LM6000 Gas Turbine - Generator Package Familiarization Training

LM6000 Gas Turbine - GeneratorPackage Familiarization Training

Lotte PPTAPakistan

2012

GE Energy

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All rights reserved by the General Electric Company. No copies permitted without the prior written consent of the General Electric Company.

The text and the classroom instruction offered with it are designed to acquaint students with generally accepted good practice for the operation or main-tenance of equipment and/or systems.

They do not purport to be complete nor are they intended to be specific for the products of any manufacturer, including those of the General Electric Company; and the Company will not accept any liability whatsoever for the work undertaken on the basis of the text or classroom instruction. The manufacturers operating and maintenance specifi-cations are the only reliable guide in any specific instance; and where they are not complete, the manufacturer should be consulted.

The materials contained in this document are intended for educational purposes only. This document does not establish specifications, operating procedures or maintenance methods for any of the products referenced. Always refer to the official written materials (labeling) provided with the product for specifications, operating procedures and maintenance requirements.

Proprietary Training Material Property of GE. Use of these materials is limited to agents and GE employees, or other parties expressly licensed by GE. Unlicensed use is strictly prohibited.

2012 General Electric Company

LM6000 Gas Turbine Generator Package Familiarization Training Lotte PPTA, Pakistan 1

GE Energy

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LM6000 Gas Turbine - Generator Package Familiarization Training

Lotte PPTA Pakistan

2012

TAB Title Module

1 BOC-FAM Course Introduction F-000-00-00-000-00

2 Turbine Basics F-000-00-10-000-00

3 Construction and Operation F-060-00-10-000-00

4 Turbine Support Systems F-060-00-20-000-00

5 Turbine Lube Oil System (Woodward Control) F-060-00-20-100-00

6 Variable Geometry System (Woodward Control) F-060-00-20-200-00

7 Start System (Woodward Control) F-060-00-20-050-00

8 Gas Fuel System, DLE F-060-00-20-301-02

9 Ventilation & Combustion Air System (Woodward Control) F-060-00-20-401-00

10 Water Wash System F-060-00-20-500-00

11 Vibration Monitoring System (Bently Nevada 3500) F-060-00-20-700-00

12 Fire Protection System F-060-00-20-800-00

13 Basic Electricity and Generation F-000-00-60-001-00

14 50 HZ Generator Construction F-000-00-30-100-01

15 50HZ Generator Lube Oil System F-060-00-30-300-01

16 Control System (Woodward Control) F-060-00-40-100-00

17 Sequences F-060-00-50-000-00

18 Appendix F-000-00-60-002-00

19 Reference Drawings

LM6000 Gas Turbine Generator Package Familiarization Training Lotte PPTA, Pakistan 2

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A Mechanical Flow and Instrument Drawings

F&ID Symbols 7232796-571231A Hydraulic Start System 7232796-571232A Ventilation and Combustion Air System 7232796-571239B Turbine Lube Oil System 7232796-571244A Fuel System 7232796-571245A Turbine HYD Sys 7232796-571247A Generator Lube Oil System 7232796-571248A Water Wash System 7232796-571262B Sprint Main 7232796-571268A Sprint Skid 7232796-571270A Auxiliary Systems 7232796-571272B Fire Protection System GA-1177

B Electrical Drawings

Elect Sym 7232796-730005A Junction Box 7232796-730012A TCP Plan and Elevation 7232796-730014A One-Line Drawing 7232796-730031C Control System Worksheet 7232796-730146B Cause & Effect 7232796-730149A

C General Arrangement Drawings

Main skid 7232796-571200B Sprint Skid 7232796-571209A Aux skid 7232796-571218A Lube oil skid 7232796-571221A

GE Aero Package Training Course Introduction

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F-000-00-00-000-00 BOC/FAM Course Introduction


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This document is intended for training use only. It is not intended to cover all possible variations in equipment or to provide for specific problems that may arise.

Technical drawings and descriptions herein are intended to illustrate conceptual examples and do not necessarily represent as-supplied system details. System users are advised to refer to drawings of current release when conducting troubleshooting, maintenance procedures, or other activities requiring system information.

GE Aero Energy Products advises that all plant personnel read this training manual and the Operation & Maintenance Manual to become familiar with the generator package, auxiliary equipment and operation.

This manual is not a replacement for experience and judgment. The final responsibility for proper, safe operation of the generator package lies with the Owners and Operators. Operation and performance of auxiliary equipment and controls not furnished by GE is the sole responsibility of the Owners and Operators.

Reproduction of this guide in whole or in part without written permission is prohibited.


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Course Objectives

This training course is designed to provide system operators with :

Understanding of basic Gas Turbine and Generator operationUnderstanding of how each of the sub systems operates, individually and as part of the total packageAbility to initiate and maintain normal system operationAbility to recognize system alarm and fault information and take appropriate actionUnderstanding of system documentationKnowledge of serviceable components and maintenance required for normal operation

This course should be considered a mandatory prerequisite for more advanced training in package mechanical maintenance or control system maintenance and troubleshooting.


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OVERVIEW OF GE ENERGY PRODUCTS

GE Energy is a leading supplier of diesel and aero-derivative gas turbine packages for industrial and marine applications, with many units operating throughout the world.

GE Energy takes single source responsibility for the total equipment package and provides field service for the equipment once it has been installed.

All of GE Energys skill and field experience is built into each unit. Customers needs are met with standardized designs, which have been proven time and time again in tropical heat, desert sand and arctic cold.

For a customer with special requirements, GE Energy adds features from a list of pre-engineered options.

GE Energy provides job-site supervision and operator training, offers total plant operation and maintenance when desired, and backs up each unit with a multi-million dollar inventory of turbine parts, as well as a service department with trained personnel ready to perform field service anywhere in the world 24 hours a day, 365 days a year.

Meeting customers requirements for quality, dependability and outstanding service is the commitment of GE Energy.


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SAFETY CONSIDERATIONS

The following are general safety precautions that are not related to any specific procedures and do not appear elsewhere in this manual. Personnel must understand and apply these precautions during all phases of operation and maintenance.

Health Hazards

Use all cleaning solvents, fuels, oil adhesives, epoxies, and catalysts in a well-ventilated area. Avoid frequent and prolonged inhalation of fumes. Concentrations of fumes of many cleaners, adhesives, and esters are toxic and cause serious adverse health effects, and possible death, if inhaled frequently. Wear protective gloves and wash thoroughly with soap and water as soon as possible after exposure to such materials. Take special precautions to prevent materials from entering the eyes. If exposed, rinse the eyes in an eyebath fountain immediately and report to a physician. Avoid spilling solvents on the skid. Review the hazard information on the appropriate Material Safety Data Sheet and follow all applicable personal protection requirements.

Environmental Hazards

The disposal of many cleaning solvents, fuels, oils, adhesives, epoxies, and catalysts is regulated and, if mismanaged, could cause environmental damage. Review Material Safety Data Sheets, product bulletin information, and applicable local, state and federal disposal requirements for proper waste management practices.


