specific brands of

oil, specific types of oil, specific engine models, and specific operating criteria. General oil

change intervals may be extended periodically and will be reflected by revisions to the Pratt and

Whitney Engine Service Bulletin 1001. Permission for extension of oil drain intervals may be

granted to operators through monitoring programs, which are conducted by most

2 - 10 Effective: 1/26/05

* NOTE * The graduations on the dipstick indicate the oil level in U.S. quarts

below maximum capacity of the oil tank. The normal cold oil level is

the Maximum Cold mark on the dipstick. The normal hot level is

Maximum Hot mark on the dipstick. A dipstick reading of 3 will indicate

that the system requires 2 U.S. quarts to replenish to normal level if

the oil is cold. If the oil is hot, it will take 3 U.S. quarts to replenish.


major oil companies that have been approved by Pratt and Whitney Canada. Service Bulletin 1001

will be revised periodically to include newly approved oils. Refer to Figure 2-8 for the locations called

out in the following procedure.

1. Place suitable containers or drip pan under the engine.

2. Remove lockwire from the main oil tank’s drain plug #6 from boss on compressor

inlet case. Remove drain plug. Discard the preformed packing. Or drain oil at drain

port on left shin skin (on aircraft equipped with quick drain).

3. Remove the rear case drain plug #2 from the six o'clock position on the rear face of

the accessory gearbox housing. Discard the preformed packing.

4. Remove the chip detector #3 from the six o'clock position on the reduction gearbox

front case. Discard the preformed packing.

5. Examine the drained oil for the presence of foreign matter.

E. Refill the oil tank by accomplishing the following procedures.

1. Install the chip detector #3 with new preformed packing on the reduction gearbox.

Torque chip detector body #3 45 to 55 lb. in. and lockwire.

2. Install rear case drain plug #2 with new preformed packing in the accessory gearbox

housing. Tighten and torque to 215 to 240 lb. in. and lockwire.

3. Install the drain plug #6 with the new preformed packing in the bottom of the air inlet

case. Or simply install a cap on the drain port on the left shin skin (on aircraft

equipped with quick drain) and lockwire.

4. Fill the oil tank with the specified oil to the level of maximum graduation on the

dipstick.

F. Install the filler cap and dipstick assembly in the oil tank. Ensure that the cap is locked securely.

1. Start the engine and run at idle for approximately 2 minutes to circulate the oil

through the system.



4. Check the oil level in the tank. Replenish, as required, to the normal level on the

dipstick.

5. Install the filler cap and dipstick assembly in the oil tank. Ensure that the cap is

locked securely.

Effective: 1/26/05 2- 11


G. If an engine is to be operated with an oil brand or type that differs from that on which it

previously operated or if the oil system has been contaminated by other than metallic matter,

the oil system should be flushed by following the steps below.

1. Place suitable containers or drip pan under the engine.

2. Remove the oil drain plug or chip detector from the reduction gearbox and the

plugs from the inlet case and the accessory gearbox housing.

3. With the drains open, place the starting control lever to cutoff and the ignition

switch to off. Motor the engine with the starter only to allow the scavenge pumps

to clear all lubricating oil.

4. Reinstall all drain plugs and the chip detector.

5. Refill the engine oil tank with new type oil.

6. Start the engine and run at idle speed for a minimum of two minutes.



9. Repeat Steps 1. through 3.

10. Remove the main oil filter. Clean or replace the filter and reinstall.

11. Remove the reduction gearbox oil strainer and clean. Reinstall the strainer.

12. Reinstall all engine drain plugs and the chip detector. Tighten, torque, and

lockwire.

13. Repeat Steps 5. through 8.

14. Check the oil levels and replenish, as necessary.

15. Install the filler cap and dipstick assembly in the filler tube. Ensure that the cap is

correctly installed and locked.

2 - 12 Effective: 1/26/05

** CAUTION **

Limit the engine rotation to a minimum time

which is required to accomplish the complete

draining. Also observe the starter operating

limitations.


A. Remove the drain plug from the six o'clock position on the

accessory gearbox.

B. Using a mirror and light, inspect the scavenge screen through the

drain hole.

C. If there is evidence of carbon, try to dislodge it with a stiff paintbrush.

D. Flush out any removed carbon.

E. If the carbon cannot be removed by the above method, the

accessory gearbox should be removed and the screen cleaned.

Refer to the Section Accessory Gearbox in the Pratt and Whitney

Maintenance Manual for the removal procedure.

FUEL SYSTEM

A. Refuel the aircraft with fueling facilities that contain filters for removing the moisture content

from the fuel. If the fueling facilities with filters are not available, filter the fuel through a good

grade chamois. The fuel tanks should be serviced after the last flight of the day to allow

maximum time for the moisture to reach the header tank. Service the aircraft with Jet A, Jet

B, JP-4, and JP-5. If jet fuel is not available, aviation gasoline MIL-G-5572 (all grades) may

be used for a maximum of 150 hours between overhauls. For the Restricted Category,

service the aircraft with Jet A, Jet B, JP-4, JP-5, and automotive diesel number 1D or 2D in

accordance with P&WACL Service Bulletin Number 1344 or 14504. If the jet fuel or diesel

fuel is not available, aviation gasoline MIL-G-5572 (all grades) may be used for a maximum

of 150 hours between overhauls. Automotive diesel fuel is approved only for flights when the

free air temperature is above +20 degrees Fahrenheit use grade #1D or +40 degrees

Fahrenheit use grade #2D.

** CAUTION **

Different formulations of the various oil brands may have varying detergent

actions. After an oil brand change, the above may cause the release of

carbon particles into the oil system which would result in the clogging of the

scavenge screen. After a change of oil brand, the main oil filter should be

inspected for carbon particles at 10-hour intervals. There should be 5

inspections for a total of 50 hours, and the filter should be checked at the

routine oil filter checks there after up to 500 hours. If an excess of the

normal amount of carbon is noted, the following steps should be

accomplished.

Effective: 1/26/05 2- 13


1. Turn all the switches off.

2. Remove the fuel filler cap. Fill the tank until the fuel level rises to the filler neck. Install the

fuel filler cap and service the opposite fuel tank.

3. After fueling is complete, check for security of both fill port caps. Wash any spilled fuel from

the wing surface with clean water.

B. Three fuel drain points are provided to allow fuel draining in order to extract the moisture and

sediment entrapped in the system. The drains are located at the bottom of each wing tank (Fig. 2-2),

the header tank (Fig. 2-3), and firewall fuel filter (Fig. 2-4). All fuel drains should be drained prior to

the first flight of the day. Drain a small quantity of fuel into a transparent container to permit

inspection for the presence of moisture or sediment. The fuel should be drained until all evidence of

moisture or sediment disappears. Also provided are two fuel vent drains located on each side of

fuselage under the wings (Fig. 2-5). The last drain port is provided to drain the residue fuel reservoir.

At engine shutdown, fuel from the flow divider/dump valve, located at the 6 o’clock position on the

engines fuel nozzle manifold, is directed to a residue fuel reservoir “EPA tank” located on the L/H cowl

shin skin. This reservoir holds approximately 3 engine shutdowns worth of fuel before the fuel will exit

the reservoirs’ vent system.

2 - 14 Effective: 1/26/05

*** WARNING ***

Ground the aircraft and the fuel servicing equipment to the aircraft.

Smoking in or around the aircraft during refueling operations is

prohibited. Fire protection equipment must be immediately

available.

* NOTE *

As the wing tanks are interconnected through the header tank, the

fuel can flow from one tank to another. Topping off both wing tanks

may be required more than one time to assure that both wing tanks

are full.


Visually check that all drain valves are closed after draining.

C. The airframe is equipped with two fuel filters, a ¼ inch mesh finger strainer is installed in the

outlet fitting from the header tank and an airframe supplied, gascolator type, 25-micron main

fuel filter located on the forward L/H side of the firewall. Inspect the ¼ inch mesh finger strainer

annually or if the fuel system is suspected or has been contaminated with foreign debris: i.e.

Main fuel filter red bypass indicator was popped, main fuel filter has contamination, foreign

debris noted in drained fuel sample container, known fuel contamination …etc. The 25-micron

main fuel filter element should be inspected, cleaned or replaced, and reinstalled every 100

hours, when the red fuel bypass indicator button has popped, or any time fuel system

contamination is suspected. Refer to chapter 5 for main fuel filter servicing procedures.

DEFUELING

During the defueling operation, jet fuel fumes are present; therefore, extreme caution must be

exercised to prevent fire hazards.

Effective: 1/26/05 2- 15

* NOTE * This reservoir should be emptied after each engine shutdown.

* NOTE *

It is common and normal after an engine compressor Water Wash

or Performance Recovery Wash to have water or soap appear in

the reservoirs’ drained waste fuel.

*** WARNING ***

If the red fuel bypass indicator button has popped out, determine

and remove the cause of the fuel obstruction before further flight.

Remove, inspect, clean or replace, reinstall the filter 25-micron

element. You may then reset the red bypass button by pressing it

in with finger pressure.


A. Ground aircraft and all defueling equipment or containers to the aircraft.

B. Place vented container of adequate capacity under the three drain points. Verify that the

containers are properly grounded to the aircraft.

C. Open the drain valves and allow all fuel to drain.

D. Close the drain valves and move the fuel containers to a safe distance from the aircraft.

E. Verify that all the drain valves are closed.

INDUCTION SYSTEM

The prime difference between the agricultural and a normal installation is the air cleaning system

incorporated in the engine air intake system. The air filter is located below the engine air inlet plenum

between the center and rear fire seals and is a washable reusable barrier type filter.

POWER PLANT INTERNAL CLEANING

Refer to Pratt & Whitney Canada Maintenance Manual for the -34 engine for proper internal cleaning.

2 – 16 Effective: 1/26/05

*** WARNING ***

Smoking on or around the aircraft is not permitted during the

defueling procedure. Fire protection equipment must be

immediately available.

** CAUTION **

Chemicals should not be allowed to remain in an engine any

longer than overnight, and a water wash should not be performed

any sooner than 45 minutes after shutdown. It may be more

convenient and practical to wash the engine before working the

next morning. This is acceptable if extremely corrosive chemicals

are not being used.


LANDING GEAR, WHEELS & BRAKES

Check all gear assemblies for general cleanliness, security of mounting, and hydraulic leaks at

prescribed inspection intervals. Lubricate all lubrication points on main and tail gear assemblies at

prescribed intervals.

TIRES

Tires should be inspected for proper inflation, breaks, cuts, and foreign objects in tread, flat spots and

exposed cord. Replace tire if there is any question of its reliability. Proper inflation is necessary for

maximum tire life. Maintain 29x11-10pr main wheel pressure at a minimum of 40 psi to a maximum of

62 psi, depending on the load and runway conditions. 5.00-5 10pr Type III tail wheel tire pressure

should be 88 psi maximum. The wheels and tires are balanced assemblies. If tires are suspected of

being out of balance, they may be balanced on automotive type balancing equipment. If aircraft is out

of service, rotate tires every seven days to prevent flat spots from developing.

BRAKE BLEEDING

Brake bleeding should be performed when air is suspected of being entrapped in brake lines. See

Section Six for brake bleeding procedures.

INSPECTION

Only the items to be inspected are listed and details as to how to check or what to check for are

generally excluded. Those checks can be found in specified Section of this manual.

INSPECTION CHECK LIST

A. Movable parts are to be checked for lubrication, servicing, security of attachment, binding,

excessive wear, Safety, proper operation, proper adjustment, correct travel, cracked fittings,

security of hinges, defective bearings, cleanliness, corrosion, deformation, sealing, and tension.

B. Fluid lines and hoses are to be checked for leaks, cracks, dents, kinks, chafing, proper radius,

security, corrosion, deterioration, obstructions, and foreign matter.

C. Metal parts are to be checked for security of attachment, cracks, and metal distortion, broken

spot welds, corrosion, condition of paint, and any other apparent damage.

Effective: 1/26/05 2- 17


D. Wiring is to be checked for security, chafing, burning, defective insulation, and loose or broken

terminals, heat deterioration, and corroded terminals.

E. Bolts in critical areas are to be checked for correct torque, or when visual inspection indicates

the need for a torque check. See (Fig 2-7) Torque Chart.

F. Filters, screens, and fluids are to be checked for cleanliness, contamination and/or replacement

at specified intervals.

This Manual contains information on aircraft systems and operating procedures required for safe and

effective maintenance. It shall not be used as a substitute for sound judgment.

Clean the aircraft prior to performing any inspections on the airframe or engine. Before removal of

detachable skins, fairings, and cowlings wash all exterior surfaces of the aircraft with plain water

and any commercial soap or detergent. Soap and detergent are organic chemicals and it is

important that all traces be removed by flushing with plain water.

*NOTE*

Certain chemicals cannot be removed

effectively by detergent solutions. Special

cleaning agents are available for that purpose. It

is suggested that the chemical suppliers be

contacted for cleaning agents that are suitable

for those special needs.

Inspection intervals are greatly influenced by particular operational priorities, operating conditions,

environment, and routine inspection results.

Due to the anticipated operating environment, servicing and overhaul interval should be in

accordance with Pratt & Whitney’s recommendations for the PT6A-41AG engine for the PT6A-41,

PT6A-41AG, and PT6A-42.

Due to the anticipated operating environment, servicing and overhaul interval should be in

accordance with Pratt & Whitney’s recommendations for the PT6A-34AG for the PT6A-34 engine.

Perform the tasks shown in the following Inspection Chart at the prescribed intervals.

2 - 18 Effective: 1/26/05


INSPECTION CHART

PROPELLER (Refer to Hartzell Manual #139 Propeller Owner’s Manual and Logbook.) D

aily

50

HR

S

100

HR

S

400

HR

S

1. Remove the spinner and check for cracks. X

2. Check the back plate for cracks and corrosion. X

3. Check for grease and oil leaks. X

4. Check the pitch rods and lock nuts. X

5. Check the condition of the reverse return springs. X

6. Check the hub bolts and balance screws of the blades for safety. X

7. Inspect the blades for nicks and cracks. Refer to the Hartzell Manual #139.

X

8. Inspect the hub parts for cracks and corrosion. X

9. Lubricate the propeller with Aeroshell 6 grease only. Remove the rear “Zerk” fitting from each blade clamp. Using a hand operated grease gun, grease each forward fitting slowly. Lubrication is complete when grease emerges in a steady flow with no air pockets or moisture, and has the color and texture of the new grease. Reinstalled the rear “Zerk” fittings.

X

Effective: 1/26/05 2- 19


Dai

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50

HR

S

100

HR

S

400

HR

S

10. Check counterweight bolts for safety. X

11. Check ring rod-end jam nuts. X

13. Re-install spinner. Rotate prop and check alignment of low pitch stop collar. (.010 max. run out) X

14. Check carbon block side clearance. New Block: .001”-.002” clearance Used Block: .010”-max allowed

X

15. Check beta control valve clevis slot end for alignment with face of cap nut. X

16. Check fuel governor reset arm for hitting stop. X

17. Check prop governor control levers for hitting stop in low and high pitch. X

18. Check reversing cable housing jam nuts and pins for safety and condition of housing. X

19. Inspect overspeed governor. X

2 - 20 Effective: 1/26/05


ENGINE EXTERNALS

REFER TO THE ENGINES’ APPROPRIATE PRATT & WHITNEY MAINTENANCE MANUAL FOR PERTINENT DETAILS ON ENGINE

INSPECTION

Dai

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50

HR

S

100

HR

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400

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S

1. Check the tubing, wiring, control linkages, and hose assemblies for security of all the accessible connections, clamps, and brackets.

X

INTENTIONALLY LEFT BLANK

2. Check tubing and hose assemblies for evidence of wear, chafing, cracks, and corrosion.

X

3. Check the tubing, wiring, control linkages, and hose assemblies for evidence of fuel and oil leakage. X

4. Lubricate interconnecting rod ball ends, where applicable.

X

5. Check the air inlet screen area for cleanliness.

X

6. Check the gas generator case for cracks, distortion, and corrosion. X

7. Check the fireseals for cracks and security of brackets and seals. X

8. Check the exhaust duct for cracks and distortion.

X

9. Check the propeller shaft seal for oil leaks.

X

10. Check security and condition of engine mounts.

X

Effective: 1/26/05 2- 21


Dai

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50

HR

S

100

HR

S

400

HR

S

11. Check the security of the accessories. X

12. Check the security of accessory linkages.

X

13. Check the security of pneumatic lines.

X

14. Check for evidence of oil and fuel leaks in accessory areas. X

16. Check security and mounting of starter/generator. Check brushes for wear.

X

17.

If starter/generator is equipped with a dry spline drive, remove the starter/generator and check the engine’s drive gear splines for wear using special P&W tool. If splines check OK, lubricate starter/generator drive shaft splines with Dow Corning M-77 Molycoat™ and reinstall starter/generator.

300

HR

S

*NOTE* As of 1998 all new AG engines will be delivered with a wet spline type starter generator gear shaft that requires no maintenance. You can identify an engine’s wet spline (female) gear shaft as follows: From the engine side (female) the splines are recessed approximately ¾ inch. Looking into the center of the gear shaft you can see right through to the AGB diaphragm wall. There will be a “0” ring installed on the starter/generator’s drive shaft (male splines).

INTENTIONAL LEFT BLANK

2 - 22 Effective: 1/26/05


ENGINE OIL SYSTEM

Dai

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S

100

HR

S

400

HR

S

** CAUTION **

Do not mix different brands or types of oil when changing oil or when replenishing the oil between oil changes.

1. Check the oil level. Oil change recommended every 400 hours. X X X

* NOTE * To avoid overfilling of oil tank, and high oil consumption, an oil level check is recommended within 30 minutes after engine shutdown. Ideal interval is 15 to 20 minutes. If more than 30 minutes has passed, and the dipstick indicates that oil is needed, start the engine and run at ground-idle (low idle) for five minutes, and recheck oil level.

2. Check condition and security of oil filler cap. X


3.

Remove, inspect, clean, and reinstall oil filter in accordance with instructions obtained in the engine’s appropriate Pratt & Whitney maintenance manual. NOTE: Do not clean ultrasonically. Elements must be discarded after 1000 hours or after heavy contamination.

X

4. Check the chip detector for continuity using a suitable ohmmeter. An open circuit condition must exist which indicates no ferrous contamination at pole tips.

X



Effective: 1/26/05 2- 23


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100

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400

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S





ENGINE FUEL SYSTEM

1. Check the fuel for presence of water. X X

2. Check the fuel pump for security and fuel leakage. X

3. Inspect, clean and reinstall high pressure fuel pump 74-micron inlet fuel filter. X

* NOTE * On new aircraft, check the filter after each flight until there is no evidence of contamination. Check the filter after the first flight or ground run when any upstream component is replaced.

4. At the fuel pump outlet, check the 10-micron filter for foreign matter and/or distortion. Install new filter every 100 hours or as service conditions indicate.

X

5. Check the drain valve for security and leakage. X X

2 - 24 Effective: 1/26/05


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100

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400

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S

*NOTE*

When a problem is found, refer to the appropriate section in the Pratt and Whitney Maintenance Manual.

6. Check the fuel control unit for security, linkages, and pneumatic tubes.

X

7. Check the fuel control unit bearing for wash-out which is indicated by blue dye - grease and fuel mixed – at the FCU vent.

X X

8.

Check the fuel manifold and nozzle assemblies with a functional test. For improved hot section durability. It is recommended that the fuel nozzle assemblies are inspected and functional tested in accordance with time limits set forth in the engines appropriate maintenance manual.

X X

IGNITION SYSTEM

1. Check the ignition exciter for security and condition. X

2. Check the ignition cable for chafing, wear, and security. X

3. Check the spark igniters for cleanliness and erosion. Perform an operation test. X

PNEUMATIC SYSTEM

1.

Clean the air compressor delivery filters (P3) every 100 hours. The maximum interval is 1000 hours for the disposable type. For the metallic cleanable type, return to an approved overhaul shop for ultrasonic cleaning every 1000 hours. After the cleaning, the element may be reused.

X

Effective: 1/26/05 2- 25


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100

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400

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AIRFRAME FUEL SYSTEM

1. Remove, inspect, clean, and re-install the airframe main 25-micron fuel strainer. X

2. Drain wing tanks, header tank; vent system (2 ea.), gascolator bowl, and residue fuel reservoir “EPA tank.” Check for any debris, sediment, or water and take corrective action if any is found.

X X

3. Turn the electric fuel pump on and check the fuel lines for leaks. X

4. Inspect the fuel lines and supports for security and signs of chafing. X

5. Check the fuel shutoff valve for leaks in the open and close position. X

6. Check the fuel tank gauges for proper operation. Rock the wings to slosh the fuel to see that the pointers are free.

X X

*NOTE*

When a problem is found, refer to section 5.

MAIN LANDING GEAR

1.

Remove, clean, and inspect main landing gear bolts, P/N AN10-33 or NAS6610D42, every 800 hours or annually, whichever occurs first, for cracks, wear, damage, corrosion, and general condition. Replace as necessary or, if no defects are noted at time of inspection reinstall. Replace main landing gear attach bolts every 1600 hours or every other annual, which ever occurs first.

*NOTE*

Prior to installing main landing gear bolts, apply a generously amount of MIL-G-81322 (Aeroshell 22) grease to main landing gear knurled bushings as well as to the bolts. Torque hardware I/A/W Torque Chart (Figure 2-7) and install cotter pins.

X Inspect

800 hours and

annually. Replaced

1600 hours or

biannually.

2.

Grease shock struts and shock strut attach to tripod zerk fittings with MIL-G-81322 (Aeroshell 22) grease. Inspect rubber shock biscuits for distortion, splits, or deterioration. If replacement of biscuits is required, see chapter 6 for pertinent data.

*NOTE*

The upper shock strut attach bolt is a close tolerance, heat treated NAS bolt. Do Not replace with an AN bolt.

X

2 - 26 Effective: 1/26/05


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100

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400

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S

3. Check the tires and tubes, wheels, and brake discs and lining for general condition. X X

4. Check the spindle for straightness and tightness. X

5. Check, inspect, lubricate with MIL-G-81322 (Aeroshell 22) grease, and reassemble all wheel bearings. (See chapter 6 for pertinent data.)

X

6. Check the master cylinders, parking brake valves, brake lines, brake calipers, all brake fittings, and brake bleeders for leakage, general condition, and security.

X

7. Check brake fluid level in each master cylinder and top off with fresh MIL-H-5606 aviation hydraulic fluid as required. X

8. Check the operation and holding ability of the pedal and parking brakes. Bleed hydraulic systems if required. X X

TAIL GEAR

1.

Remove, clean, and inspect leaf spring forward attach bolt P/N NAS6207-38D every 100 hours. Upon reassembly lubricate bolt and leaf spring hole with Snap-on™ General Purpose Antiseize or equivalent or MIL-G-81322 (Aeroshell 22) grease. Torque to specifications I/A/W Torque chart (figure 2-7). Replace MS24665-300 cotter pin each inspection.

X

2.

Inspect all bolts holes for elongation. As a general rule, replace components with holes that are out-of-round by 0.005” or more. Replacement of the leaf spring forward attach P/N NAS6207-38D bolt (Inspect every 100 hours) with a larger diameter is not approved. The leaf spring may not be “drilled out” for a larger bolt.

X

3.

Inspect main spring leaf for corrosion and cracks. Check aircraft maintenance records to be sure spring leaf P/N 5079-1 has not exceeded its five thousand (5,000) flight hour life limit. Replace leaf spring as needed.

X

Effective: 1/26/05 2- 27


Dai

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HR

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100

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400

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S

4. Inspect P/N95207-1 Acetal (Delrin®) lower support block spacer for wear and cracks. X

5.

Inspect upper and lower leaf spring support blocks, and attachment hardware for wear, corrosion, and cracks. Ensure that the leaf spring support blocks grips the leaf spring tightly to prevent leaf spring movement fwd. and aft. Ensure flexible sealant around contact edges of support blocks, lower support block spacer and leaf spring is intact to prevent collection of potential corrosive material in this area. Lubricate 2 ea. Trunnion Zerk (grease) fittings with MIL-G-81322 (Aeroshell 22).

X

6. Check locking cable for security and free movement, grease cable and wheel with MIL-G-81322 (Aeroshell 22) grease, and assure wheel bearing is completely greased.

X

7. Check for any loose play in tail wheel. X

8. Inspect the tire, wheel body and bearings, spindle, and the fork for general condition. X

9. Check the housing for cracks and corrosion. X

10. Check the taper bearings and spindle-shaft for corrosion and wear. X

11. Inspect the lock pin and plate for wear at the ends for correct operation. Check the lock pin cable and spring for corrosion and correct operation.

X

*NOTE* After the components have been installed, seal the contact edges where the spring, P/N 5079-1 (replace every 5,000 hours) upper support block P/N 94131-9, lower support block P/N 94131-11 and spacer P/N 95207-1 come together with a high quality flexible silicone sealant or fuel tank sealant CS3204 BS (AMS-S-8802 formerly MIL-S-8802) to help block the collection of potential corrosive contaminants in this area.

2 - 28 Effective: 1/26/05


FUSELAGE SKINS

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400

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1. Inspect all panels and cowlings for cracks, chaffing, and security of fasteners. X

2. Check the camloc receptacles for corrosion, wear, and locking action. X

HOPPER

1. Inspect the hopper baffles for security and condition. X

2. Check the hopper lid for condition of seal and security of latches. X

3. Inspect the hopper for evidence of leaks and for general condition. X

4. Check the gate for evidence of leaks and for proper operation. X

5. Check the hopper vent tube for corrosion and security. X

6. Check the gaskets on both the return and outlet lines. X

7. Check the hopper gate handle and the push rod for cracks around the welds. Check the condition of the push rod boot. X

8. Check emergency shut-off valve for leaks and proper operation X X

WINGS

1. Inspect the aileron brackets for cracks and security. X

Effective: 1/26/05 2- 29


Dai

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100

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400

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S

2. Check the boots at the aileron push rod entrance to the wing root for condition and security. X

3. Check for deposits of chemicals around and behind the wing center section and all attachment fittings. Check closely for corrosion. Keep clean.

X


WINGS

4. Inspect the wing skins for cracks, loose rivets, general condition of the paint, and corrosion.

X

5. Inspect the front and rear spar flanges, ribs, and other structures for cracks and corrosion. X

6. Check the spray booms attach points for security. X

7. Check the pitot line in the right wing for security and for air leaks. Drain the low spots. X

8. Inspect the wing/fuselage attach angles for signs of cracks and corrosion. X

9.

Annually inspect the lower spar splice blocks (P/N 22508T001 upper half and P/N 22508T002 lower half) as follows: Visually inspect splice blocks with a 10X magnifying glass or dye penetrant. Inspect for external cracks around the ¼ inch and 5/16 inch hole locations. If no cracks are detected this portion of the wing inspection is complete. If cracks are found remove the splice blocks before next flight and inspect the lower spar cap for cracks in accordance with Thrush Aircraft Inc. Service Bulletin SB-AG-39. If cracks are found in spar cap contact Thrush Aircraft Inc for possible repair or replacement. If no cracks are found in spar cap, replace the cracked splice blocks with new units. Refer to Section VIII “Wing Removal” for splice block removal and installation.

Annu

ally

2 - 30 Effective: 1/26/05


FUSELAGE FRAME

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400

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1. Inspect the fuselage tubing for signs of corrosion or cracks, particularly around welds and in the hopper area. X

2. Check for elongated holes in the engine mount fittings and bellcranks. X

3. Inspect all spring gear attachment fittings, main gear support beam, and beam end plates for security, cracks, and corrosion.

X

4. Check the condition of the paint and refinish, if necessary. X

CONTROL SYSTEMS

1. Check all turnbuckles for corrosion and for proper lock wiring. X


2. Inspect all cables and end fittings for wear. Check for correct tension. X

3. Check all push rods for loose bearings, endplay, straightness and paint condition. X

4. Check idlers and bellcranks for binding or for slack. X

5. Inspect the rudder pedals and the support brackets for general condition. X

6. Inspect the attachment of the control stick to the main torque tube for slack and bearing wear. X

Effective: 1/26/05 2- 31


2 – 32 Effective: 1/26/05

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400

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7. Check control stick to main torque tube bolt for proper torque (65 to 70 in. lbs.) X

8. Check the aileron control stops for tightness and for condition of fittings. X

9.

Inspect all push-pull tubes rod-end jam nuts for security. Inspect all witness/inspection holes with a piece of .032” safety wire to insure that all rod-ends are screwed far enough onto the push-pull tubes.

X

10. Inspect the push rods for clearance to the structure. X

11. Inspect the trim systems for correct operation and for general condition. X

12. Remove control stick from main the torque tube bolt, inspect and replace bolt as required.

X 500

Hours


METAL EMPENNAGE

1.

Check the travel of the movable surfaces. Elevator up---------------27 degrees ±1 degree Elevator down-----------17 degrees ±1 degree Rudder--------------------24 degrees ±1 degree Tab up----------------------8 degrees ±1degree Tab down------------------22 degrees ±1 degree

X

2. Check for warped contours of the fixed surfaces due to improperly tightened brace struts. X X


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100

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400

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S

3. Inspect horizontal stabilizer “V” struts, fittings, and hardware for security, cracks and corrosion. X

4. INTENTIONALLY LEFT BLANK

5. Inspect all hinges for wear. Replace sealed bearings, if needed. X

6. Check security of all bolts. X

7. Check the external skins for general condition. X

8. Check the drain holes for obstruction. X


AILERONS AND FLAPS

1.

Check the control movements.

Aileron up 21 degrees ±1 degree

Aileron down 17 degrees ±1 degree

Flap down 15 degrees ±1 degree

X

2. Check the security of the counterweights, which are installed in the leading edges of the ailerons. X


Effective: 1/26/05 2- 33


2 – 34 Effective: 1/26/05

Dai

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S

100

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S

400

HR

S

3. Inspect all the skins and ribs for cracks, loose rivets, general condition, and corrosion. X

4. Inspect the flap push rods, mounting brackets, torque tube, and bearing housings. X X

5. Inspect the flap actuator motor and worm drive for general condition and freedom of travel. Grease Zerk fitting with (MIL-G-81322 (Aeroshell 22) on bottom of unit.

X

6.

Aileron servo tabs

a. Check security of hinges

b. Check for looseness of rod ends and bolts.

c. Check for freedom of travel.

X




COCKPIT

1. Check the condition of the instrument markings and the placards. X X

2. Check the instrument lines for leaks, security, and chafing. X

3. Check the hopper for leaks and security of mechanism. X

4. Check the security and condition of the seat belts, shoulder harness, and inertia reels. X


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100

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400

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5. Check the seat for security and proper adjustment operation. Check the seat fabric for general condition. X

6. Check the windshield and windows for cracks, crazing or scratches, and missing screws. X

7. Check the doors for security of hinges and for correct operation of door locks. X

8. Check operation of flight & engine controls to ensure proper operation and installation. X X

ELECTRICAL SYSTEM

1. Check the battery charge and water level. X X




2. Check battery relays, spike diodes, regulator, fuses, and switches for security. X

3. Check all wiring for chafing and clamping. X

4. Check all terminals for security and corrosion X

5. Check the battery’s vent hoses for security and deterioration. X

Effective: 1/26/05 2- 35


AIRFRAME MAINTENANCE

CORROSION CONTROL

The lower part of the aircraft is painted with ultra gloss polyurethane. The forward upper glare shield

part is painted with flat black polyurethane. The fuselage frame is painted with a primer, and then

painted with a gray ultrathane.

All repairs involving refinishing should be painted to the original specifications. The following

procedures should be carried out step by step.

A. Sand part to bare metal using 180 paper or finer. Avoid removal of cladding with the Alclad

parts, whenever possible.

B. Thoroughly clean area with isopropyl alcohol, a solvent, or thinner. Allow time to dry.

C. Apply one thin spray coat of Epoxy primer with Epoxy hardener. Allow time to dry.

D. Mix the required quantity of Polyurethane (use the directions on the can) with the prescribed

amount of activator. Spray a smooth and even coat directly onto the primed surfaces. Apply at

least two coats and allow time for drying between the coats.

A regular and thorough cleaning of both the interior and exterior of the aircraft is a major part of

corrosion control. All areas of the aircraft are accessible for cleaning by removal of the panels. The

cleaning procedure that follows is recommended for general purposes.

A. Wash all exterior surfaces of the aircraft with plain water and any commercial soap or detergent.

Soap and detergent are organic chemicals, and it is important that all traces be removed by

flushing with plain water.

B. Detach all removable panels from the aircraft. Wash down the rear fuselage aft of the wing

trailing edge. Tube joints, skin bends, and so forth should receive particular attention. Remove

excess moisture after flushing.

2 - 36 Effective: 1/26/05

* NOTE * Certain chemicals cannot be removed effectively by detergent

solutions. Special cleaning agents are available for that purpose. It is

suggested that the chemical suppliers be contacted for cleaning agents

that are suitable for those special needs.


C. The forward fuselage and engine section should not be cleaned with water unless close attention

is made to avoid removal of lubricants and to avoid possible rusting of components and

hardware. A general purpose, non-corrosive cleaning agent is preferred in those areas.

D. Particular attention should be given to the wing center splice fittings and the attachments of the

oil cooler, hopper and engine mount.

E. Hopper cleaning should be accomplished at the end of each working day. A good commercial

detergent should be used and followed by a thorough flush with water. Leave the hopper door

and gate open for thorough drying.

WINDSHIELD

An anti-static type of plastic cleaner, such as Mirror Glaze or equivalent, is recommended for best

cleaning. The side windshields are plastic and should not be cleaned with gasoline, alcohol, acetone,

and lacquer thinner, or window cleaning spray. Those fluids will soften the plastic and cause crazing.

