The PW200 Turboshaft
By Serge Richer
In the late 1980’s, Pratt & Whitney Canada (P&WC) launched the PW200, a new line of helicopter turboshaft engines. Today, the PW200 series engine is installed in a broad range of new generation, light-twin helicopters around the world, including the Agusta A109E, Bell Model 427, Eurocopter EC135P1, MDHI MD Explorer and the Kazan Ansat.
The PW206 engine family, rated at a nominal 600 shaft horsepower (shp), evolved to the PW207 engine, with a nominal 700 shp. The PW207 engine delivers increased power through component and material improvements and by incorporating compressor treatment for enhanced handling capability. In addition, its digital engine control system incorporates built-in engine condition monitoring system (ECMS) functions. P&WC certified the first helicopter with a full authority digital engine control (FADEC) system on the PT6B-36, and with the first integrated ECMS on the PW207. This evolution is in keeping with the maintenance approach promoted by P&WC, and used effectively over the years, which is based on a philosophy of preventive maintenance through engine condition monitoring.
Principles of operation
The PW207 is a lightweight, free turbine turboshaft engine incorporating a single stage centrifugal compressor driven by a single stage turbine. A single stage power turbine drives a reduction gearbox comprised of two stage helical gear train (6,000 RPM).
Metered fuel from the fuel control is sprayed into a reverse flow annular combustion chamber through 12 individual fuel nozzles mounted around the gas generator case. A high voltage ignition unit and dual spark igniters start combustion. A single channel, digital electronic control system, with a mechanical backup fuel control, ensures accurate control of engine output speed and fast response to changes in power demand. An electrical torque motor located within the fuel metering module (FMM) works in conjunction with the electronic engine control (EEC).
Inlet air enters the engine through a radial inlet plenum chamber, formed by the compressor inlet case, and is directed rearward to the centrifugal impeller. High-pressure air from the impeller passes through diffuser tubes, which turn the air 90 degrees in direction and convert velocity to static pressure. This high-pressure air surrounds the combustion chamber liner.
The combustion chamber liner consists of two annular weldments bolted together at the rear dome-shaped end and an outer liner, which incorporates an integral large exit duct. The liner assembly has perforations of various sizes that allow entry of compressor delivery air. The flow of air changes direction 180 degrees as it enters and mixes with fuel.
Fuel is injected into the combustion chamber liner through 12 fuel nozzles arranged evenly around the gas generator case for ease of starting. The nozzles are supplied with fuel by a flow divider and fuel manifolds. The air/fuel mixture is ignited and the resultant gases then expand from the combustion chamber, reverse direction in the exit duct zone, and pass through the compressor turbine stator vanes to the single stage compressor turbine.
On to the turbines
The stator vanes ensure that the expanding gases impinge on the turbine blades at the correct angle, with a minimum loss of energy.
The still-expanding gases continue rearward to the power turbine stator and turbine. The exhaust gas from the power turbine is then directed through an annular exhaust plenum into the atmosphere.
The power turbine drives the output shaft through a two-stage reduction gearbox located at the front of the engine. The gearbox incorporates a phase shift torque meter device with sensors, which provides inputs to the EEC for control and an accurate cockpit indication of engine torque.
All engine-driven accessories are mounted on the reduction gearbox. A gas generator (Ng or N1) speed sensor is mounted on the upper left section of the reduction gearbox front housing. This sensor passes downward through the housing and exits directly over the starter generator drive spur gear. When the spur gear teeth pass by the probe, they create impulses, which are transmitted to the EEC. The operation of the power turbine (Nf, or N2) sensor is the same as the NI speed sensor, but it receives its signal from the torque shaft.
Fuel is supplied by an engine-driven fuel pump and its flow to the fuel manifold is controlled by the EEC, fuel control, and a fuel flow divider. The EEC controls fuel flow through inputs to the fuel-metering module (FMM). The FMM provides a full capability manual backup in the event of EEC degradation.
A single channel, digital electronic control system, with a mechanical backup fuel control, ensures accurate control of engine output speed and fast response to changes in power demand. An electrical torque motor located within the fuel metering module (FMM) works in conjunction with the electronic engine control (EEC).
The engine oil supply is contained in an integral oil tank with a total capacity of 1.35 U.S. gallons (5.12 litres) and includes a filler, dipstick and oil level sight glass.
ECMS: Designed for light-twins
The ECMS system is designed to meet the requirements of light-twin engine helicopters where great emphasis has been placed on size, weight and cost. To achieve this goal, a data collection unit (DCU) was incorporated in the engine bill of material to replace the analog temperature and torque trim units. The EEC and the DCU are used to monitor usage and the condition of the engine by monitoring selected parameters and storing necessary information. A portable ground based computer, which uses P&WC ground based software (GBS-PWC™), processes the maintenance data. The ECMS use the same information available to the maintenance staff and the operating crew, thereby ensuring that the helicopter’s "dispatchability" will not be affected. In addition, the maintenance manual contains backup instructions in case the ECMS is unserviceable.
The EEC provides for input signal conditioning, computations, data storage, and communication with ground based computer systems for maintenance trend monitoring. The EEC monitors information and stores it in the engine-mounted DCU memory to provide low cycle fatigue measurement, engine parameter exceedance monitoring, control system fault logging, and event logging.
