72 Engine Turboprop

 

TTP	B1-L3	ATA 72
Beech 90 Series	B2-L1	Engine Turbo Prop

Engine Turbine/Turboprop

00   General

10   Reduction Gear and Shaft Section

20   Air Inlet Section

30   Compressor Section

40   Combustion Section

50   Turbine Section

60   Accessory Drives

70   Bypass Section

The engine on the King Air C90A/B/GT aircraft is the PT6A-21 and is rated at 550-shaft horsepower (SHP). The engine on the C90GT/GTi aircraft is a PT6A-135A and is also rated to 550.

 

The nomenclature stands for: P—Propeller

 

T—Turbine

 

6—6th design series

 

A—Two-stage reduction gearbox

 

21 or 135A—The sequence number in the design series

 

The PT6 was developed in 1958 for use in fixed- and rotary-wing aircraft. The first certificated engine (PT6A-6 used on the early model King Air 90s) entered service in 1962 and was rated at

450-shaft horsepower. Since then, the output of the PT6A engine has more than tripled with little outward changes. To date, the largest PT6A is the –68 and is rated at 1,600 gross shaft horsepower.

 

A unique feature of the PT6A engine is that it is a reverse-flow engine. Internally, the airflow makes two 180° turns during passage through the engine. The air enters the inlet at the rear of the engine and flows forward through the compressor, discharges into the diffuser case, and makes a 180° turn into the combustion chamber, then another 180° turn occurs as it leaves the combustion chamber. The heat-charged air then passes through both the compressor turbine and the power turbine assemblies and exits through the exhaust ducts at the front of the engine to the atmosphere. This design results in a compact, lightweight, and relatively powerful engine.

Description and Operation 

The PT6A Series power plant is a lightweight free turbine engine. The 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. 

    The 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 compressor consists of three axial stages combined with a single centrifugal stage, assembled as an integral unit. 

     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 converts velocity to static pressure. The diffused air then passes through straightening vanes to the annulus surrounding the combustion chamber liner and the gas generator case. 

      The combustion chamber liner consists of an annular weldment having 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 ducted through another 180° turn, and the gas flow is 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 arranged for ease of starting. Fuel is supplied by a dual manifold consisting of primary and secondary transfer tubes and adapters. 

During the starting sequence, fuel only flows through the primary nozzles, but as the engine rpm increases through approximately 35% N1, the secondary nozzles flow fuel. This sequenced fuel flow makes for a smooth and even-temperature engine start. Two spark igniters that protrude into the liner, initially ignite the fuel/air mixture. Once combustion begins, it is self-sustaining and the igniters are turned off. The resultant combustion gases expand and flow from the liner, reverse direction in the exit duct zone, and pass through the compressor turbine inlet guide vanes to the compressor turbine. The guide vanes ensure the expanding gases impinge on the compressor turbine blades at the correct angle, with minimum loss of energy. The still expanding gases are then directed forward to drive the power turbine. The power turbine section has no mechanical connection to the compressor section.

     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 directed through an annular exhaust plenum to atmosphere via twin opposed exhaust ports provided in the exhaust duct. 

    Interturbine temperature (T5) is monitored by a cold junction thermocouple system comprising a bus-bar, probes and harness assembly installed between the compressor and power turbines with the probes projecting into the gas path. A terminal block mounted in the gas generator case provides a connection point to cockpit instrumentation and to a T5 trim thermocouple mounted externally in the air inlet zone.

PTP  Beech 90 Series	B1	LOC
	B2