Avoid rubbing the plastic surface with a dry cloth, as that can cause scratches and build up an

electrical charge (static) which will attract dust particles. If scratches are visible after removing the

dust accumulation, finish the plastic with a quality grade of commercial wax. Apply the wax in a thin,

even coat and carefully buff out with a soft cloth. Do not buff or polish in one area for more than a

brief period of time. The heat generated by rubbing the surface may soften the plastic and may

produce visual distortion.

The middle section of the windshield is safety plate glass for better resistance to scratching. It is

enclosed in an aluminum frame.

HOPPER REPAIR

Hopper repair may be accomplished in accordance with the instructions contained in Section IX.

FUEL TANK REPAIR

Fuel tank repair may be accomplished in accordance with the instructions contained in Section V.

BATTERY MAINTENANCE

The 24-volt batteries are installed in the engine compartment between the engine and firewall.

Access is gained to the batteries by removal of a cowling. Battery servicing involves adding distilled

water to maintain electrolyte level of 3/16 inch over the separators, checking the cable connections,

and neutralizing or cleaning any spilled electrolyte or any corrosion. Use bicarbonate of soda and

clean water to neutralize corrosion. Follow with a thorough flushing of clean water and wipe dry.

Effective: 1/26/05 2- 37


Clean the cable and terminal connections with a wire brush and coat with petroleum jelly to minimize

corrosion.

A hydrometer test of the battery’s solution should be made each 50 hours of operation, or more often

in hot weather. If the specific gravity tests 1.240, the battery should be removed and recharged. The

solution levels should be examined and, when necessary, add distilled water to maintain the level of

3/16 inch over the separators. If distilled water is added, do it just prior to recharging so that the

added water mixes with the solution. When the recharging is completed, the specific gravity should

be between 1.275 and 1.300.

The battery should be checked for grounding to the case. A voltmeter can be used to check between

the positive cell and the case. A ground fault exists if there is a reading on the voltmeter. A dated

service record shall be attached or stamped on the terminal side of the battery to indicate that the

battery has been capacity tested. Refer to Section Ten for recharging procedures.

LUBRICATION

For the lubrication requirements, refer to Figure 2-6 Lubrication Chart (8 sheets). Before adding

grease to fittings, wipe the fittings clean. Lubricate the fittings and wipe off the excess lubricant.

Lubricate the hinges with a squirt can or a brush moistened with oil. Wipe off the excess oil to prevent

accumulation of dirt and grit.

2 - 38 Effective: 1/26/05

** CAUTION **

Do not allow the bicarbonate of soda to enter the

battery filler openings, as it will neutralize the

electrolyte, which could permanently damage the

batteries.


Effective: 1/26/05 2- 39

Figure 2-1


Figure 2-2

2 - 40 Effective: 1/26/05


Effective: 1/26/05 2- 41

Figure 2-3


2 - 42 Effective: 1/26/05

Figure 2-4


Effective: 1/26/05 2- 43


** WARNING ** The drawings of Figure 2-6, sheets 1 thru 8, are for lubrication reference only. They do not show proper assembly details and may not be used as assembly reference. Refer to the appropriate parts manual for details concerning parts assembly.

APPLICATION SYMBOL SPECIFICATIONS AND TYPE OF LUBRICATION

HAND PACK

MIL-G-81322 (AEROSHELL 22) AIRCRAFT GREASE

LUBRICATION GUN

MIL-G-81322 (AEROSHELL 22) AIRCRAFT GREASE

OIL CAN

MIL-L-22851 (AEROSHELL OIL W 15W50) OR

EQUIVALENT – LUBRICATING OIL

* NOTE*

Use only Aeroshell 6 in propeller.

2 - 44 Effective: 1/26/05

Figure 2-6 Lubrication Chart

Sheet 1 of 8



Sheet 2 of 8

Effective: 1/26/05 2- 45



Sheet 3 of 8

2 - 46 Effective: 1/26/05



Sheet 4 of 8

Effective: 1/26/05 2- 47



Sheet 5 of 8

2 - 48 Effective: 1/26/05



Sheet 6 of 8

Effective: 1/26/05 2- 49



Sheet 7 of 8

2 - 50 Effective: 1/26/05


Effective: 1/26/05 2- 51


Sheet 8 of 8


FINE THREAD SERIES

BOLTS BOLTS Steel Tension Steel Tension AN 3 thru AN 20

AN 42 thru AN 49 AN 73 thru AN 81 AN 173 thru AN 186 AN 20033 thru MS 20046 MS 20073 MS20074 AN509 NK9 MS 24604 AN 525 N K525 MS27039

MS 20004 thru MS 20024 NAS 144 thru NAS 158 NAS 333 thru NAS 340 NAS 583 thru NAS 590 NAS 624 thru NAS 644 NAS 1103 thru NAS 1120 NAS 1202 thru NAS 1210 NAS 1303 thru NAS 1320 NAS 6203 thru NAS 6220 NAS 6603 thru NAS 6620 NAS 172 NAS 174 NAS 517

Steel Shear NAS 454 NUTS NUTS Steel Tension Steel Shear Steel Tension Steel Shear AN 310

AN 315 AN 363 AN 365 NAS 1021 MS 17825 MS 21045 MS 20365 MS 20500 NAS 679 MS 21042 MS 21044N MS 21046

AN 320 AN 364 NAS 1022 MS 17826 MS 20364 MS 21083N MS 21245

AN 310 AN 315 AN 363 AN 365 MS 17825 MS 20365 MS 21045 NAS 1021 NAS 679 NAS 1291 MS 21042 MS 21044N MS 21046

AN 320 AN 364 NAS 1022 MS 17826 MS 20364 MS 21083N MS 21245

Nut-bolt size

Torque Limits in-lbs


Nut-bolt size



Min. Max. Min. Max. Min. Max. Min. Max. 8-36 12 15 7 9 8-36 ----- ----- ----- -----

10-32 20 25 12 15 10-32 25 30 15 20 ¼ -28 50 70 30 40 ¼ -28 80 100 50 60

5/16 - 24 100 140 60 85 5/16 - 24 120 145 70 90 3/8 – 24 160 190 95 110 3/8 – 24 200 250 120 150 7/16-20 450 500 270 300 7/16-20 520 630 300 400 ½ - 20 480 690 290 410 ½ - 20 770 950 450 550

9/16 – 18 800 1,000 480 600 9/16 – 18 1,100 1,300 650 800 5/8 – 18 1,100 1,300 660 780 5/8 – 18 1,250 1,550 750 950 ¾ - 16 2,300 2,500 1,300 1,500 ¾ - 16 2,650 3,200 1,600 1,900

7/8 – 14 2,500 3,000 1,500 1,800 7/8 – 14 3,550 4,350 2,100 2,600 1 – 14 3,700 4,500 2,200 3,300 1 – 14 4,500 5,500 2,700 3,300

1 1/8 -12 5,000 7,000 3,000 4,200 1 1/8 -12 6,000 7,300 3,600 4,400 1 ¼ - 12 9,000 11,000 5,400 6,600 1 ¼ - 12 11,000 13,400 6,600 8,000

Figure 2-7 Torque Chart

2 - 52 Effective: 1/26/05


Figure 2-8

Effective: 1/26/05 2- 53


Section 3 HYDRAULICS

TABLE OF CONTENTS SECTION THREE............................................................................................................................................. 1 HYDRAULICS................................................................................................................................................... 1 HYDRAULIC SYSTEM ..................................................................................................................................... 2 GENERAL DESCRIPTION...............................................................................................................................2

Effective: 1/26/05 3- 1


HYDRAULIC SYSTEM

SECTION THREE GENERAL DESCRIPTION

The S2R-T34 aircraft has two individual hydraulic systems using MIL-H-5606 fluid. The main landing gear utilizes a master brake cylinder for the operation of the landing gear brakes and parking brakes. The master brake cylinder is connected to the disc type brake calipers by brake lines that are supported by and clamped to the airframe structure forward of the master brake cylinder. The hydraulic brake lines are of rigid steel tubing, except for the flexible hoses on the landing gear assembly. The master brake cylinder is installed aft of the rudder-brake pedals and is actuated by toe pressure on the pedals. As toe pressure is applied to the pedals, the push rod, piston and spring are pressed into the master brake cylinder. This compresses hydraulic fluid in the lines and applies pressure to the appropriate brake. Operate individual parking brakes as follows: ON – Depress rudder pedal, pull parking valve lever, take pressure off of rudder pedal. OFF – Depress rudder pedal, valve will deactivate and lever will pop in.

Effective: 1/26/05 3 - 2


Section 4 POWERPLANT AND PROPELLER

TABLE OF CONTENTS POWERPLANT AND PROPELLER .............................................................................................…2

GENERAL DESCRIPTION ...........................................................................................................2 AIR CLEANING SYSTEM.............................................................................................................2 POWER PLANT............................................................................................................................2

DESCRIPTION AND OPERATION...........................................................................................2 ENGINE BUILDUP ....................................................................................................................4 ENGINE REMOVAL ..................................................................................................................6

ENGINE INSTALLATION............................................................................................................10 PROPELLER...............................................................................................................................13

DESCRIPTION AND OPERATION.........................................................................................13 PROPELLER AND BETA FEEDBACK MECHANISM ............................................................13 CONSTANT SPEED UNIT (CSU)...........................................................................................14 PROPELLER OVERSPEED GOVERNOR .............................................................................18 ENGINE POWER AND PROPELLER CONTROLS (FCU).....................................................18 PROPELLER SPEED SELECT AND FEATHERING CONTROL...........................................20 PROPELLER SETTINGS........................................................................................................20

PROPELLER MAINTENANCE ...................................................................................................21 PROPELLER REMOVAL ........................................................................................................21 PROPELLER INSTALLATION ................................................................................................22 CONSTANT SPEED UNIT (CSU) REMOVAL.........................................................................24 CONSTANT SPEED UNIT (CSU) INSTALLATION ................................................................24 PROPELLER CSU HIGH RPM ADJUSTMENT......................................................................25 PROPELLER OVERSPEED GOVERNOR REMOVAL ..........................................................26 PROPELLER OVERSPEED GOVERNOR INSTALLATION ..................................................26

ENGINE CONTROLS .................................................................................................................26 RIGGING INSTRUCTIONS.....................................................................................................26 AIRFRAME CONTROL LINKAGES ........................................................................................26 PROPELLER REVERSING INTERCONNECT LINKAGE ......................................................26 FRONT LINKAGE ...................................................................................................................27 REAR LINKAGE......................................................................................................................27 CONDITION LEVER LINKAGE...............................................................................................28 PROPELLER RIGGING ..........................................................................................................29 ENGINE RIGGING CHECKS AND ADJUSTMENTS..............................................................30 LOW (GROUND) IDLE ADJUSTMENTS ................................................................................31 HIGH (FLIGHT) IDLE ADJUSTMENTS ..................................................................................31 PROPELLER GOVERNOR CHECK .......................................................................................32 MAX PROPELLER SPEED CHECK .......................................................................................32 OVERSPEED GOVERNOR CHECK (PROP TEST) ..............................................................32 REVERSE MAX POWER CHECK ..........................................................................................34 MAXIMUM GAS GENERATOR SPEED CHECK....................................................................34 ENGINE PERFORMANCE CHECK........................................................................................35 COCKPIT ENGINE CONTROL QUADRANT..........................................................................36 ENGINE LIMITS ......................................................................................................................37 ENGINE DATA PERTINENT TO THRUSH AIRCRAFT INC INSTALLATION .......................37

Effective: 1/26/05 4 - 1


POWERPLANT AND PROPELLER

SECTION ONE

GENERAL DESCRIPTION

The S2R-T34 Turbo Thrush agricultural airplane utilizes the following Powerplants:

STANDARD: Pratt & Whitney Canada PT6A-34AG OPTIONAL: Pratt & Whitney Canada PT6A-34 Pratt & Whitney Canada PT6A-36 (Dry Configuration Only) Pratt & Whitney Canada PT6A-41AG Pratt & Whitney Canada PT6A-41 Pratt & Whitney Canada PT6A-42

AIR CLEANING SYSTEM

The prime difference between the agricultural application and a normal installation is the air

cleaning system incorporated in the engine air intake system. Air is inducted into the engine by a

screened opening in the lower cowling or by an optional pitot inlet built into the lower nose bowl.

The air filter panel is a K & N cleanable barrier filter. It provides high efficiency, maximum reliability,

long service life and low overall cost.

The barrier filter unit is made of a cotton mesh with a light coat of K & N special red oil to assist in

collecting dust. The filter can be removed, cleaned and reserviced I/A/W cleaning instructions from

K & N P/N 99-5000 (aerosol) or P/N 99-5050 Recharger filter care service kit, obtained locally.

POWER PLANT DESCRIPTION AND OPERATION

The PT6A-34AG (See Figure 4-0) series power plant is a lightweight free turbine engine. Each

engine utilizes two independent turbine sections: one driving the compressor in the gas generator

section and the second driving the propeller shaft through a reduction gearbox. The engine is self-

sufficient, since its gas generator driven oil system provides lubrication for all areas of the engine,

pressure for the torquemeter and power for propeller pitch control.

Inlet air enters the engine through an annular plenum chamber, formed by the compressor inlet

case where it is directed forward to the compressor. The PT6A-34AG compressor consists of three

axial stages combined with a single centrifugal stage, assembled as an integral unit.

The engine is equipped with a wash ring at the compressor air inlet screen. A line running from this

wash ring to a port on the outside of the cowling gives the capability to cleanse the compressor

section without engine cowling removal.

Effective: 1/26/05 4 – 2


A row of stator vanes, located between each stage of compression, diffuses the air, raises its static

pressure and directs it to the next stage of compression. The compressed air passes through

diffuser tubes, which turn the air through ninety degrees in direction and convert 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 is an annular, heat resistant alloy; domed at the front end where it is

supported inside the gas generator case by the 14 fuel manifold adapter sheaths and both igniters.

The rear end of the combustion chamber is open and is supported by the large and small exit ducts.

The liner assembly has perforations of various sizes that allow entry of compressor delivery air.

The flow of air changes direction 180 degrees 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 liner

eliminates the need for a long shaft between the compressor and the compressor turbine, thus

reducing the overall length and weight of the engine.

Fuel is injected into the combustion chamber liner through 14 simplex nozzles supplied by a dual

manifold consisting of primary and secondary transfer tubes and adapters. Two spark igniters that

protrude into the liner ignite the fuel/air mixture. 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 impinge on the

turbine blades at the most optimum angle, with minimum loss of energy.

The still expanding hot gases from the gas generator are still directed forward to the power turbine

inlet guide vane which directs, at the most optimum angle, the gas flow onto the power turbine

which drives the propeller shaft via a two-stage reduction gear box.

The compressor and power turbines are located in the approximate center of the engine with their

respective shafts extending in opposite directions. This feature provides for simplified installation

and 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.

Interturbine 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 cockpit

instrumentation.

All engine-driven accessories, with the exception of the propeller governor, overspeed governor

and NP tachometer-generator, are mounted on the accessory gearbox at the rear of the engine.

4 – 3 Effective 1/26/05


These components are driven by the compressor by means of a coupling shaft, which extends

the drive through a tube at the center of the oil tank. The rear location of accessories provides for

a clean engine and simplifies maintenance procedures. 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 total capacity of 2.3 US gallons and is provided with a

dipstick.

An engine-driven fuel pump further pressurizes fuel supplied to the engine from an external source

and the fuel control unit (FCU) controls its flow to the fuel manifold.

The power turbine drives a propeller through a two-stage planetary reduction gearbox located at the

front of the engine. The gearbox embodies an integral torquemeter device, which is instrumented

to provide an accurate indication of engine power. A chip detector is installed at the bottom of the

gearbox.

The propeller reversing installation is comprised of a single-acting hydraulic propeller that is

controlled by a propeller governor which combines the functions of a normal constant speed unit

(CSU), a reversing valve and a power turbine (Nf) governor. A mechanical linkage between the

propeller governor Beta control valve and the air bleed link enables the FCU and the propeller

governor to modify engine power to maintain power turbine speed at a speed slightly less than the

selected rpm when operating in the Beta control range.

ENGINE BUILDUP

Engine build-up consists of the removal of accessories and equipment from the old engine and

installing them on the new engine. Consult the Engine Maintenance Manual for removal and

replacement procedures. After all accessories and equipment have been installed on the new

engine, proceed as follows:

Effective: 1/26/05 4 – 4

** CAUTION **

Consult the Engine Maintenance Manual before removing the

new engine from the shipping container.


A. Remove the engine control brackets and supports from the old engine and install on the

new engine. Consult the Engine Maintenance Manual for the proper torque values.

4 – 5 Effective: 1/26/05

** CAUTION **

If the old engine is being removed because of oil

contamination or of the possibility of oil contamination, scrap

the following items: (a) oil cooler and (b) all oil carrying lines

and hoses.

If the old engine has oil contamination, the following items

must be sent to an appropriate maintenance facility for

disassembly and flushing to remove all contaminants or they

must be replaced: (a) overspeed-governor, (b) propeller, (c)

fuel/oil heat exchanger (d) propeller governor (C.S.U.).

(NOTE: The fuel/oil heat exchanger and propeller governor

normally comes with the new engine.) Failure to comply with

the above will prove to be false economy, as the new engine

will be contaminated by old impurities.

* NOTE *

Tag or identify all hoses, bolts, nuts, and electrical connector

plugs and note harness clamp locations for installation on the

new engine. Cap all open hoses and engine ports to prevent

contamination.


B. Remove the engine mounts from the old engine and install on the new engine, using

the same bolts, washers, and gaskets. Torque the bolts to 250-325 inch pounds and

secure with safety wire. (Ref. Figure 4-4)

C. Remove the exhaust stacks from the old engine and install on the corresponding (left or

right) exhaust ports of the new engine. Torque the bolts to 50-70 inch-pounds.

ENGINE REMOVAL

A. Preliminary steps:

1. Turn fuel shut off handle to close fuel shut off valve.

2. Make sure all electrical power to the aircraft is disconnected.

3. Provide suitable containers under the engine to catch fuel and oil spillage.

B. Remove engine cowlings.

Effective: 1/26/05 4 – 6

* NOTE *

If the engine mounts are removed for replacement, they must

be all the same part numbers. Torque the engine mount to

engine mount truss bolts to 480-600 inch-pounds.

** CAUTION **

To prevent damage to internal mechanisms, engines expected

to be idle for more than seven days, due to maintenance or

other reasons, should be preserved in accordance with the

engine manufacturer's recommendations as outlined in the

Engine Maintenance Manual.


C. Disconnect battery.

D. Remove propeller.

E. Remove exhaust ducts.

F. Remove cannular inlet cover (3 places) from compressor inlet.

G. Securely cover the engine compressor inlet screen to prevent entry of foreign material.

H. Disconnect the following tube and hose assemblies at the locations noted:

1. Oil cooler hoses.

2. Gas generator case front drain valve hose.

3. Torque system lines at forward fire seal.

4. Gas generator case rear drain valve line.

5. Delta “P-Lines.”

6. Fuel inlet, outlet hoses, and vent line (three places) at the engine driven fuel boost

pump.

7. Oil cooler hoses (2) at the engine.

8. Compressor wash ring tube assembly at the union forward of aft fireseal.

4 – 7 Effective: 1/26/05

* NOTE * Tag and identify all tube and hose assemblies to facilitate

and ensure correct installation of the engine. Cap and plug

all openings to prevent contamination.


9. Fuel inlet manifold adaptors dump tube at front fireseal.

10. Fuel inlet hose at the oil-to-fuel heater.

11. Fuel purge hose at the start control unit or rear of F.C.U. on engines not equipped

with a start control.

12. Oil pressure line at engine.

13. Disconnect the engine overboard breather hose.

14. High-pressure fuel pump drain.

15. Fuel pressure line from rear of oil to fuel heater.

16. Torque indicating systems hoses at aft fireseal.

17. P2.5 overboard bleed ducting from engine belly-band and aft fireseal.

I. Disconnect the electrical leads and connector plugs at the locations noted. Remove

electrical harness clamps, as necessary, to allow engine removal.

1. ITT harness at the T5 terminal block.

2. Overspeed governor prop test solenoid.

3. Prop beta micro switch.

4. Np tachometer generator.

Effective: 1/26/05 4 – 8

* NOTE *

Tag or identify all electrical leads and connector plugs. Note

harness clamp locations to facilitate and ensure correct

installation. Cap all plugs and receptacles to prevent

contamination.


5. Tq pressure transmitter (NOTE: This not installed on aircraft with direct reading Tq

gauge.).

6. Ng tach generator.

7. Oil temperature sending unit.

8. Ignition leads at exciter box.

9. Starter/generator terminal block.

10. Fuel flow transducer (If equipped).

11. Engine ground cable from rear of engine driven boost pump.

12. P3 heat terminal.

J. Disconnect the engine controls.

1. Disconnect the propeller rod end at the propeller governor control lever and

remove the cable from the forward fireseal.

2. Remove the prop cable from the forward fireseal.

4 – 9 Effective: 1/26/05

* NOTE *

Tag and retain all attaching control cable parts for engine

installation. Note clamp locations to facilitate control cable

installation.

* NOTE * After wiring harness has been removed from above items,

remove the grommet at basket assembly aft. Close out

and carefully pull the harness aft and clear of basket.

Secure harness to prevent damage until ready to re-install.


3. Disconnect the condition lever control push-pull tube rod end at the lever on the

start control unit or on top of F.C.U. on engines not equipped with start control

units.

4. Disconnect the power control cable rod end at the power-input lever.

K. Remove engine mount cuffs at aft fireseal (8 locations).

L. Remove forward and rear fireseals.

M. Remove forward engine mount basket assembly.

N. Remove the engine unit from the aircraft as follows:

1. Attach the engine sling to the engine hoisting lugs. Position a suitable hoist directly

over the engine and attach to the engine sling.

2. Raise the hoist sufficiently to take the weight of the engine.

3. Remove the cotter pins and attaching hardware, which attaches the engine

vibration, mounts to the mounting, trusses.

4. Remove the bolts and washers attaching the mounts to the engine mount truss.

O. Hoist the engine unit clear of the fuselage nose section and install in a suitable stand.

Remove the engine sling.

ENGINE INSTALLATION

A. Install the engine unit in the aircraft as follows:

1. Attach the engine sling to the engine hoisting lugs. Position a hoist directly over

the engine and attach to the engine sling.

2. Remove the engine from the stand and carefully position in the engine mount.

3. Align the boltholes of the engine vibration mounts with those of the engine mounts.

Install the attaching hardware. Torque the bolts to 480-600 inch-pounds and install

cotter pins. (Refer to Figure 4-4.)

4. Install forward engine mount basket assembly.

5. Install forward and rear fireseals.

6. Install engine mount cuffs at aft fireseal (8 locations).

7. Seal all mating joints to assure proper sealing of cannular inlet and filter area with

RTV sealant.

Effective: 1/26/05 4 – 10

** CAUTION **

Before hoisting the engine unit clear of the fuselage nose

section, check that all wiring, cables, and tube and hose

assemblies are disconnected and free from snagging.


B. Connect the following tube and hose assemblies at the locations noted:

1. Oil cooler augmentation lines and hoses.

2. Gas generator case front drain valve hose.

3. Torque system lines at forward fireseal.

4. Gas generator case rear drain valve line.

5. Fuel inlet, outlet hoses, and vent line (three places) at the engine driven fuel boost

pump.

6. Oil cooler hoses (2) at the engine.

7. Compressor wash ring tube assembly at the union forward of aft fireseal.

8. Fuel inlet manifold adaptor dump tube at front fireseal.

9. Fuel inlet hose at the oil-to-fuel heater.

10. Fuel purge hose at the start control unit or rear of F.C.U. on engines not equipped

with a start control.

11. Oil pressure line at engine.

12. Engine overboard breather hose at engine.

13. High-pressure fuel pump drain.

14. Fuel pressure line from rear of oil to fuel heater.

15. Torque indicating systems hoses at aft fireseal.

16. P2.5 overboard bleed ducting to belly-band and aft fireseal.

C. Connect the electrical leads and connector plugs at the locations noted:

1. Engine ground cable to rear of engine-driven boost pump.

2. Fuel flow transducer (if equipped).

3. Starter/generator terminal block.

4. Ignition leads at exciter box.

5. Oil temperature-sending unit.

6. Ng tach generator.

7. Tq pressure transmitter (if equipped).

4 – 11 Effective: 1/26/05

* NOTE *

Route forward electrical harness through forward

close out. Install grommet in slot provided. Secure

harness to basket structure as previously noted.

Connect harness to items as follows.


8. Np tach generator.

9. Prop beta micro switch.

10. Overspeed governor prop test solenoid (Be sure to install the two ground wires on

the mounting stud.)

11. ITT harness at the T5 terminal block.

12. P3 heat terminal.

D. Connect the engine controls.

1. Attach the propeller control cable housing to the forward fireseal and connect the

propeller control cable rod end to the propeller governor control lever.

2. Connect the fuel condition control cable rod end at the FCU condition lever or start

control’s control lever on engines equipped with a start control unit.

3. Connect the power control cable rod end at the FCU power input lever.

E. Install the propeller.

F. Rig the engine controls.

G. If necessary, refer to the Engine Maintenance Manual for depreservation of the engine oil

and fuel systems.

H. Service the engine oil system.

I. Remove the cover from the compressor inlet screen.

J. Install the engine cowling.

K. Perform the engine ground test and checks. (Refer to procedures outlined later in this

section and Pratt & Whitney Maintenance Manual.)

Effective: 1/26/05 4 – 12

* NOTE * Clean terminal ends and torque the ITT harness

connections in accordance with the Engine

Maintenance Manual.

** CAUTION **

Prior to engine run-up, ensure the engine air inlet

plenums are free of foreign objects.


PROPELLER DESCRIPTION AND OPERATION

This section describes the function of the following:

- Propeller and Beta Feedback Mechanism

- Constant Speed Unit (CSU)

- Propeller Overspeed Governor

- Engine Power and Propeller Controls (FCU)

- Propeller Speed Select and Feathering Control

- Propeller Settings

*NOTE* The Fuel Control Unit (FCU) is not included as

it is covered in the Engine Maintenance Manual

PROPELLER AND BETA FEEDBACK MECHANISM

The propeller has three blades mounted on a hollow hub in the front end of which is a servo-piston

that moves forward under servo-oil pressure or rearward under feather returns spring pressure.

(Ref. Fig. 4-5) There are three links from the servo-piston. One goes to each blade root, and these

links transmit forward motion of the servo-piston to the blade roots and pivot the blades in the

decrease pitch direction. When servo-piston pressure is relieved, the servo-piston moves rearward

under feather return spring pressure and pivots the blades in the increase pitch direction. This

action is assisted by centrifugal force of the counterweight on each blade root.

Servo-oil is supplied from the constant speed unit (CSU). It flows through oil passages in the engine

reduction gear case through a transfer tube between the reduction gear case and propeller oil

transfer housing; then via the propeller oil transfer housing, the engine shaft, the hollow hub, and

the internal oil ports in the servo-piston. Refer to Figure 4-5.

The beta feedback mechanism has three low pitch stop rods (Fig. 4-7, #32) that are screwed into

the propeller feedback ring (Fig. 4-3, #4). These three rods slide fore and aft in small bushings

mounted in a flange integral with the hollow hub.

Near the forward end of each low stop rod is a beta nut (Fig. 4-7, #27). Ahead of these is the ring

rod end (Fig. 4-7, #33) which steadies the low stop rods. As the servo-piston moves forward, it

picks up on the beta nuts at a certain preset blade pitch. From that instant the propeller feedback

4 – 13 Effective: 1/26/05


ring (Fig. 4-7, #4) moves forward with the servo-piston. As it moves, the reverse return springs (Fig.

4-7, #34) are compressed. During the return motion, when the servo-piston moves rearward, the

reverse return springs maintain contact between the beta nuts (Fig. 4-7, #27) and the servo-piston

by pushing aft on small plates attached to each low stop rod.

This forward and reverse movement of the propeller feedback ring is used to monitor blade pitch

change during beta and reverse. The motion is transmitted to the beta control valve in the CSU via

the carbon block (Fig. 4-7, #2) and the propeller reversing lever (Fig. 4-7, #3).

As the propeller reversing lever pivots back and forth, it opens or closes the beta control valve (Fig.

4-7, #18) which is attached to the middle of the propeller reversing lever.

The beta feedback mechanism has two uses.

A. It enables the aircraft pilot to select blade angle directly during beta and reverse.

B. It allows provision of a hydraulic low pitch stop during flight.

CONSTANT SPEED UNIT (CSU)

For clarity and ease of understanding, the CSU is described in five different sections:

A. Servo-oil Supply

B. Constant Speed Section

C. Power Turbine Governing Section

D. Beta Control Valve Section

E. Feathering

A. Servo-oil Supply

The servo-oil which is used to vary the propeller blade angle is supplied by the CSU. Refer to

Figure 4-6. An oil pump in the base of the CSU boosts the engine oil pressure to approximately

385-PSI. The oil is then routed past a pressure relief valve (Fig. 4-6, #7) through the beta control

valve port (Fig. 4-6. #11) to a chamber formed by the hollow drive shaft (Fig. 4-6, #15) and the

lower part of the pilot valve plunger (Fig. 4-6, #16). Here it is ready for delivery to the propeller

servo-piston. Excess oil pressure and flow is bypassed via the relief valve back to the pump inlet.

During normal constant speed operation the beta control valve port (Fig. 4-6, #11) is always open.

The beta control valve (Fig. 4-6, #10) plays no part on the propeller blade angle control.

Effective: 1/26/05 4 – 14


B. Constant Speed Section

The constant speed section maintains constant propeller speed during takeoff, climb, and cruise by

controlling the flow of servo-oil to and from the propeller servo-piston.

A hollow drive shaft (Fig. 4-6, #15) is driven by a bevel gear on the engine propeller shaft. On top

of the drive shaft there are two rotating flyweights (Fig. 4-6, #8) that pivot outward. This action

provides an upward force proportional to propeller RPM. The feet of the flyweights tend to lift the

pilot valve plunger (Fig. 4-6, #16) and the force of the speeder spring (Fig. 4-6, #6) tends to push

the pilot valve plunger down. The interaction of these two forces controls the propeller speed.

The lower end of the pilot valve plunger (Fig. 4-6, #16) covers the ports in the hollow shaft (Fig. 4-6,

#15) in the CSU body. This mechanism directs the servo-oil to the propeller. When the upward

force of the flyweight’s equals the downward force of the speeder spring, the ports are covered and

no servo-oil flows to or from the propeller. The propeller blades remain at constant pitch. This is

termed "on speed" condition.

The propeller RPM at “on speed” condition may be selected by the operator. He may vary the

downward force on the speed spring (Fig. 4-6, #6) by actuating the speed select lever (Fig. 4-2, #3)

which is connected to the propeller control lever on the throttle quadrant.

If the operator selects a low speeder spring force, it follows that only a low flyweight force is needed

to lift the pilot valve plunger into the "on speed" condition. This is achieved at low flyweight and low

propeller RPM. The converse occurs if the operator selects high speeder spring force.

The CSU maintains selected propeller RPM automatically and compensates for "overspeed" and

"underspeed". When the propeller RPM is higher than the selected speed, the "overspeed"

condition occurs. The "underspeed" condition results when the propeller RPM is lower than the

selected speed. These conditions are described in detail below.

1. If the propeller RPM drops below the selected speed, the flyweight force decreases

and the force of the speeder spring pushes the pilot valve plunger down. This

process provides oil to the propeller servo-piston. The servo-piston moves

forward, which fines out the blades. The propeller RPM will then increase. As the

propeller RPM reaches the selected speed, the flyweight force lifts the pilot valve

back to the "on speed" condition.

2. If the propeller RPM rises above the selected speed, the flyweight force increases

and overcomes the force of the speeder spring to lift the pilot valve. The oil is

dumped from the propeller, which causes the blades to coarsen pitch. The

4 – 15 Effective: 1/26/05


propeller RPM will then decrease. As the propeller RPM reaches the selected

speed, the speeder spring force pushes the pilot valve back to the "on speed"

condition.

C. Power Turbine Governing Section

The Nf governor or fuel-topping governor of the power turbine governing section of the CSU has

two functions in the propeller speed control.

1. The first function is during the constant speed operation of takeoff, climb, and

cruise when it acts as a safety in the "overspeed" condition only. If a malfunction

occurs which allows the propeller RPM to exceed selected RPM by 6%, the Nf

governor bleeds Py air from the fuel control unit (FCU) to limit power.

2. The second function is during reverse propeller control when it will start to bleed Py

air from the fuel control unit (FCU) to keep the propeller and therefore the Nf power

turbine from overspeeding. This will limit propeller RPM 4% - 6% (88 – 132 RPM)

below the propeller RPM selected on the speeder spring which is 2200 RPM,

because the propeller control lever is still in full forward position. This will in turn

limit max reverse propeller RPM to 2068 – 2112 RPM. During beta operations the

propeller control lever on the throttle quadrant is at the max RPM (full forward

position). The speeder spring (Fig. 4-6, #6) is exerting its maximum downward

force so that it will always exceed the upward force of the flyweights in order to

keep the pilot valve plunger down at all times during the beta and reverse. The oil

passage to the propeller will then be wide open, and the oil flow is now controlled

only by the beta control valve which is upstream of the pilot valve plunger.

3. The components used in the Nf governor include the air bleed lever (Fig. 4-6, #2),

the orifice lever (Fig. 4-6, #22), and the fuel governor reset arm (Fig. 4-6, #19).

4. If a malfunction causes propeller "overspeed" that cannot be controlled by the CSU

constant speed section during the constant speed operation of takeoff, climb, and

cruise; then the top of the pilot valve plunger (Fig. 4-6, #16) lifts the air bleed lever

(Fig. 4-6, #2) The air bleed lever tilts and allows the orifice lever (Fig. 4-6, #22) to

bleed Py air from the FCU. The governing action begins when the propeller RPM

is approximately 6% [(132 RPM Np) or 2332 RPM maximum propeller speed]

above that selected on the speeder spring. In this phase of the operation the fuel

Effective: 1/26/05 4 – 16


governor reset arm (Fig. 4-6, #19) is against the maximum stop.