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Fire Hazards

Keep all cleaning solvents, oils, esters and adhesives away from exposed-element electric heaters, sparks or flame. Do not smoke when using flammable materials, in the vicinity of flammable materials, or in areas where flammable materials are stored. Provide adequate ventilation to disperse concentrations of potentially explosive fumes or vapors. Provide approved containers for bulk storage of flammable materials, and approved dispensers in the working areas. Keep all containers tightly closed when not in use.

Electrical Hazards

Use extreme care when working with electricity. Electricity can cause shock, burns or death. Electrical power must be off before connecting or disconnecting electrical connectors. Lethal output voltages are generated by the ignition exciter. Do not energize the exciter unless the output connection is properly isolated. Be sure all leads are connected and the plug is installed. All personnel should be cleared to at least 5 feet before firing the exciter.

Compressed Air Hazards

Air pressure used in work areas for cleaning or drying operations shall be regulated to 29 psi or less. Use approved personal protective equipment (goggles or face shield) to prevent injury to the eyes. Do not direct the jet of compressed air at yourself or other personnel so that refuse is blown onto adjacent work stations. If additional air pressure is required to dislodge foreign materials from parts, ensure that approved personal protective equipment is worn, and move to an isolated area. Be sure that the increased air pressure is not detrimental or damaging to the parts before applying high-pressure jets of air.


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Procedural Hazards

Observe all specified and logical safety practices when assembling or disassembling the engine. Wear safety glasses or other appropriate eye protection at all times. Do not allow safety wire or wire clippings to fly from the cutter when removing or installing wire. Do not use fingers as guides when installing parts or checking alignment of holes. Use only correct tools and fixtures. Avoid shortcuts, such as using fewer-than-recommended attaching bolts or inferior-grade bolts. Heed all warnings in this manual and in all vendor manuals, to avoid injury to personnel or damage to gas turbine parts.

Tooling Hazards

Improperly maintained tools and support equipment can be dangerous to personnel, and can damage gas turbine parts. Observe recommended inspection schedules to avoid unanticipated failures. Use tooling only for its designed purpose and avoid abuse. Be constantly alert for damaged equipment, and initiate appropriate action for approved repair immediately.


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Gas Turbine Operational Hazards

The outside surfaces of the engine are not insulated; therefore, adequate precautions shall be taken to prevent operating personnel from inadvertently coming into contact with these hot surfaces.

The gas turbine is a source of considerable noise. It is necessary for personnel working on the gas turbine or in its vicinity to wear proper ear protection equipment when it is operating.

The gas turbine is a high-speed machine. In case of component failure, the skid housing would contain compressor and turbine blade failures, but might not contain major compressor or turbine disk failures. Operating personnel shall not be permanently stationed in or near the plane of the rotating parts.

Low-pressure, high-velocity airflow created by the compressor can draw objects or personnel into the engine. Although an inlet screen is used, personnel should not stand in front of the inlet while the engine is operating.

When entering the gas turbine enclosure, the following requirements must be met:

The gas turbine will be shut down or limited to core idle power.

The fire extinguishing system will be made inactive.

The enclosure door shall be kept open. If the gas turbine is operating, an observer shall be stationed at the enclosure door, and confined space entry procedures will be followed.

Avoid contact with hot parts, and wear thermally insulated gloves, as necessary.

Hearing protection (double) will be worn if the gas turbine is operating.

Do not remain in the plane of rotation of the starter when motoring the gas turbine.

When performing maintenance on electrical components, turn off electrical power to those components, except when power is required to take voltage measurements. Lock out all controls and switches, if possible; otherwise, tag electrical switches Out of Service to prevent inadvertent activation. Tag the engine operating controls Do Not Operate to prevent the unit from being started during a shutdown condition.


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Cleanliness and FOD/DOD

FOD/DOD (foreign object damage/domestic object damage) is the single major cause of premature gas turbine failure. Prevention is the only practical means of protecting against FOD, and adherence to the following guidelines cannot be over-emphasized.

Empty pockets of all lose objects.

Keep maintenance area clean and organized.

Keep FOD containers in the work area to receive bits of safety wire, used gaskets, O-rings and other similar types of debris. USE THEM.

Do not use the gas turbine as a shelf to hold parts and tools during maintenance.

Install protective covers and caps on all exposed openings during maintenance.

Remove protective caps and covers only when required to install a part or make a connection.

After protective caps and covers are removed, inspect all openings and cavities for foreign objects and cleanliness.

After maintenance, thoroughly clean and inspect work area. Account for all tools, parts, and materials used during maintenance.

GE Energy g Turbine Basics

F-000-00-10-000-00 Turbine Basics

TURBINE BASICS



OVERVIEW

The major components of the engine are a compressor section, combustion section, and a turbine. The turbine is mechanically coupled and drives the compressor by a drive shaft.

The compressor, combustor, and turbine are called the core of the engine, since all gas turbines have these components. The core is also referred to as the gas generator (GG) since the output of the core is hot exhaust gas.

The gas is passed through an exhaust duct to atmosphere. On some types of applications, the exhaust gas is used to drive an additional turbine called the power turbine which is connected to a piece of driven equipment (i.e. generators, pumps, process compressors, etc).

Because of their high power output and high thermal efficiency, gas turbine engines are also used in a wide variety of applications not related to the aircraft industry. Connecting the main shaft (or power turbine) of the engine to an electro-magnet rotor will generate electrical power. Gas turbines can also be used to power ships, trucks and military tanks. In these applications, the main shaft is connected to a gear box.



TURBINE BASICS

The balloon drawings above illustrate the basic principles upon which gas turbine engines operate.

Compressed inside a balloon, as in (A) above, exerts force upon the confines of the balloon. Air, which has weight andoccupies space, by definition, has mass. The mass of the air is proportional to its density, and density is proportional totemperature and pressure. The air mass confined inside the balloon, accelerates from the balloon, creating a force as it isreleased (B). This force increases as mass and acceleration increase, as stated in Newtons second law; force equals masstimes acceleration (F = MA).

The force created by the acceleration of the air mass inside the balloon results in an equal and opposite force that causesthe balloon to be propelled in the opposite direction, as stated in Newtons third law (for every action, there is an equal andopposite reaction). Replacing the air inside the balloon, as in (C) sustains the force and, although impractical, allows a loadto be driven by the force of the air mass accelerating across and driving a turbine, as in (D).

In (E) a more practical means of sustaining the force of an accelerating air mass used to drive a load is illustrated. Ahousing contains a fixed volume of air, which is compressed by a motor driven compressor. Acceleration of the compressedair from the housing drives a turbine that is connected to the load.

In (F) fuel is injected between the compressor and the turbine to further accelerate the air mass, thus multiplying the forceused to drive the load.

In (G) the motor is removed and the compressor is powered by a portion of the combustion gas, thus making the engineself-sufficient as long as fuel is provided.

In (H) a typical gas turbine-engine operation is represented. Intake air is compressed, mixed with fuel and ignited. The hotgas is expanded across a turbine to provide mechanical power and exhausted to atmosphere.