Low cycle fatigue measurement is provided for the compressor impeller, compressor turbine and power turbine disks. Throughout the life of the engine, these three rotating components accumulate cycle counts for all three rotating components. Total cycle counts are stored in the DCU for retrieval by maintenance personnel equipped with appropriate ground support computers. This information is also available on the helicopter display systems. This automatic low cycle fatigue measurement offers a more accurate assessment of component usage, leading to more reliable life calculation. It provides a reduction in operating costs for the majority of operators flying missions lighter than those used by P&WC for rotor analysis system. It is also very convenient for the operating crew since the manual recording of engine operating parameters, such as the number of flights, number of engine starts and excursions to contingency power, is no longer required.
The EEC provides exceedance monitoring to calculate the exact operating time in each time-limited power band. This function is always active for the following engine parameters: Ng, measure gas temperature, engine torque, and Nf. The EEC monitors and records the time spent in the different rating bands and it indicates excursion as well as impending and actual time limit expiration.
In order to perform post-flight analysis of excursions, the EEC provides excursion-logging functions. These functions are always active and record a snapshot of various engine and control system parameters at the moment any exceedance threshold is crossed.
These snapshots, or data frames, are recorded in a "circular buffer" in the DCU and are stamped with an identification code indicating the parameter(s) and the threshold(s) crossed.
In order to facilitate fault investigation and retrieval of fault information, the control system maintains two independent "circular buffers" for faults in the EEC, EEPROM, and in the DCU.
When a fault occurs, a snapshot of specified EEC parameters is taken and saved in a fault data frame in both "circular buffers." A fault code is also included in order to indicate the triggering fault.
To facilitate flight investigation, the control system maintains two independent "circular buffers" for several types of events in the EEC, EEPROM, and in the DCU. The events are defined as follows:
• Peak values of NG, MGT, Torque and Npt during flight that are above specified thresholds
• Commanded reversion to Manual or Auto mode
• Detection of unexpected flameout
• Peak MGT value during start
• Detection of EEC/DCU mismatch
• New/reprogrammed DCU configuration
• EEC Power Supply Reset with engine running
When an event occurs and all the relevant criteria are met, a snapshot of specified EEC parameters is taken and saved in an event data frame in both "circular buffers." An event code is also included in order to indicate the triggering event.
The DCU is the electronic storage device, which is attached to the engine and remains with the engine at all times. Its primary purpose is to provide storage for engine trim data, LCF counting data and exceedance excursion counters. It also provides features, that may be expanded without modification to the DCU, to facilitate storage for other engine specific information, and to maintain a history of the engine to which it is attached. Pratt & Whitney Canada supplies its GBS-PWC™ ground based software, which uses this data to perform the following functions:
• Record of Fault Code data for quick and simple troubleshooting
• Download of exceedance data
• Download/upload running times and cycles of engines
• Analysis tools to display and plot data
• Real time fault and engine data display and recording
Exceedance recording, event logging and control system fault logging ensures that correct maintenance action can be taken following any exceedances of engine maintenance manual limitations. This limits unnecessary removals based on imprecise information. Control system fault logging reduces troubleshooting time by providing system fault code recording and display, along with automatic engine parameter recording when an EEC fault is detected.
Preventative maintenance philosophy
P&WC’s proven approach to maintenance is based on a philosophy of preventive maintenance through engine monitoring and inspection. To assist this approach, the PW200 engine is equipped with diagnostic features such as an integrated engine condition monitoring system, a chip detector, oil and fuel filter impending bypass indicators, and convenient borescope access provisions, which permit compressor and hot-section inspection without engine disassembly.
Scheduled PW200 line maintenance requirements were determined following a Maintenance Steering Group (MSG-III) style analysis in conjunction with the respective helicopter manufacturers. The Maintenance Manual incorporates the results of this analysis.
Unscheduled maintenance activity is prompted by a detected fault condition in the engine or the engine control system. A fault condition will be detected by the engine control fault detection system, the Helicopter Engine Condition Trend Monitoring (HECTM™), and/or by scheduled maintenance activity. The troubleshooting procedures contained in the Maintenance Manual were verified and expanded during the flight test program to assist in fault isolation. These procedures will evolve with continued field experience.
By monitoring engine performance parameters (i.e. Ng & Nf), the HECTM™ program will enable operators to schedule a hot-section inspection (HSI). The PW200 design features modularity and accessible hot-section components for improved HSI and shop maintenance capability.
The types of heavy maintenance, which can be performed in the field, include:
• Removal/replacement of gearbox output and accessory drive seals
• Removal/replacement of either gearbox or turbomachine modules
• Inspection and replacement of hot-section components
A program is currently in place for the modular repair and overhaul of PW200 engines. The engine is separated into gearbox and turbomachine modules. A modular logbook ensures traceability of modules and components.
The engine operates to a 3,000-hour overhaul interval (3,500-hour for mature engines) with a minimum of required maintenance. The compressor impeller and two turbine disks feature a generous life limit of 15,000 cycles for the compressor impeller, and the power turbine disk and the CT disk life is 10,000 cycles. This declared life limit, combined with the fractional cycle counting method performed by the EEC, provides for two overhaul intervals in typical commercial service.
Pratt & Whitney Canada provides tuition-free training to operators and their local service facilities to allow operators to handle PW200 line maintenance, hot-section inspections and all other maintenance between overhauls. Trained P&WC Service Centre specialists are available to provide on-site assistance, upon request.