5. During beta and reverse the pilot valve plunger (Fig. 4-6, #16) is always in a

lowered position. Therefore, in order that the air bleed lever (Fig. 4-6, #2) can

contact the pilot valve plunger, the fulcrum point (Fig. 4-6, #1) of the air bleed lever

is lowered by lowering the rest post (Fig. 4-6, #20). This action is performed by the

NF governor reset arm (Fig. 4-6, #19). As the aircraft operator commands the beta

operation, the fuel governor reset arm (Fig. 4-6, #19 and Fig. 4-7, #28) moves off

the maximum stop (Fig. 4-6, #18) by the fuel governor interconnection rod (Fig. 4-

7, #14). This action continually lowers the reset post (Fig. 4-6, #20) to lower the

RPM from its normal overspeed protection duty of being set at 106% Np to a

setting of 96% Np. This will keep the propeller from never exceeding 96% Np

(2112RPM) as the aircraft operator chooses beta and reverse operations by

bleeding Py pressure (pneumatic governor servo pressure). This causes a

decrease in Py pressure at the computing section of the FCU (fuel control unit),

causing the fuel metering valve to move in a closing direction, thus reducing fuel

flow and consequently Ng and Nf speeds.

D. Beta Control Valve Section

The beta control valve (Fig. 4-6, #10) performs two functions in the propeller control.

1. The first function during takeoff, climb, and cruise is to act as a hydraulic low pitch

stop by limiting the finest blade angle possible in flight to the low blade angle. As

power is reduced, the constant speed section maintains selected propeller speed

by fining the propeller blade angle until the servo-piston picks up the beta nuts.

The beta feedback mechanism starts to close the beta control valve by moving it

forward. As the blades fine out further, the valve closes completely at the low

blade angle. Because the beta control valve is upstream of the pilot valve plunger,

the constant speed section can no longer select finer blade angles because its

supply is cut off.

Except for a malfunction, the hydraulic low pitch stop is normally achieved in

descent only. It is available only as a safety during takeoff, climb, and cruise.

Normally in those configurations the blades are much coarser than the angle at

which the servo-piston picks up the beta nuts.

4 – 17 Effective: 1/26/05


2. The second function of the beta control valve is to enable direct control of the

propeller blade angle in beta and reverse. After the hydraulic low pitch stop is

reached, finer blade angles through flat pitch to reverse can be selected by the

aircraft operator after landing. If the beta control valve is opened again by

rearward movement, the servo-oil flows to the propeller and moves the blades to a

finer angle. This can be continued to the maximum reverse blade angle. The beta

feedback mechanism will limit the blade angle reached in beta or reverse to that

desired by the aircraft operator. It does this by reclosing the beta control valve.

E. Feathering

Feathering is accomplished by raising the override rod (Fig. 4-6, #5). This pulls the pilot valve

plunger up to dump the servo-oil from the propeller. The blades feather automatically under the

action of the counter-weights and feather springs. Refer to Figure 4-5.

PROPELLER OVERSPEED GOVERNOR

The propeller overspeed governor (Fig 4-5) is installed in parallel with the propeller governor and

mounted at the approximate 10 o'clock position on the front case of the reduction gearbox. The

governor is incorporated to control any propeller overspeed condition by immediately bypassing

pressure oil from the propeller servo to the reduction gearbox sump. The governor consists of

conventional type flyweights mounted on a hollow-splined shaft and driven by the accessory drive

gearshaft. The hollow shaft embodies ports, which are normally closed by a pilot valve installed in

the shaft centerbore and held in position by the governor speeder spring. The spring tension acts in

opposition to the centrifugal force of the rotating flyweights.

When a propeller overspeed condition occurs, the increased centrifugal force sensed by the

governor flyweights overcomes the speeder spring tension and lifts the pilot valve to bypass

propeller servo oil back to the reduction gearbox sump via the governor hollow drive shaft. This

allows the combined forces of the blade counterweights and the return springs to move the

propeller blades toward a coarse pitch position, thereby absorbing engine power and reducing

propeller rpm. A solenoid-operated valve is incorporated to facilitate functional testing of the

overspeed governor. When operated, the valve resets the governor below its normal overspeeds

setting. (See overspeed governor check later in this section.)

ENGINE POWER AND PROPELLER CONTROLS (FCU) (REF FIG. 4-18)

The cockpit power lever is connected to the engine power lever (Fig. 4-7, #5). This operates the

cam follower pin as shown in Figure 4-7.

Effective: 1/26/05 4 – 18


Connected to the beta control cam (Fig. 4-7, #23) is the push-pull control cable which runs forward

on the engine to connect to the top end of the propeller reversing lever (Fig.4-7, #3), via the fuel

governor interconnection rod (Fig. 4-7, #14), and connects to the fuel governor reset arm. Refer to

the right-hand detail Figure 4-7.

The FCU is operated by the FCU actuating lever (Fig. 4-7, #8), the FCU control rod (Fig. 4-7, #7),

and the FCU arm (Fig. 4-7, #9).

In all forward configurations, which include low idle, takeoff, climb, and cruise, the power lever

control performs only one function - the function of scheduling fuel. When the cockpit power lever is

advanced, the cam follower pin (Fig. 4-7, #6) moves forward and pushes the FCU arm (Fig. 4-7, #8)

forward to schedule more fuel.

The extension of the cam follower pin rides in the track of the beta control cam (Fig. 4-7, #23). In all

forward configurations the path taken by the cam follower pin exactly matches the cam track.

Therefore, the beta control cam does not move; the push-pull control cable is inoperative; the top

end of the propeller reversing lever does not move; and the fuel governor reset arm remains on the

maximum stop on the CSU.

In beta after touchdown the power lever has two functions. It schedules the blade angle directly,

and it resets the Nf governor down. After the blades have passed zero pitch, the power lever

begins its third function in reverse. That function is to schedule the fuel flow as well.

After touchdown the aircraft operator presses the override button which is located on the power

lever and moves it rearward. The cam follower pin loses contact with the FCU actuating lever, and

the FCU will stay at flight idle (68% Ng), because the condition control rod (Fig. 4-7, #13) will

prevent any further lowering of gas generator speed. As the cam follower pin moves rearward, it

picks up the cam track of the beta control cam and starts to move it rearward. This action pulls the

push-pull control cable as well. This action also pulls the propeller reversing lever and the fuel

governor-interconnecting rod.

The FCU remains at flight idle while the blades fine out until the cam follower pin picks up on the

dead band adjustment screw (Fig. 4-7, #37). This moves the FCU reversing lever, which starts to

schedule more fuel in reverse. From this instant the cockpit power lever is performing three

functions. The functions are scheduling the blade angle directly, scheduling the fuel flow, and

setting the Nf governor down. The action continues up to maximum reverse blade angle.

4 – 19 Effective: 1/26/05


PROPELLER SPEED SELECT AND FEATHERING CONTROL

The cockpit propeller lever has two functions:

A. The first function is to select the propeller RPM in takeoff, climb, and cruise configurations.

B. The second function is to feather the propeller when it is required.

The cockpit propeller lever is connected to the speed select lever on the CSU.

The first function is performed by varying the speeder spring pressure (Fig. 4-6, #6) by rotating the

propeller speed select lever (Fig. 4-6, #3) toward the propeller speed max stop (Fig. 4-7, #36). The

second function is performed by rotating the propeller speed select lever (Fig. 4-6, #3) toward the

feathering stop. This action will cause the override rod (Fig. 4-6, #5) to pull the pilot valve plunger

(Fig. 4-6, #16) upward, therefore allowing servo oil to be dumped from the propeller servo piston.

This action will cause the propeller blades to travel to the feather position, by the feather-return

spring pressure acting on the propeller servo piston.

PROPELLER SETTINGS (HARTZELL HC-B3TN-3C (or D)/T10282N or T10282N+4)

A. Maximum RPM . . . 2,200 RPM

B. Cruise Power . . . Approximately +35° @ 30 in. station

C. Full Feathered Angle . . . 87.0± 1.0° @ 30 in. station

D. Mechanical Reverse Pitch Stop . . . -8.0°± 0.5° @ 30 in. station

E. Angle at which servo piston just touches the three low pitch stop rod beta nuts (which move

the propeller beta feedback ring) is 18.0°± 0.1° when blades are held toward decrease

position at the 30 inch station. The hydro-mechanical low pitch stop occurs at a blade

angle of approximately 11° when the propeller dome has traveled sufficiently to fully close

the beta valve and shut off the oil flow to the propeller.

The specific low-pitch blade-angle determined through aircraft flight and ground test, during

which the controllability of the aircraft is checked, is approximately 11° for the Hartzell HC-

B3TN-3D/T10282N+4 propeller installed on the Turbo Thrush.

The way to set the specific low-pitch blade-angle is to adjust the hydraulic low pitch stop.

This is accomplished by proper adjustment of the three beta nuts, by using the 2000 RPM

Np Torque Setting Chart (Figure 4-8).

Effective: 1/26/05 4 – 20


PROPELLER MAINTENANCE

PROPELLER REMOVAL

A. Remove the forward cowl from the engine.

B. Remove the spinner dome by removing the attaching screws from around the rear

circumference.

C. Disconnect the front fork-end from the propeller-reversing lever. Disconnect the pivot bolt

securing the reversing lever to the propeller governor actuating lever and lift the reversing

lever free of the collar prior to pulling the low pitch stop collar fully forward.

D. Install the feedback ring-puller and pull the low pitch stop collar fully forward.

E. Remove the safety wire from the propeller mounting bolts. Using a 5/8" box head wrench,

remove the eight bolts securing the propeller in place and remove the propeller from the

airplane.

4 – 21 Effective: 1/26/05

** CAUTION **

The procedures in the step above must be accomplished to

avoid damaging the propeller governor.

** CAUTION **

Make sure that the tool is not cocked to avoid damaging the

propeller. Take the precautions necessary to avoid bending or

otherwise damaging the three spring-loaded rods and the beta

feedback ring.

* NOTE *

Mark propeller hub flange and the engine shaft flange so that

the propeller can be reinstalled in its original position. This will

prevent disturbing the propeller/engine combination dynamic

balancing if the same propeller is to be reinstalled.


PROPELLER INSTALLATION

A. Place the new O-ring seal over the engine shaft.

B. Pull the beta ring fully forward with the puller.

C. Install the propeller on the engine by inserting the two dowel pins on the propeller flange in

the appropriate holes on the propeller shaft flange.

D. After assuring that complete and true surface contact between the flanges has been

established, apply (MIL-PRF-83483, Hartzell P/N A-3338-1 or latest upgrade) antiseize

compound to mounting bolt threads and washer surfaces (and remainder of bolt if desired).

For the HC-B3TN-3D propeller install eight (8) P/N B-3339 bolts and eight (8) A-2048-2

washers through engine flange into the propeller flange.

Effective: 1/26/05 4 – 22

** CAUTION **

Make sure the tool is not cocked to avoid damaging the

propeller. Take the precautions necessary to avoid bending

or otherwise damaging the spring-loaded rods and the beta

feedback ring (brass ring).

*** WARNING ***

Chamfer of washer must face bolt head at installation.

* NOTE *

The propeller will fit on the engine in two positions, 180°

from each other. Either position is permissible to use. If

the same propeller is being reinstalled, install in the original

position as previously marked. This will prevent disturbing

the propeller/engine combination dynamic balancing.


E. Using (Hartzell P/N AST-2877) special torquing adapter and a standard torque wrench,

torque all eight bolts according to instructions as outlined in the latest edition of Hartzell

Propeller, Inc. Owner's Manual & Log Book No. 139.

F. Safety all mounting bolts in an airworthy manner with .032-inch minimum diameter stainless

steel wire.

G. Remove the feedback ring puller and connect the propeller reversing lever to the propeller

control linkage.

H. Check the propeller reversing linkage on the front end of the engine for proper rigging.

I. Reinstall the spinner dome and engine cowling.

J. Perform the necessary engine run-up checks.

4 – 23 Effective: 1/26/05

* NOTE *

Do not add more than four (4) balance weights (P/N A-1305) in

any one stack. A maximum total of eight (8) weights are

allowed on any one clamp half.

** CAUTION **

With the carbon block assembly held against one side of

the beta feed back ring, check the side clearance (Refer

to Figure 4-9). Clearances can be established by

dressing the block(s) side(s) as required.

* NOTE *

Thrush Aircraft Inc. recommends that the propeller be dynamically

balanced to the engine whenever a new propeller or an overhauled

propeller is installed, or any time there is a question of the propeller’s

balance. Following the instruction of the propeller balancing equipment

(Chadwick Helmuth Vibrex or equivalent equipment), set the amplitude of

vibration given in IPS (inches per second) on the balancer's meter to a

level of .2 or less at 1,500 rpm Np by adding weights to the light blade(s)

or spinner bulkhead in accordance with Hartzell Propeller Owners

manual P/N 139 chapter 6.


CONSTANT SPEED UNIT (CSU) REMOVAL

A. Remove the forward engine cowling.

B. Remove bolt that secures the propeller control cable to the governor's speed select lever.

C. Remove the cotter pin, castellated nut, and washer and bolt securing the Nf governor-

interconnecting rod to the Nf governor reset arm.

D. Remove the cotter pin, castellated nut, washer, bolt and spacer securing the front clevis

end to the propeller-reversing lever.

E. Remove the cotter pin, washer, clevis pin and bushing securing the propeller-reversing

lever to the beta valve. Remove reversing arm.

F. Disconnect coupling nut of pneumatic (Py) front tube from straight nipple on propeller

governor.

G. Remove the four nuts and washers anchoring the governor to the mounting pad on the

reduction gearbox case.

H. Remove governor and governor mounting pad gasket.

I. If CSU is to be replaced by a new or overhauled unit, remove the straight nipple from Py

port on governor. Remove "O" ring and retain nipple for reuse on the replacement unit.

CONSTANT SPEED UNIT (CSU) INSTALLATION

A. Install a new gasket over the four studs on the governor-mounting pad.

B. If a new or overhauled propeller governor is to be fitted, install straight nipple as follows:

1. Lubricate new "O" ring with clean engine oil and install on nipple.

2. Install nipple in Py port on propeller governor. Tighten and torque nipple to 65 to

70 lb. in.

C. Lightly coat the splined shaft of the governor with clean engine oil.

Effective: 1/26/05 4 – 24

** CAUTION **

Make sure the gasket is placed on the mounting pad with the

raised side of the screen up so that it will fit into the recess on

the base of the governor.


D. Position the governor on the mounting pad and secure it in place with the four attaching

washers and nuts. Torque the nuts to 125 to 135 inch-pounds.

E. Secure the propeller-reversing lever to the beta valve with the attaching bushing, clevis pin,

washer and cotter pin.

F. Secure the front clevis end to the propeller reversing lever with the attaching spacer, bolt,

washers, castellated nut and cotter pin.

G. Secure the Nf governor-interconnecting rod to the Nf governor reset arm with the attaching

bolt, washer, castellated nut and cotter pin.

H. Secure the governor speed select lever to the propeller speed control cable with the

attaching bolt, washer and nut.

I. Ensure the governor's stop plate contacts both the high RPM stop screw and the feathering

stop screw, when the propeller control lever in the cockpit is operated. Ensure that there is

sufficient cushion at both positions on quadrant. If linkage will not allow proper travel,

adjust the control linkage at either rod end or move speed select lever on the governor to

obtain necessary travel.

J. Connect coupling nut of the pneumatic (Py) front tube to propeller governor. Tighten nut;

torque to 90 to 100 lb. in., and lock wire.

K. Check engines front linkage rigging.

L. Accomplish propeller governor operational checks in accordance with the appropriate Pratt

& Whitney Maintenance Manual.

M. Install the forward engines cowling.

PROPELLER CSU HIGH RPM ADJUSTMENT (REF TO FIG. 4-14)

If a high RPM adjustment is required, turn the high RPM stop screw on the governor head

clockwise to decrease or counter clockwise to increase RPM as required to obtain 2,200 RPM

propeller speed (NP). After adjustment, ensure there is sufficient cushion at both the feathered and

high RPM positions at the propeller lever on the throttle quadrant. Lockwire the high RPM stop

screw after adjustment.

4 – 25 Effective: 1/26/05

** CAUTION **

Ensure drive splines are completely engaged by checking

that flange of governor rests squarely on gasket with no gap.

Rotate propeller to assist engagement, if necessary.


PROPELLER OVERSPEED GOVERNOR REMOVAL

A. Remove the forward engines cowling.

B. Remove the safety-wire and disconnect the electrical plug from the governor solenoid

valve.

C. Remove the four self-locking nuts and plain washers securing the governor and remove the

governor from the left side of the reduction gear housing.

PROPELLER OVERSPEED GOVERNOR INSTALLATION

A. Install a new gasket on the mounting pad.

B. Apply clean engine oil to the governor splined drive.

C. Position the governor on the mounting pad and install the four plain washers and self-

locking nuts. Apply a torque of 125 to 135 inch-pounds to the mounting nuts. (Make sure

you have the two ground wires under one of the nuts.)

D. Connect the electrical plug to the governor solenoid valve.

E. Reinstall the forward cowling.

F. Perform overspeed governor check (prop test) and adjust as necessary.

ENGINE CONTROLS RIGGING INSTRUCTIONS

The following instructions will produce nominal settings of the engine’s operating parameters. If an

engine is installed, fuel control, propeller or propeller governor replaced or any time the adjustment

of these units is disturbed, the engine controls rigging should be checked.

AIRFRAME CONTROL LINKAGES

Proper engine/airframe control system rigging is a prerequisite in order to achieve satisfactory

engine operation. The airframe control system will provide the required throws, travel limits, and

etc. necessary for the engine controls operation.

PROPELLER REVERSING INTERCONNECT LINKAGE

Details for assembly and disassembly of engine push-pull cable are contained in the Engine

Maintenance Manual.

Effective: 1/26/05 4 – 26


FRONT LINKAGE (Figure 4-10)

A. Disconnect rear clevis (7) from the beta control cam (6). Refer to Figure 4-11.

Disconnect fuel governor interconnection rod (3) from the push-pull cable terminal (1).

Refer to Figure 4-10.

B. Pull hard forward on propeller reversing lever (2) so that push-pull control terminal (1) is

against stop inside adjustable stop (16).

C. Check that rear of clevis slot end (14) is flush with front face of cap nut.

Adjust lot-pitch stop adjuster (17) if required. To achieve this:

Turn (17) in if clevis slot end is forward of cap nut.

Turn (17) out if clevis slot end is aft of cap nut.

D. Disconnect push-pull control terminal (1) and check that its travel is within limits of 1.00 to

1.250 inches; then reconnect.

E. Pull hard forward on propeller reversing lever (2); hold fuel governor reset arm (5) on

max. stop (4); adjust fuel governor interconnection rod (3) until retaining bolt exactly

aligns with the lower hole of the fuel governor reset arm (5); shorten one rod end on (3)

one half turn and connect.

F. Push cockpit power lever into high power range; pull hard forward on propeller reversing

lever (2); adjust rear clevis (7) (See Fig. 4-11) to exactly align hole thru (7) with the

second hole from the top in the beta control cam (6); lengthen push-pull control ½ turn on

(7); install retaining pin, washer and cotter. Operate power lever from idle to max; check

for free movement.

REAR LINKAGE (Figure 4-11)

A. Disconnect rear clevis (7). With forward pressure on beta control cam (6), move cam

follower lever (16) rearwards as far as possible without beta control cam (6) moving. This

should result in the cam follower pin (9) being at the track point (17) of the beta control cam

(6).

B. With the interconnect rod (10, Fig. 4-11) and the start control rod (12) disconnected from

F.C.U. arm (11), rotate F.C.U. arm (11) softly clockwise until pick-up point is felt (inside

F.C.U.). Check that F.C.U. arm (11) is 16 degrees (Fig. 4-12) below horizontal.

4 – 27 Effective: 1/26/05

** CAUTION ** Never attempt to move the power lever into reverse with

engine shutdown without first removing the pin at the rear

clevis. (Refer to Figure 4-11, item #7)


*NOTE*

(Ref. Detail D) to reset the FCU arm position, loosen the FCU arm

extension and adjust the serrated spacer. There are 24 serrations

on the inner face of the spacer, i.e. each serration equals 15.0

degree increments. On the outer face of the spacer there are 25

serrations, i.e. each serration equals 14.4 degree increments.

Therefore, difference equals 0.6 degree. After adjustment, tighten

the FCU arm extension to a torque of 25 to 35 lb. in., and lockwire.

C. Hold the F.C.U. arm (11) at the pick-up point, with the cam follower pin (9) in the track

point of the reversing cam (idle position) (Detail B). Adjust the length of the interconnect

rod (10) until the holes in the rod ends align with the second from the top hole of the FCU

actuating lever (8) and the inner hold of the FCU arm (11).

D. Lengthen the interconnect rod (10) an additional one and a half turns to ensure that the

FCU arm (11) is approximately 2 degrees under the pick-up point.

*NOTE* These 2 degrees of dead band travel will ensure that the fuel

control unit is at the idle position while allowing minor idle

adjustment without appreciably affecting the pick-up point.

E. With the input lever (18) at the idle position (track point) and the cockpit power lever at idle

(against the beta and reverse lockout), adjust and connect the cockpit power lever control

cable to the input lever (18) in the inboard hole. Verify that when the cockpit power lever is

at the full power position that the max. Ng speed adjuster (19) contacts the max. stop on

the F.C.U. (20) and that there is sufficient cushion (1/8” – 1/4”) between cam follower pin

(9) and of track on beta control cam (6).

F. With the input lever (18) at the idle position (track point), adjust the dead band screw (Detail

B, Fig. 4-11) until a 5/32” gap is achieved.

G. Operate cockpit power lever from idle to max. and check that there is no binding, that the

F.C.U. max. stop contacts when cockpit power lever is at the max. and that F.C.U. arm (11)

is clear of pick-up point when cockpit power lever returns to idle.

H. When rigging is completed, ensure all cotter pins and safety wires are correctly installed.

CONDITION LEVER LINKAGE (Figure 4-13) (FOR ENGINES EQUIPPED WITH START CONTROL UNITS)

A. Place the condition control lever in ground idle position on the cockpit control quadrant.

Effective: 1/26/05 4 – 28


B. Position start control lever (4) so that rigging hole (3) is located between the 45° and 72°

rigging slots (5).

C. Adjust rod ends on push-pull cable or push-pull cable housing as necessary, and connect

to start control lever.

D. Operate the condition control lever on the cockpit control quadrant throughout its full range

to ensure freedom of movement and check the following:

1. When in the idle cutoff position, the start control lever (4) is contacting the cutoff

and dump stop (1).

2. When in the ground idle position, the rigging hole (3) in the start control lever (4) is

anywhere between the 45° and 72° rigging slots (5).

3. When in the flight idle position, the start control lever (4) is contacting the flight idle

max. stop (2).

CONDITION LEVER LINKAGE (Figure 4-2 and 4-3) (FOR ENGINES NOT EQUIPPED WITH START CONTROL UNITS) A. Place condition lever on engine control quadrant (Fig. 4-19) in the LOW IDLE (ground idle)

position.

B. On the FCU check that the cutoff valve is now open and the top of the Cut-off and High-idle

input lever (on very top of the FCU) is parallel with the top of the FCU (this lever will be

approximately midway between cutoff and High idle positions.

C. Adjust rod end on push-pull cable or push-pull cable housing to line up rod end with inboard

(2nd from top) hole in the Cut-off and High-idle input lever on the FCU. Install attaching

hardware and ensure that rod end stays in “witness” by checking with a 0.020” piece of

safety wire.

D. Operate condition lever on engine control quadrant and throughout its full range to ensure

freedom of movement and check the following:

1. When in the idle CUT-OFF position, the Cut-off and High-idle input lever is

contacting the cut-off stop on top of the FCU and there is still “cushion” at the

engine control quadrant.

2. When in the LOW IDLE the flat top of the Cut-off and High-idle lever is parallel with

top of the FCU.

3. When in the HIGH IDLE position, the Cut-off and High-idle input lever is contacting

the flight idle max. stop on FCU and there is still “cushion” at the engine control

quadrant.

4 – 29 Effective: 1/26/05


PROPELLER LEVER LINKAGE (Figure 4-14)

A. Place propeller control lever in the full increase position on the cockpit control quadrant.

B. Turn propeller speed select lever (1) clockwise until the levers stop plate contacts propeller

speed max. stop (2).

C. Adjust rod end on propeller control cable and connect to propeller speed select lever (1).

* NOTE * It may be necessary to rotate the propeller speed select lever (1)

around the serrations provided on the C.S.U. to obtain the

necessary travel and/or clearance.

D. Operate the propeller control lever on the cockpit control quadrant throughout its full range

to ensure freedom of movement and check the following:

1. When in the full increase position, that the propeller speed select lever (1) is

contacting the propeller speed max. stop (2).

2. When in the feathering position, that the propeller speed select lever is contacting

the propeller feathering stop (3).

Upon completion of rigging and prior to engine running, a functional check of the system’s operation

should be carried out. This check should include the operation of all controls throughout their entire

operation range and checking for freedom of all movement, freedom from binding, security and

safety. Ensure that there is sufficient cushion (1/8” – ¼” min.) at the forward and rearward position

of each control lever on the cockpit control quadrant. Control lever movement must be halted by

stops engaging at the engine control units and not by control levers contacting stops at cockpit

control quadrant.

** CAUTION ** Never attempt to move the power lever into reverse with engine

shutdown without first removing the pin at the rear clevis. (Fig 4-

11, #7)

ENGINE RIGGING CHECKS AND ADJUSTMENTS

This section details various engine and propeller functional checks, which are performed after

engine control rigging and engine run up.

Effective: 1/26/05 4 – 30


** CAUTION ** Before verification of any engine indicator reading, allow engine,

engine sensors, and engine gauges to warm to normal operating

temperatures. After proper warm up, allow a five minute

stabilization period at each check point. Also, lightly tap gauges

with fingertips to avoid erroneous instrument readings. Ensure that

none of the engine operating limits is exceeded at any time.

LOW (GROUND) IDLE ADJUSTMENTS (Ref. Fig. 4-3) A. Run engine to bring oil temperature within normal operating ranges (38° C minimum). B. Set condition lever to LOW (Gnd.) IDLE and power control lever in IDLE. Check Ng

tachometer for a reading of 52.5% +1%, -0% Ng. If it is not, proceed as follows:

1. If Ng is less than 52.5%, remove safety wire; hold idle speed adjustment screw

(Fig. 4-15, #1) and loosen jam nut. Turn adjustment in a clockwise direction, as

required, to achieve 52.5%, +1%, -0% Ng. Tighten jam nut and safety.

*NOTE*

Fine Adjustment: 1/16 turn of adjustment screw

will vary Ng approximately 2% Ng.

2. If Ng is more than 53.5%, adjust idle speed adjustment screw (Fig. 4-15, #1) in a

counter clockwise direction, as required, to achieve 52.5%, +1%, -0% Ng.

3. If there is no response to adjustment in (1) and (2) preceding, ensure that engine

rear linkage rigging is correct and that the F.C.U. speed setting shaft and F.C.U.

arm (Fig. 4-15, #5) is approximately 2 degrees under the pick-up point (inside

F.C.U.). After verification of proper engine rear linkage rigging and there is still

no response to adjustments, the fuel control unit must be replaced.

HIGH (FLIGHT) IDLE ADJUSTMENTS (FIGURE 4-11) (FOR ENGINES EQUIPPED WITH START CONTROL UNITS)

A. With engine running, set condition lever to HIGH (Flt) idle position and power control lever

in idle. Check Ng tachometer for a reading of 68% Ng +0%, -0%. If it is not, proceed as

follows:

4 – 31 Effective: 1/26/05


1. If Ng is less than 68% Ng, adjust upper, lower and/or both rod ends to lengthen the

telescopic condition control rod (Fig. 4-11, #12) as required to achieve 68% Ng.

2. If Ng is more than 68% Ng, adjust upper, lower and/or both rod ends to shorten the

telescopic condition control rod (Fig. 4-11, #12) as required to achieve 68% Ng.

3. If you run out of adjustment on the telescopic condition control rod, adjust the flight

idle max stop (Fig. 4-13, #2) on the start control unit as required to achieve 68%

Ng. This is a super sensitive adjustment that requires very little movement to

obtain a large result. Fine adjustment should be made by adjusting the rod ends

on the telescopic condition control rod (Fig. 4-11, #12) as outlined in steps (1) and

(2).

HIGH (FLIGHT) IDLE ADJUSTMENTS (FIGURE 4-3) (FOR ENGINES NOT EQUIPPED WITH START CONTROL UNITS)

A. With engine running and stabilized (5 minutes), set condition lever to HIGH (Flight) IDLE

position and power control lever in idle, check Ng tachometer for a reading of 68% +0%, -

0%. If it is not, proceed as follows:

1. Turn HIGH IDLE adjustment screw CW to increase %rpm.

2. Turn HIGH IDLE adjustment stop screw CCW to decrease %rpm.

MAX PROPELLER SPEED CHECK

A. Set propeller control lever to obtain 100% Np. Check Np tachometer for a reading of 2200

rpm If it is not, proceed as follows.

B. Adjust max. Np adjustment screw (Fig. 4-14, #2) on C.S.U. to obtain 2,200 +/-0 RPM Np.

Turning the adjustment screw clockwise will decrease propeller RPM. Turning the

adjustment screw counter clockwise will increase propeller RPM.

OVERSPEED GOVERNOR CHECK (PROP TEST) The overspeed unit, mounted on the left side of the reduction gear housing, is preset at the

factory to govern at 104 percent Np speed (approximately 2288 rpm) and should not normally

require re-adjusting. This setting can be reduced to 2025 +20 rpm for testing purposes by means

of a testing solenoid. Perform the overspeed governor check as follows:

A. Place the propeller lever in full increase rpm position (forward).

B. Increase propeller RPM with the power lever until propeller RPM is 2100 RPM.

C. Turn prop test switch ON.

Effective: 1/26/05 4 – 32


* NOTE *

This switch is spring loaded to the off position. Forward pressure

must be held for the switch to remain on.

D. The propeller RPM should drop to 2025 +20 RPM or 92% Np.

E. Turn prop test switch OFF. RPM should increase back to 2100.

F. Reduce power back to idle.

If RPM is not governed at 2025 +20 RPM with prop test switch on, adjust the overspeed governor

as follows:

A. Turn adjustment screw on overspeed governor clockwise two complete turns. (This

assures that the overspeed governor is not affecting primary (CSU) governor).

B. Start engine and check max propeller rpm. Adjust primary (CSU) governor accordingly to

achieve 100% Np (2200 rpm).

C. Turn the overspeed governor adjustment screw counter clockwise two turns that had

previously been added.

D. Turn the adjustment screw counter clockwise an additional one turn for preliminary setting.

** CAUTION **

In no case should any engine limitations be exceeded.

E. Start engine.

F. Place prop lever in full forward position.

* NOTE *

Prop test switch must be off during this procedure.

G. Advance power lever until overspeed governor governs.

H. Adjust overspeed governor until 2120 rpm is achieved. (Turning screw clockwise will

increase rpm. Turning screw counter clockwise will decrease rpm.)

I. Shut down engine.

J. Turn overspeed governor adjustment screw clockwise one full turn. (This increases the

overspeed governing rpm from the established 2120 rpm to 2288 rpm or 104% Np). Note:

One full turn of overspeed governor adjustment screw equals 168 rpm.

K. Perform overspeed governor check to confirm proper operation of overspeed governor.

4 – 33 Effective: 1/26/05


* NOTE * Overspeed units that fail to govern at the prescribed settings

should be replaced and returned to a qualified repair station for

service.

REVERSE MAX POWER CHECK

A. Fuel Condition Lever – Flight idle

B. Propeller Control Lever – Full INCREASE

C. Power Lever - Pull slowly from IDLE to Full REVERSE

D. Torque should be 35 PSI minimum. If not, check engine rigging.

MAXIMUM GAS GENERATOR SPEED CHECK (FIGURE 4-15)

Although fuel control units are calibrated for maximum governing during bench check (overhaul),

the maximum gas generator speed setting should be checked after engine installation or

component replacement. The part power trim stop (Fig. 4-15, #4) is a movable spacer placed

between the maximum Ng stop and the power lever anvil, and represents 1,700 rpm Ng speed

decrement. This enables setting of max. gas generator speed over a wide range of ambient

conditions without exceeding engine limitations. After adjustments have been completed, the part

power trim stop is returned to the stowed position on the FCU. To adjust, proceed as follows:

** CAUTION ** During adjustments and normal operation, ensure that engine

operating limits are not exceeded.

** CAUTION ** Fill hopper and hold the elevator control firmly full up during all

high power ground operations to keep aircraft from nosing over.

A. Position part power trim stop (4) so as to limit power control lever travel.

B. Carry out satisfactory start.

C. Advance power control level to part power trim stop. Allow engine to stabilize for minimum

of two minutes.

D. Check that maximum gas generator speed (Ng) is 97.1% and adjust max. governing speed

adjustment screw (3) as required.

E. Shut down engine.

Effective: 1/26/05 4 – 34


F. Stow trim stop (4) and safety wire attachment screw.

G. Safety wire maximum speed stops screw (3) and locknut.

ENGINE PERFORMANCE CHECK

The installed engine performance check curves enable the engine performance to be check on the

ground over a wide range of ambient conditions without over torquing or over temperaturing the

engine. With acquisition of a new aircraft or following installation of a new or an overhauled engine,

Thrush Aircraft Inc and Pratt and Whitney of Canada recommend that the operator institute and

maintain an engine condition monitoring program. The procedure compares the manufacturer’s

predicted parameters with the actual readings during a high power ground check. The data is

further compared with previously recorded data and can reflect a very accurate trend in the engine

performance.