Gas Turbine Operation Vs.Reciprocating Engine Operation



COMPRESSION COMBUSTION EXPANSION EXHAUST

Four processes occur in gas turbine engines, as illustrated above. These processes, first described by George Brayton and called the Brayton cycle, occur in all internal combustion engines. The Brayton steps are as follows:

Compression occurs between the intake and the outlet of the compressor (Line A-B). During this process, pressure and temperature of the air increases.

Combustion occurs in the combustion chamber where fuel and air are mixed to explosive proportions and ignited. The addition of heat causes a sharp increase in volume (Line BC).

Expansion occurs as hot gas accelerates from the combustion chamber. The gases at constant pressure and increased volume enter the turbine and expand through it. The sharp decrease in pressure and temperature (Line C-D).

Exhaust occurs at the engine exhaust stack with a large drop in volume and at a constant pressure (Line D-A).

The number of stages of compression and the arrangement of turbines that convert the energy of accelerating hot gas into mechanical energy are design variables. However, the basic operation of all gas turbines is the same.



CONVERGENT AND DIVERGENT DUCTS

Compressors in gas turbine engines use convergent and divergent ducts to generate the high pressures necessary to (a) provide a wall of pressure, preventing expanding hot gas from exiting through the engine inlet, as well as, through the exhaust; and (b) provide the proper ratio of air-to-fuel for efficient combustion and cooling of the combustion chamber.

Pressure decreases through convergent ducts and increases through divergent ducts, a phenomenon which is demonstrated in paint spray equipment. Compressed air, forced through a convergent duct, generates a lower pressure through the narrow section to draw in paint.

Expansion through a divergent section then increases pressure and air volume, dispersing the paint in an atomized mist.



INLET GUIDE VANES

Inlet guide vanes direct, or align, airflow into the first rotating blade section where velocity is increased by the addition ofenergy. The following stator vane section is divergent, providing an increase in static pressure and a decrease in air velocity. Airflow then enters the second stage at a higher initial velocity and pressure than at the inlet to the preceding stage. Each subsequent stage provides an incremental increase in velocity and static pressure until the desired level of pressure and velocity is reached.

Some compressor stator vanes are designed to move, changing their divergence, allowing regulation of compressor outlet pressure and velocity to achieve the proper ratio of air for fuel combustion and cooling versus engine speed and power output.



Axial Flow Compressor Centrifugal Flow Compressor



COMPRESSORS

Compressors in gas turbine engines use convergent and divergent ducts to generate the high pressures necessary to (a)provide a wall of pressure, preventing expanding hot gas from exiting through the engine inlet as well as through the exhaust; and (b) provide the proper ratio of air-to-fuel for efficient combustion and cooling of the combustion chamber.

Pressure decreases through convergent ducts and increases through divergent ducts, a phenomenon which is demonstrated in paint spray equipment. Compressed air, forced through a convergent duct, generates a lower pressure through the narrow section to draw in paint. Expansion through a divergent section then increases pressure and air volume, dispersing the paint in an atomized mist.

All turbine engines have a compressor to increase the pressure of the incoming air before it enters the combustor. Compressor performance has a large influence on total engine performance. There are two main types of compressors: axial and centrifugal.

In the illustration, the example on the left is called an axial compressor because the flow through the compressor travels parallel to the axis of rotation. An apparent contradiction in the operation of the axial-flow compressor is that high pressureis generated, although the overall divergent shape would appear to cause a lower output pressure. Output pressure is increased by divergence in each static inter-stage section. Rotating compressor blades between each static stage increases the velocity that is lost by injecting energy.

The compressor on the right is called a centrifugal compressor because the flow through this compressor is turned perpendicular to the axis of rotation. Centrifugal compressors, which were used in the first jet engines, are still used on small turbojets and turbo-shaft engines. Modern large turbojet, turbofan, and turbo-shaft engines usually use axial compressors.



COMPRESSOR STALL

A stall can happen within the compressor if the air moves from its general direction of motion (also known as the angle of attack). At this point, the low pressure on the upper surface disappears on the stator blade. This phenomenon is known as astall. As pressure is lost on the upper surface, turbulence created on the backside of the stator blade forms a wall that willlead into the stall. Stall can be provoked if the surface of the compressor blade is not completely even or smooth. A dent in the blade, or a small piece of material on it, can be enough to start the turbulence on the backside of the blade, even if the angle of attack is fairly small. Each stage of compression should develop the same pressure ratio as all other stages. Whena stall occurs, the front stages supply too much air for the rear stages to handle, and the rear stage will choke.

High Angle of AttackIf the angle of attack is too high, the compressor will stall. The airflow over the upper airfoil surface will become turbulent and

destroy the pressure zone. This will decrease the compression airflow. Any action that decreases airflow relative to engine speed will increase the angle of attack and increases the tendency to stall.

Low Angle of AttackIf there is a decrease in the engine speed, the compression ratio will decrease with the lower rotor velocities. With a

decrease in compression, the volume of air in the rear of the compressor will be greater. This excess volume of air causes a choking action in the rear of the compressor with a decrease in airflow. This in turn decreases the air velocity in the front ofthe compressor and increases the tendency to stall.



Can Type Combustor Annular Type Combustor



COMBUSTORS

All turbine engines have a combustor, in which the fuel is combined with high pressure air and burned. The resulting high temperature exhaust gas is used to turn the turbine and produce thrust when passed through a nozzle.

The combustor is located between the compressor and the turbine. The combustor is arranged like an annulus, or a doughnut, as shown by illustrations above. The central shaft that connects the turbine and compressor passes through the center hole. Combustors are made from materials that can withstand the high temperatures of combustion. The liner is often perforated to enhance mixing of the fuel and air.

There are three main types of combustors, and all three designs are found in gas turbines:

The combustor at the right is an annular combustor with the liner sitting inside the outer casing which has been peeled open in the drawing. Many modern combustors have an annular design.

The combustor on the left is an older can or tubular design. Each can has both a liner and a casing, and the cans are arranged around the central shaft.

A compromise design (not shown) is a can-annular design, in which the casing is annular and the liner is can-shaped. The advantage to the can-annular design is that the individual cans are more easily designed, tested, and serviced.

Turbine blades exist in a much more hostile environment than compressor blades. Located just downstream of the combustor, turbine blades experience flow temperatures of more than a thousand degrees Fahrenheit. Turbine blades must be made of special materials that can withstand the heat, or they must be actively cooled. In active cooling, the nozzles andblades are hollow and cooled by air which is bled off the compressor. The cooling air flows through the blade and out through the small holes on the surface to keep the surface cool.



FLAME-STABILIZING AND GENERAL-FLOW PATTERNS

The flame stabilizing and general-flow patterns are illustrated above for a typical can-type combustion chamber. Although modern engines use one continuous annular combustion chamber, the can-type simplifies illustration of the cooling and combustion techniques used in all combustion chambers.

The temperature of the flame illustrated in the center of the combustor is approximately 3200F at its tip when the engine is operating at full load. Metals used in combustion chamber construction are not capable of withstanding temperatures in this range; therefore, the design provides airflow passages between the inner and the outer walls of the chamber for cooling and flame shaping.