All forms of engine deterioration are accompanied by an increase in interturbine temperature and

fuel flow at a given power. Compressor deterioration is, in most cases, due to dirt deposits and

causes an increase in gas generator speed at a given power setting. This form of deterioration can

be remedied by field cleaning, i.e. daily washes and power recovery washes. (Refer to applicable

Pratt & Whitney Maintenance Manual for detailed compressor cleaning procedures.) The following

procedure is for the PT6A-34AG installed in the Turbo Thrush.

A. Obtain the ambient air temperature in degrees Celsius (°C).

B. Obtain field barometric pressure in inches of mercury (“HG.). In order to get field

barometric pressure, set hands of altimeter to zero (0) feet. Read inches of mercury in the

altimeter setting window. This is field barometric pressure in inches of mercury.

C. From the appropriate ground run performance check curve (PT6A-34AG Figure 4-7);

record the following for the prevailing ambient conditions.

1. Torque Pressure (PSI)

2. Fuel Flow (pounds/hour)

3. Gas Generator Speed (%)

4. ITT (°C.)

5. Propeller Speed (2200 RPM)

* NOTE * This check should be accomplished without bleed air or power

extraction for operation of aircraft ancillaries. The aircraft should

be faced into the wind.

4 – 35 Effective: 1/26/05


D. Perform the pre-start check.

E. Accomplish a satisfactory engine start.

F. Set the propeller control lever output speed to provide 100% Np (2200 RPM) and set the

power control lever to target torque setting previously determined in Step C. Allow sufficient

warm-up time for stabilization of the engine parameters and instrumentation.

G. Observe and record:

1. Target Torque is set to chart value. Propeller speed is set to 2200 RPM.

2. Fuel flow if the aircraft is equipped with gauge.

3. Gas generator speed (%).

4. ITT (°C)

5. Ensure that the parameters observed during the check are within the following

limits.

a. +15 lb/hr fuel flow. If fuel flow is more than 75 lb/hr below chart value,

check instrumentation.

b. Maximum gas generator speed is not exceeded.

c. Maximum interturbine temperature line is not exceeded. If temperature is

more than 75°C below target temperature, check instrumentation.

* NOTE *

The limits that are shown on the check chart cover the normal

engine deterioration, but deviations should not necessarily be a

cause for engine rejection until all possible troubleshooting has

been completed.

The observed values of fuel flow, gas generator speed, and interturbine temperature should show

margins from the acceptance values previously recorded from the performance check curves. The

magnitude of the margins will depend on the condition of the engine. When the performance

standards is such that the gas generator speed and/or interturbine margins are zero and are

unaltered by compressor washing, the engine is described as “fully performance deteriorated.” It

will just meet the specification of minimum performance conditions.

COCKPIT ENGINE CONTROL QUADRANT

The cockpit engine control quadrant is located on the left hand cockpit side wall. The engine

controls (Ref. Fig. 4-19) consist of the following (a) power control lever (throttle), (b) fuel control

lever (mixture), and (c) propeller control lever (propeller lever). Refer to Figure 4-21 to see

relationship of cockpit engine control levers to engine control units.

The cockpit control quadrant requires almost no maintenance, other than occasional light oil

lubrication. Refer to Figure 4-20 for an exploded view of cockpit engine control quadrant.

Effective: 1/26/05 4 – 36


ENGINE LIMITS

Refer to chart (Figure 4-23) for the Pratt & Whitney PT6A-34AG equipped model Thrush for engine

limitations. Refer to the appropriate section of the Pratt & Whitney Canada PT6A-34AG

Maintenance Manual P/N 3021242 for correct procedure, should any engine limitation ever be

exceeded.

ENGINE DATA PERTINENT TO THRUSH AIRCRAFT INC INSTALLATION

Low (Ground) Idle 52.5% +1%, -0% Ng

High (Flight) Idle 68% Ng

Fuel Boost Pump Pressure 20 ±1 psig

Oil Pressure 85-105 psig (PT6A-34AG)

Part Power Trim 97.1% Ng

Max. Reverse 35 psi Torque Pressure (minimum)

Prop Test (O/S Governor) 2025 ±20 RPM NP

Max. Propeller RPM 2200 RPM NP

Hydraulic Low Pitch Stop Check in accordance with 2000 RPM NP Torque Setting Chart (Fig 4-8)

4 – 37 Effective: 1/26/05


Figure 4-1

PT6A-34AG - ENGINE ASSEMBLY

Effective: 1/26/05 4 – 38


Figure 4-2

4 – 39 Effective: 1/26/05


Figure 4-3

Effective: 1/26/05 4 – 40


Figure 4-4

4 – 41 Effective: 1/26/05


Effective: 1/26/05 4 – 42

Figure 4-5 Propeller Control System


4 – 43 Effective: 1/26/05

Figure 4-6 Constant Speed Unit


Effective: 1/26/05 4 – 44

Figure 4 – 7 Propeller Reversing Mechanism


Figure 4-8 2000 RPM NP Torque Setting Chart

Effective: 1/26/05 And Beta Nut Adjustment 4 – 45


Effective: 1/26/05 4 – 46

Figure 4-9 Carbon Block Assembly Replacement Guidelines


Front Linkage Figure 4-10

4 – 47 Effective: 1/26/05


Rear Linkage Figure 4-11

Effective: 1/26/05 4 – 48


Figure 4-12 Engine Power Lever Setting Template

4 – 49 Effective: 1/26/05


Figure 4-13

Starting Flow Control

Effective: 1/26/05 4 – 50


Figure 4-14 Constant Speed Unit (C.S.U.) Adjustment

4 – 51 Effective: 1/26/05


Figure 4-15

Fuel Control Unit (F.C.U.) Adjustments

Effective: 1/26/05 4 – 52


This Page

Intentionally Left

Blank

Figure 4-16

4 – 53 Effective: 1/26/05


Figure 4-17

Installed Engine Performance Checking Curve Lower Cowl Air Cleaner Panels

Effective: 1/26/05 4 – 54


Figure 4-18 Cockpit Power Lever (Throttle)

To Propeller Blade Angle Relation

Effective: 1/26/05 4 – 55


Figure 4-19 Cockpit Engine Control Quadrant

4 – 56 Effective: 1/26/05


Figure 4-20 Cockpit Engine Control Quadrant

Exploded View

Effective: 1/26/05 4 – 57


Figure 4-21 Cockpit Engine Control Levers to Engine Control Units Relation

Effective: 1/26/05 4 – 58


This Page

Intentionally Left

Blank

Figure 4-22

4 – 59 Effective: 1/26/05


ENGINE LIMITS PT6A-34AG, -34, -36 (PT6A-34AG shown)

GENERAL NOTE: The operating parameters in this Table are individual limits for each engine parameter and do not apply simultaneously. See appropriate engine maintenance manual for more detailed pertinent data. *These values are time limited to two (2) seconds. **These values are time limited to twenty (20) seconds. ***Torque limit for Max Np = 2000 RPM ****Torque limit for Np = 2200 RPM

GENERAL NOTE: The operating parameters in this Table are individual limits for each engine parameters and do not apply simultaneously. See appropriate engine maintenance manual for more detailed pertinent data. *These values are time limited to five (5) seconds. **Time limited to ten (10) minutes at any operating condition. ***Torque limit shown for Np = 1600-2000 RPM †For PT6A-42 790ºC. ††For PT6A-42 765ºC. †††For PT6A-42 100 to 135 STARTER – Maximum starter engagement duration is 30 Seconds followed by a 60 Seconds cool down

period. Total of 3 cycles to be followed by 30 minutes starter cool down. Figure 4-23

Effective: 1/26/05 4 – 60

POWER SETTING SHP TORQUE

(PSI) ITT (ºC) (%) Ng Np (RPM) OIL

PRESSURE (PSIG)

OIL TEMP. (ºC)

Takeoff & Max continuous 750 64.5***

58.7**** 790 101.6 2200 85 TO 105 10 TO 99

Max Climb 700 60.2*** 54.7**** 765 101.6 2200 85 to 105 0 to 99

Max Cruise 700 60.2*** 54.7**** 740 2200 85 to 105 0 to 99

Idle 685 40 Min. -40 to 99

Starting 1090* -40 Min.

Acceleration 68.4** 850* 102.6* 2420 0 to 99

Max. Reverse 750 64.5*** 61.4**** 790 101.6 2100 85 to 105 0 to 99

ENGINE LIMITS PT6A-41AG, -41, -42 (PT6A-41AG shown)

POWER SETTING SHP

*** TORQUE

(PSI) ITT (ºC) (%) Ng Np (RPM)

OIL PRESSURE

(PSIG)

OIL TEMP. (ºC)

Takeoff & Max continuous 750 64.5 750† 101.5 2000 105 to 135††† 0 to 99

Max Climb 700 60.2 725†† 101.5 2000 105 to 135††† 0 to 99 Max Cruise 700 60.2 725† 2000 105 to 135††† 0 to 99 Idle 60 Min. -40 to 99 Starting 1000* -40 Min. Acceleration 68.4* 850 102.6 2205 99 to 104** Max. Reverse 750 64.5 750 101.5 2000 105 to 135††† 0 to 99


Section 5 FUEL SYSTEM

TABLE OF CONTENTS

FUEL SYSTEM

GENERAL DESCRIPTION ...........................................................................................................................2 MAINTENANCE PRECAUTIONS.................................................................................................................3 SUB-SYSTEMS AND COMPONENTS.........................................................................................................3

FUEL QUANTITY INDICATING SYSTEM ................................................................................................3 FUEL QUANTITY INDICATOR .................................................................................................................4 TRANSMITTER.........................................................................................................................................4 REMOVAL..................................................................................................................................................4 INSTALLATION..........................................................................................................................................4 CALIBRATION ...........................................................................................................................................5 AUXILIARY ELECTRIC FUEL PUMP .......................................................................................................6 AUXILIARY ELECTRIC FUEL PUMP REMOVAL .....................................................................................6 AUXILIARY ELECTRIC FUEL PUMP INSTALLATION.............................................................................6 ENGINE-DRIVEN FUEL PUMP ................................................................................................................7 FUEL FILTER............................................................................................................................................7 OPTIONAL: FUEL FLOW.........................................................................................................................8 FUEL LINE MAINTENANCE.....................................................................................................................9

SEALING COMPOUNDS..............................................................................................................................9 LEAK SEALING.......................................................................................................................................10 RESEALING AFTER COMPLETE SKIN REMOVAL..............................................................................11

TROUBLESHOOTING................................................................................................................................11 FUEL SYSTEM TROUBLE SHOOTING CHART....................................................................................12

ACTIVATING HOPPER (FERRY) FUEL SYSTEM ....................................................................................16 Operating instructions for the P/N 60167 ferry fuel system ........................................................................17

Effective: 1/26/05 5 -1


FUEL SYSTEM

SECTION FIVE

GENERAL DESCRIPTION

A 230 U.S. gallon fuel supply is available for the Turbo Thrush. In each wing, fuel is contained

inside integral wing tanks (wet wing fuel tanks) just outboard of the fuselage. The left wing and right

wing fuel tanks are interconnected through a 5 U.S. gallon header tank that is located in the

fuselage. The fuel supply lines, to the engine, are routed from the header tanks outlet finger screen

through a fuel shutoff (on/off) valve to an electric driven fuel boost pump. The electric driven fuel

boost pump discharge is then routed through a 25-micron main fuel filter to an engine driven fuel

boost pump. The electric driven fuel boost pump serves two purposes, first as a backup system to

provide continuous fuel pressure to the engines high pressure fuel pump in case the engine driven

fuel boost pump fails and secondly to provide boosted fuel pressure to the engines high pressure

fuel pump during engine starting. The aircraft’s fuel system is equipped with two fuel filters, a ¼

inch mesh finger strainer is installed in the outlet fitting from the header tank and a 25-micron,

airframe supplied, main fuel filter located on the forward L/H side of the firewall. Fuel from the

aircraft fuel system enters the engines high pressure fuel pump which has two fuel filters, an 74-

micron inlet filter and a 10-micron discharge filter (Refer to the engines appropriate maintenance

manual for pertinent maintenance details for the engine supplied filters and fuel systems). The fuel

tank vent system is designed to keep the fuel spillage to a minimum. The fuel tanks are vented

through tubing connected at both the inboard and outboard ends of the individual fuel tanks to the

centrally located vent system in the fuselage. Ram air enters a vent scoop, on the fuselage, under

the left wing and pressurizes the vent system to maintain positive pressure on the fuel tanks. The

vent system is provided with two quick drain, located on the fuselage under each wing to drain any

fuel that might happened to have got in the tanks outboard vent lines. At engine shutdown, fuel

from the flow divider/dump valve, located at the 6 o’clock position on the engines fuel nozzle

manifold or start control unit on older engines equipped with a start control unit, is directed to a

residue fuel reservoir “EPA tank” mounted inboard on the L/H aft shin skin. This reservoir holds

approximately 3 engine shutdowns worth of fuel before the fuel will exit the reservoirs’ vent system.

(NOTE: This reservoir should be emptied after each shutdown.) (NOTE: It is common and normal

after an engine compressor Water Wash or Performance Recovery Wash to have water or soap

appear in the reservoirs’ drained waste fuel.) The fuel quantity gauge is located on the lower left

instrument panel. The fuel quantity indicated system consists of two transmitters, one indicator

gauge, and a L/H or R/H tank fuel quantity selector switch. A transmitter, installed in each wing

tank transmits an electrical signal to the single fuel quantity indicator. The instrument reads both

the left and right fuel tanks singularly as chosen by the electrical control switch, adjacent to the fuel

Effective: 1/26/05 5 -2


quantity indicator gauge on the instrument panel. The two fuel tanks are serviced through filler

ports located on the top of both wings. The filler ports incorporate security chains to prevent the lost

of the fuel caps. Service the aircraft from refueling facilities that utilize proper ground handling

equipment and filter systems to remove impurities and water accumulations from the bulk fuel. If

filtering facilities are not available, filter the fuel through a quality high-grade chamois. Fuel tanks

should be serviced after the last flight of each day to reduce condensation and allow any entrapped

water accumulations to settle to the fuel system drains, to be removed, prior to the next flight.

MAINTENANCE PRECAUTIONS

The establishment of safe maintenance procedures is necessary to ensure safety of personnel and

prevent damage to the aircraft when performing fuel system maintenance. The principle

precautions that should be enforced are enumerated as follows:

A. Perform fuel system maintenance in an approved work area.

B. Ground aircraft and maintenance stands to a common ground; ground- attaching surfaces

must not be painted.

C. Remove external power sources and disconnect batteries.

D. Suspend all maintenance except fuel system maintenance, unless area is declared safe from

explosive vapors.

E. Assure that fire-extinguishing equipment is readily available.

F. Use air-driven power tools only.

G. Use explosive-proof electric lights or flashlights.

H. Wear cotton clothing to avoid possible static electricity discharge.

I. Service, defuel, and refuel aircraft as outlined in Section II.

J. Do not remove components from the fuel system until replacement components or covers are

available for exposed openings.

K. Always replace O-rings, seals, etc. when re-installing fuel system components.

SUB-SYSTEMS AND COMPONENTS

FUEL QUANTITY INDICATING SYSTEM

The fuel quantity indicating system consists of a fuel quantity indicator, located in the left instrument

panel and electrically connected to a fuel quantity transmitter installed in each fuel tank. The fuel

quantity indicating circuit is provided with variable resistors within the transmitters. These resistors

vary the current flow through the indicating circuit. As the current flow varies, the needle on the fuel

quantity indicator will indicate the level of fuel sensed by the fuel quantity transmitter.

5 -3 Effective: 1/26/05


FUEL QUANTITY INDICATOR

One fuel quantity indicator is installed in front of the pilot in the left instrument panel. This

instrument is a single reading indicator that serves either the left of right fuel tank by operation of an

electric fuel tank-selector switch on the left instrument panel. A transmitter installed in each fuel

tank sends input signals to the indicator that gives the proper level of fuel in each tank. The

instrument face is marked in increments from empty to full. Refer to Section VIII for additional

information.

TRANSMITTER

The fuel quantity transmitters are installed in the inboard aft corner of the wing fuel tanks. Access to

the transmitter is gained by removing the inboard cover plate. As the fuel level increases, the float

arm is repositioned. This produces a minimum resistance through the transmitter, permitting

maximum current flow through the fuel quantity indicator and maximum pointer deflection. As the

fuel level is lowered, resistance in the transmitter is increased, producing a decreased current flow

through the fuel quantity indicator and consequently a smaller pointer deflection on the fuel quantity

indicator.

Removal

Removal of the fuel quantity transmitter can be accomplished through the inboard cover plate on

the upper surface of the wing.

A. Defuel aircraft as outlined in Section II.

B. Remove inboard cover plate.

C. Disconnect electrical leads at the transmitter.

D. Remove attaching screws, washers and bushings, and carefully remove transmitter

assembly.

Installation

The transmitter can be installed by reversing the removal procedures. Do not damage float or bend

float arm when placing the transmitter into the tank or incorrect readings will result.

Effective: 1/26/05 5 -4


Calibration

The fuel quantity transmitter and indicator have been calibrated at the factory and should not

require recalibration. However, if for some reason the system requires recalibration, the electrical

system should be carefully checked prior to recalibration. When necessary, the fuel quantity

indicating system is calibrated as follows:

A. Defuel aircraft as outlined in Section II.

B. Connect an APU (auxiliary power unit) to the external power connector.

C. Turn APU on and adjust to 27.5 volts.

D. Turn battery switch ON. Readjust APU to 27.5 volts, if necessary.

E. Place fuel quantity switch to L.H. MAIN tank.

F. With the transmitter float resting on the bottom of the fuel tank, set indicator needle to the

empty mark by adjusting the screw on front of indicator.

G. Raise float to touch top of fuel tank and set indicator needle to the full mark by adjusting

trimmer screw on back of indicator.

H. Place fuel quantity switch to R.H. MAIN tank.

I. Repeat steps F and G for right fuel tank.

J. Turn battery switch OFF.

K. Turn APU OFF.

L. Level aircraft as outlined in Section II.

M. Disconnect right fuel tank fuel lines to header tank at wing outlet ports. Cap fuel tank fittings

and disconnected fuel lines.

N. Fill header tank and lines with fuel.

O. Add one (1) U.S. gallon of fuel to each wing tank.

P. Turn APU ON.

Q. Turn Battery switch ON.

R. Check fuel quantity indicator for correct reading of each tank. Indicator should read empty

(O) at one gallon.

S. Complete the calibration (see Table below on this page).

T. After completion of the preceding steps, the calibration should be correct. If not, check

transmitter float arm for correct down (empty) position and correct by bending float support

arm as needed. Recalibrate system and check for correct reading.

U. If the system is still out of calibration, remove and replace the transmitter and/or the

indicator.

5 -5 Effective: 1/26/05


V. Turn off and disconnect APU.

W. Turn battery OFF.

X. Restore the fuel system to its original configuration.

CAPACITY (U.S. GALLONS)

EMPTY 1/2 FULL 82

1 Gallon

40 (-3, +5)

Gallons

Per Tank

82 (-0, +8)

ABOVE 82

UNGAUGEABLE

AUXILIARY ELECTRIC FUEL PUMP

The auxiliary fuel pump is installed on the underside, left side of aircraft cockpit aft of the fuel

header tank. A two-position switch labeled AUX FUEL PUMP on the switch panel electrically

controls this pump. The pump is a positive displacement vane type with a balanced-type relief

valve, and provides a fuel pressure of 20± psi. This pump provides positive fuel pressure for engine

starting and may be used for continuous engine operation in the event of engine-driven fuel pump

failure. Maintenance and disassembly of this pump is not authorized. Therefore, the servicing is

limited to the removal and replacement of the pump.

AUXILIARY ELECTRIC FUEL PUMP REMOVAL

A. Close fuel shutoff valve.

B. Remove drain plug and drain pump.

C. Disconnect electrical connector from motor.

D. Remove hose from pump and cap hose.

E. Remove attaching hardware and remove pump assembly from support bracket.

AUXILIARY ELECTRIC FUEL PUMP INSTALLATION

A. Install pump to support brackets and tighten hardware.

B. Connect hose to pump

C. Open fuel shutoff valve.

D. Connect electrical connector to pump motor.

E. Operate fuel pump and check for fuel leaks at lines and fittings.

Effective: 1/26/05 5 -6


ENGINE-DRIVEN FUEL PUMP

The engine-driven fuel pump installed on the lower right rear portion of the engine is provided with a

relief valve that will allow fuel to pass through from the airframe pump to the engine in the event of

pump failure.

FUEL FILTER

The fuel strainer in the filter should be removed, inspected and cleaned every 100 hours of

operation or sooner if improper fuel circulation is suspected. (See Figure 5-1 and 5-2 fuel filter.)

AIRBORNE 1J18 FUEL FILTER SERVICE INSTRUCTIONS

** CAUTION **

The following procedures must be followed in the order of

steps given to avoid damage to the components and to

assure proper functioning of the unit.

Refer to illustration (Fig 5-2) for identification of parts during disassembly and re-assembly.

A. Turn airframe fuel shutoff valve to “OFF” position. Cut, remove and discard safety wire (not

shown) securing filter bowl assembly.

B. Using 13/16” wrench unscrew hex nut, (Item 1) bowl retainer. (Right hand threads.)

C. Pull filter bowl (Item 2) straight off filter housing stud.

D. Using one thin ½” open end wrench, hold filter retaining nut (Item 3) while loosening jam nut

(Item 4) with second ½” wrench.

** CAUTION **

DO NOT twist or bend stud. Stud is not a removable item.

DO NOT pry on filter element.

E. Remove retaining and jam nuts (Items 3 and 4).

F. Filter element (Item 5) will now drop off stud.

5 -7 Effective: 1/26/05


G. Seal central tube opening of filter element with suitable size rubber plug, to keep inside of

filter element from getting contaminated during the cleaning process. Gently clean filter

element by rinsing in new/unused solvent (Safty-kleen SK-105, Varsol, MIL-PRK-680 Type

II, Odorless mineral spirits or equivalent) and blowing debris off surfaces using a low-

pressure (up to 30 PSI Max) clean compressed air source.

* * CAUTION * *

DO NOT scrape, pry or poke mesh surfaces with sharp objects.

DO NOT attempt to separate segments of filter elements.

H. Replace filter element (Item 5) on stud and secure with retaining nut (Item 3) tightened

moderately.

I. While holding retaining nut with thin ½” open-end wrench tighten jam nut (Item 4) with

second ½” wrench.

** CAUTION **

DO NOT ALLOW STUD TO TWIST.

J. Replace filter bowl “0” ring. Apply a light coating of Vaseline to new filter bowl “0” ring seal,

and then locate new seal in groove on inside lip of filter housing.

K. Push filter bowl (Item 2) into housing taking care not to cock sideways.

L. Replace fuel bowl retaining nut “0” ring. Apply light coating of Vaseline to “0” ring on filter

bowl retainer nut (Item 1) and install on stud with 50 to 60 inch pounds torque.

M. Secure Filter bowl retainer nut with .032” stainless steel lockwire.

N. Turn airframe fuel shutoff valve to “ON” position. Turn electric fuel boost pump on and

observe 20 PSI on fuel pressure gauge. Observe fuel filter assembly for leaks prior to

closing filter access panel.

OPTIONAL: FUEL FLOW

Some aircraft are equipped with a fuel flow unit. The Shadin Company Inc. Miniflo Digital fuel

management system incorporates an indicator and transducer. The transducer is installed in the

fuel line between the engine’s FCU and the fuel flow divider/dump valve. (See Shadin Company

Inc. Miniflow maintenance manual for troubleshooting and repairing data for the fuel flow system.)

Effective: 1/26/05 5 -8


FUEL LINE MAINTENANCE

CS 3204 A2 may be used as a thread lubricant or to seal minor connection leaks throughout the

fuel system. Apply sparingly to male fittings only. Always insure that a sealing compound or

residue from a previously used compound, or any other foreign matter does not enter the fuel

system.

Any repair that breaks the fuel tank seal will necessitate resealing of that area of the tank. Repair

parts that need sealing must be installed and riveted during the sealing operation.

SEALING COMPOUNDS

CS 3204 A2 or B2 meets AMS-S-8802 (formerly Mil-S-8802) standards. It is a fuel resistant sealant

use on integral “wet wing” fuel tanks as well as other areas subject to contact with aircraft fuels,

lubricants, oils, agriculture chemicals, water and/or weathering. Thrush Aircraft Inc uses two

grades; CS 3204 A2 which is thin, brush able, and self leveling liquid and CS 3204 B2 which is a

thixotropic paste that will not flow or sag on overhead or vertical surfaces. Thrush Aircraft Inc

recommends the use of “Semkit®,” which are easy-to-use pre-measured 6 oz. Plastic tubes with a

4.5 oz. Fill of product. When mixing materials packaged in bulk or when only a small quantity is

required, stir 10 parts by weight of the part “B” component into 100 parts by weight of the part “A”

component. Mix and stir both components until a uniform gray color is achieved. There should be

no white or black streaks in the properly blended material. Blend the components slowly, as violent

stirring will entrap air in the cured sealant. Do not thin CS 3204 with solvents. Thoroughly clean all

surfaces to which CS 3204 is to be applied immediately prior to sealant application. Cleaning shall

be accomplished with clean lint-free paper or cloth towels or small paintbrushes soaked with

Acetone or Methyl Ethyl Ketone and wiped clean. Always clean an area longer and wider than the

width of the finally applied sealant to insure maximum bonding. CS 3204 is also used to make and

seal all exposed stressed skin joints and overlap fillets, fiberglass to aluminum overlap fillets and

seal cockpit windows to prevent water and agriculture chemical entry into these vital structures. CS

3204 is used to seal all bolts in hold-down and carry-through duty in the chemical hopper. CS 3204

can be painted when cured. Alternate sealers for CS 3204 class A & B are PR-1422 class A & B

and PR-1750 class A & B. For fast set up times (20 minutes application life and cure time) “Quick

Set” CS-3204 B1/4 or PR-1435 may be used as a alternate sealer anywhere on the aircraft. CS

3600 (Mil-S-4383C) is used by Thrush Aircraft Inc as a topcoat for all the above listed polysulfide

5 -9 Effective: 1/26/05

** CAUTION **

Protect all drain openings and fuel outlet screens when

applying sealant.


sealers inside of the Thrush’s integral fuel tanks. It is one part (no mixing), has the consistency of

thin syrup and can be painted on top of all previously sealed internal fuel tank seams. If CS 3600 is

used, it must be allowed to air dry for 4 days minimum before being exposed to fuel. “Semkit®” pre-

measured cartridges can best be utilized by use of either a Semco® model 250 pneumatic or model

850 hand operated application gun. Thrush Aircraft Inc factory uses the following nozzles:

Semco® model Nos. 252 (2.5”, 1/16” orifice), 410 (4”, 1/32” orifice), or 440 (4”, 1/8” orifice). The

plastic nozzle tip can be cut with a razor knife to enlarge or modify the tips orifice size and shape to

control the size and shape of the material bead.

LEAK SEALING

Determine the approximate location of the leak by visual inspection through the cover plates in the

lower surface of the wing. After leak area is determined, drain all fuel from affected tank. See

Section II for defueling procedures.

***WARNING***

Refer to and adhere to all measures and precautions

obtained from the applicable Material Safety Data Sheet

(MSDS) prior to using or removing CS 3204 and any

other chemicals, adhesives, materials, oils, fuels,

sealers, cleaners, or solvents listed in this manual.

A. Remove the cover plates on upper surface of wing to repair the tank leak. Sealing can be

accomplished through these openings.

B. Clean the general area of the leak with clean paper towels. Apply an even coating of CS

3204 A2 with a stiff clean brush. Catalyst is furnished and should be carefully mixed

according to instructions on the container.

C. Allow the sealer to dry overnight.

D. After drying, the sealer should be checked for air bubbles or thin spots. Additional sealer

should be applied where necessary.

E. Reinstall the cover plates on wing upper surface.

Effective: 1/26/05 5 - 10

**CAUTION**

Assure that the leak is not being caused by structural

cracks, loose seams or any source other than a pinhole

from around a properly installed fastener.


RESEALING AFTER COMPLETE SKIN REMOVAL

To reseal the fuel tanks after removing or repairing the wing skin, proceed as follows:

A. Prior to installing the wing tank skin, all surfaces that will receive sealant shall be cleaned

and etched.

B. Apply CS 3204 B2 mixed compound to all areas of contact between the skin and rib

structure.

C. Rivet the wing tank skin in place and allow the sealer to dry until tacky to the touch.

D. After adequate drying, the sealer should be checked for air bubbles or thin spots. Apply

additional CS 3204 A2 sealer as necessary.

E. Reseal cover plates and fuel quantity transmitter mounting with CS 3204 B2.

F. Vacuum tank area thoroughly to remove all particles of dried sealant, dirt or other foreign

matter.

G. Allow the sealant to cure for 16 hours or more.

H. Pressure check fuel tank from 38 to 44 inches of water-manometer, for 3 (+1) minutes.

**CAUTION** Do not attempt to apply pressure to the tank without first sealing off all lines

and vents, and without an adequate regulator to control pressure. Do not

pressurize the tank in excess of 1.589 psi (44.0 inches of water-manometer)

or damage may occur.

I. To prevent water and chemical entry into wing and empennage skin joints and edges,

make fillets by applying a small bead of CS 3204 B2 to all skin edges, joints, and overlaps.

The fillets can be painted after sealer has dried.

TROUBLESHOOTING

The trouble-shooting figure in this section discusses symptoms, which can be diagnosed and

interprets the results in terms of probable causes and the appropriate corrective remedy to be

taken. Review all probable causes given and check other listings of troubles with similar symptoms.

Items are presented in sequence, but not necessarily in order of probability.

5 -11 Effective: 1/26/05


FUEL SYSTEM TROUBLE SHOOTING CHART

TROUBLE PROBABLE CAUSE REMEDY

Fuel tanks empty.

Check fuel quantity. Service

with proper grade and amount

of fuel.

Fuel quantity indicator circuit

breaker open or defective.

Check visually. If not open,

check continuity. Reset.

Replace if defective.

No fuel quantity indication.

Defective fuel quantity

indicator or transmitter.

Disconnect wire from

transmitter at indicator not

registering and attach it to an

indicator that is registering. If

indicator does not register,

transmitter is defective. If the

new indicator registers, the

existing indicator is defective.

Replace defective transmitter

or indicator.



Effective: 1/26/05 5 -12



Loose connections or open

circuit.

Check connections and wiring.

Tighten connections; repair or

replace wiring.

Left and right fuel quantity

indicator switch defective.

Check continuity and replace if

defective.

No power to gauge.

Check power to gauge. If no

power, check for defective

circuit breaker.

No fuel quantity indication.

continued.

Power, ground and

transmitter checks good.

Circuit board on rear of gauge

defective. (Replace board) or

entire gauge.

Fuel indicated full at all times. Open ground between gauge

and transmitter.

Check ohms to transmitter.

Check for broken wire.

Transmitter should read 0

ohms when fuel tank is empty

and 33 ohms when fuel tank is

full.

No fuel flow to engine-driven

fuel pump. Fuel tanks empty.

Check fuel quantity. Service

with proper grade and amount

of fuel.

5 -13 Effective: 1/26/05



Fuel line disconnected or

broken.

Inspect fuel lines. Connect or

repair fuel lines.

Header tank outlet fuel

strainers plugged.

Disconnect fuel lines from tank

outlets. No fuel indicates

plugged strainers. Remove

and clean strainers and flush

out tanks.

Fuel filter element plugged. Inspect filter element. Clean

or replace filter element.

No fuel flow to engine-driven

fuel pump. continued.

Fuel line plugged.

Starting at fuel pump inlet,

disconnect fuel lines

successively until plugged line

is located. Clean out or

replace fuel line.

Partial fuel flow from the

preceding causes.

Use the preceding isolation

procedures, checking for

sufficient rate of flow. Using

the preceding remedies. Fuel starvation after starting.

Malfunction of engine-driven

fuel pump.

Check pump outlet during

starting. Replace fuel pump.

See Section IV.

Effective: 1/26/05 5 -14



Fuel starvation after starting.

continued. Fuel vents plugged.

Pressure check each vent line.

Clean or replace vent line.

Defective electric auxiliary

fuel pump switch.

Check continuity of switch.

Replace defective switch.

Open or defective circuit

breaker.

Check visually. If not open,

check continuity. Reset.

Replace if defective.

Loose connections or open

circuit.

Check connections and wiring.

Tighten connections; repair or

replace wiring.

Defective auxiliary fuel pump.

Disconnect outlet line. With

proper fuel supply to pump,

fuel under pressure should

flow from outlet. Replace

defective pump.

No fuel flow when auxiliary

pump is turned on.

Defective engine-driven fuel

pump by-pass valve.

Check pump outlet during

starting. See Section IV and

replace fuel pump if by-pass

valve is defective or installed

backwards.


5 -15 Effective: 1/26/05



Fuel flow indicator

inoperable. No voltage to indicator.

Check voltage and ground

wire. If voltage is present and

ground is good, replace

indicator, maintaining the

same K factor.

Indicator comes on but will

not show fuel flow.

Bad wires to transducer or

defective transducer.

To check transducer, remove

four screws holding wire

housing to flow vane housing.

With battery power on, pass

screwdriver back and forth

over wire housing pickups.

You should get a reading on

indicator. If no reading,

replace units.

*NOTE*

Any time you have to replace either the fuel flow indicator or the transducer, you must be sure to

have unit calibrated to same K factor as set by manufacturer. This will cause bad indications if

mismatched K factors are installed together.

ACTIVATING HOPPER (FERRY) FUEL SYSTEM

A. Remove spray pump and spray pump discharge line to spray valve. Tie-rap any wires or

cables to upper portion of pump mount.

B. Open hopper gatebox dump gate.

C. Assure hopper has been cleaned thoroughly and there is no presence of water or

chemicals.

D. Assure side loading plumbing has been cleaned and there is no presence of water or

chemicals.

E. Install the 2" camloc female cap on the spray valve return inlet fitting located inside the

hopper gatebox left hand side and lockwire.

Effective: 1/26/05 5 -16


F. Install cap on hopper outlet fitting and lockwire.

G. Hook up fuel line from cap to fuel selector valve. Make sure that the fuel line is secured to

aircraft structure and will not foul any moveable controls.

H. Service hopper with approved fuel.

** CAUTION ** Operation instructions must be followed to operate

aircraft using ferry fuels.

Operating instructions for the P/N 60167 ferry fuel system

A. Securely attach these instructions in the cockpit on the hopper, directly in

front of the pilot's face at the time of installation of the ferry fuel system.

B. In the United States, an aircraft with this ferry fuel system installed and

connected to the normal fuel system must be operated on a special flight

authorization (ferry permit) regardless of whether the ferry fuel system is

actually used on any particular flight.

C. Due to vapor lock considerations, use of aviation gasoline as an alternate

fuel is prohibited in either the wing tanks or the hopper.

D. Do not use hopper fuel for takeoff, landing, or flight at low altitude.

Use hopper fuel only for level cruising flight above 3000 feet above ground

level. Always operate the electric fuel pump and the ignition switch while

changing the fuel selector in flight. Always switch fuel at or below cruise

power settings.

E. Except in emergency, do not dump hopper fuel in flight or on the ground

with the engine running.

F. Drain the hopper sump and all other normal fuel system sumps prior to

flight.

5 -17 Effective: 1/26/05


Important Note: with the ferry fuel selector in hopper position, drain all

trapped air from the hopper fuel line by operating the fuel filter drain

located on L/H shin skin forward of firewall (see Fig. 5-1). Unless this

procedure is followed after each refueling, the engine may flameout

when hopper fuel is selected in flight.

G. Never use the hopper as a fuel tank unless it is completely clean and dry.

H. Remove these instructions from the cockpit only after removal of the ferry

fuel system from the aircraft.

Effective: 1/26/05 5 -18


Figure 5-1

5 -19 Effective: 1/26/05


Figure 5-2

Effective: 1/26/05 5 -20


Section 6 LANDING GEAR, WHEELS & BRAKES

TABLE OF CONTENTS LANDING GEAR, WHEELS AND BRAKES .................................................................................................... 2 GENERAL DESCRIPTION .............................................................................................................................. 2

REMOVAL .................................................................................................................................................... 2 CLEANING, INSPECTION AND REPAIR OF MAIN GEAR......................................................................... 2 CLEANING, INSPECTION AND REPAIR OF SHOCK STRUTS................................................................. 2

TAIL LANDING GEAR ..................................................................................................................................... 3 SPRING STEEL............................................................................................................................................ 3 REMOVAL .................................................................................................................................................... 3 CLEANING, INSPECTION AND REPAIR OF TAIL GEAR .......................................................................... 3 DISASSEMBLY OF SPINDLE HOUSING ASSEMBLY................................................................................ 4 CLEANING, INSPECTION AND REPAIR OF TAIL GEAR SPINDLE HOUSING ASSEMBLY. .................. 4 INSTALLATION ............................................................................................................................................ 5 TAIL GEAR RIGGING................................................................................................................................... 5

TAIL WHEEL REMOVAL AND DISASSEMBLY.............................................................................................. 6 TO REMOVE AND DISASSEMBLE TAIL WHEEL/TIRE, PROCEED AS FOLLOWS:................................ 6 INSPECTION OF TAIL WHEEL ASSEMBLY............................................................................................... 6 TAIL WHEEL REASSEMBLY AND INSTALLATION (5.00—5 10PR) ......................................................... 7

WHEELS AND BRAKES.................................................................................................................................. 8 GENERAL DESCRIPTION .............................................................................................................................. 8 MAIN WHEEL REMOVAL AND DISASSEMBLY............................................................................................. 8

DIVIDED TYPE WHEEL (CLEVELAND) ...................................................................................................... 8 INSPECTION OF MAIN WHEEL ASSEMBLY ............................................................................................. 9 REASSEMBLY AND INSTALLATION ........................................................................................................ 10

SERVICING ................................................................................................................................................... 11 MEASURING BRAKE LINING WEAR........................................................................................................ 11 REMOVAL OF LININGS FROM CALIPERS .............................................................................................. 11 REPLACEMENT OF ORGANIC LININGS.................................................................................................. 12 REASSEMBLY OF ORGANIC LININGS TO CALIPER.............................................................................. 12 BRAKE LINING CONDITIONING PROCEDURES .................................................................................... 12 T34 THRUSH NONASBESTOS ORGANIC LININGS................................................................................ 13 BRAKE REMOVAL AND DISASSEMBLY.................................................................................................. 13 BRAKE REASSEMBLY AND INSTALLATION........................................................................................... 15 REMOVAL OF BRAKE MASTER CYLINDERS ......................................................................................... 15 DISASSEMBLY AND REPAIR ................................................................................................................... 15 INSTALLATION .......................................................................................................................................... 16 BRAKE BLEEDING..................................................................................................................................... 16

TROUBLESHOOTING CHART...................................................................................................................... 18

Effective: 1/26/05 6-1


LANDING GEAR, WHEELS AND BRAKES

SECTION SIX

GENERAL DESCRIPTION Each main landing gear installation consists of a landing gear assembly, shock strut assembly and a

wheel and brake assembly. The landing gear assembly is bolted to the fuselage frame at two locations

and to the shock strut assembly at one location. The shock strut assembly is in turn bolted to the

fuselage frame. Lubrication fitting are provided for the pivot points and for the shock strut assembly.

Lubrication should be applied sparingly and all parts wiped clean to prevent collection of dirt (refer to

lubrication Chart, Section 2). All landing gear hinge points should be carefully inspected for wear and

damage during each landing gear check. Trouble shoot the landing gear by using the charts at the

back this section, and always places the aircraft on jacks prior to performing maintenance procedures

on the landing gear system.

REMOVAL

A. Jack aircraft as outlined in Section 2.

B. Remove fuselage skins as required.

C. Disconnect flexible hydraulic brake line at top of landing gear assembly.

D. Remove bottom bolt from shock strut assembly.

E. Remove the bolts attaching landing gear assembly to fuselage.

F. Remove top bolt on shock strut assembly.

CLEANING, INSPECTION AND REPAIR OF MAIN GEAR

A. Clean all parts with a suitable type cleaning solvent.

B. Inspect all bolts, bearings and bushings for excess wear, corrosion and damage.

C. Check all welds for cracks.

D. Repair of the landing gear is limited to reconditioning of parts, such as replacing

components, bearings and bushings, smoothing out minor nicks and scratches and

repainting areas where paint has chipped or peeled

CLEANING, INSPECTION AND REPAIR OF SHOCK STRUTS

A. Remove top and bottom attaching bolts, and remove complete shock strut assembly from

aircraft.

B. Support strut under fork end and slide a 1-1/2 inch ID sleeve over slotted end. Apply light

pressure to sleeve, sufficient to relieve pre-load from biscuits.

C. Remove bolt holding biscuit retainer and disassemble unit.

D. Clean all parts with a suitable type cleaning solvent.

E. Inspect rubber shock biscuits for distortion, splits or deterioration. Replace as required.



F. Inspect welds for cracks.

G. Inspect bolt holes for elongation.

H. Inspect all areas for evidence of corrosion.

I. Repair of shock struts is limited to replacement of parts, smoothing out minor nicks and

scratches and repainting areas where paint has chipped or peeled.

*NOTE* Upper shock strut attach bolt is close tolerance, heat treated NAS bolt. Do

not replace with AN type bolt.

TAIL LANDING GEAR

SPRING STEEL

The tail gear consists of a 1-inch thick alloy spring steel, tail gear sub-assembly fork assembly and wheel

assembly. The tail wheel is a locking type and is actuated from the elevator bell crank by a cable.

Centering springs align the tail wheel, and a pin engages and locks the wheel in the trailing position.

Pushing the control stick fully forward disengages the locking pin and the wheel is free to caster for taxiing.

REMOVAL (REFER TO FIG. 6-1 AND 6-1a)

A. Remove fuselage skins as required.

B. Using a suitable Jack. Jack and secure tail of aircraft, using jackpoint.

C. Remove outer dust cover (hubcap).

D. Remove cotter pin and axle castellated nut, and spacer P/N CA84106-05-2, and then remove the

tire/wheel assembly.

E. Disconnect flex control lock cable at pivot arm and cable hold down clamp.

F. Disconnect centering springs from tail wheel centering arm assembly by removing attach bolt.

* NOTE * Do not alter lock cable or elevator travel stops. Alteration of tail gear

lock cable or elevator travel stops will require re-rigging of tail wheel

locking system.

G. If disassembly of tail wheel/tire assembly is necessary, follow steps highlighted on page 6-5

and 6-6.

H. Remove main leaf spring assembly by removing NAS6207-38D bolt (inspect every 100 hours)

holding spring to trunnion assembly. Remove two-(2) each NAS6606-54 bolts that holds the lower

spring support block to upper support block. Note how many 90056-26 washers were located on

each side between support blocks.

I. Remove trunnion assembly from fuselage by removing trunnion attach shaft.

CLEANING, INSPECTION AND REPAIR OF TAIL GEAR


6-3 Effective: 1/26/05


B. Remove, clean, and inspect leaf spring forward attach P/N NAS6207-38D bolt every 100 hours.

Upon reassembly lubricate bolt and leaf spring hole with Snap-on™ General Purpose Antiseize or

equivalent or MIL-G-81322 (Aeroshell 22) grease. Torque to specifications I/A/W Torque chart

(figure 2-7). Replace MS24665-300 cotter pin each inspection.

C. Inspect all bolts holes for elongation. As a general rule, replace components with holes that are out

of round by 0.005” or more. Replacement of the leaf spring forward attach P/N NAS6207-38D bolt (inspect every 100 hours) with a larger diameter is not approved. The leaf spring may not be “drilled out” for a larger bolt.

D. Inspect main spring leaf for corrosion and cracks. Check aircraft maintenance records to be sure

spring leaf P/N 5079-1 has not exceeded its five thousand (5,000) flight hour life limit. Replace leaf

spring as needed.

E. Inspect spindle housing assembly welds for cracks.

F. Inspect spindle housing assembly for cracks and corrosion.

G. Inspect lock pin and upper and lower lock plates for wear, corrosion, cracks, and proper operation.

H. Inspect centering springs for corrosion, wear at ends, and for correct operation.

I. Inspect lock pin flexible cable and spring for corrosion and correct operation.

J. Inspect P/N95207-1 Acetal (Delrin®) lower support block spacer for wear and cracks.

K. Inspect upper and lower leaf spring support blocks, and attachment hardware for wear, corrosion,

and cracks. Ensure that the leaf spring support blocks grips the leaf spring tightly to prevent leaf

spring movement fwd. and aft. Ensure flexible sealant around contact edges of support blocks,

lower support block spacer and leaf spring is intact to prevent collection of potential corrosive

material in this area. Lubricate 2 ea. Trunnion Zerk (grease) fittings with MIL-G-81322 (Aeroshell

22)

L. Repair of the tail landing gear is limited to replacement of component parts, bearings, bushings,

smoothing out minor nicks and scratches, repainting chipped or peeled areas.

DISASSEMBLY OF SPINDLE HOUSING ASSEMBLY (REFER TO FIG. 6-1)

A. If desired, remove bolts, nuts and washers that bolt tail wheel fork to spindle.

B. Remove bolts, nuts, and washers that bolt centering arm to top of spindle and remove centering

arm. Note orientation for proper reassembly.

C. Remove red plastic cap plug (dustcover).

D. Remove cotter pin, castellated nut, tongue washer, grease cup washer, and cone bearing.

E. Remove spindle assembly and thrust washer from spindle housing. Do not remove upper bearing

cup or bottom bronze bushing unless replacement is indicated by inspection.

CLEANING, INSPECTION AND REPAIR OF TAIL GEAR SPINDLE HOUSING ASSEMBLY.


B. Inspect all bolts, bearings and bushings for excessive wear, corrosion and damage.



C. Inspect spindle assembly for cracks, excessive wear, corrosion and damage.

D. Inspect spindle housing for cracks, excessive wear, corrosion and damage.

E. Inspect lock pin lower plate and lock pin top plate assembly for cracks, corrosion and damage.

F. Repair of tail gear sub-assembly is limited to reconditioning of parts such as replacing bearings and

bushings, smoothing out minor nicks and scratches, repainting chipped or peeled areas and

replacement of component parts.

INSTALLATION (REFER TO FIG. 6-1 & 1a)

The tail gear may be installed by reversing the removal procedures. Ensure that trunnion is straight down

(6 O’clock position) and that leaf spring support blocks grips the leaf spring tightly to prevent movement

fwd. or aft. (Add or subtract P/N 90056-26 washers/spacers (.063”) between upper and lower support

blocks to achieve a tight grip of leaf spring after bolts are properly torqued.) All bolts shanks and bolt

holes are to be coated with Snap-on™ General Purpose Antiseize lubricant or equivalent before

installation. Lubricate all bearings, bushings, and Zerk (grease) fittings with MIL-G-81322 (Aeroshell 22)

grease. Torque all hardware in accordance with TORQUE CHART (figure 2-7) with the exception of the

top spindle castellated nut and wheel/tire axle castellated nut, which should be torqued as follows:

A. For spindle castellated nut: While manually rotating spindle, torque spindle castellated nut to 20

inch-pounds, continue rotating spindle and back off to zero inch-pounds. While manually rotating

spindle, torque nut to 10 inch pounds. If not in locking position, advance nut to next position, not to

exceed 30º, and install cotter pin. Bend ends of cotter pin around spindle castellated nut. Note:

Spindle must rotate freely without perceptible play.

B. For tail wheel axle castellated nut: While manually rotating wheel/tire, torque axle castellated nut to

80 inch-pounds, continue rotating wheel and back off to zero inch-pounds. While manually rotating

wheel/tire, torque to 30 to 40-inch pounds. Rotate axle castellated nut (clockwise or

counterclockwise) to nearest slot and cotter pin hole, and insert cotter pin. Bend ends of cotter pin

around axle nut. Note: Wheel/tire must rotate freely without perceptible play.

C. After the components have been installed, seal the contact edges where the spring P/N 5079-1

(replace every 5,000 hours), upper support block P/N 94131-9, lower support block P/N 94131-11

and spacer P/N 95207-1 come together with a high quality flexible silicone sealant or fuel tank

sealant CS3204 B2 (AMS-S-8802 formerly MIL-S-8802) to help block the collection of potential

corrosive contaminants in this area.

D. Carefully lower aircraft to ground and remove Jack.

E. Recheck tire inflation pressure (5.00-5 10pr Type III is 88psi) and install dust cover (hubcap).

TAIL GEAR RIGGING

Rigging will be required if lock cable or elevator travel stops have been altered in any way. Rig as follows:

A. Place elevator in a 17 (±1) degrees down position.

B. Connect lock cable to pivot arm.



C. Assure lock pin is flush with bottom of lock pin cylinder when making final adjustment to lock

cable turn barrel.

D. Adjust top plate as required to assure straightforward travel of aircraft tail wheel when tail gear is

locked.

TAIL WHEEL REMOVAL AND DISASSEMBLY

TO REMOVE AND DISASSEMBLE TAIL WHEEL/TIRE, PROCEED AS FOLLOWS: (REFER TO FIG. 6-1)

A. Using a suitable Jack. Jack and secure tail of aircraft, using provided jackpoint.

B. Remover dustcover (hubcap) and deflate tire by depressing the schrader valve stem plunger until

air can no longer be heard escaping from the tire.

C. Remove schrader valve core.

D. Remove cotter pin and axle castellated nut. Rock wheel/tire slightly, then remove wheel/tire

assembly from axle, and be sure to capture the spacer P/N CA 84106-05-2 located under nut.

E. From each side of wheel; carefully remove snap ring, felt grease seal retainer, felt grease seal,

grease seal ring and cone bearing. Store the cone bearings. Label the bearings for reinstallation

into position from which it was removed.

F. With the tire completely deflated, removing the wheel through-bolts will separate the wheel

halves. Pull the wheel halves from the tire by removing the wheel half opposite the valve stem

first. Mark wheel halves to note relationship to each other for reassembly.

INSPECTION OF TAIL WHEEL ASSEMBLY

A. Visually check all parts for cracks, corrosion, distortion, defects and excessive wear.

B. Inspect felt grease seals. Replace if surface is hard or contaminated, or shows evidence of

excessive wear. Lightly saturate grease seal felts with SAE 10wt. Oil (3-in-ONE oil) (do not soak).

C. Inspect tire for cuts, anomalies, internal damage and deterioration.

D. Inspect inner tube for cuts, wrinkles, anomalies and deterioration. Note: Do not use a used inner

tube with a new tire. Tubes grow in service, taking a permanent set of about 25% larger than

original size. This makes a used tube too large to use in a new tire, which could cause a wrinkle and

lead to tube failure.

E. Inspect wheel bearing grease for contamination and solidification at each periodic inspection.

Repack bearings with MIL-G-81322 (Aeroshell 22) or equivalent grease. Note: Do not exceed 500

wheel miles or on annual inspection whichever comes first between repacking intervals.

F. Clean and inspect bearing cups and cones. Note: Do not spin dry bearings or handle bearing

components with bare hands. The bearing cup should not be removed except when replacement

is necessary due to scratches, nicks, pitting, spalling, corrosion, brinelling, or evidence of

overheating. Note: If bearing cup is replaced, its companion bearing cone must also be replaced.



1. Bearing cup removal: Heat wheel half in an oven not exceeding 212°F for 15 minutes.

Remove wheel half from heat source and immediately remove bearing cup by carefully

tapping out evenly from the inside with a fiber drift.

2. Bearing cup installation: Place wheel half in oven not exceeding 212°F for 15 minutes. Chill

new bearing cup in an atmosphere of -25°F to -65°F for no less than 4 hours. Chilling can

also be accomplished by placing the bearing cup in dry ice for a minimum of 15 minutes.

Dry cup thoroughly and installed chilled bearing cup into bore of heated wheel half using a

thin coat of zinc chromate primer as a protectant/lubricate. Tap gently into place with fiber

drift making sure bearing cup is evenly seated against shoulder of wheel half. Avoid

cocking bearing cup during installation. If bearing cup will not seat properly in wheel half,

repeat above said procedures or replace wheel half assembly.

G. Replace any wheel casting that is distorted, corroded, or has visible cracks.

TAIL WHEEL REASSEMBLY AND INSTALLATION (5.00—5 10pr)

To assemble and reinstall tail wheel, refer to Figure 6-1 and proceed as follows:

A. Wipe tire and tube (serviceable or new) with denatured alcohol, followed by soap and water, then

dry thoroughly.

B. Inflate the inner tube just enough to round it out, dust tube lightly with tube talc.

* NOTE * Tires and tubes are balanced as individual units and marked at time of

manufacture. The tire balance mark is a red dot. The tube balance mark

is a yellow stripe on the base of the tube. Always assemble tire and tube

with marks aligned.

C. Place tube in tire and align balance marks. If tube has no balance mark, place valve stem

adjacent to tire balance red dot.

D. Install tire and tube on the wheel half containing the valve stem hole and then the opposite.

E. Install the wheel through-bolts with bolt heads opposite valve stem side, tighten nuts evenly and

torque to 90 inch-pounds.

*** WARNING *** Uneven or improper torque may cause a bolt or wheel failure. Inflate tire

until tire beads are sealed, remove schrader valve core, and allow tube

to completely deflate. Install the valve-core and inflate 5.00-5 10pr tire to

88 psi. Assure schrader valve does not leak before replacing valve cap.

F. Repack bearing cones with MIL-G-81322 (Aeroshell 22) grease or equivalent.

G. On each side of wheel; apply a thin coating of grease on bearing cups, installed freshly repacked

bearing cones, install flat grease seal ring, install felt grease seal retainer with felt seal installed



Note: Lightly saturate grease seal felts with SAE 10wt. Oil (3-in-ONE oil) (do not soak), and

carefully install snap ring. Install the two- (2) P/N95435-11 spacers, one on each side of wheel

assembly.

H. Install dust cover with center hole on opposite valve stem side of wheel.

I. Inspect tail wheel axle for anomalies, then apply a light coating of grease.

J. Install tail wheel/tire assembly onto tail wheel axle with valve stem side facing outboard.

K. Install tail wheel axle castellated nut: While manually rotating wheel/tire, torque axle castellated nut

to 80 inch- pounds, continue rotating wheel and back off to zero inch-pounds. While manually

rotating wheel/tire, torque to 30 to 40-inch pounds. Rotate axle castellated nut (clockwise or

counterclockwise) to nearest slot and cotter pin hole, and insert cotter pin. Bend ends of cotter pin

around axle nut. Note: Wheel/tire must rotate freely without perceptible play.

L. Carefully lower aircraft to ground and remove Jack.

M. Recheck tire inflation pressure (5.00-5 10pr is 88psi) and install dust cover (hubcap).

WHEELS AND BRAKES

GENERAL DESCRIPTION

The divided type wheels (including tail wheel) are machined castings, consisting of two sections called

wheel halves. The wheel halves, which are secured together by bolts and nuts, are interchangeable, and

the complete wheel assemblies are interchangeable according to wheel size. The wheels operate on

tapered roller bearings that rotate in hardened steel races pressed into each wheel half. A brake disc

assembly is bolted to the wheel and turns with the wheel. Applying pressure to the rudder-brake pedals

individually controls the hydraulic brakes attached to the main landing gear. Movement of a rudder-brake

pedal operates the corresponding master brake cylinder, attached to the aft side of the rudder pedals, and

applies pressure to the appropriate brake. The brakes are self-adjusting, easily checked for wear, and can

be quickly overhauled by field activities.

MAIN WHEEL REMOVAL AND DISASSEMBLY

DIVIDED TYPE WHEEL (CLEVELAND)

To remove and disassemble a main landing gear wheel, proceed as follows:

A. Jack aircraft as outlined in Section Two.



B. Remove valve-core and deflate tire completely.

C. Remove bolts and washers from back plates of brake assembly and remove back plates.

D. Remove hubcap snap ring, hubcap, cotter pin, nut, washer, bearing and wheel assembly from

landing gear.

E. Break tire bead from wheel by using a mallet (do not use tire irons).

F. Remove bolts, washers and nuts and separate wheel halves. Guard valve stem to avoid damage

while removing tire and tube.

G. Remove brake disc from brake side of wheel. If disk sticks, pry out disc using non-metallic

instrument.

F. Remove bearing retainer snap ring, grease seal ring, and grease seal, spacer and bearing cone

from inboard side of wheel.

* NOTE *

Wheel halves can be replaced individually. Wheel

sets no longer have to be replaced as match pairs.

INSPECTION OF MAIN WHEEL ASSEMBLY (CLEVELAND, DIVIDED TYPE)

A. Clean all parts in cleaning solvent and dry thoroughly. A soft bristle brush may be used to remove

hardened grease, dust or dirt.

*** WARNING ***

Cleaning solutions are toxic and volatile. Use in a well-

ventilated area. Avoid contact with skin or clothing. Do not

inhale vapors.

B. Inspect bearing cones for nicks, scratches, water staining, spalling, heat discoloration, roller wear,

cage damage, cracks or discoloration.

C. Inspect wheel-bearing grease for contamination and solidification (see Inspection Intervals Chart in

Section II). When repacking wheel bearings, use MIL-G-81322 (Aeroshell 22).

D. Inspect wheel halves for cracks, corrosion and other damage. A cracked or badly corroded casting

should be replaced. Small nicks scratches or pits can be blended out using fine 400-grit sandpaper.



E. Inspect snap rings and grease seals for distortion or wear. Replace parts, if damage or deformed.

Saturate grease seal felts with SAE 10 oil (do not soak).

F. Inspect bearing cups for looseness, scratches, pitting, corrosion, or evidence of overheating. The

bearing cups are pressed into the wheel halves and should not be removed unless replacement is

necessary due to the above conditions. If replacement is necessary, proceed as follows:

1. Insert wheel half into boiling water for one (1) hour or place it in an oven at 250 degrees

Fahrenheit for 30 minutes.

2. Remove wheel half from source of heat and invert wheel half. If bearing cup does not drop

out, tap the bearing cup evenly from the axle bore with a fiber drift pin or suitable arbors

press.

3. When replacing a bearing cup, repeat step 1, and chill bearing cup in dry ice for a minimum

of 15 minutes.

4. Remove wheels half from source of heat and bearing cup from the dry ice.

5. Dry the chilled bearing cup and coat its contacting surfaces with zinc chromate primer.

6. Install the chilled bearing cup into the bearing bore of the heated wheel half. Tap bearing

cup gently and evenly into place, using a fiber drift pin or suitable arbor press.

G. Inspect wheel brake disc assembly for cracks, excessive wear or scoring, rust and corrosion.

Remove corrosion and blend out small nicks using fine (400 grit) sandpaper. Replace brake disc if

worn below wear limit of .395 inch (see Fig. 6-2). Coning of disc in excess of 0.015 inch is cause

for replacement of disc (see Fig. 6-3).

H. Inspect self-locking nuts for self-locking feature. Replace nuts if they can be turned onto the bolt

past the self-locking section by the finger.

REASSEMBLY AND INSTALLATION (CLEVELAND DIVIDED TYPE)

Tires and tubes are balanced as individual units and marked at the time of manufacture. The tire balance

mark is a red dot. The tube balance mark is a yellow stripe on the base of the tube. The following

procedure is suggested as a guide for mounting the tires in balance and installing the wheels.

A. Reassemble cone bearings, grease seals, felts and snap ring into the proper wheel halves.

Lubricate bearings. See Inspection of Main Wheel Assembly.

B. Inflate tube sufficiently to round it out.

C. Dust tube with a small amount of tube talc.

D. Insert tube into tire so that balance mark (yellow or white band) is radically aligned with the tire

balance mark (red dot).

E. Place outer wheel half into tire and pull tube valve stem through valve hole.

Effective: 1/26/05 6-10


F. Turn tire and wheel half over and place inner wheel half into the tire and align the bolt holes with the

outer wheel half.

G. Place brakes discs into the inner wheel half and align bolt holes.

H. Install bolts through the inner wheel half and washers and nuts on the outer wheel half.

I. Tighten nuts evenly and torque to 150 inch-pounds.

*** WARNING ***

Uneven or improper torque may cause Bolt or wheel failure.

J. Inflate tube until beads seat on wheel flanges. Remove valves core and allow tube to deflate.

K. Install valves core and inflate tires from 40 to 60 psi. Check to assure valve stem does not leak

before installing valve cap.

L. Lubricate washer and axle nut (see Section II Servicing). Install wheel assembly on axle and

secure with washer and axle nut.

M. While manually rotating wheel, torque axle nut to 80 inch-pounds, continue rotating wheel and back

off to zero inch-pounds. While manually rotating wheel, torque to 40-inch pounds. If nut is not to

locking position, advance to next position, not to exceed 30 degrees, and install cotter pin.

N. Install hubcap and hubcap retaining ring.

O. Install brakes back plate assembly and torque bolts to 60 inch-pounds. These bolts are self-

locking and should be inspected for the self-locking feature. Replace bolts if the self-locking

feature is damaged or destroyed.

P. Wheels may be repainted if the parts have been repaired and thoroughly cleaned. Paint exposed

areas with one coat of zinc primer and one coat of aluminum lacquer.

* NOTE * Do not paint working surfaces of the bearing cups.

SERVICING

MEASURING BRAKE LINING WEAR

(See latest edition of Cleveland Manual number AWBCMM0001-5 for pertinent details.)

The minimum wear thickness for replacement of metallic linings is 0.100 inch (refer to Fig. 6-2).

REMOVAL OF LININGS FROM CALIPERS

A. Remove backing plate attaching bolts and washers, and remove back plates and insulator shim.

A. Carefully slide brake caliper out of torque plate bushing.

B. Slide pressure plate assembly (lining carrier) off anchor bolts.

6-11 Effective: 1/26/05


REPLACEMENT OF ORGANIC LININGS

A. Old organic linings may be removed by using a small drift pin or carefully drilling out the rivets with a

1/8-inch diameter drill. Use care to prevent elongating the rivet holes. Deburr the surface adjacent

to the lining to allow lining to set flush.

B. Clean pressure plate and back plate surfaces of dirt, grease, etc. before installing new linings.

C. Inspect pressure plate and back plate for excessive corrosion, visible damage, or excessive

warping. Straighten pressure plate too less than 0.010 inch (0.254mm) flatness.

D. Align new factory authorized replacement lining segments on pressure plate/back plates and install

P/N 105-0200 rivets, using Cleveland’s rivet set, P/N 199-1, or appropriate riveting tool.

E. Check to be sure lining is tight and movement free with no distortion of parts.

F. With tubular rivets, splits may result from the clinching operation. Refer to rivet sketch (figure 6-4)

for acceptance criteria.

REASSEMBLY OF ORGANIC LININGS TO CALIPER

A. Carefully wipe dirt, grease, etc. from cylinder, pressure plate, and portions of piston extending

beyond cylinder face, and push piston back into cylinder.

B. Slide pressure plate with new lining over anchor bolts and install brake caliper into torque plate.

For equipment that is operated in an amphibious environment, or in extremely wet climates,

lubricate the anchor bolt with Lubriplate. For equipment used in a non-amphibious environment,

or in extremely wet climates, lubricate the anchor bolt with a dry film lubricant (silicon spray). DO

NOT USE GREASE OR OIL. These materials will attract dirt enhance the wear of the anchor

pins.

C. Install back plate attachment bolts and washers in brake caliper.

D. Install insulator shims (typically used with metallic lining) and spacers as applicable.

E. Slide back plates between brake disc and wheel/tire and install back plate attachment bolts and

washers into back plates.

F. Torque brake assembly back plate tie bolts to 60 inch/pounds. Two different types of back plate

tie bolts are used. The patch lock bolt (nylon material embedded in threaded end) will required

replacement 6 to 8 installations or whenever the bolts can be run in past the locking feature by

use of fingers only. Bolts with drilled heads require safety wire after torquing.

BRAKE LINING CONDITIONING PROCEDURES

When new linings have been installed, it is important to condition them properly to obtain the service life

designated into them.

Effective: 1/26/05 6-12


T34 THRUSH NONASBESTOS ORGANIC LININGS

A. Taxi aircraft for 1500 feet with engine at 1700 rpm applying brake pedal force as needed to

develop a 5-10 mph taxi speed.

B. Allow the brakes to cool for 10 – 15 minutes.

C. Apply brakes and check for restraint at high static throttle. If brakes hold, conditioning is

complete.

D. If brakes cannot hold aircraft during static run-up, allow brakes to completely cool and repeat

steps A through C.

*** WARNING ***

Due to the efficiency of these brakes, extremely hard braking on aircraft

with tail wheels could result in lifting the tail from the ground, creating a

nose over condition.

*** WARNING ***

Use extreme caution to prevent aircraft from nosing over when running

engine at high static throttle (power lever) settings, hopper and fuel

system should be full to help keep aircraft from nosing over.

This conditioning procedure will wear off high spots and generate sufficient heat to create a thin layer of

glazed material at the lining friction surface. Normal brake usage should generate enough heat to

maintain the glaze throughout the life of the lining.

Properly conditioned linings will provide many hours of maintenance free service. A visual inspection of

the brake disc will indicate the lining condition. A smooth surface, one without grooves, indicates the

linings are properly glazed. If the disc is rough (grooved), the linings must be reglazed. The conditioning

procedure should be performed whenever the rough disc condition is observed. Light use, such as in

taxiing, will cause the glaze to be worn rapidly.

BRAKE REMOVAL AND DISASSEMBLY (CLEVELAND DISC TYPE)

A. Release parking brake.

B. Jack aircraft as outlined in Section II.

C. Disconnect and cap brake hydraulic supply line at brake housing.

D. Remove back plate assemblies from calipers.

6-13 Effective: 1/26/05


E. Remove caliper assemblies.

F. Remove pressure plate assembly.

G. Clean all metal surfaces with denatured alcohol and dry thoroughly. All “0” rings are to be replace.

Remove pistons by injecting air into the caliper ports (15 to 20 psi) maximum pressure.

Use caution when blowing pistons out the caliper cylinders

with air, as pistons could fly out at high velocity. It is

suggested that the caliper be turned over so that pistons

face table working surface. Use a rag to cushion piston

contact with table surface to prevent piston damage. Make

sure to wear all applicable personal protective safety

equipment.

The brake caliper pistons on the T34 Turbo Thrush are

equipped with a friction spring (drag ring) on the piston tail.

It is recommended that this ring NOT be removed unless it

is damaged or corroded.

* NOTE *

*** WARNING ***

H. Inspect brake cylinders for cracks, nicks, corrosion and damaged threads. Inspect inlet and outlet

hydraulic ports for foreign contaminates. Examine cylinder walls for scoring or excessive wear.

Blend and polish light scratches in piston cavities with fine emery cloth, 600 grit. Castings that

are cracked or have damaged threads should be replaced.

I. Inspect anchor bolts for cracks, corrosion, permanent set and excessive wear. Replace bolts that

are bent, cracked or severely corroded.

J. Inspect pistons for cracks, nicks, burrs, or excessive wear. Remove burrs and blend out nicks,

using fine emery cloth 600 grit, and clean thoroughly.

K. Inspect pressure plate assembly for cracks, damaged pins and excessive warped contours.

Replace pressure plate if cracked or severely deformed. Replace cracked or deformed pins.

L. Inspect brake cylinder bolts for cracks, damaged threads, and self-locking feature. Replace bolts

that are cracked, bent or have damaged threads.

M. Inspect brake linings for cracks, edge chipping, and surface deterioration. Linings should be

replaced when worn to a thickness of 0.100 inch.

Effective: 1/26/05 6-14


*NOTE*

Clean outside of pistons before inserting

pistons into brake caliper housing assembly.