Air flowing into the inner chamber is directed through small holes to shape the flame centering it within the chamber, to prevent its contact with the chamber walls. Approximately 82% of the airflow into combustion chambers is used for cooling and flame shaping; only 18% is used for fuel combustion. Regulation of fuel flow determines engine speed. Stator vane control in the compressor controls pressure and velocity into the combustion chamber as a function of compressor speed.



TURBINE

All gas turbine engines have a turbine located downstream of the combustor to extract energy from the hot flow and turn the compressor. Work is done on the turbine by the hot exhaust flow from the combustor.

Since the turbine extracts energy from the flow, the pressure decreases across the turbine. The pressure gradient helps keep the boundary layer flow attached to the surface of the turbine blades. Since the boundary layer is less likely to separate on a turbine blade than on a compressor blade, the pressure drop across a single turbine stage can be much greater than the pressure increase across a corresponding compressor stage. A single turbine stage can be used to drive multiple compressor stages. Because of the high pressure change across the turbine, the flow tends to leak around the tips of the blades. The tips of turbine blades are often connected by a thin metal band to keep the flow from leaking.

Turbine blades exist in a much more hostile environment than compressor blades. Sitting just downstream of the combustor, the blades experience flow temperatures of more than a thousand degrees Fahrenheit. Turbine blades must be made of special materials that can withstand the heat, or they must be actively cooled. In active cooling, the nozzles and blades are hollow and cooled by air which is bled off the compressor. The cooling air flows through the blade and out through the small holes on the surface to keep the surface cool.



TURBINE (Continued)

The compressor drive turbine is an impulse reaction-type designed for maximum efficiency in converting hot-gas flow into rotational mechanical energy. A first-stage fixed nozzle directs flow into the first-stage of rotating blades. The impulse of expanding hot gas upon the lower surface of each rotating blade propels motion in the upward direction.

Hot gas flow above the following blade creates a lower pressure above the blade as above an aircraft wing, causing additional rotational force. Subsequent stages operate identically, multiplying the rotational force. Compressor and load-driving turbines consist of a varying number of stages, depending upon the load being driven and other design considerations.



.

Single Shaft Twin Shaft

Concentric Shaft with Power Turbine

Concentric Shaft



TURBINE SHAFTS

The figure above shows the standard gas turbine shaft arrangements. Single shaft illustration is the traditional single shaftassembly. It consists of the axial flow compressor; Turbine and Power Turbine are all mechanically linked. If we add to this shaft the generator and gearbox, we have a shaft system with a high moment of inertia. This is the favored configuration forelectrical generation because this provides additional speed (Frequency) stability of the electrical current during large load fluctuations. This configuration is typical of heavy-duty industrial frame turbines, such as the MS7001.

The twin shaft illustration shows the standard two shaft arrangement with the compressor and turbine only connected, and an unconnected power turbine and output shaft that will rotate independently. This configuration is favored for variable speed-drive packages, such as pumps and compressors, because the gas generator or gas producer can run at its own optimum speed for a given load. The LM2500 utilizes this configuration and has been applied to both electric power generation and a variety of mechanical drive applications.

Aircraft jet engines have for many years been adapted for industrial use as shown in the diagrams above. The concentric shaft illustration, above left, shows a more complicated aero-derivative industrial turbine arrangement. This, too, is still essentially a two shaft configuration but the gas generator core (an original jet-engine) was designed with two spools, a Low Pressure Shaft and a High Pressure Shaft. This engine configuration allows the load to be driven from either the exhaust end or the compressor air intake end. This is the configuration used by the LM6000

The concentric shaft with power turbine illustration is essentially a two shaft arrangement with a gas generator originally designed for propulsion. An independently rotating Power Turbine, manufactured especially to match the flow of the jet engine, is added to the gas path as the power/torque producer. This configuration is found in the LM1600 and the LMS100.



NOx CONTROLOxides of Nitrogen result from the thermal fixation of molecular nitrogen and oxygen in the combustion air. Its rate of formation is extremely sensitive to local flame temperature and, to a lesser extent, to local oxygen concentrations. Virtually all thermal NOx is formed in the region of the flame at the highest temperature. Maximum thermal NOx production occurs at a slightly lean fuel-to-air ratio due to the excess availability of oxygen for reaction within the hot flame zone. Control oflocal flame fuel-to-air ratio is critical in achieving reductions in thermal NOx.

Combustion ControlsReduction of Nox emissions are accomplished by: Injection of water or steam at the fuel nozzle in order to reduce combustion temperature Specially designed Dry Low Emissions (DLE) combustors and fuel systems

The injection of water or steam into the flame area of a turbine combustor provides a heat sink, which lowers the flame temperature and thereby reduces thermal NOx formation. Water or steam injection, also referred to as "wet controls," have been applied effectively to both aeroderivative and heavy duty gas turbines, and to all configurations. Reduction efficiencies of 70 to 85+ percent can be achieved with properly controlled water or steam injection, with NOx emissions generally higher for oil-fired turbines than for natural gas-fired units. The most important factor affecting reduction efficiency is the water-to-fuel ratio. In general, NOx reduction increases as the water-to-fuel ratio increases; however, increasing the ratio increases carbon monoxide and, to a lesser extent, hydrocarbon emissions at water-to-fuel ratios less than one. Further, energy efficiency of the turbine decreases with increasing water-to-fuel ratio.

Post-Combustion ControlsThe major type of post-combustion control used in gas turbines is Selective Catalytic Reduction (SCR). Applications use SCR to supplement reductions from steam or water injection, or combustion modifications. Carefully designed SCR systems can achieve NOx reduction efficiencies as high as 90 percent. The Selective Catalytic Reduction (SCR) process reduces NOx emissions by using ammonia in the presence of a catalyst. Vaporized ammonia is injected into the flue gas at the appropriate temperature. The ammonia functions, in the presence of the NOx removal catalyst, as a reducing agent to decompose nitrous oxides NOx in the flue gas into nitrogen gas and water vapor.

LM6000 Construction and Operation

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F-060-00-10-000-00 LM6000 Construction and Operation

LM6000 CONSTRUCTION and OPERATION


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3.1 ENGINE OVERVIEW

Developed from CF6-80C2 turbofan engine Liquid, Gas and Dual Fuel packages available Steam or Water Injection and Dry Low Emissions combustor systems available Most efficient simple-cycle gas turbine in class


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The General Electric LM6000 gas turbine is a stationary gas turbine that is derived from the family of CF6 jet engines. The aircraft version of the engine is called the CF6-80C2 turbofan engine and is used to drive several types of wide body commercial aircraft, including the Boeing 747-400.

The experience and technology of the CF6-80C2 and the well-proven LM2500 have been applied to the LM6000 to make it one of the best engines on the market today.

Although the LM6000 gas turbine was developed recently (first application in 1992), General Electric was one of the first developers of the aero-derivative (a gas turbine designed first as a flight engine, then redesigned for industrial use) with more than 30 million running hours. General Electric engines have an availability of 99.6% overall.

The LM (Land and Marine) series of gas turbines has the following gas turbines: LM500, LM1500, LM1600, LM2500, LM2500+, LM5000, LM6000 ranging in power output from 14 to 50 megawatts (MW).


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The following changes were made to convert the CF6-80C2 to the LM6000:

Front fan removed and inlet guide vanes added LP compressor from the CF6-50 / LM5000 used Front and rear frames adapted Output shafts added to the front of the LPC and the back of the LPT Bearing 7R added New industrial fuel system added Balancing disk added to the LPT Hydraulic control system for the variable geometry added

Since its introduction in 1992, the original LM6000PA was followed by introduction of the model PB, the dry low emissions (DLE) version.

In 1998, the PC model was introduced and incorporated design changes to the LPC, HPC, LPT, balance piston system and the fuel system. These design changes increased shaft power output by approximately 3.4 MW, and engine efficiency by approximately 2%.

The LM6000 PD is the LM6000 PC modified with the Dry Low Emission Combustion System (DLE). This model made its appearance in mid-1998. DLE system requires changes to be made to the fuel nozzles and the annular combustion chamber.


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Dual Rotor, concentric drive-shaft design

Hot or Cold end drive configurations

5-stage low-pressure compressor (LPC), 2.4:1 compression ratio

14-stage variable-geometry high-pressure compressor (HPC), 12:1 compression ratio

Variable Inlet Guide Vanes (Optional), Variable Bleed Valves and Variable Stator Vanes

2-stage high-pressure turbine (HPT)

5-stage low-pressure turbine (LPT)


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The LM6000 gas turbine is a dual-rotor, concentric drive-shaft, gas turbinecapable of driving a load from the front and/or rear of the low-pressure (LP)rotor. The main components consist of a variable inlet guide vane (VIGV)assembly or inlet frame assembly, a 5-stage low-pressure compressor (LPC), a14-stage variable-geometry high-pressure compressor (HPC), an annularcombustor, a 2-stage high-pressure turbine (HPT), a 5-stage low-pressureturbine (LPT), an accessory gearbox (AGB) assembly, and accessories.

The LP rotor consists of the LPC and the LPT that drives it. Attachment flanges are provided on both the front and the rear of the LP rotor for connection to the packager-supplied power shaft and load. The high-pressure rotor consists of the 14-stage HPC and the 2-stage HPT that drives it. The high-pressure (HP) core consists of the HPC, the combustor, and the HPT. The high- and low-pressure turbines drive the high- and low-pressure compressors through concentric drive shafts.


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Air enters the gas turbine at the IGV/VIGVs and passes into the LPC. The LPCcompresses the air by a ratio of approximately 2.4:1. Air leaving the LPC is directedinto the HPC. Variable bypass valves (VBVs) are arranged in the flow passagebetween the two compressors to regulate the airflow entering the HPC at idle and atlow power. To further control the airflow, the HPC is equipped with variable statorvanes (VSVs).

The HPC compresses the air to a ratio of approximately 12:1, resulting in a total compression ratio of 30:1, relative to ambient. From the HPC, the air is directed into the single annular combustor section, where it mixes with the fuel from 30 fuel nozzles. An igniter initially ignites the fuel-air mixture then, once combustion is self-sustaining, the igniter is turned off. The hot gas that results from combustion is directed into the HPT that drives the HPC. This gas further expands through the LPT, which drives the LPC and the output load.


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3.2 ENGINE STATIONS


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As in the aircraft industry, determine the left and right of the engine by looking intothe air flow or upstream. From this vantage point specific areas can be describedusing their clock hour positions, such as 3 oclock for the right side and 9oclock for the left side, etc.

Various signals measured on the LM6000 gas turbine are called after the so calledengine stations, which are engine locations, numbered in the direction of airflow,from 0 to 8. Station 0 (zero) is the LP compressor inlet; station 8 is the powerturbine exhaust. Typical signal names refer to the stations. Station numbers maybe subdivided, using alphabetical character or a decimal as a suffix.


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Complete list of LM6000 stations:

1 VIGV inlet2 LPC inlet2.3 LPC discharge2.4 LPC bleed2.5 HPC inlet2.6 HPC bleed 7th stage2.7 HPC bleed 8th stage2.8 HPC bleed 11th stage3 HPC discharge3.6 Fuel nozzle4 HPT inlet (nozzle)4.1 HPT 1st stage blade4.2 HPT exhaust4.8 LPT inlet5 LPT exhaust5.5 LPT rear frame exhaust5.6 LPT exhaust diffuser

Items in bold denote engine instrumentation locations.


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3.3 BEARINGS AND SUMPS

Roller bearings take radial loadsBall bearings take radial and axial (thrust) loads

Each rotating system uses one ball bearing The LP system uses the 1B bearing for axial positionThe HP system uses the 4B bearing for axial position


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Sump A houses the No. 1B, No. 2R, and No. 3R bearings. The No. 1B bearing is a ball-type thrust bearing that carries the thrust loads for the LP rotor (LPC and LPT). The No.2R bearing supports the low-pressure compressor rotor (LPCR) and the No. 3R bearingsupports the high-pressure compressor rotor (HPCR) forward shaft.

The B and C sump houses the No. 4R bearing, the No. 4B bearing and the No. 5R bearing. The No. 4R bearing supports the aft shaft of the HPCR. The No. 4B bearing carries the thrust loads for the HPR (HPC and HPT). The No. 5R bearing supports the high-pressure turbine rotor (HPTR) at its forward shaft.

The D and E sump houses the No. 6R and No. 7R bearings. The No. 6R bearing supports the forward end of the low-pressure turbine rotor (LPTR) shaft. The No. 7R bearing supports the aft end of LPTR shaft and the balance piston system.


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Synthetic lube oil is supplied to the bearings and scavenged out of the sumps by aseven (7) element pump assembly. A single supply element provides lubricating oil to allthe bearings and gearboxes. The remaining six elements are utilized to scavenge oilaway from the bearing sumps and gearboxes. The sump-A scavenge oil drains to thetransfer gearbox (TGB) through the 6:00 oclock compressor front frame (CFF) strut thathouses the radial driveshaft. Oil is then scavenged through the transfer gearbox. TheNo. 4R/4B and No. 5R bearing zones of the sump-B and sump-C are individuallyscavenged, as are the No. 6R and No. 7R bearing zones of the D and E sump. All sumpsemit oil mist-carrying air that is vented to a packager-supplied air-oil separator.


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Dry Sump Construction (Simplified)


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The gas turbine design uses the dry sump system to provide lubrication to the gas turbine main bearings. The dry sump system employs five subsystems:

Oil Supply - Oil is delivered to the bearings through jets pressurized by a supply pump deliver oil onto the bearings. Oil Scavenge - Oil scavenge is accomplished when suction, created by the pumping action of a scavenge oil pump, is applied to a port in the lowest point of the oil-wetted cavity. Seal Pressurization - Bleed air, directed to the sump cavity by ports or tubes in the engine structure, pressurizes seals. Sump Vent - By venting the oil-wetted cavity out the top to ambient air pressure, a positive flow of pressurizing air to the sump is maintained. Cavity Drain - Oil leaked from the seals (sump B and sump C) is carried to an overboard dump location.


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.