N. Inspect torque plate for cracks, nicks, burrs, rust, excessive wear and brinelling in bolt holes.

Replace torque plate if cracked or severely deformed.

O. Clean repaired surfaces and areas of the brake assembly from which paint has been removed.

P. Paint exposed areas with one coat of zinc primer and one coat of aluminum lacquer.

** CAUTION **

Do not paint pistons or piston bores in the brake

housing. Keep paint off of brake linings.

Q. Check the wheel brake disc. See procedures under Inspection of Main Wheel Assembly.

BRAKE REASSEMBLY AND INSTALLATION (CLEVELAND DISC TYPE)

A. Install friction spring on piston assembly (if removed).

B. Lubricate large O-ring with MIL-H-5606 hydraulic fluid and install in groove in brake housing bore

area.

C. Install piston assembly in brake housing.

D. Install pressure plate assembly on anchor bolts.

E. Install brake assembly to torque plate.

F. Install back plate assemblies with bolts and washers. Torque bolts to 60 inch-pounds.

REMOVAL OF BRAKE MASTER CYLINDERS

A. Disconnect and cap hydraulic lines.

B. Remove master cylinder retaining bolts.

C. Remove master cylinder.

DISASSEMBLY AND REPAIR

Repair is limited to replacement of parts, cleaning and adjustment. Use clean hydraulic fluid MIL-H-5606 as

a lubricant during re-assembly of the cylinders (see Fig. 6-5).

6-15 Effective: 1/26/05


INSTALLATION

To install the brake master cylinders, reverse the removal procedures and fill and bleed brakes as outlined

in this section.

BRAKE BLEEDING

To bleed the brakes proceed as follows:

A. Place parking brake control in OFF position.

Keep master cylinder reservoir full of the proper type

fluid throughout bleeding operation.

* NOTE *

B. Prepare a piece of 5/32” clear plastic (preferred) or rubber (any color) tubing at least 12 inches long.

Remove bleeder screw dust cap. Install one end of hose onto bleeder screw.

C. Place free end of hose in a clean glass receptacle containing enough hydraulic fluid to cover end of

hose. End of bleeder hose must be submerged at all times to properly check for air bubbles and

prevent entry of air into hydraulic system.

D. Apply brake pressure and open bleeder screw approximately 1/3 to ½ turn, close bleeder screw

before releasing brake pressure to avoid reentry of air into brake system. Repeat this procedure

until system is free of air.

E. Tighten bleeder screw, remove rubber hose and replace dust cap.

F. Repeat bleeding procedure for opposite brake.

Effective: 1/26/05 6-16


THIS

PAGE

INTENTIONALLY

LEFT

BLANK

Effective: 1/26/05 6-17


TROUBLESHOOTING CHART


Worn or loose wheel bearings.

Jack tail, remove wheel and inspect bearings. Replace with new lubricated bearing if necessary. Tail wheel shimmy.

Tire imbalance. Jack tail and remove tire for balance check. Rebalance.

Incorrect tire pressure. Pressure check tire. Inflate to recommended pressure.

Excessive/uneven tire wear.

Tail gear sub-assembly bearings worn or loose.

Jack tail; remove tail gear sub-assembly. Repair or replace as required.

Lock cable out of adjustment or broken. Adjust or replace as required.

Tail wheel fails to lock or unlock.

Lock pin or lock pin spring broken or damaged. Repair or replace as required.

Tire imbalance. Jack aircraft and remove tire for balance check. Rebalance.

Main landing gear shimmy.

Worn or loose wheel bearings.

Jack aircraft and check wheels for end play. Replace with new lubricated bearings if necessary.

Dragging brakes. Parking brake valve holding. Check parking brake valve. Release parking brake valve.



Effective: 1/26/05 6-18



Restriction in hydraulic lines or restriction in parking brake valve.

Have someone apply and then release brakes. Wheel should rotate freely as soon as brake is released. If wheel fails to rotate freely, loosen brake line at brake housing to relieve any pressure trapped in line. If wheel now turns freely, the brake line is restricted. Drain all brake lines and clear the inside of brake line. If cleaning the lines fails to give satisfactory results, the parking brake valve may be faulty and should be repaired.

Worn, scored or warped brake disc (see Fig. 6-2 and 6-3).

Visually check disc. Replace brake disc and lining if required.

Damage or accumulated dirt restricting free movement of wheel brake parts.

Check parts for freedom of movement. Clean and repair or replace parts as necessary.

Dragging brakes. (Cont.)

Leak in system.

Check entire hydraulic system for leaks. If hydraulic reservoir, parking brake valve, or wheel brake assemblies are leaking, they must be repaired or replaced.

Air in system. Bleed system.

Lack of fluid in brakes. Check hydraulic reservoir fluid level. Fill and bleed if necessary.

Brakes are spongy or fail to operate.

Brake assemblies’ defective. Repair or replace as required.

6-19 Effective: 1/26/05


Effective: 1/26/05 6-20

Figure 6-1


Figure 6-1a

Effective: 1/26/05 6-21


Figure 6-2

Effective: 1/26/05 6-22


Figure 6-3

6-23 Effective: 1/26/05


Rivet Acceptance Criteria

1 The split shall not occur inside the crest of the clenched surface.

2 No more than two splits shall occur in a 90° area.

3 A total of no more than three splits shall be allowed.

Figure 6-4

Effective: 1/26/05 6-24


Figure 6-5

6-25 Effective: 1/26/05


Section 7 FLIGHT CONTROLS

TABLE OF CONTENTS FLIGHT CONTROLS ....................................................................................................................................... 2

GENERAL DESCRIPTION........................................................................................................................... 2 MAINTENANCE OF FLIGHT CONTROLS .................................................................................................. 2 FLIGHT CONTROL SYSTEMS.................................................................................................................... 3

CONTROL STICK..................................................................................................................................... 3 REMOVAL OF BEARINGS FROM CONTROL STICK, FORK AND TORQUE TUBE ............................. 3 BEARING INSTALLATION IN THE CONTROL STICK FORK AND ON THE TORQUE TUBE............... 3 AILERONS................................................................................................................................................ 4 AILERON REMOVAL ................................................................................................................................ 4 AILERON INSTALLATION ........................................................................................................................ 4 AILERON RIGGING .................................................................................................................................. 4 AILERON TRIM TABS............................................................................................................................... 5 WING FLAPS............................................................................................................................................ 5 FLAP JACKSCREW REMOVAL ............................................................................................................... 5 FLAP REMOVAL ....................................................................................................................................... 6 FLAP INSTALLATION ............................................................................................................................... 6 FLAP RIGGING ......................................................................................................................................... 6 RUDDER................................................................................................................................................... 7 RUDDER REMOVAL................................................................................................................................. 7 RUDDER INSTALLATION......................................................................................................................... 7 RUDDER PEDAL REMOVAL.................................................................................................................... 7 CONTROL CABLES REMOVAL ............................................................................................................... 7 CONTROL CABLES INSTALLATION ....................................................................................................... 8 RUDDER RIGGING................................................................................................................................... 8 RUDDER TRIM TAB ................................................................................................................................. 8 BALANCE CABLE RIGGING .................................................................................................................... 8 ELEVATORS ............................................................................................................................................ 9 ELEVATOR BALANCE SPRING............................................................................................................... 9 ELEVATOR REMOVAL............................................................................................................................. 9 ELEVATOR INSTALLATION..................................................................................................................... 9 ELEVATOR RIGGING............................................................................................................................. 10 ELEVATOR TRIM TABS ......................................................................................................................... 10 ELEVATOR TRIM TAB REMOVAL......................................................................................................... 10 ELEVATOR TRIM TAB RIGGING........................................................................................................... 11 EMPENNAGE ......................................................................................................................................... 11 EMPENNAGE REMOVAL ....................................................................................................................... 11 EMPENNAGE INSTALLATION............................................................................................................... 12 WING REMOVAL .................................................................................................................................... 13 WING INSTALLATION ............................................................................................................................ 14

TROUBLESHOOTING CHART.................................................................................................................. 19 AILERON SYSTEM ................................................................................................................................ 19 FLAP SYSTEM ....................................................................................................................................... 20 RUDDER SYSTEM................................................................................................................................. 21 ELEVATOR CONTROL SYSTEM........................................................................................................... 23 ELEVATOR TRIM CONTROL SYSTEM ................................................................................................ 24

Effective: 1/26/05 7 - 1


FLIGHT CONTROLS

SECTION SEVEN GENERAL DESCRIPTION

The aircraft is equipped with flight control surfaces consisting of ailerons, elevators, rudder, wing flaps,

elevator tabs, rudder trim tabs and aileron trim tabs. The ailerons and flaps are an all-metal construction.

The empennage is of an all-metal construction consisting of horizontal stabilizer, vertical stabilizer, rudder

and elevators. Control of the ailerons, elevators and rudder are provided through a control stick and rudder

pedals. A switch located on the back of the throttle quadrant controls the electrically actuated flaps. A lever

located on the left side of the cockpit manually controls the elevator trim tabs. Fixed, ground adjustable trim

tabs are located on the rudder and both ailerons. The control stick and rudder-brake pedals are

mechanically interconnected to the push tubes, push rods, bellcranks, cables and torque tube which actuate

the primary flight controls. Control cable pulley brackets are provided with guards to prevent the cable from

jumping the pulley groove. The all-metal, electrically actuated wing flaps provide additional lift for shorter

takeoff distances and slower landing speeds. Wing flaps may be positioned at any setting between up and

down by intermittent operation of the flap switch.

MAINTENANCE OF FLIGHT CONTROLS

Special care must be exercised when performing control system maintenance. Emphasis shall be given to

security of attachments, correct alignment of rod ends, use of correct hardware, and proper safetying of

materials. Control cables must be free of kinks and pulleys must be aligned with the cables. Position cable

pulleys and route cables to avoid contact with the aircraft structure. Inspect work areas for mislaid tools or

parts with could foul the controls, and perform a functional check of the controls prior to replacement of

access covers. It is recommended that a test flight be accomplished before the aircraft is released for

routine operation when a control system component has been replaced or aircraft rigging has been altered.

Re-rigging the control systems will seldom be necessary if correct maintenance technique is employed

when system components are removed and replaced. Do not disturb position of rod end fittings when

control system components are removed, unless absolutely necessary. When deemed necessary, record

the amount of change required. This to return the fittings to original their original position when the

maintenance or repair action is complete. When control system components are being removed, carefully

note location and position of attaching parts and hardware and return to original location or position when

installing new components and parts. Rigging instructions are provided in succeeding paragraphs for the

empennage and each flight control system. Read these instructions carefully before starting the rigging

operation. Select and accomplish only those rigging steps applicable to the job

Effective: 1/26/05 7 - 2


requirement. The following procedures should be followed when rigging control cables.

Rigging should be accomplished in a hangar. When necessary to rig aircraft in the open, it should be

accomplished during coolest part of the day with tail of the aircraft pointing toward sun. If aircraft is moved

into a hangar for rigging, allow 90 minutes for control cables to adjust to hangar temperature.

The ailerons, elevators, and rudder are all balanced control surfaces and their static balance must be

checked in accordance with the limits show in table 7-19 after repaint or repair.

*** WARNING *** Failure to stay within control surfaces static balance limits could lead to control surface flutter, which could lead to loss of aircraft, life and/or property.

FLIGHT CONTROL SYSTEMS CONTROL STICK

The control stick, which is supported by bushings and bearings, is attached to the torque tube located on top

of the cockpit floor. A series of push tubes, push rods and bellcranks from solid connections between the

control stick and the ailerons. The control stick activates the elevators through push tubes, bellcrank, idler

and elevator horns. The control stick forks and torque tube may be removed for replacement of bearings

and bushings.

REMOVAL OF BEARINGS FROM CONTROL STICK, FORK AND TORQUE TUBE

Remove the control stick dust cover base assembly and the side skins under the cockpit. Use the following

procedures to replace the bearings.

A. Disconnect the aileron push rods and elevator push tube.

B. Remove the attaching hardware securing the control stick fork to the torque tube.

C. Withdraw controls stick and fork from aircraft.

D. Remove the bolts securing the torque tube to the pillow blocks. (See Figure 7-1)

E. Remove the torque-tube from aircraft.

INSTALLATION OF BEARINGS IN THE CONTROL STICK FORK AND ON THE TORQUE TUBE

A. Install bearings on the torque tube and in the control stick forks as required.

B. Install torque-tube in the pillow blocks.

C. Tighten pillow block hardware per torque values in Section II.

D. Install control sticks fork on torque tube and tighten hardware per torque values in Section II.

E. Check freedom of movement on control stick and torque tube.

F. Lubricate bearings per Section II of this manual.

G. Connect the elevator push-tube and aileron push-rods and check for proper operation of

control system.

7 – 3 Effective: 1/26/05


H. Replace control stick dust cover base assembly and aircraft side skins.

AILERONS

An all-metal aileron is installed outboard of each wing flap. Each aileron operates on bearing hinges and is attached to the aft wing spar at three points. Two balance weight is installed in the outboard leading edge of each aileron to prevent flutter.

The aileron control is driven by a single push rod from the control stick torque tube to a vertical bellcrank at the right side of the fuselage. A short push rod connects the bellcrank to a vertical idler in the left side of the fuselage. In each wing, the inboard push tubes connect between the bellcrank and idler in the fuselage to the aft side of a bellcrank near the inboard end of the aileron. From the forward side of this bellcrank, the outboard push tube connects to the forward arm of the drive bellcrank located at the aileron mid span. The short arm of the drive bellcrank is connected to a push rod that drives the aileron. The ailerons are also connected to the rudder controls by spring-loaded cables that enable the ailerons to be activated in conjunction with the rudder. This provides a safety factor. In case the aileron system becomes inoperative, the rudder system will lift the aileron. (Fig. 7-2)

AILERON REMOVAL A. Disconnect push rod at aileron. Do not change position of rod end on push rod. B. Remove aileron hinge bolts. C. Remove aileron from aircraft.

AILERON INSTALLATION Installation of the aileron is the reverse of the removal procedure. In the event push rod length has been altered, streamline trailing edge of opposite aileron with trailing edge of wing and flap and secure with a temporary lock. Adjust push rod length to align attaching bolt hole with hole in aileron hinge fitting, when aileron is in neutral position. Recheck aileron rigging.

AILERON RIGGING

Assure the ailerons are attached and the system push tubes are assembled, except for the two lateral push rods in the fuselage. Ensure that flaps have been rigged. Rig the ailerons as follows:

A. Clamp the ailerons at the trailing edge of the wing tip in the neutral position. Ailerons are in neutral

when ailerons are 1/8" below flap trailing edge.

B. Adjust the length of the push rod from aileron outboard wing bellcrank until inboard wing bellcrank is

perpendicular to the rear spar, both sides. This can be checked through the inspection holes just

forward of the rear spar.

Effective: 1/26/05 7 - 4


C. Attach the lower, lateral fuselage push rod between the left fuselage idler bellcrank and the right

fuselage bellcrank, adjusting the length of the rod to fit those items.

D. Install the upper, lateral fuselage push rod from right fuselage bellcrank to control stick torque-tube

fitting. Adjust the length of this push rod to center the control stick.

E. Clamp control stick in center position and free clamps on aileron. Set trailing edges of ailerons

0.125-inch below trailing edge of flap trailing edge by lengthening push rod from aileron to outboard

wing bellcrank.

F. Adjust and lock the aileron stops, accessible through the inspection holes forward of the aileron, for

the required travel. Aileron up travel should be 21 (±1) degrees and down travel should be 17 (±1)

degrees.

G. Go back through system and lock all check nuts.

H. To adjust the springs in the rudder-aileron interconnect system, clamp the rudder and ailerons in the

neutral position and adjust the turnbuckles until the springs are the same length.

AILERON SERVO TRIM TABS

A variable position trim tab is attached to each aileron. A wing high attitude of either wing may be corrected

by adjusting the applicable trim tab down. Adjusting the tab up will correct a low wing attitude. Begin with

both tabs in neutral position (straight with trailing edge of aileron).

A. Electric Aileron Trim Tabs. Thrush Aircraft Inc has made available an electric aileron trim tab

normally installed on left aileron. Initially, the tab is rigged the same as servo-trim tabs, then the

tab can be electrically adjusted to obtain level flight. Compensating for the wind and in normal

turns, the tab will act as a servo-tab.

WING FLAPS

Wing flaps installed on the S2R-T34 are of an all-metal construction and hinged on ball bearings. Each flap

extends outboard from the fuselage to the aileron and is attached to the aft wing spar by four (4) stainless

steel hinges. A switch located on the aft of the throttle quadrant electrically controls the flap operation.

Movement of the flaps is by a torque tube located below the cockpit floor and rotated by an electric motor-

driven jackscrew. Push rods attached to the arms of the torque tube move the flaps to the desired position.

The flaps have been completely sealed against chemical spray.

FLAP JACKSCREW REMOVAL

A. Remove left side fuselage cowling under cockpit door and hopper.

B. Remove wing root fairings to gain access to the flap push road and jackscrew attach bolts.

Disconnect each flap from the push rod and allow flap to swing and hang under wing.

C. Disconnect the electrical connections, connecting the micro-switches and motor. Identify the wires

and locations for installation reference.

7 – 5 Effective: 1/26/05


D. Remove one attach bolt at the motor end, loosen the other bolt slightly and remove the four bolts

connecting the jackscrew to the flap torque tube.

E. Install new flap jackscrew assembly into fuselage and connect with hardware which was removed or

new hardware. NOTE: Rigging must be checked after installation of new jackscrew or pushrods.

Refer to flap rigging.

F. Reconnect the electrical wires and test flap motor for proper operation.

G. Connect pushrods and after flaps are rigged properly ensure that all bolts are tight and wires are tied

off.

H. Reinstall fairings and cowling.

FLAP REMOVAL

A. Disconnect flap push rod at flap. Do not change position of rod end on push rod. (See Figure 7-3)

B. Remove flap hinge bolts.

C. Remove flap from aircraft.

FLAP INSTALLATION

Installation of the flap is the reverse of the removal procedure. In the event push rod length has been

altered, the flap will have to be completely re-rigged.

FLAP RIGGING A. With the master switch “ON,” fully retract the flaps (up flaps) with the flap switch.

B. Disconnect the flap push – pull rods at the torque tube arms.

C. Hold a straight edge on the wing lower surface at wing station 49.0 (approximately 24 inches

outboard of the fuselage side). In the properly rigged flap “up” position, the straight edge should

contact the lower surface of the wing, front spar, the flap trailing edge and the lower surfaces.

D. Adjust the flap push-pull rods to the proper length and connect them to the torque tube arms.

E. Shorten the length of the maximum travel limit bolt located on the right side of the fuselage adjacent

to the torque tube.

F. With the flap switch, lower the flaps to the fully extended position.

G. Using a propeller protractor or equivalent instrument to measure the flap angular travel, adjust the

down micro-switch located on the vertical shaft adjacent to the jackscrew to achieve 15 (+/-1)

degrees. Retract and extend the flaps after each adjustment to verify proper adjustment.

H. With the flaps in the fully extended position, adjust the maximum travel stop bolt so that there is a

0.060” to 0.100” gap between the bolt head and the stop pad.

I. Retract the flaps with flap switch and turn the master switch “OFF.”

Effective: 1/26/05 7 - 6


J. Tighten and torque all hardware to the specifications called out in Section II of this manual.

RUDDER

The metal-covered rudder is attached to the vertical stabilizer at three hinge points. The rudder control

cable is connected directly from the rudder horn to the rudder pedal adjustment channels. The left and right

rudder cables route from the adjustment channels aft around pulleys where they pass through the fuselage

side skins and attach to adjustment straps on the rudder horn. A spring-loaded balance cable is routed

between the pedal adjusting channels and forward around pivoted

pulleys located on the hopper rear wall. The rudder controls are interconnected by springs to the aileron

system so that a wing may be lifted with rudder alone. This feature provides a convenience during cross-

country flight and is an added safety feature in case the aileron system becomes inoperative. (Fig. 7-6)

RUDDER REMOVAL A. Disconnect rudder cables from rudder horn.

B. Remove attaching hardware from rudder hinge points.

C. Remove the rudder from the aircraft.

RUDDER INSTALLATION A. Place rudder on hinge points.

B. Install the hardware in the hinge.

C. Attach rudder cables to rudder horn.

D. Check rudder operation to determine that no friction or binding is evident.

E. Readjust control cables and rudder stops as required per rigging instructions.

RUDDER PEDAL REMOVAL

Use Figure 7-7 as a guide when removing or installing rudder pedals.

CONTROL CABLES REMOVAL

A. Disconnect the aft cables from forward side of shackles.

B. Remove skins from side of fuselage.

C. Disconnect cables at turnbuckles.

D. Remove all cable guards from the rudder cable pulleys and disconnect the aft cables from rudder

horn. The cables from the turnbuckles aft are free for removal.

E. Disconnect cables from adjustment channel. The cables from the turnbuckles forward to

the pedals are free for removal.

7 – 7 Effective: 1/26/05


F. Remove the balance cables.

G. Remove rudder-aileron cable from aileron vertical bellcrank.

CONTROL CABLES INSTALLATION (Fig. 7-6)

A. Install the cables in reverse order of the removal procedures.

B. Check rigging per rigging instructions.

C. Assure all cables, cable guards and turnbuckles are installed properly and safety wired. Replace all

skins removed for access.

RUDDER RIGGING

A. Position the rudder pedals at mid-adjustment position in the adjustment channel.

B. Center and lock the rudder.

C. Adjust the turnbuckles in each rudder cable, at fuselage station 175.12, to bring the rudder pedals

approximately 11.00 inches from the back of the hopper.

D. Safety-wire the turnbuckles with 0.041 stainless steel wire.

E. Adjust and lock the rudder stop bolts, located at the base of the rudder post, to limit the travel of the

rudder to 24 (±1) degrees left and right of center. (Fig. 7-8)

RUDDER TRIM TAB (Fig. 7-10)

A fixed-position trim tab is attached to the lower edge of the rudder. An out-of-trim rudder can be trimmed

by bending the metal trim tab. Use forming blocks when bending tab and do not bend more than 0.50 inch

deflection in either direction. BALANCE CABLE RIGGING (Fig. 7-11)

The Thrush incorporates a rudder-aileron balance cable/spring system. The cables are attached to the

rudder pedals and routed out of the cockpit and to the spring which is attached to the opposite aileron

bellcrank. Adjustments are accomplished with the turnbuckles located on each cable. The forward pulley

retaining bracket should be located on the tube 62.37 fuselage station at L/H 6 1/4 ±1 R/H 5 ±1 inches up

Effective: 1/26/05 7 - 8

* NOTE * If the turnbuckles run out of adjustment, relocate adjustment strap

on rudder horn to shorten cable.


The system is correctly adjusted when the rudder and ailerons simultaneously align in the neutral position.

Ensure there is no contact between balance springs.

ELEVATORS

Each elevator is attached to the rear spar of the horizontal stabilizer at three hinge points. The control stick

is connected to the elevators through the use of a bellcrank, idler, push tubes and

elevator horn. The right and left elevators are attached to a common elevator horn. (Fig. 7-12) ELEVATOR BALANCE SPRING

The elevator system has a balance spring attached to the forward elevator bellcrank with a cable on top

and spring on the other end attached to bellcrank support bracket. The cable and spring are connected with

a turnbuckle for final adjustments. With the flaps up and the elevator in neutral and the forward stand

assembly is clamped on flap transfer tube at 48° from center on fwd side. Rig the cable to obtain a spring

length of 39". The dual cockpit elevator balance spring is connected to the lower portion of the forward

elevator bellcrank and to the left lower longeron fuselage station 193.43 inches with a turnbuckle for

adjustment. With flaps up and elevator neutral, rig spring to 36 inches in length.

ELEVATOR REMOVAL A. Disconnect aft push tube from elevator horns.

B. Disconnect the trim tab push rod at elevator trim-tab.

C. Remove hardware attaching both of the elevators horns together.

D. Remove all hinge bolts from leading edge of elevator.

E. Remove the elevator from aircraft.

ELEVATOR INSTALLATION

A. Position elevator on aircraft.

B. Attach the elevator to the horizontal stabilizer.

C. Bolt the two elevator horns together.

D. Connect aft push tube to elevator horns.

E. Connect trim-tab push rod at elevator trim.

7 – 9 Effective: 1/26/05


ELEVATOR RIGGING

The aft push tube will have to be disconnected from elevator horns for adjustment.

A. Set the forward stop on the control stick so the stick is approximately seven inches from the hopper

when in full forward position.

B. Set the elevator to its full down travel of 17 (±1) degrees and adjust the aft end of push tube at

the elevator horn to match that position. Connect push tube to elevator horn.

C. Set the elevator at the full UP travel of 27 (±1) degrees and adjust the aft stop on the control stick to

match this position.

D. Tighten the lock nut against the rod end bearing at the elevator horn.

ELEVATOR TRIM TABS

Controllable trim tabs, located on the inboard trailing edge of each elevator, are operated by an elevator trim

tab control lever located on the left side of the cockpit. Linkage between the elevator trim tab control lever

and the elevators consists of push rods, bellcranks and fairleads. The push rod leading from the trim tab

control lever to the trim tab assembly runs along the left side of the fuselage and is guided at intervals by

four fairleads. The aft end of this push rod attaches to a bellcrank. This bellcrank has arms at each end

permitting two short push rods to be routed back to bellcranks located on the inboard side of the horizontal

stabilizer. A short push rod leads from these bellcranks to horns on the trim tabs. (Fig. 7-13)

ELEVATOR TRIM TAB REMOVAL

A. Disconnect push rod from trim tab.

B. Remove rivets attaching trim tab hinge to elevator and remove trim tab.

Effective: 1/26/05 7 - 10

** CAUTION **

Rigging of the elevators will require that tail gear be checked for

proper operation. See Section VI, Landing Gear.

*NOTE *

Assure the inspection hole in the rod end is covered by the

push tube threads. It may be necessary to let the control

stick come back slightly to achieve coverage.


ELEVATOR TRIM TAB RIGGING

A. Place the trim tab control lever in cockpit in the neutral position.

B. Adjust push rods to position both the aft fuselage and stabilizer bellcranks in a center (neutral)

position.

C. Place elevator in neutral position. Adjust the length of push rod, between the two bellcrank

assemblies.

D. Tighten all bolts.

E. Loosen bolts attaching trim control lever stop and adjust the stop to provide proper trim tab travel.

The trim tab travel should be 8 (±1) degrees up and 22 (±1) degrees down.

F. Measure free-play of the tab at the trim tab horn attaching point. The total maximum free play

should not exceed 0.125-inch.

EMPENNAGE

The vertical stabilizer, rudder, horizontal stabilizer and elevators are constructed of Alclad aluminum. All

stabilizers are connected to the fuselage structure by bolts and supported by adjustable struts. Rudder and

elevators are attached to the stabilizers by hinges containing sealed bearings. See Section VII, Flight

Controls, for rigging instructions.

EMPENNAGE REMOVAL

A. Remove rudder and elevators from stabilizer as outlined in this section of the manual.

B. Remove horizontal struts, being sure to mark left and right. (The top of the struts can be identified

by a small rectangular section of weld line at the V-end of the aft tube.)

7 – 11 Effective: 1/26/05

*NOTE *

When measuring trim tab travel, the elevator should be in the

neutral position.

** CAUTION **

The forked ends of the struts are torqued to align with horizontal

attachments. Movement of the fork will require re-torquing as

outlined in the installation instructions.


EMPENNAGE INSTALLATION

A. Install horizontal stabilizer with AN6-46A bolts on forward attaches with one each bushing (P/N

9040-018) (9/16” long) between horizontal stabilizer and forward fuselage attach fitting. Place one

each bushing (P/N 9040-108) (1 ¼” long) between the aft stabilizer attach fitting and the fuselage

fitting and install AN6-44A bolts. Torque bolts to 408 ± 15 inch-pounds of torque. Check that the

cord line of the horizontal stabilizer is -0.50º ± .25º (this means nose down) relative to the leveling

longeron used for weight and balance (under the cockpit) (see figure 7-18). Use up to 3 ea. AN960-

616 washers under either the fwd or aft bushings to achieve the required incidence angle.

B. Install left and right struts using the strut/plates and AN5-6A bolts on lower attach (strut to fuselage)

and AN6-12A bolts on upper attach (strut to horizontal). Install rudder lock plate on left lower strut

attach; tighten all bolts and nuts.

When installing a new strut or new strut parts, accomplish as follows:

With new strut(s) adjust fork ends as required to bring the strut as close as possible to lower

fuselage attach without touching. The strut should be centered fore and aft with fork ends. The

forks should be shimmed with a P/N 40024-3 spacer and P/N 21194 washers as required (different

thicknesses are available) and at least one P/N 21194-C copper crush washer. Then torque to at

least 100 inch-pounds and not more than 190 inch-pounds and align with attach points

simultaneously. The lower strut/plates can now be trimmed to fit if needed and drilled with a .312

(5/16") drill bit and bolted into place using AN5-6A bolts. If only re-torque is required and torque

cannot be achieved with old shims, the replacement of the copper crush washer only should be

sufficient to regain correct torque and proper angle for alignment.

C. Install left and right elevator using P/N 40065-1 spacer, AN4-12A bolt and AN4-11A bolts in center

and outboard hinges. Connect elevator control arm and check travel 27° ±1° up and 17°±1° down.

D. Connect elevator trims tabs and check for proper travel.

E. Vertical Stabilizer: Install forward attachment loosely with NAS6207-68 bolt. Using either no shims

(normal) or if a gap exist, use one or more of the following P/N 21209T001 (.125”) and/or P/N

21209T002 (.250”) shim(s) at upper attach and P/N 21208T001 (.125”) and/or P/N 21208T002

(.250”) shim(s) at lower attach. Install hardware upper and lower and navigation light ground wire

(lower attach); tighten all vertical fin hardware per torque table (fig. 2-7). Using a string pulled tight

through upper rudder hinge and lower rudder hinge, check hinges for alignment fore, aft, left and

right. It is permissible to add (1) P/N 40207T005 (.050”) or 40207T007 (.063”) between the center

hinge bearing housing and vertical fin rear spar to achieve proper alignment.

Effective: 1/26/05 7 – 12


*NOTE * A TAPERED shim(s) P/N 90220T001 (.125” to .080”)

upper attach shim or P/N 90221T001 (.100” to .075”)

lower attach shim may be required on top and/or lower

attach to property align hinges during the string alignment

check.

F. Install the wire deflector cable and allow sufficient turnbuckle travel to permit tensioning of the

deflector cable. Attach and tension cable to 35 ±3 Lbs.

G. Install the rudder using AN4-11A bolts, connect navigation light ground wire to rudder horn bolt, and

connect navigation light power wire. Install rudder cables and check travel 24° left and right from

center.

WINGS WING REMOVAL (Figures 7-9, 7-10, 7-16 and 7-17) A. Park the aircraft in a closed door hangar and secure the aircraft for maintenance.

B. Disconnect the batteries and external power sources.

C. Gain access to the wing splice area by removing the wing root fairings and the necessary aircraft

side and belly panels. Disconnect the electrical, fuel, spray and flight control systems at points

appropriate for wing removal. Remove the spray pump and bracket.

D. Support the weight of the wings at the jack points located approximately eight feet from the wing

tips and under various wing ribs to prevent wing movement when the attach bolts are removed.

E. Remove the bolts holding the rear spar to the fuselage, one place on each wing.

F. Remove the ¼ inch bolts securing the left and right wing, inboard and outboard attach angles to the

spar webs, 12 places on each wing.

G. Back the locknuts off of the NAS bolts in the splice fittings far enough to conceal the end of the bolt

threads. Spray the bolt shanks with WD-40 or an equivalent lubricant.

H. Remove the eight each ¼ inch bolts securing the two each U-shaped Clevis (part number 22506-

11) to the tube nuts and remove the two clevises. Loosen the lower splice fitting tube nuts. As the

nuts are loosened, the bolts will back out of the holes. Once the tube nuts are off of the lower bolts,

loosen the upper tube nuts to back them out of the holes. If one of the bolts does not move,

rethread the tube nut onto this bolt and place a spacer between the opposite end of the tube and

the spar cap (or NAS bolt). Proceed, once again, to loosen the tube nut and back the bolt out of the

hole. Remove tube nuts after obtaining sufficient clearance.

7 – 13 Effective: 1/26/05

*NOTE *

Adjusting the angle of the wings with the wing jacks may assist in

wing bolt removal.


I. Place an aluminum block on the smallest nut and with a 4X rivet gun, drive the nut flush to the

splice fitting. Proceed in a similar manner with the remainder of the bolts working from the smallest

to the largest.

J. Remove all of the nuts and tube nuts from the NAS bolts. Place the aluminum block on the

threaded end of the smallest bolt and with the 4X rivet gun, drive it flush with the splice fitting.

Proceed in a similar manner with the remainder of the bolts working from the smallest to the largest.

K. Using a suitable phenolic or soft metal drift, drift out the NAS bolts from the lower splice fittings

(upper and lower) and lower spar cap with a 2X rivet gun working from smallest size bolt to largest

size.

L. Pry the upper splice fitting (lower half) off the NAS bolts by tapping phenolic, hard plastic or

aluminum wedges between the upper splice fitting (lower half) and the upper spar cap. Remove the

lower half of the upper splice fitting.

M. Place the aluminum block on the threaded end of the smallest bolt in the upper splice fitting and

with a 2X rivet gun, drive it flush with the spar cap. Proceed in a similar manner with the remainder

of the bolts working from the smallest to the largest.

N. Pry the upper splice fitting (top half) off of the upper spar gap by tapping aluminum wedges between

the two. Remove the top half of the upper splice fitting along with the NAS bolts

O. Remove the bolts securing the center wing splice plate to the right wing. The splice plate will

remain attached to the left wing.

P. Slide the wings directly away from the fuselage lifting the wing roots sufficiently to clear the fuselage

lower longerons.