Bearing Oil Seals


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Bearing Oil Seals

When a fault occurs and oil leaks across the oil seals, it must not be allowed to become a fire hazard or to contaminate the customer bleed air. Therefore, a drain is provided to the pressurization chamber. The drainage line is directly connected to an overboard drain port without shutoff so that, whenever the gas turbine is running, there is a flow of air out the drain. Scavenge pumps are connected by tubes to a low drain point in each sump. Whenever the gas turbine is running, the scavenge pumps are working to remove the oil from the sump drains.

The rotating seal provides multiple serrations machined to a knife edge. The stationary shroud portion of the seal provides a surface opposite the knife edges. The seals reduce the leakage from one cavity to the other. Sump pressurizing airflow supply is a volume and pressure great enough to maintain a flow radially inward to the sump cavity across the oil seals and outward to the gas turbine cavity across the air seals. The airflow inward to the sump sweeps with it any oil that may be on the seals keeping the oil contained in the sump. The inflowing air is removed by both the vent system and the scavenge oil system.

The Sump design uses pressurized labyrinth type oil seals between the sump housing and the shaft to contain the oil within the sump, and pressurized labyrinth venting seals to maintain pressurizing air separate from the primary gas turbine airflow.


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3.4 MAJOR COMPONENTS


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3.4 MAJOR COMPONENTS

Inlet Volute Variable inlet guide vane (VIGV) assembly Low-pressure compressor (LPC) assembly Low-pressure compressor bypass-air collector Variable bypass valve system Low-pressure compressor front frame assembly High-pressure compressor (HPC) assembly Compressor rear frame assembly Combustor assembly High-pressure turbine assembly Low-pressure turbine assembly Turbine rear frame assembly Accessory gearbox


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3.4.1 AIR INLET VOLUTE

Inlet Volute-ALF LP Compressor Mounting Face


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The Air Inlet Volute provides for a smooth transition of airflow from the air filter enclosure into the first stage of the low pressure compressor. The volute changes the airflow direction from a vertical to a horizontal flow. The air inlet casing assembly comprises an external casing, approximately rectangular in shape, and forms a circular internal casing to which the low pressure compressor mounts. The generator drive shafts then runs through the center of the volute to the generator.

A flexible joint of Neoprene rubber polymer is fitted between the inlet volute and the enclosure air ducting to accommodate relative movements. A trash screen (FOD screen) is also included for additional protection against debris in the inlet system.

Mounted on the forward end (ALF) of the inlet volute are the online and offline water wash manifolds. The LP SPRINT manifold is mounted on the rear (ALF) of the volute. Located on the bottom of the volute is a drain line with check valve that is plumbed to the customer provided waste fluid tank.


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3.4.2 INLET GUIDE VANE ASSEMBLY


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3.4.2 INLET GUIDE VANE ASSEMBLY

The air intake section is designed to interface with a radial inlet duct, which allows inlet air to be drawn from the side or top or with an axial inlet system, which draws air from the front. The radial inlet duct is compatible with either forward or rear drive installations, while the axial inlet can be used only in rear drive installations.

The Optional Variable Inlet Guide Vane Assembly (VIGV) is located at the front of the LPC. It allows flow modulation at partial power, resulting in increased engine efficiency. The VIGV system consists of 43 stationary, leading-edge vanes and variable trailing flaps. The variable flaps can be rotated from 10 degrees open to +60 degrees closed by means of an actuation ring, which is driven by twin hydraulic actuators at the 3 oclock and 9 oclock positions. Both actuators are equipped with linear variable-differential transformers (LVDTs).


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Normal engine operation is approximately 5 degrees open (full power) to +35 degrees closed (idle power). The flaps will also close during large power reductions in order to quickly reduce the LPC flow rate and maintain the LPC stall margin. The packager-supplied control is designed to provide excitation and signal conditioning for both LVDTs. It also controls VIGV position by means of closed-loop scheduling of the VIGV actuator position, based on LPC inlet temperature (T2) and HPC discharge static pressure (PS3) corrected to gas turbine inlet pressure conditions (P0).

The VIGV system improves performance for both simple cycle and heat-recovery cycles. It also helps minimize the variable bypass valve (VBV) flow and pressure levels, thereby reducing associated flow noise. A pressurized rotating seal between the VIGV hub and the LPC rotor prevents ingestion of unfiltered air into the flow path. The LM6000 PC engine can be provided with or without the VIGV assembly. LM6000 PC models without a VIGV assembly have a 43-strut inlet frame.


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3.4.3 LOW-PRESSURE COMPRESSOR (LPC) ASSEMBLY


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3.4.3 LOW-PRESSURE COMPRESSOR (LPC) ASSEMBLY

The forward end of the low-pressure compressor is mounted to the IGV/VIGV assembly, while the rear mounts to the Compressor Front Frame (CFF).

The LM6000 LPC is a 5-stage, axial-flow compressor with a 5-stage fixed stator. The LPC stator case contains the stator vanes for the LPC rotor. The case is horizontally split to facilitate repair.


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LPC Rotor

Blade Locking Lugs


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LPC Rotor

Individual disks are used in stages 0 and 1. Stages 2 thru 4 of the LPC rotor are an integral spool. Stages 0 and 1 blades have been modified to include squealer tips.

Stage 0 blades are individually retained in the axial dovetail slots of the disk by a one-piece blade retainer. Stages 1 thru 4 LPC blades are retained in circumferential slots in the stage 1 disk and stages 2 thru 4 spool. The blade-retention features permit individual blade replacement. Blades in stages 0 thru 3 can be removed without removing the rotor. As the compressor rotates, the blades load centrifugally and become tight fitting.


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Low Pressure Compressor Casing and Stators


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LPC Stator Vanes

The stages 0 thru 2 stator vanes are individually replaceable. The vanes are shrouded to reduce vane response to aerodynamic forces. Wear strips are utilized between the vane dovetails and the LPC casing slots. The stage 3 casing is a full-circumferential case and is lined with honeycomb material over the rotor blade tips. Stage 3 vanes are bolted to the stage 3 case forward flange. The stage 4 stator vanes are mounted in the front frame and supported on the inside diameter by a support structure that is bolted to the engine front frame.


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3.4.6 LOW PRESSURE COMPRESSOR BYPASS AIR COLLECTOR

The LPC bypass-air collector is a duct attached to the front frame. It collects LPC discharge air, vented through the LPC bypass doors, and directs it overboard through packager-provided ducting.


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Variable Bleed Valves


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Variable Bypass Valve System

The variable bypass valve (VBV) system is located in the front frame assembly. This system is used to vent LPC discharge air overboard through the LPC bypass-air collector in order to maintain LPC stall margin during starting, partial power operation, and large power transients. The VBV system consists of 12 variable-position bypass valves, 6 VBV actuators (two with LVDTs) Linear Variable Differential Transformer, 6 actuator bell cranks, 12 VBV doorbell cranks, and an actuation ring.

Actuators are installed at the 1 oclock, 3 oclock, 5 oclock, 7 oclock, 9 oclock, and 11 oclock positions on the engine. The six actuators are positioned with one VBV door on each side of each actuator. Bell cranks and pushrods mechanically link the actuators, the actuation ring, and the VBV doors. The actuator positions the actuation ring, which opens and closes the VBV doors. The 5 oclock and 11 oclock position actuators are equipped with integral LVDTs for position indication. The packager-supplied control is designed to provide excitation and signal conditioning for both LVDTs and, to control VBV position by means of closed-loop scheduling of VBV actuator position, based on LPC inlet temperature (T2) and high-pressure (HP) rotor speed corrected to inlet conditions (XN2.5R2).


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3.4.7 LOW PRESSURE COMPRESSOR FRONT FRAME ASSEMBLY


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3.4.7 LOW PRESSURE COMPRESSOR FRONT FRAME ASSEMBLY

The front frame is a major structure that provides support for the LPC rotor and the forward end of the HPC rotor through the No. 1B, No. 2R, and No. 3R bearings. The frame also forms an airflow path between the LPC and the HPC inlet. Front engine mount provisions are located on the front frame 3 oclock and 9 oclock positions. One pad is included on the frame outer case for mounting HPC inlet temperature sensors T2.5 and HPC pressure sensor P2.5. The sensors provide control information to the fuel management system.

The front frame is made from a high-strength stainless steel casting. Twelve equally spaced radial struts are used between the hub and outer case to provide support for the inner hub. Twelve variable-position bypass valve doors are located on the outer wall for LPC discharge bleed.

The front frame contains the engine A-sump, which includes a thrust bearing (1B) and roller bearing (2R) that support the LPC rotor, and a roller bearing (3R) that supports the forward end of the HPC rotor. Lubrication oil supply and scavenge lines for the A sump are routed inside the frame struts. The inlet gearbox is located in the A sump with the radial drive shaft extending outward through the strut located at the 6 oclock position.


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Inlet Gearbox

Radial Drive Shaft


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Radial Drive Assembly

The radial drive shaft assembly is located in the 6 oclock CFF strut. The shafts serve to transmit torque from the Inlet Gearbox (IGB) to the Transfer Gearbox (TGB).The drive shaft assembly consists of three machined, tubular steel shafts, housing, and bearings.

The upper radial shaft is splined at the upper end to the IGB and at the lower end to the radial mid-shaft. The shaft is enclosed by the front frame and supported by a ball bearing at its lower end. The radial mid-shaft is splined at the upper end to the upper shaft and at the lower end to the lower shaft. The mid-shaft is enclosed in a housing and supported by a ball bearing at its lower end. The lower radial shaft is splined at the upper end to the mid-shaft and at its lower end to the TGB. The lower shaft is enclosed by the radial adapter portion of the TGB.


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HPC CASE(LOWER HALF)

HPC ROTOR

HPC CASE(UPPER HALF)

G-66-04

3.4.8 HIGH PRESSURE COMPRESSOR (HPC) ASSEMBLY


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3.4.8 HIGH PRESSURE COMPRESSOR (HPC) ASSEMBLY

The LM6000 HPC is a 14-stage, axial-flow compressor. It incorporates VIGVs and variable stators in stages 05 to provide stall-free operation and high efficiency throughout the starting and operating range. Provisions for customer-use bleed air are available at stage 8 and at the compressor discharge. On earlier PA/PB model turbines the seventh and eleventh stages bleed air is utilized, while, later versions (PC/PD) use eighth and eleventh stage bleed air. Compressor discharge air is extracted for cooling and pressurization of the engine components.


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High Pressure Compressor Rotor Layout


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HPC Rotor

The HPC rotor is a bolted assembly of five major structural elements consisting of a stage 1 disk, a stage 2 disk with an integral forward shaft, stages 39 spool, a stage 10 disk, and stages 1114 spool with an integral rear shaft. These structural elements are connected through fully rabbeted joints at stage 2 and stage 10. On newer model HPC there are only four major structural elements. In these versions, the 10thstage disk has been deleted and added as an integral component of the 10--14 stage spool assembly.


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Typical Blade Profiles


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Disk 1 and 2 Loading

Stages 1 and 2 blades are individually retained in axial dovetail slots, and the remaining blades are held in circumferential dovetail slots. These features allow individual stage 1 blade replacement without disassembly of the rotor.

Stage 1 blades are shrouded at mid-span for the purpose of reducing vibratory stress. All other blades are cantilevered from the rotor structure.


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High Pressure Rotor Assembly


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G-145-04HP STATOR

CASE (LOWER)

HP STATORCASE (UPPER)

VARIABLE STATOR VANES

VIGV

STAGE 1VANES

HPC Stator Casing


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HPC STATOR

The HPC stator consists of a cast stator case that contains the compressor stator vanes. The inlet guide vanes and the stages 15 vanes can be rotated about the axis of their mounting trunnions to vary the pitch of the airfoils in the compressor flow path. Vane airfoils in the remaining stages are stationary. All fixed and variable vanes are non-interchangeable with other stages to prevent incorrect assembly. The casing is split along the horizontal split-line for ease of assembly and maintenance. The inlet guide vanes and the stages 1 and 2 vane shrouds also support interstage rotor seals. The shrouds are designed to allow the removal of either half of the compressor casing. There are 14 axial stations provided for borescope inspection of blades and vanes.


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HPC Stator Casing and Vane Assembly


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Variable Stator Vane Assembly


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VARIABLE STATOR VANE ASSEMBLY

The VSV assembly, an integral part of the HPC stator, consists of two VSV actuators and levers, actuation rings, and linkages for each VSV stage. Stator vane position is vital to stable, efficient operation of the engine. While the HPC is designed for peak aerodynamic efficiency at full power and full speed, it must also operate at lower speeds. At these lower speeds, the later stages of the compressor cannot consume all the air delivered by the earlier stages. The variable stators accommodate this situation by limiting the compression ratio of the first six stages of the compressor at low speeds and changing the compression at higher speeds.

This is accomplished with two hydraulic actuators, one at the 3:00 oclock position and one at the 9:00 oclock position. Each actuator uses an LVDT for position feedback to the control system. The control system is designed to provide excitation and signal conditioning for both LVDTs, and to control VSV position by means of closed-loop scheduling of VSV actuator position, based on corrected HP rotor speed (XN2.5R) and inlet temperature (T2.5).


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Variable Stator Vane Assembly

Variable Stator Vane Actuation Rings

Actuator


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Compressor Rear Frame Assembly


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Compressor Rear Frame Assembly

The compressor rear frame (CRF) assembly connects the compressor-casing flange to the high-pressure turbine nozzle assembly and consists of an outer case, 10 struts, and the B- and C-sump housings. The outer case supports the combustor, fuel manifolds and fuel nozzles, two ultraviolet flame detectors for flame sensing, an accelerometer, discharge static (P3) and HPC discharge temperature sensor (T3). The hub provides support for a thrust bearing (4B) and two roller bearings (4R and 5R) to support the midsection of the HP rotor system.

Bearing axial and radial loads, and a portion of the first-stage nozzle load, are transmitted through the hub and 10 radial struts to the case. The hub, struts, and outer casing are a one-piece casting. The casting is welded to the fuel embossment ring and bolted to the aft case. This serves as the structural load path between the compressor casing and the HPT stator case. Seven borescope ports are provided for inspection of the combustor, pre-mixers, and HPT. B-sump and C-sump service lines are contained in, and pass through, the CRF struts.