Q. Discard all used nuts, bolts and washers.

WING INSTALLATION (Figure 7-9, 7-10, 7-16 and 7-17) To install the wings, proceed as follows using all new nuts, tube nuts, bolts and washers:

Effective: 1/26/05 7 - 14

*** WARNING ***

Under no circumstances should the bolts be turned while the

threads are in the spar cap. This could damage the hole surface

finish which could result in a stress concentration.

* NOTE * All bolt shanks and bolt holes are to be coated with Snap-On™ General Purpose Antiseize lubricant or equivalent before installation.


A. If the landing gear is not installed, support the fuselage at a convenient height using jacks at the

landing gear attach points, left and right, and at the aft jack point.

B. On the left wing, install the outboard wing attach angle on the main spar web with NAS1104-16

bolts, ¼ inch AN960 and MS20002C washers and MS21042 nuts. Torgue the MS21042 nuts per

Table 7-3, six places.

C. On the left wing, install the center splice plate if removed, to the spar web and install the 3/16” and

¼” bolts and c/s screws, AN960 and MS20002C washers and MS21042 nuts per Figure 7-9.

Torque the MS21042 nuts per Table 7-3, 39 places.

D. Elevate the left wing and place it in position. Butt the wing attach angle squarely against the

fuselage vertical wing attach tube assembly. Locate the rear spar fitting into the fuselage attach

fitting and install the 7/8” bolt, AN960 washers and MS21044N nut. Bring nut up snug but not to

final torque.

E. Support the wing at the wing jack point on the front spar, approximately 8 feet inboard of the wing

tip.

F. Rest the inboard end of the main spar on the lower longeron with a ½ inch temporary spacer

between.

G. Install the left wing inboard attach angle on the main spar web with NAS ¼ inch bolts, AN960 and

MS20002C washers and MS21042 locknuts. Torque the locknuts per Table 7-3, 6 places.

H. Align the bolt holes in the wing attach angles with the holes in the fuselage vertical wing attach tube.

Install one AN 5/16” bolt in the top hole and secure with an AN960 washer and AN365 locknut.

Install locknut finger tight.

I. On the right wing, install the outboard wing attach angle on the main spar web with NAS ¼ inch

bolts, AN960, and MS2002C washers, and MS21042 locknuts. Torque the MS21042 nuts per

Table 7-3, 6 places.

J. Elevate the right wing and place in position with the 1/2 inch temporary spacers as was done with

the left wing. Align the holes in the right wing spar web with the holes in the center splice plate.

Install the 3/16” and ¼” bolts and nine each countersunk screws with heads forward and secure

with AN960 and MS20002C washers and MS21042 locknuts. Bring the nuts up snug and torque

only those which will be inaccessible when the splice fittings are installed.

K. Install the right wing inboard attach angle on the main spar web with NAS ¼ inch, AN960 and

MS20002C washers and MS21042 locknuts. Torque the locknuts per Table 7-3, six places.

7 – 15 Effective: 1/26/05

** CAUTION **

On all NAS bolts described hereinafter, insure that at least two bolt

threads are showing beyond the edge of the fiber locknut and that

the nut has not bottomed out on the bolt threads.


L. Lubricate the 12 NAS wing splice attach bolts and the holes in the upper splice fittings, (top and

bottom halves) with Antisieze. Position the upper plate part number 20242-2 on top of the top

splice fitting and insert the two large NAS bolts with MS20002C countersunk washers. Be sure to

mate the countersunk portion of the washer with the head of the bolt. Insert the remaining 10

NAS bolts with their respective holes and position all 12 bolts flush with the bottom of the fitting.

M. Apply Antiseize to the matching holes in the upper spar cap and place the top splice fitting into

position. Move the wings as required and press the bolts through the holes in the upper spar cap.

N. Place the upper splice fitting (bottom-half) into position and press it upward over the bolts. Install

MS20002 (no countersink) washers (12 places) and bring the 22506-7 tube nuts (two places) and

the MS21044N nuts (10 places) up snug but not to full torque.

O. Correctly position the lower plate part number 22514-1 on bottom of lower splice fitting (bottom)

and insert two large one inch NAS bolts with MS2002C countersunk washers with chamfer

towards the bolts hexagonal head. Place (upper) lower splice fitting on top of lower spar cap.

Push lower bottom splice fitting with large bolts until flush with lower wing spar and through upper

splice fitting. Install and snug tube nuts but not full torque. Install the remaining NAS bolts in the

lower splice fittings starting with the largest bolts working to the outboard.

P. Align the holes in the wing attach angles (left and right wings) with the holes in the fuselage

vertical wing attach tubes and install the remaining seven AN 5/16 inch bolts with AN960 washers

and MS21044N nuts. Bring the locknuts up snug but not to full torque.

Q. Torque all locknuts on the NAS ¼ inch bolts through the spar web and splice plates per Figure 7-

3, 39 places.

R. Torque the tube nuts and the MS21044N locknuts on the NAS bolts through the upper and lower

splice fittings per table (Figure 7-3), four places and 24 places, respectively.

Effective: 1/26/05 7 - 16

** CAUTION **

Be sure to orient the MS20002C washers such that each countersink

is mated with the radius in the underside of the NAS bolt heads.

** CAUTION **

Use extreme care to avoid damaging the spar cap and the surface

finish of the holes through the spar cap.

** CAUTION **

Use extreme care to avoid damaging the spar cap and the surface

finish of the holes through the spar cap.


S. The upper and lower tube nuts are joined together by u-shaped clevises (part number 22506-11).

The tube nut connection will require the use of Thrush Aircraft Inc tool part number ESK681-1 or

equivalent in drilling and locating the .25” diameter holes. After drilling deburr all holes, and install

clevises with eight each. NAS1104-26 bolts, MS20002C4 chamfered washers, under the head,

AN-960-416 washers, and MS21044N4 locknuts. Torque locknuts per table 7-3.

T. Remove the ½ temporary spacers from between the wing spars and the longerons, two places.

U. Torque the locknuts on the AN5 bolts through the left and right wing attach angles and the

fuselage vertical tubes per Table 2-7, eight places.

V. Torque the locknuts on the AN7 bolts through the left and right wing rear spar attach fittings per

Table 2-7, two places.

W. Release the wing jacks and check the wing dihedral for 3 ½ degrees.

X. Install the spray pump and bracket and torque all nuts and bolts.

Y. Complete the wing installation by making the necessary connections in the electrical, fuel, pitot,

and flight control systems.

Z. Reconnect the batteries.

MODIFIED WING INSTALLATION PROCEDURE FOR S2R AIRCRAFT WHICH HAVE TO HAVE THE WING ATTACH ANGLES REPLACED. A. With the wings supported, position the wing roots in the fuselage and place a ½ inch temporary

spacer between the lower longeron and the main spar cap of each wing. (This will properly set

the wings angle of incidence.)

B. Install the spar web splice plate and the upper and lower main spar cap splice fittings in

accordance with the normal installation procedure. Torque all nuts to specification.

C. Install the wing attach bolts at the left and right rear spars and torque to specification.

D. On the 20243-3 wing attach angles (outboard only), draw heavy black line from top to bottom so

as to bisect the area of the forward face when it is positioned on the aircraft. This line will be on

the outside of each angle and will mate with the back side of the main spar web.

E. Position these outboard angles so that they mate flush with the bushings through the vertical

fuselage tube, flush with the spar web and contact the top of the lower spar cap. Now raise the

angle so that there is approximately a 3/16 inch clearance between the angle and the spar cap.

Clamp the angle to the vertical fuselage member.

F. Sight through the bolt holes in the main spar web and position the wings so that black lines drawn

in step “D” are visible and that they appear symmetrical between left and right wings. This

centers the wings with respect to the fuselage. Repeat step “E” if the 3/16 inch clearance has

changed.

7 – 17 Effective: 1/26/05


G. Transfer the location of each of the holes through the vertical fuselage member to the attach

angle by sliding a 5/16 inch drill bit (preferably one that has a pilot tip), through each of the four

holes in succession. Do not drill these holes because edge distance must be checked.

H. Remove the angle and check to confirm that the mark for the bottom hole is at least 5/8 (2X hole

diameter) inch away from the edge of the angle. If it is not, file the bottom edge of the angle

which comes into contact with the lower spar cap in such a manner as to allow the angle to be

lowered and yet provide clearance with the lower spar cap. Repeat the hole transfer procedure

and recheck for edge distance.

I. Drill the four 5/16 inch holes through the outboard wing attach angles.

J. Place the inboard attach angles back to back against the outboard angles. Align and clamp the

two angles and transfer drill the 5/16 inch holes through the inboard angles.

K. Temporarily install the four wing attach angles to the fuselage down tubes and torque the 5/16

inch nuts and bolts to specification.

L. Again sight through the ¼ inch bolt holes through the main spar web and confirm that the black

lines appear symmetrical between the left and right wings.

M. Transfer the location of the ¼ inch holes to the four wing attach angles, using the main spar web

as an index. A few holes may be drilled and bolts installed to keep the spar web flush with the

forward face of the angles for transfer accuracy. Use caution not to enlarge the holes through the

spar web.

N. Remove the four attach angles and drill and ream the ¼ inch holes at the marked locations with a

drill press or milling machine (.250”/.254”). Debur all holes.

O. Reinstall the wing attach angles and torque all nuts and bolts to the specification called out in

Figure 2-7.

P. Remove the ½ inch temporary spacers between the lower longeron and the lower spar cap.

Q. Release the wing supports and check the wing dihedral for 3 ½ degrees.

R. Complete the wing installation by making the necessary connections in the electrical, fuel, pitot,

and flight control systems.

Effective: 1/26/05 7 - 18



AILERON SYSTEM

TROUBLE CAUSE REMEDY

Control stick bearings dry or

worn.

Check control stick

bearings for

lubrication, excessive

wear and cleanliness.

Torque tube bearings dry or

worn.

Check bearing for

lubrication, excessive

wear and cleanliness.

Resistance to control stick

movement.

Bent aileron. Repair or replace

aileron.

Aileron push rods and tubes

out of rig.

Rig in accordance

with aileron rigging

procedures. Incorrect aileron travel.

Aileron bellcrank stops

incorrectly

Rig in accordance


procedures.

Incorrect rigging of push rods

and tubes.

Rig in accordance


procedures. Correct aileron travel cannot

be obtained by adjusting

bellcrank stops. Incorrect rigging of

bellcranks.

Rig in accordance


procedures.

7 – 19 Effective: 1/26/05


FLAP SYSTEM


Circuit breaker out. Reset circuit breaker.

Defective flap switch. Replace flap switch.

Defective flap motor. Replace flap motor.

Defective electrical circuit. Replace defective wires.

Stripped or broken jackscrew

on flap motor.

Replace jackscrew

assembly.

Flaps do not extend (down) or

retract (up).

Defective microswitch. Replace microswitch.

Flaps fail to retract (up)

completely. Incorrect rigging of push rods.

Rig in accordance with

rigging procedures.

Flaps fail to extend (down)

completely. Incorrect rigging of push rods.


rigging procedures.

Flaps not synchronized or fail

to fit evenly when retracted

(up).

Incorrect adjustment of push

rods.

Adjust in accordance with

rigging procedures.

Effective: 1/26/05 7 - 20



Bent push rods. Straighten or replace. Flaps not synchronized or fail

to fit evenly when retracted.

Continued. Bent flap. Repair or replace flap.

Broken arm on torque tube or

broken push rod. Replace broken parts.

Flaps on one side fail to

operate.

Disconnected push rod.

Connect push rod and

recheck rigging

procedures.

RUDDER SYSTEM


Cables loose. Adjust in accordance with

rigging procedures.

Broken pulley. Replace pulley. Lost motion between rudder

pedals and rudder.

Bolts attaching rudder horn to

rudder loose. Tighten bolts.

7 – 21 Effective: 1/26/05



Cables too tight.

Adjust cables in

accordance with rigging

procedures.

Pulleys binding or rubbing.

Provide proper clearance

if pulleys are rubbing

pulley brackets or cable

guards.

Rudder binding caused by

bent rudder horn. Replace rudder horn.

Rudder pedal needs

lubrication. Lubricate as required.

Cables not in place on

pulleys. Install cables correctly.

Excessive resistance to

rudder pedal movement.

Bent rudder. Repair or replace rudder.

Rudder pedals not neutral

when rudder is streamlined.

Rudder cables incorrectly

rigged.


rigging procedures.

Incorrect rudder travel. Rudder horn stops incorrectly

adjusted.

Adjust in accordance with

rudder rigging

procedures.

Effective: 1/26/05 7 - 22


ELEVATOR CONTROL SYSTEM


Pulley binding or rubbing.

Provide proper

clearance if rubbing

pulley bracket or guard.

Binding control stick bearings.

Lubricate bearings.

Repair or replace

elevator horns.

Resistance to control stick

movement.

Elevator hinges need

lubrication.

Lubricate hinges as

required to give free

movement.

Incorrect elevator travel.

Elevator bellcrank, idler, and

push tubes incorrectly

adjusted.

Adjust in accordance

with rigging procedures.

Correct elevator travel cannot

be obtained by adjusting

bellcrank, idler and push

tubes.

Control stick stops incorrectly

rigged.

Adjust control stick

stops in accordance


7 – 23 Effective: 1/26/05


ELEVATOR TRIM CONTROL SYSTEM


Push rods binding.

Check push rods at

fairings for free

movement. Trim control lever moves with

excessive resistance.

Trim tab hinge binding.

Lubricate hinge. If

necessary replace

hinge.

Incorrect trim tab travel. Incorrect adjustment of push

rods.

Adjust in accordance


Effective: 1/26/05 7 - 24


7 – 25 Effective: 1/26/05

Figure 7-1


BOLT SIZE NUT TORQUE

Figure 7-2

Effective: 1/26/05 7 - 26


3/16 – 32 25 – 30 in. – lbs.

1/4 – 28 80 – 100 in. – lbs.

5/16 – 24 120 – 145 in. – lbs.

3/8 – 24 200 – 250 in. – lbs.

7/16 – 20 520 – 630 in. – lbs.

1/2 – 20 770 – 950 in. – lbs.

5/8 – 18 1,250 – 1,550 in. – lbs.

3/4 – 16 2,650 – 3,200 in. – lbs.

1 – 12 4,500 – 5,500 in. – lbs.

7 – 27 Effective: 1/26/05

Torque Table for NAS Fasteners Through Wing Center Splice Fittings,

Splice Plates, and Wing Attach Angles

Figure 7-3


Effective: 1/26/05 7 - 28

Figure 7-4


Figure 7- 5

7 – 29 Effective: 1/26/05


Effective: 1/26/05 7 - 30

Figure 7-6


7 – 31 Effective: 1/26/05

Figure 7-7


Effective: 1/26/05 7 - 32

Figure 7-8


Figure 7-9

7 – 33 Effective: 1/26/05


Effective: 1/26/05 7 - 34

Figure 7-10


7 – 35 Effective: 1/26/05

Figure 7-11


Effective: 1/26/05 7 - 36

Figure 7-12


7 – 37 Effective: 1/26/05

Figure 7-13


Effective: 1/26/05 7 - 38


7 – 39 Effective: 1/26/05

Figure 7-15


Effective: 1/26/05 7 - 40

Figure 7-16


\

7 – 41 Effective: 1/26/05

Figure 7-17


Figure 7-18

Effective: 1/26/05 7 - 42


FLIGHT CONTROL STATIC BALANCE LIMITS After repaint or repair of balanced control surfaces, they must be checked for proper balance using the limits found in the charts below. Stripping of paint and repainting may be necessary if control fails to be within limits.

*** WARNING *** Failure to stay within control surfaces static balance limits could lead to control surface flutter, which could lead to loss of aircraft, life, and/or property.

“LOW SPEED” RUDDER ASSY P/N 40226T100: INCH-POUNDS OF IMBALANCE FROM HINGE LINE, TRAILING EDGE HEAVY.

INCH POUNDS

CONDITION MINIMUM MAXIMUM

MANUFACTURING 45 100

FIELD REPAIR 45 110 “LOW SPEED” ELEVATOR ASSY P/N 40058T505 “L/H” or T506 “R/H”: INCH-POUNDS OF IMBALANCE FROM HINGE LINE, TRAILING EDGE HEAVY.

INCH POUNDS


MANUFACTURING 40 53

FIELD REPAIR 40 57 “LOW SPEED” AILERON ASSY P/N 52081T061 “L/H or “R/H”: INCH-POUNDS OF IMBALANCE FROM HINGE LINE, TRAILING EDGE HEAVY (AILERON INVERTED “FLAT SIDE FACING UP”).

INCH POUNDS


FIELD REPAIR NO

MINIMUM VALUE

41.4

TABLE 7-19

7 – 43 Effective: 1/26/05


Section 8 INSTRUMENTS

TABLE OF CONTENTS INSTRUMENTS ............................................................................................................................................... 2

GENERAL DESCRIPTION ....................................................................................................................... 2 INSTRUMENT SYSTEM MAINTENANCE ............................................................................................... 2 FLIGHT INSTRUMENTS.......................................................................................................................... 3 PITOT-STATIC SYSTEM .......................................................................................................................... 3 MAINTENANCE........................................................................................................................................ 3 INSPECTION AND LEAKAGE TEST........................................................................................................ 3 ALTIMETER............................................................................................................................................... 5 AIRSPEED INDICATOR............................................................................................................................ 5 MAGNETIC COMPASS............................................................................................................................. 5 MAGNETIC COMPASS COMPENSATION .............................................................................................. 5 BANK INDICATOR .................................................................................................................................... 6 POWER PLANT INSTRUMENTS ............................................................................................................. 6 MISCELLANEOUS INSTRUMENTS ........................................................................................................ 6 FUEL QUANTITY INDICATOR ................................................................................................................. 6 VOLTMETER............................................................................................................................................. 6 AMMETER................................................................................................................................................. 7 HOPPER QUANTITY ................................................................................................................................ 7 CALIBRATION OF REMOTE GAUGE ...................................................................................................... 7 TROUBLESHOOTING CHART ................................................................................................................ 9 AIRSPEED INDICATOR............................................................................................................................ 9 ALTIMETER............................................................................................................................................. 10 MAGNETIC COMPASS........................................................................................................................... 11 ENGINE OIL PRESSURE GAUGE ......................................................................................................... 12 ENGINE FUEL PRESSURE GAUGE...................................................................................................... 14 TACHOMETER........................................................................................................................................ 15 FUEL QUANTITY INDICATOR ............................................................................................................... 15 ENGINE OIL TEMPERATURE GAUGE.................................................................................................. 17 HOPPER QUANTITY SYSTEM .............................................................................................................. 18 INSTRUMENT MARKINGS..................................................................................................................... 19

Effective: 1/26/05 8 - 1


INSTRUMENTS

SECTION EIGHT GENERAL DESCRIPTION

The standard instruments are located on three panels in the cockpit. An upper panel, a left lower panel, and a

right lower panel. The left lower panel contains a clock, oil temperature gauge, oil pressure gauge, fuel pressure

gauge, air filter Delta “P” gauge, hour meter, airframe related electrical switches and fuel quantity gauge. The right

lower panel contains the voltmeter, ammeter, and circuit breakers. The upper instrument panel contains the gas

generator percent tachometer (Ng), propeller tachometer (Np), torque pressure gauge, ITT (T5) indicator, boom

pressure gauge, air speed indicator, altimeter, fluid compass, engine warning lights, stall warning light and bank

indicator. All instruments are lighted with a post light or internally lighted and controlled with rheostats located on

the left lower panel.

Optional instruments and gauges are available upon request. A few of the optional instruments are hopper

quantity, Shadin Miniflo™ fuel flow, Micronair™ chemical flow meter, Crophawk™ chemical flow meter, encoding

alt., artificial horizon, electric turn and bank, vertical speed, and directional gyro.

INSTRUMENT SYSTEM MAINTENANCE

Unless otherwise specified, field maintenance of instrument systems is limited to removal and replacement of

defective instruments and transmitters; authorized in-service adjustment of transmitters and instruments; and

repair of instrument systems between the instrument and signal source (transducer). Reliability of the various

instruments and related systems can be sustained by routine inspection of electrical wiring for chafing, and

electrical connections for security. All fluid pressure, pitot pressure, and static line connections must be tight at all

times and lines must be correctly routed and secured. Electrical wiring must be free from chafing, properly

connected and secured. Instrument ports and lines disconnected during system maintenance must be capped or

plugged immediately to prevent the entrance of foreign material and consequent instrument malfunction.

Maintenance procedures pertaining to a specific instrument or system are contained in subsequent paragraphs.

As a general rule, it is recommended that the instrument signal source and means of transmission to the

instrument be rung out before changing an instrument. If a new instrument or a transducer is available, it may be

expedient to utilize them in the system as required to determine if the malfunction is in the instrument, signal

source or interconnecting line.

Effective: 1/26/05 8 - 2


FLIGHT INSTRUMENTS

Flight instruments consist of the magnetic compass, airspeed indicator, altimeter, and bank indicator. The pitot-

static system provides pitot (impact) and static (atmospheric) air pressure to the airspeed indicator and static air

pressure to the altimeter indicator.

PITOT-STATIC SYSTEM

The pitot head installed near the wing tip of the right wing lower surface provides pitot pressure. The pitot

pressure line is connected to the airspeed indicator. A static pressure line is connected to the altimeter and to the

airspeed indicator. The static pressure ports are located on both sides of the aft fuselage where they are

connected through a yoke to a tube that runs forward along the left side of the fuselage to the instruments.

MAINTENANCE

Flight instruments utilizing pitot-static pressure are highly sensitive to pressure variations. Therefore, all tubing

and line connections must be absolutely airtight to prevent erratic indications. Moisture drains for the system are

installed in the lines at two different locations; the most in-board end of tubing in wing and in the aft fuselage just

aft of the static ports. Drain the pitot-static system periodically and whenever the system operates erratically. If

after draining, and any of the pitot-static instruments are still inoperative or erratic, clear the pitot-static vent lines of

any remaining restrictions with dry, low-pressure compressed air. Disconnecting the static line at the altimeter and

applying two to four psi air pressure to the static line may purge the lines. Disconnecting the line from the

airspeed indicator and applying two to four psi pressures to the line may purge the pitot pressure line. Cap

instrument inlets before attempting to clear lines.

INSPECTION AND LEAKAGE TEST

** CAUTION ** Be sure air pressure is directed towards the pitot head and not

toward the instruments when purging the system.

The following procedure outlines inspection and testing of the static pressure system, assuming the altimeter has

been tested and inspected in accordance with current Federal Aviation Administration Regulations.

A. Ensure the static system is free from entrapped moisture and restrictions.

B. Ensure no alterations or deformations to the static lines have occurred.

C. Attach a source of suction to static pressure source opening.

D. Slowly apply suction until altimeter indicates a 1000-foot increase in altitude.

8 – 3 Effective: 1/26/05


** CAUTION ** When applying or releasing suction, do not exceed the

range of the vertical speed indicator or airspeed indicator.

E. Cut off the suction source to maintain a closed system for one minute. Leakage shall not exceed 100-foot of

altitude loss as indicated on altimeter.

F. If leakage rate is within tolerance, slowly release suction source.

* NOTE * If leakage rate exceeds the maximum allowable, first

tighten all connections then repeat the leakage test. If

leakage rate still exceeds the maximum allowable, use

the following procedure.

G. Disconnect static pressure line from airspeed indicator to altimeter. Cap tee at altimeter so that the altimeter

is the only instrument still connected to static pressure system.

H. Repeat the leakage test to check whether the static pressure system or the removed instruments are the

cause of leakage. If instruments are at fault, they must be repaired by an appropriately rated repair station or

replaced. If the static pressure is at fault, use the following procedure to locate the system leakage.

I. Attach a source of positive pressure to the static source opening.

** CAUTION ** Do not apply positive pressure with the airspeed indicator

connected to the static pressure system.

J. Slowly apply positive pressure until altimeter indicates a 1000-foot decrease in altitude and maintain this

altimeter indication while checking for leaks. Coat line connections, static pressure fittings and static

source external port opening with solution of mild soap and water, watching for bubbles to locate leaks.

K. Tighten leaking connections. Repair or replace any parts found defective.

L. Reconnect airspeed indicator to the static pressure system and repeat leakage test per steps C. through

F.

Effective: 1/26/05 8 - 4


ALTIMETER

The altimeter is equipped with three concentrically arranged pointers with a range of 0 - 100,000 feet. The

intermediate hand indicates altitude in hundreds of feet in 20-foot increments. The shortest hand indicates altitude

in thousands of feet and the longest pointer in tens of thousands of feet. A moveable barometric scale, visible

through a small window in the main dial, indicates the barometric pressure in inches of Hg and millibars. An

adjusting knob provides a means of adjusting the three pointers and barometric scale simultaneously to correct for

changes in atmospheric pressure and to establish the proper reference to sea level. Barometric pressure is

sensed through the instrument static system.

AIRSPEED INDICATOR

The airspeed indicator registers airspeed in miles-per-hour and/or knots. The indicator is operated by the

pressure differential between impact air pressure from the pitot tube and barometric pressure sensed through the

static system.

MAGNETIC COMPASS

The magnetic compass is a semi-floating cylinder encased in a liquid filled case with expansion provisions to

compensate for temperature changes. The compass is mounted on the instrument panel, is internally lighted, and

is equipped with compensating magnets that are adjustable from the front of the case. Covers on the face of the

compass allow access to adjust the compensating magnets. The compass should be swung and compensated at

regular intervals and at any time equipment installations are made that could cause compass deviation.

MAGNETIC COMPASS COMPENSATION

Locate the aircraft in area suitable for the method of magnetic compass compensation to be used. Close doors

and place flaps in a retracted position. Set the throttle at cruise position with engine operating. Place all electrical

switches, alternator, radio and other equipment in a mode normally used in flight and proceed with the following:

A. Set adjustment screws of compensating magnets to zero. Zero position is when the dot on the screw is

lined up with the dot on the compass frame.

B. Position aircraft in a magnetically north direction. Adjust north-south adjustment screw until compass

reads exactly north.

C. Position aircraft in a magnetically east direction. Adjust east-west adjustment screw until compass reads

exactly east.

D. Position aircraft in a magnetically south direction. Notice the resulting south error. Adjust north-south

adjustment screw so that one-half of the error has been removed.

8 – 5 Effective: 1/26/05


E. Position aircraft in a magnetically west direction. Notice the resulting west error. Adjust east-west

adjustment screw so that one-half of the error has been removed.

F. Position aircraft in successive magnetically 30-degree directions and record all errors on the deviation

card furnished with the compass.

BANK INDICATOR

The bank indicator, installed in the center of the upper instrument panels is a curved, fluid-filled tube containing a

ball. The gravitational and centrifugal forces position the ball within the tube to indicate correct lateral altitude for

the degree of banking.

POWER PLANT INSTRUMENTS

This group consists of the oil temperature gauge, oil pressure and fuel pressure gauge, ITT indicator, torque

indicator, propeller RPM indicator, percent gas generator speed indicator, and fuel quantity indicator. These

instruments are operated by fluid pressure, variation in electrical resistance created by a float operated transmitter,

variations in electrical resistance from a temperature probe, and by electrical variations from a tach-generator.

See Figure 8-1 for instrument markings.

MISCELLANEOUS INSTRUMENTS


A fuel quantity indicator registers the amount of fuel in the system up to a maximum of 164 U.S. gallons. Fuel

from 165 to 230 U.S. gallons is un-gaugeable. The indicator is basically a millivoltmeter that receives input signals

from the fuel quantity transducers (liquid level senders). The face of the fuel quantity indicator is marked in

increments from empty to full. The indicator is used in conjunction with two float-operated variable-resistance

transducers, one installed in each tank. The full tank position of the transducer float produces a minimum

resistance through the transducer, permitting maximum current flow through the fuel quantity indicator and

maximum pointer deflection. As the fuel level of the tank is lowered, resistance in the transducer is increased,

producing a decreased current flow through the fuel quantity indicator and a small pointer deflection. The fuel

quantity indicating system is calibrated by adjusting the fuel quantity transducer float arms and the indicator as

outlined in Section V.

VOLTMETER

A voltmeter displays electrical system voltage when the master switch is on and allows the pilot to monitor bus bar

voltage. Normal voltmeter readings must be within the green arc (24.0 to 30.5 volts). Insufficient voltage or

overcharging is indicated by a lower red arc (minimum) 16.0 to 22.5 volts, and an upper red arc (maximum) 30.5

Effective: 1/26/05 8 - 6


to 36.0 volts respectively. Continuous operation over 30.5 volts is detrimental to the life of the battery and could

cause loss of electrical power. A yellow arc from 22.5 to 24.0 volts indicates a caution range.

AMMETER

The ammeter displays current flow, in amperes, from the aircraft generator to the battery, or from the battery to the

electrical system. With the engine operating, the ammeter should indicate the on charge side unless there is an

aircraft generator malfunction, or if the electrical load demand exceeds the aircraft generator output, the ammeter

will indicate the discharge side. Continuous operation on the discharge side will be detrimental to battery life and

may cause loss of electrical power.

HOPPER QUANTITY (OPTIONAL EQUIPMENT)

The hopper quantity consists of three parts -- the level sensing element in hopper, FA-A control box normally

located on left side of cockpit, and quantity gauge located in instrument panel. The gauge is adjustable between 0

gallons to 360 gallons and incorporates two lights, one amber light for low quantity, and one red right for hopper

empty. The system can be calibrated as per the following instructions.

* NOTE *

A scale is provided to show the relationship of the remote

gauge, % scale on analog control unit type FA-A, and

number of inches the bottom of the floating ball is away

from the top of the lower stop collar. The column labeled

ground shows how many gallons are in the hopper with the

aircraft in the ground attitude in relationship to the remote

gauge reading.

CALIBRATION OF REMOTE GAUGE

A. The remote hopper level gauge markings are an indication of hopper load in level flight.

B. A screw type adjuster located on the face of the gauge at the six o’clock position adjusts the remote

gauge 0 mark. Adjustment of the screw CW or CCW will move pointer left or right.

C. With floating ball against lower stop collar, adjust screw adjuster until pointer is aligned with 0 mark on

gauge.

8 – 7 Effective: 1/26/05


D. The 360-gallon mark is adjusted by turning a screw head located on a 20 K ohm potentiometer on a

circuit board attached to rear of gauge.

E. With floating ball against the top stop of the sending unit; adjust the potentiometer until pointer is aligned

with 360 mark on gauge.

* NOTE * Unit must be on to adjust 360 side of gauge.

* NOTE *

The small % scale on analog control unit Type

FA-A will move in direct relationship with the

remote gauge.

F. Once steps 3 and 5 are completed, the unit is in calibration.

* NOTE *

Hopper loads above 360 gallons are un-

gaugeable.

G. Also provided are two hopper-level warning lights -- one amber and one red. They both have a push-to-

test feature and a dimming capability. The lights are adjusted to come on at any position (hopper level)

that you may desire by potentiometers located under pop-off caps on the face of the analog control unit

Type FA-A. The amber light is adjusted to come on by adjusting Pot 1 labeled set point 1, and the red

light is adjusted to come on by adjusting Pot 2 labeled set point 2.

H. The amber and red lights can be set at any position you may desire. Thrush Aircraft Inc recommends

setting the amber light to come on at 25 gallons, or 8 3/8" from top of lower stop collar to bottom of floating

ball, and the red light to come on at 6 1/2 gallons, or 2 1/2" from top of lower stop collar to bottom of

floating ball.

Effective1/26/05 8 - 8



AIRSPEED INDICATOR


Hand fails to respond

Pitot pressure connection not

properly connected to

pressure line from pitot tube.

Test line and connection

for leaks. Repair or

replace damaged line,

tighten connections.

Pitot or static lines clogged.

Check line for

obstructions. Blow out

lines.

Leak in pitot or static line.

Test lines and

connections for leaks.

Repair or replace

damaged lines, tighten

connections.

Defective mechanism.

Substitute known-good

instrument and check

reading. Replace

instrument.

Incorrect indication or hand

oscillates.

Leaking diaphragm.


instrument and check

reading. Replace

instrument.

Excessive vibration.

Check instrument

mounting screws.

Tighten mounting

screws. Hand Vibrates.

Excessive tubing vibration.

Check clamps and lines

connections for security.

Tighten clamps and

connections.

8 – 9 Effective: 1/26/05


ALTIMETER


Static line plugged.

Check line for obstructions. Check

static source. Blow out lines. Clean

static source. Instrument fails to operate.

Defective mechanism. Substitute known-good altimeter and

check reading. Replace instrument.

Hands not carefully set. Reset hands with know.

Leaking diaphragm Substitute known-good altimeter and

check reading. Replace instrument. Incorrect indication.

Pointers out of calibration. Compare reading with known-good

altimeter. Replace instrument.

Static pressure irregular.

Check lines for obstructions or leaks.

Check static source. Blow out lines,

tighten connections. Clean static

source. Hands oscillate.

Leak in airspeed or vertical

speed indicator installations.

Check other instruments and system

plumbing for leaks and obstructions.

Blow out lines, tighten connections.

Effective: 1/26/05 8 - 10


MAGNETIC COMPASS


Compass not properly

compensated

Swing compass and

compensate.

Excessive card error

External magnetic interference Locate and eliminate

interference.

Insufficient liquid Replace compass.

Excessive card oscillation

Excessive vibration of compass Remove cause of vibration.

8 – 11 Effective: 1/26/05



Card element not level,

sluggish

Compass excessively

compensated

Back compensating screws off

to remove all compensation,

then recompensate compass.

Liquid leakage from case

Leaking float chamber due to

broken cover glass or case, or

defective sealing gaskets, weak

or detached card magnets, pivot

friction, or broken jewel

Replace compass.

ENGINE OIL PRESSURE GAUGE


Worn or bent movement Replace instrument.

Dirty or corroded movement Replace instrument. Gauge has erratic

operation

Pointer bent and rubbing on dial,

dial screw or glass Replace instrument.

Effective: 1/26/05 8 - 12



Pressure line clogged

Check line for

obstructions. Clean

line.

Leak in pressure line

Check line for leaks

and damage. Repair or

replace damaged line.

Pressure line broken

Check line for leaks

and damage. Repair or

replace damaged line.

Pointer loose on staff Replace instrument.

Gauge does not register

Damaged gauge movement Replace instrument.

Gauge pointer fails to return to

zero Foreign matter in line

Check line for

obstructions. Clean

line.

Gauge does not register properly Faulty mechanism Replace instrument.

8 – 13 Effective: 1/26/05


ENGINE FUEL PRESSURE GAUGE


Gauge inoperative or erratic.

Low pressure or flow registered. Restricted, broken or leaking line

Clear and clean line.

Tighten fittings or

replace, if necessary.

Vapor in fuel line

Start and run engine

until instrument

registers normally.

Fuel pressure or fuel flow

registered is high, low or erratic Faulty relief valve in engine-

electric driven pump(s) or

defective pump(s)

See fuel pump in

Power Plant Section for

replacement or relief

valve adjustment

instructions.

Effective: 1/26/05 8 - 14


TACHOMETER


Tachometer

generator/tachometer defective

Test generator for

output. Overhaul or

replace if necessary.

Test instrument and

replace, if necessary. Tachometer registers low,

erratically, or no reading.

Tachometer generator shaft

sheared

Replace tachometer

generator.



Fuel tanks empty

Check fuel quantity.

Service with proper

grade and amount of

fuel.

Failure to indicate

No power to indicator or

transmitter. (Pointer stays below

E).

Check circuit breaker.

Inspect for open circuit.

Reset breaker, repair

or replace defective

wire.

8 – 15 Effective: 1/26/05



Grounded wire. (Pointer stays

above F).

Check for partial

ground between

transmitter and

indicator. Repair or

replace defective wire.

Low voltage

Check voltage at

indicator. Correct

voltage.

Defective indicator


indicator. Replace

indicator.

Failure to indicate. continued

Defective sending unit Replace sending unit

Defective indicator


indicator. Replace

indicator. Sticky or sluggish indicator

operation

Low voltage

Check voltage at

indicator. Correct

voltage.

Registers either full or empty Float arm stuck Free float arm.

Effective: 1/26/05 8 - 16


ENGINE OIL TEMPERATURE GAUGE


Defective indicator or transmitter Replace instrument or

transmitter.

Gauge has erratic operation

Grounded wire

Check for partial ground

between transmitter and

indicator. Repair or

replace defective wire.

Gauge does not register No power to indicator or defective

instrument or transmitter

Check circuit breaker.

Inspect for open circuit.

Reset breaker. Repair or

replace instrument or

transmitter.

8 – 17 Effective: 1/26/05


HOPPER QUANTITY SYSTEM (OPTIONAL EQUIPMENT)


Bad sensing element

Disconnect wires and check

resistance. Refer to Fig. 8-

2. No indication on indicator

Resistance check bad.

Replace element (non-

repairable).

No indication and resistance

checked good Bad indicator

Check power and ground

connections. Observe

meter on FA-A control unit.

If meter works with

movement of ball, indicator

or wires to indicator are bad.

No indication and resistance

checked good Bad FA-A control

Check wiring. Wiring OK,

replace FA-A control. Pins

5, 8, 11 and 15 are power

pins on FA-A unit. Pin 15 is

attached to 1A C/B. Pins 8

and 11 are jumped to pin

15. Pin 5 goes to + on rear

of tank quantity gauge.

Grounds are located on pins

4 and 16. Pin 4 is

connected to ground on rear

of tank quantity gauge. Pin

16 is chassis ground.

Effective: 1/26/05 8 - 18


INSTRUMENT MARKINGS Limits Features Meaning

OIL TEMP ( °C) -40 -40 to10 10 to 99 99

Red Radial Yellow Arc Green Arc Red Radial

MINIMUM CAUTION NORMAL MAXIMUM

OIL PRESSURE (PSI) PT6A-34AG, -34, -36

40 40 to 85 85 to 105 105



OIL PRESSURE (PSI) PT6A-41AG,-41 60 60 to 105** 105 to 135 135



OIL PRESSURE (PSI) PT6A-42 60 60 to 100*** 100 to 135 135



FUEL PRESSURE (PSI) 5 None 5 to 50 50



ITT (°C) PT6A-34AG, -34, -36, -42 None None 400 to 790 790



ITT (°C) PT6A-41AG, -41

None None 400 to 750 750



TORQUE (PSI) PT6A-34AG, -34, -36 None None 0 to 58.7 58.7*


MIMIMUM CAUTION NORMAL MAXIMUM

TORQUE (PSI) PT6A-41AG, -41, -42 None None 0 to 64.5 64.5



Figure 1

8 – 19 Effective: 1/26/05


PROPELLER RPM (Np) PT6A-34AG, -34, -36




PROPELLER RPM (Np) PT6A-41AG, -41, -42




GAS GENERATOR SPEED (% Ng)

None None 50 to 101.5 101.5



AIRSPEED 66 to 123 MPH CAS 70 to 126 MPH CAS 129 to 159 MPH CAS 159 MPH CAS

White Arc Green Arc Yellow Arc Red Radial

Flaps Operating Range Normal Operating Range Caution Range Never Exceed

* In addition to a red radial limit marker at 58.7 PSI, the torque meter is marked with a red diamond at 64.5 PSI. (58.7 PSI is the torque limit for 2200 RPM, the normal takeoff RPM. 64.5 PSI is the torque limit for 2000 RPM, an optional RPM setting for in-flight use.) ** The 60-105 oil pressure caution range is for idling and for emergency completion of a flight using 36 PSI torque or below. *** The 60-100 oil pressure caution range is for idling and for emergency completion of a flight using 36 PSI torque or below.

Figure 1a

Effective: 1/26/05 8 – 20


LEVEL SENSING UNIT

BLUE

BLACK

BROWN

WIRES COMING FROM TUBE

APPROXIMATE RESISTANCE WITH FLOAT IN CENTER OF TUBE

1. BLACK

BLUE 3000 – 5000 OHMS

2. BLACK

BROWN 2000 – 3000 OHMS

3. BROWN

BLUE 2000 OHMS

4. BLACK

BLUE

SMALLER OR EQUAL TO BLACK/BROWN PLUS

RESISTANCE BETWEEN BROWN/BLUE

Figure 8-2

8 - 21 Effective: 1/26/05


Section 9 DISPERSAL SYSTEMS

TABLE OF CONTENTS DISPERSAL SYSTEMS ................................................................................................................. 2

GENERAL DESCRIPTION ......................................................................................................... 2 HOPPER ................................................................................................................................. 3 HOPPER CARE ...................................................................................................................... 3 HOPPER REPAIR................................................................................................................... 3 HOPPER GATE BOX REMOVAL ........................................................................................... 4 HOPPER GATE BOX INSTALLATION ................................................................................... 4 HOPPER ADAPTER BOX REMOVAL.................................................................................... 5 HOPPER ADAPTER BOX INSTALLATION............................................................................ 5 DISPERSAL EQUIPMENT...................................................................................................... 5 AGITATOR AND SPREADER................................................................................................. 6

Effective: 1/26/05 9 - 1


DISPERSAL SYSTEMS SECTION NINE

GENERAL DESCRIPTION

A reinforced fiberglass hopper is the principal part of both the solid and spray units. The hopper top

forms the cowling from the cockpit forward to the firewall. The hopper gate box is designed to be

liquid as well as dust tight. Emergency jettison controls permit the entire liquid load to be dumped in

approximately 6.5 seconds for the 510 gal. Hoppers.

The dispersal system has been designed to handle a wide range of dispersal equipment, and to

allow for a quick, easy changeover from one type of equipment to another. All dispersal plumbing is

externally mounted and equipped with quick-disconnects to allow for ease of maintenance and

cleaning. The streamlined aluminum extrusion spray booms are located below the wing trailing

edge and utilize the downwash from the wing to increase penetration. The booms are fitted with

spraying system diaphragm type nozzles and normally will use 35 nozzles for low volume output

and 70 nozzles for high volume output. In addition, the spray booms have large end plugs that can

be removed to aid in flushing the system.

The spray pump is located under the fuselage between the main landing gear struts. A three-way

suck-back spray valve located at the left, underside of the fuselage, controls the spray pressure and

flow. The valve is actuated from the cockpit to obtain the desired operating pressures for various

spray applications. Spray pressure is indicated by a gauge mounted on the upper instrument panel

and is controlled by a vernier adjustment on the liquid spray-operating handle. The spray pump is a

wind-driven fan type, and is controlled from the cockpit by means of a cable to adjust the fan blade

pitch to increase or decrease pump pressure.

Effective: 1/26/05 9 - 2


HOPPER (Figure 9-2) HOPPER CARE

Regardless of the materials used in the construction or coating of the hopper, it should be

thoroughly washed after each day’s work. Use cold, clean water and any domestic detergent.

Inspect the interior of the hopper daily, for evidence of chemical attack, such as surface roughness

or deterioration of the resin. Look for cracks that may have started in the areas of highest stress,

such as attach points and stiffener center portions. Repairs may then be made at the beginning of

the problem, rather than after it has progressed to a serious degree.

* NOTE *

After washing, it is very important that the door and gate be left

open for good ventilation and complete drying. It is good practice

to rinse the hopper with cold water after use with chemicals, even

if the idle period ahead is going to be only a few hours.

HOPPER REPAIR (Figure 9-1)

Hopper repair may be accomplished as follows:

A. Fiberglass surfaces must be clean, dry and free of oil, wax or other foreign matter. If

chemical erosion is evident, sand rough areas and wash with any good domestic detergent.

Rinse with clean water. Sand all surfaces that are to receive a polyester coating. Use

Ashland Specialty Chemical Company’s 7241 T15 AROPOL™ polyester resin or

equivalent for the hopper repair.

B. If damage consists only of surface cracks, excessive abrasion or chemical erosion, sand all

affected surfaces smooth. Extend the prepared surface six inches beyond the damaged

area.

C. If damage consists of cracks or holes extending completely through the wall, sand the

surfaces on both sides deep enough to expose the first layer of cloth.

D. Surface damage requires repairs only to the eroded or cracked side. Damage extending

through the wall requires repairs to both the inner and outer surfaces of the hopper. The

number of layers in either case should equal the original basic wall thickness. (Figure 9-1)

Highly stressed areas, such as attach points, require an extra layer of cloth and mat on

each side, in addition to the basic wall thickness.

9 – 3 Effective: 1/26/05


E. Curing temperature is 70°F minimum. Higher temperatures accelerate curing. A maximum

of 150°F for four hours is recommended followed by ten or more hours at 70°F.

F. Brush the resin generously over the entire area. Apply alternate layers of fiberglass cloth

and mat. Each layer should overlap the preceding layer approximately one inch. After

each layer is in place, use a squeegee and/or roller to remove excess resin and air voids.

HOPPER GATE BOX REMOVAL (Figure 9-4)

Remove the aircraft skins to gain access to hopper gatebox bolts and nuts. Disconnect dump fork,

spray tube, adapter box and vent, emergency shut off cable and spray pump. Remove the nuts and

bolts and pry off adaptor box.

** CAUTION **

If bolts do not drive out easily, turn bolts to break glue, and then drive bolts out.

Clean off all old gasket material by scraping, being cautious not to gouge mating surfaces.

HOPPER GATE BOX INSTALLATION (Figure 9-4)

Before installing gate box, be sure that all mating surfaces are clean and dry. For maximum

strength, apply 3M Scotch-Weld ™ DP-190 Translucent Epoxy adhesive evenly to both mating

surfaces and both sides of gaskets thoroughly. Using alignment pins to hold gaskets in place and

to help align gate box, install bolts using large area washer on hopper side and tighten nuts.

* NOTE *

Excessive uncured adhesive can be cleaned up with keytone type

solvents. (When using solvents, extinguish all ignition sources and

follow the manufacturer’s precautions and directions for use for

handling such materials.) Application of adhesive to substrates

should be made within 75 minutes after mixing. Working life is 80

minutes. Higher temperatures will reduce these times.

Allow 24 hours for sealer to cure before putting back into service. Reinstall tubes, controls and

cables and pump. Fill hopper and check for leaks. Water should be allowed to stay in hopper for a

minimum of two hours with no leaks.

Effective: 1/26/05 9 - 4


HOPPER ADAPTER BOX REMOVAL (Figure 9-4) Remove all bolts that attach adapter box to hopper throat. Pry the adapter box from the hopper to

release the grip created by the sealer used during assembly. Clean off all old gasket material by

scrapings being cautious not to gouge the fiberglass hopper lip or adapter box mating surface.

HOPPER ADAPTER BOX INSTALLATION (Figure 9-4) Reverse the removal procedure.

DISPERSAL EQUIPMENT

Because of the variety of solid dispersal equipment available and the wide variety of dusts, seeds,

pellets and granular material that can be dispersed by the spreader, it is advised that the

manufacturer's instructions for the various types of dusting equipment be carefully followed for best

results. In conjunction with the manufacturer's instructions for maintenance of the spray dispersal

system, it is recommended that a periodic interval be established for accomplishment of the

following:

A. Inspect the hopper baffles for security and condition.

B. Inspect hopper lid for condition of seal and security of latches.

C. Inspect the hopper for indications of leaks and general condition.

D. Inspect hopper gate for evidence of leaks and proper operation.

E. Inspect hopper vent tube for evidence of corrosion and security.

F. Inspect emergency and 3-way valve handles and controls rods for cracks around welds.

Check condition of control rod boot.

G. Inspect liquid lines for leaks and hose deterioration.

H. Inspect all line supports and clamps for security or corrosion.

I. Drain and clean spray strainer.

J. Inspect the pump, fan and brake assemblies for security and proper operation.

9 – 5 Effective: 1/26/05


K. Refer to manufacturer's data for pump lubrication.

L. Inspect emergency on/off control and valve for security and proper operation.

M. Inspect 3-way pressure control valve for security and proper operation.

N. Inspect both booms and the support for each boom for security and evidence of corrosion.

O. Inspect all nozzle diaphragms for deterioration.

P. Inspect all fan blades for cracks or nicks.

Q. Inspect all nozzles for orifice erosion. Replace as necessary.

AGITATOR AND SPREADER (Figure 9-5)

A. Inspect gearbox for proper oil level (refer to manufacturer's data).

B. Inspect fan, gearbox, drive shaft, agitator and coupling for security and proper operation.

C. Inspect agitator shaft seal at hopper for evidence of leaks.

D. Inspect spreader unit for cracks, loose rivets, loose or missing vanes and security to

airframe.

E. Support spreader, connect rear support tubes. Raise front connection camloc fasteners to

hopper sump and side latches. Spreader should be level, centered, and clear of door

control arms.

Effective: 1/26/05 9 - 6


Figure 9-1

9 – 7 Effective: 1/26/05


A-A

A

A

Effective: 1/26/05 9 – 8 Figure 9-2



9 – 9 Effective: 1/26/05

Figure 9-3


Figure 9-4

Effective: 1/26/05 9 - 10


Figure 9-

9 – 11 Effective: 1/26/05


Figure 9-6

Effective: 1/26/05 9 - 12


Section 10 ELECTRICAL SYSTEM

TABLE OF CONTENTS ELECTRICAL SYSTEM....................................................................................................................................2

GENERAL DESCRIPTION ...........................................................................................................................2 POWER DISTRIBUTION ..........................................................................................................................2 BATTERY AND EXTERNAL POWER ......................................................................................................2 BATTERY SERVICING..............................................................................................................................2 SERVICING BATTERY INSTALLED IN AIRCRAFT .................................................................................4 GENERATOR SYSTEM.............................................................................................................................4 DIAGRAMS ................................................................................................................................................4 BATTERY OPERATION............................................................................................................................4 BATTERY REMOVAL................................................................................................................................4 BATTERY INSTALLATION........................................................................................................................5 VOLTAGE REGULATION.........................................................................................................................5 VOLTAGE REGULATOR REMOVAL........................................................................................................5 VOLTAGE REGULATOR INSTALLATION................................................................................................5 STARTER - GENERATOR MAINTENANCE ............................................................................................6 STARTER - GENERATOR REMOVAL......................................................................................................6 STARTER-GENERATOR INSTALLATION................................................................................................7 TROUBLESHOOTING CHART.................................................................................................................8 BATTERY SYSTEM...................................................................................................................................8 STARTER/GENERATOR - GENERATOR PHASE .................................................................................10 STARTER/GENERATOR - STARTER PHASE .......................................................................................12

Effective: 1/26/05 10 - 1


Effective: 1/26/05 10 - 2

ELECTRICAL SYSTEM

SECTION TEN GENERAL DESCRIPTION

The aircraft 24-volt DC electrical system is designed to provide the utmost in reliability. Two 24-volt

storage batteries provide electric current for engine starting and a reserve source of electrical power

in the event of generator failure. A D.C. power receptacle provides a means for connecting external

power to the aircraft electrical system. To conserve battery life, external power should always be

used for starting engines when temperature is below 40°F or when performing maintenance

requiring electrical power. A generator installed on the engine supplies the primary source of

electrical power to the main bus. A voltage regulator protects the electrical system, reverse current

relay and circuit breakers. If generator output voltage is below bus voltage, the battery supplies the

busloads. The D.C. ammeter, installed in the instrument panel, indicates the discharge or charge

on the battery after the engine is started. All electrically operated motors, lighting systems and

other electrical component circuits are protected by push button thermal circuit breakers. Switches

and instruments required for operation of the aircraft electrical system are installed in the instrument

panel and engine control switch panel.

POWER DISTRIBUTION

The 24-volt D.C. electrical system depends upon electrical power from three different sources:

battery, external power and the generator. With the engine operating and the generator on the line,

electric power from the generator is provided through a circuit breaker to the main bus.

BATTERY AND EXTERNAL POWER

Two 24-volt storage batteries provide power to the circuit breaker through relays. A two-position

(BAT OFF-ON) switch located on the engine control switch panel controls the relays. Placing the

battery switch in the ON position closes the relay to supply power to the circuit breaker bus from the

battery or external power. Placing the battery switch in the OFF position de-energizes the battery

relay and terminates the supply of power to the electrical system.

BATTERY SERVICING

INITIAL SERVICING OF A DRY CHARGE GE50C BATTERY IS AS FOLLOWS:


10 - 3 Effective: 1/26/05

A. Remove seals (if present) from cells.

B. Fill each cell with 1.285 specific gravity sulfuric acid to bottom of split ring. Use only glass,

rubber or plastic materials for containing battery electrolyte fluid during servicing and wear

protective clothing and rubber gloves when handling electrolyte to prevent personal injury.

Use a solution of baking soda and water to neutralize any acid spilled on clothing, skin or

any damageable surface.

C. Sway the battery from side to side to release any trapped air. Re-adjust the electrolyte as

necessary.

C. Let battery sit unused for one hour.

E. Check and re-adjust electrolyte level as necessary by adding more electrolytes to obtain

proper level as stated in procedure B.

F. Install vent plugs tightly into each cell.

G. Clean and neutralize any spilled electrolyte on battery.

H. Charge battery until all cells are gassing freely and the charge voltage and specific gravity

of electrolyte are constant over three successive readings taken at one-hour intervals.

(This procedure may take 18 - 24 hours with a constant current charger.) During the period

of charging, the electrolyte temperatures shall be maintained between 60°F and 110°F

(15.6°C and 43.3°C). Charge rate is 3 amps. Reduce rate by 1/2 when cells start gassing.

I. When the battery is completely charged, the specific gravity should read between 1.285

and 1.295. At this point, if electrolyte level needs to be adjusted, remove or add electrolyte

to proper level and recharge for one hour.

** CAUTION **

Gasses given off by a battery under charging conditions are flammable.

* NOTE *

For more detailed instructions, see Gill Service Manual.


Effective: 1/26/05 10 - 4

SERVICING BATTERY INSTALLED IN AIRCRAFT

The 24-volt battery is installed aft of the engine on the engine mount lower longerons and is

accessible through the removable cowling skins. Check the battery electrolyte level frequently,

especially during hot weather. If a visual check shows low cell level, add distilled water to bring the

cell(s) up to the proper lever. (See battery-servicing instructions)

1.285 - 1.295 CHARGED

1.275 - Less RECHARGE

GENERATOR SYSTEM

The generator system consists of a generator; voltage regulator, reverse current relay and circuit

breaker (See electrical diagrams). The generator is connected to the circuit breaker bus and will

supply the current demands when output voltage exceeds battery voltage.

DIAGRAMS

The Electrical Diagrams of the Model S2R-T34 aircraft is at the end of this section.

BATTERY OPERATION

Battery operation is controlled by a battery switch, placarded BATT-ON-OFF, located on the switch

panel in the cockpit. The battery is capable of assuming the complete electrical load for a limited

time at 70 amps max.

The batteries are located on the battery plate assembly on the engine mount aft of the engine.

They are installed with two-battery hold down rods through the battery cover. The battery case is

vented overboard to dispose of any electrolyte or hydrogen gas fumes discharged during normal

charging operation. Air enters the battery compartment from an air scoop located in the left cowl

shin skin, circulates throughout the battery compartment, and exists through a vent in the battery

and drains overboard through a vent located on the belly skins.

BATTERY REMOVAL

A. Verify that the BATT-ON-OFF switch is off. Disconnect external power.


10 - 5 Effective: 1/26/05

B. Remove R.H. aft cowl skin.

C. Disconnect the quick disconnect from the battery and remove all safety wire.

D. Disconnect vent tubes.

E. Remove nuts from battery hold down rods and remove batteries from compartment.

BATTERY INSTALLATION

Reverse battery removal procedure.

VOLTAGE REGULATION

The generator output voltage is regulated by the voltage regulator circuitry. By using an integrated

circuit comparator amplifier with a regulated reference voltage, and difference between the

reference voltage and the generator voltage is amplified and supplied to the comparator circuit,

which controls the shunt field excitation of the generator. Prior to installation, the voltage regulator

is adjusted under NO load condition to maintain 26.5+-.2 volts DC generator output voltage. After

installation, the generator over voltage control should be adjusted to 27.5 VDC generator output

voltage at the bus with normal systems turned on.

VOLTAGE REGULATOR REMOVAL

Gain access to the voltage regulator by removing R.H. aft cowl skin. If removal is necessary,

proceed as follows:

A. Verify that the battery switch is OFF, that the external power is disconnected, and that the

batteries are disconnected.

B. Disconnect the retaining clips from the voltage regulator and remove voltage regulator from

voltage regulator base.

VOLTAGE REGULATOR INSTALLATION

A. Verify that the battery switch is OFF.


Effective: 1/26/05 10 - 6

B. Brighten up all electrical contacts, both on the voltage regulator and the 6 each fingers on

the voltage regulator base, with Scotch-brite™ 07448 Ultra fine abrasive pad or

equivalent. Place the voltage regulator into position and snap it to the regulator base with

the retaining clips.

D. Connect the battery.

E. After installing, re-check voltage regulator for 27.5 VDC with engine running and normal

systems operating. Adjust as necessary.

STARTER - GENERATOR MAINTENANCE

STARTER - GENERATOR REMOVAL

A. Verify that the battery switch is OFF and that external power is disconnected. Disconnect

the batteries.

B. Open the upper aft engine cowling to gain access to the starter-generator.

D. Loosen the quick-disconnect clamps securing the starter-generator to the mounting adapter

and remove the starter-generator.

* NOTE *

Refer to the starter-generator maintenance manual for specific

maintenance instructions.

** CAUTION **

It is mandatory that the starter-generator be fully supported from the time the retaining clamp is loosened until the unit is removed from the engine. The starter-generator must never be allowed to support its own weight through the splined shaft engagement. If this precaution is not observed, damage to the shaft shear section or engine internal starter generator drive/breather impeller

carbon seal will result.


10 - 7 Effective: 1/26/05

STARTER-GENERATOR INSTALLATION

A. Verify that the battery switch is OFF and that external power is disconnected. Disconnect

batteries.

B. On “wet type” spline engines (approx. 1998 and newer) install new “0” ring on starter-

generator drive shaft. Lubricate “wet-type” splines with engine oil. On engines equipped

with “dry type” splines, lubricate splines with Dow Corning™ M-77 molycoat.

C. Position the starter-generator on the mounting adapter and secure it in place with the quick-

disconnect clamp.

D. Close the clamp hinge over the T-bolt. Check with a mirror to make certain the clamp

groove fully captures both the flange on the quick-disconnect adapter and the flange on the

starter-generator around its entire circumference.

E. When the clamp is properly positioned and the hinge and T-bolt are closed, tighten the T-

bolt nut to a torque of 70 inch pounds. Tap circumference of clamp lightly with

plastic/rudder mallet. Re-torque T-bolt nut to 70 inch pounds and repeat until you achieve

70 inch pound of torque without nut moving.

F. Connect the electrical leads to the starter-generator as previously marked.

F. Secure the upper aft engine cowling and connect the batteries. Run the engine at idle

speed for at least two minutes. Shut down the engine and recheck the quick-disconnect

clamp for proper torque.


Effective: 1/26/05 10 - 8


BATTERY SYSTEM

TROUBLE PROBABLE CAUSE REMARKS

Battery fails to hold charge Battery defective Replace battery

Battery will not come up to

full charge Charging rate to low

Check voltage regulator

and adjust to 27.5± .2

VDC

Battery consumes water

rapidly Charging rate too high

Check voltage regulator

adjust to 27.5 ±.2

Electrolyte level too high

Remove excess

electrolyte & adjust

specific gravity

Excessive charging rate Check voltage regulator

for correct voltage

Electrolyte runs out drain

tube

Vent caps loose or broken Tighten or replace caps

Standing too long Remove battery and

recharge

Equipment left on Remove battery and

recharge Battery discharged

Short circuit in wiring Check wiring and correct


10 - 9 Effective: 1/26/05


Charging rate too high Adjust voltage regulator

to 27.5


Battery discharged, defective

or disconnected

Check battery, recharge

or replace battery

Blown C/B in battery control

circuit

Check C/B and reset if

necessary

Defective wiring in battery

control circuit

Continuity. Check

circuitry and repair as

necessary

Defective battery relay

Check relay for proper

operation and replace as

necessary

No power indicated with

battery switch on

Battery switch defective

Check switch for proper

operation. Replace if

necessary

Power on with battery switch

in OFF position Shorted or sticking contacts

Check switch and relay

for proper operation.

Replace if necessary.


Effective: 1/26/05 10 - 10

STARTER/GENERATOR - GENERATOR PHASE


Loose connection Check connections

throughout system

Open or shorted field circuit in

generator or defective

armature

Test resistance of field.

Check field circuit

connections. Replace

starter-generator if

defective.

Brushes not contacting

commutator

Clean brushes and

holders with a clean,

dry, lint-free cloth.

Replace weak springs.

Brushes worn Replace brushes

Zero or low voltage indicated

Dirty commutator Clean commutator




10 - 11 Effective: 1/26/05

Zero or low voltage

indicated. continued

Defective voltage regulator

circuit

Adjust or replace

regulator

Generator circuit breaker

tripped or 130 amp buss limit

fuse blown

Check for short circuit

and reset circuit breaker

and / or replace buss

limit fuse

Improper connections Check connections

against wiring diagram

Defective generator control

switches Continuity.

Check switch for proper

operation and replace if

necessary.

Defective Reverse Current

Relay

With engine running,

check for 27.5vdc at

both GEN and BATT

terminals. If 27.5vdc at

GEN and 24vdc at

BATT, reverse current

relay is faulty.

No generator output

Defective generator

Check generator with

an ohmmeter and

replace starter-

generator if necessary

Defective wiring

Continuity. Check

wiring and repair as

necessary Volt-ammeter does not

indicate

Defective meter Replace meter


Effective: 1/26/05 10 - 12

STARTER/GENERATOR - STARTER PHASE


Circuit breaker tripped Check for short circuit

and reset circuit breaker

Low battery

Check battery. Service

and recharge as

necessary

Starter relay inoperative Check relay for operation

and replace if necessary

Battery relay inoperative Check relay for operation

and replace if necessary

Loose connection or faulty

ground in starter power circuit

Continuity. Check and

repair starter as

necessary

Starter inoperative

Defective starter motor

Check brushes, springs

and condition of

commutator continuity.

Check starter windings

for open or short circuit.

Repair or replace




10 - 13 Effective: 1/26/05


Low battery

Check battery. Service

and recharge if

necessary

High resistance starter circuit

Check resistance of each

connection. Maximum

resistance at any

connection is 0.001 ohm.

Inspect connections for

evidence of heating.

Clean and tighten

connections as

necessary.

Starter produces low Ng

Defective starter motor

Check brushes, springs

and commutator.

Replace brushes and

springs and clean

commutator as

necessary.


Effective: 1/26/05 10 - 14

Figure 10-1 Standard Equipment Flap System Wiring


10 - 15 Effective: 1/26/05

Figure 10-2 Standard Equipment

Fuel Quantity


Effective: 1/26/05 10 - 16


Hour Meter


10 - 17 Effective: 1/26/05


Low Oil Light & Stall Warning


Effective: 1/26/05 10 - 18

THIS

PAGE

INTENTIONALLY

LEFT

BLANK.

Figure 10-5


10 - 19 Effective: 1/26/05


Windshield Washer/Wiper


Effective: 1/26/05 10 - 20

Figure 10-7

Standard Equipment Strobe Lights


10 - 21 Effective: 1/26/05

Figure 10-8a Standard Equipment

Navigation & Instrument Lighting


Effective: 1/26/05 10 - 22

Figure 10-8b Standard Equipment

Navigation & Instrument Lighting


10 - 23 Effective: 1/26/05

Figure 10-9 Engine / Pratt & Whitney

Beta & Chip Detector


Effective: 1/26/05 10 - 24


Aux Fuel Pump, Igniters & Prop Test


10 - 25 Effective: 1/26/05


ITT


Effective: 1/26/05 10 - 26


Power Distribution


10 - 27 Effective: 1/26/05

Figure 10-13 Quick Disconnect


Effective: 1/26/05 10 - 28


QDA


10 - 29 Effective: 1/26/05


QDB


Effective: 1/26/05 10 - 30


QDC


10 - 31 Effective: 1/26/05


QDD


Effective: 1/26/05 10 - 32


QDE


10 - 33 Effective: 1/26/05


QDF


Effective: 1/26/05 10 - 34


QDG


10 - 35 Effective: 1/26/05


QDH


Effective: 1/26/05 10 - 36


QDJ


10 - 37 Effective: 1/26/05


QDL


Effective: 1/26/05 10 - 38


QDN


10 - 39 Effective: 1/26/05


QDP


Effective: 1/26/05 10 - 40


QDR


10 - 41 Effective: 1/26/05

Figure 10-26a Quick Disconnect

Wire Harness Routing


Effective: 1/26/05 10 - 42

Figure 10-26b Quick Disconnect

Wire Harness Routing


10 - 43 Effective: 1/26/05

Figure 10-27 Optional

Auto Flagger


Effective: 1/26/05 10 - 44


Boom Pressure, Turn & Bank, DG


10 - 45 Effective: 1/26/05


Low Voltage Light, D.G. & Artificial Horizon


Effective: 1/26/05 10 - 46


Smoker & Map Light


10 - 47 Effective: 1/26/05


Zee Air Conditioner


Effective: 1/26/05 10 - 48


Hopper Rinse & Pitot Heat


10 - 49 Effective: 1/26/05


Micronair Flowmeter, Hopper Light & Agitator


Effective: 1/26/05 10 - 50


Avionics Buss


10 - 51 Effective: 1/26/05


Electric Fan Brake & Crop Hawk


Effective: 1/26/05 10 - 52


Hopper Quantity & Fuel Flow


10 - 53 Effective: 1/26/05


Hi Capacity Windshield Washer


Effective: 1/26/05 10 - 54


Night Working Light Control


10 - 55 Effective: 1/26/05

Figure 10-39a Optional Night Working Lights


Effective: 1/26/05 10 - 56

Figure 10-39b Optional Night Working Lights


10 - 57 Effective: 1/26/05

Figure 10-40

Standard Equipment Oil Cooler Augmentation


Section 11 AIRWORTHINESS LIMITATIONS

TABLE OF CONTENTS AIRWORTHINESS LIMITATIONS ................................................................................................................... 2

STRUCTURAL LIMITATIONS...................................................................................................................... 2 STRUCTURAL INSPECTION LIMITATIONS .............................................................................................. 2

11 - 1 Effective: 1/26/05


AIRWORTHINESS LIMITATIONS The life limited parts on the airframe are listed in the chart below and must be replaced at the flight hours shown.

STRUCTURAL LIMITATIONS

PART DESCRIPTION PART NUMBER LIFE LIMIT

Spar Cap Assy, Left hand lower 22507T001 29,000 hours

Spar Cap Assy, Right hand lower 22507T002 29,000 hours

STRUCTURAL INSPECTION LIMITATIONS

PART DESCRIPTION PART NUMBER LIFE LIMIT

Spar Splice Block, lower (upper half) 22508T001 Annual*

Spar Splice Block, lower (lower half) 22508T002 Annual*

*NOTE* Visually inspect splice blocks with a 10X magnifying glass or dye penetrant. Inspect for external cracks

around the 1/4 inch and 5/16 inch hole locations. If no cracks are detected, this portion of the wing

inspection is complete. If cracks are found remove the splice blocks before next flight and inspect the lower

spar cap for cracks in accordance with Thrush Aircraft, Inc. Service Bulletin SB-AG-39. If cracks are found

in spar cap contact Thrush Aircraft, Inc. for possible repair or replacement. If no cracks are found in spar

cap, replace the cracked splice blocks with new units. Refer to Section VII “Wing Removal” for splice block

removal and installation.

Effective: 1/26/05 11 - 2

Documents

Ayres S2R-T Turbo Thrush (2005)