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Combustor Assembly


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Combustor AssemblyThe LM6000 gas turbine uses a singular annular combustor and is furnished with 30 externally mounted fuel nozzles for liquid distillate fuel, natural gas fuel, or dual fuel, depending upon the fuel system specified by the customer. Fuel systems may also be equipped for water or steam injection for NOx suppression. This combustion system is a high-performance design that has consistently demonstrated low exit temperature pattern factors, low-pressure loss, low smoke, and high combustion efficiency at all operating conditions.

SINGULAR ANNULAR COMBUSTORKey features of the singular annular combustor are the rolled-ring inner and outer liners; the low-smoke emission, swirl-cup dome design and the short burning length. The short burning length reduces liner cooling air consumption, which improves the exit temperature pattern factor and profile. The swirl-cup dome design serves to lean-out the fuel-air mixture in the primary zone of the combustor. This eliminates the formation of the high-carbon visible smoke that can result from over-rich burning in this zone.


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Combustion Liner Assembly

Swirler with Liquid Fuel Nozzle

SAC

Outer Liner

DLE


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Combustion Liner Assembly

The combustion liner assembly is supported entirely at the aft end. The support ring on the outer liner is trapped in a groove on the compressor rear frame (CRF) aft end with the high pressure turbine case. The inner liner is supported by the inner flow path of the CRF. The combustion assembly consists of an inner cowl, an outer cowl, a dome, and an inner and outer liner.

COWLThe cowl consists of 2 parts, the inner and outer cowls separated by the dome. Its purpose is to form a smooth leading-edge which splits and meters the incoming air flow to the combustion assembly.


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DOME The dome is a fabricated component consisting of 30 vortex inducing swirl assemblies consisting of two counter-rotating primary and secondary swirlers. Their purpose is to provide flame stabilization and complete mixing of the fuel air mixture. The primary swirler floats on the face of the secondary swirler to allow growth difference for the fuel nozzles. The entire surface of the dome is swept by a film of cooling air.

LINERSThe inner and outer liners are composed of a series of circumferentially rolled ring strips joined together by resistance welding. They are protected from convective and radiant heat by continuous circumferential film cooling. Combustion zone dilution and mixing air entry is provided by a pattern of various sized circular holes in each ring. These holes provide recirculation for flame stabilization andshape the exit gas profile. Ports and tube assemblies have been located at the 3:00 and 5:00 o'clock positions for the igniter plugs. The liners and dome have a thermal barrier coating applied to the hot side.


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IGNITION SYSTEM

The energy level of the ignition system is lethal, and personnel should never contact output from the ignition exciters, leads or igniter plugs.


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IGNITION MODULES


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IGNITION SYSTEMThe ignition system produces the high-energy sparks that ignite the fuel-air mixture in the combustor during starting. The system consists of high-energy spark igniters, a high-energy capacitor-discharge ignition exciter, and an interconnecting cable. The ignition cables interconnect directly between the package-mounted exciters and the igniters, which are mounted on the engine compressor rear frame. During the start sequence, fuel is ignited by the igniter, which is energized by the ignition exciter. Once combustion becomes self-sustaining, the igniter is de-energized at 400 F (204 C).

Proper installation of the igniter plug on the combustion chamber is essential for long operating life. The igniter plug has a special distance (packing) ring that must be installed between the plug and compressor rear frame. The correct distance of the plug in the rear frame is important, both for operation and cooling, and it can be adjusted with the distance ring. Cooling is achieved with compressor air flowing alongside the igniter plug tip. Also, 12 holes in the plug tip are present for cooling purposes and, finally, 6 holes provide cooling air for the igniter plug shank.


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The ignition system is normally energized only during the starting sequence. However, the circuit should be arranged so that the ignition system can be manually operated for maintenance and testing.

To ensure a successful light off, the ignition system is comprised of two independent ignition systems. Due to already increased air temperature from compression through the compressor, and fuel atomization from the fuel nozzle, it is possible to achieve ignition with only one igniter. Running two independent systems ensures the ability to maintain normal operations even with the complete loss of one system. Because of this configuration it is necessary to check the operation of the igniter system on a routine basis in accordance with the maintenance work package.


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Typically the igniters should be checked when a turbine fails to light-off and all other primary start requirements are met. Such as:Proper acceleration of the HPC (XN2.5)Proper CDP pressure (P3)Proper fuel valve Position

This type of failure is due to loss of both igniters. The only igniter indication that the operator can monitor is the logic state change on the Turbine Overview Screen. The operator screen change is a function of an energized relay coil. If there is a failure in the ignition system, the screen may indicate proper operation but, in reality, the system is inoperable. Because of the high voltage generated by the exciter module, there is no feedback of the igniter output to give a true indication of proper operation of the circuit.

Duty cycle is: 90 seconds ON max, 2 start cycles in a 30 minute periodPower input is:106-124 volt AC, Requirement at 60 Hz or 50 Hz


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Igniter Location

Igniter Mounting Detail

Igniter 3 Oclock Location

Igniter 5 Oclock Location


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High-Pressure Turbine Assembly


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High-Pressure Turbine Assembly

The LM6000 HPT is an air-cooled, two-stage design with demonstrated high efficiency. The HPT system consists of the HPT rotor and the stage 1 and stage 2 HPT nozzles. The HPT assembly drives the HPC rotor by extracting energy from the hot-gas path stream.

HPT ROTORThe HPT rotor assembly consists of the stage 1 disk and integral shaft, a conical impeller spacer with cover, a thermal shield and a stage-2 disk. Forward and aft rotating air seals are assembled to the HPT rotor and provide air-cooled cavities around the rotor system. An integral coupling nut and pressure tube is used to form and seal the internal cavity. The rotor disks and blades are cooled by a continuous flow of compressor discharge air. This air is directed to the internal cavity of the rotor through diffuser vanes that are part of the forward seal system.


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The stage 1 disk/shaft design combines the rotor forward shaft and stage 1 disk into a one-piece unit. Torque is transmitted to the compressor rotor through an internal spline at the forward end of the disk/shaft. The stage 1 blades fit into axial dovetail slots in the disk. The stage 2 disk incorporates a flange on the forward side for transmitting torque to the stage 1 disk. An aft flange supports the aft air seal and the integral coupling nut and pressure tube. Stage 2 blades fit into axial dovetail slots in the disk.

Internally cooled turbine blades are used in both stages. Both stages of blades are cooled by compressor discharge air flowing through the blade shank into the airfoil.

The cone-shaped impeller spacer serves as the structural support between the turbine disks. The spacer also transmits torque from the stage 2 disk to the stage 1 disk. The catenary-shaped thermal shield forms the outer portion of the turbine rotor cooling air cavity and serves as the rotating portion of the interstage gas path seal.


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High-Pressure Turbine Blade Cooling


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High-Pressure Turbine Blade CoolingStage 1 High-Pressure Turbine BladesFirst-stage turbine blades, contained within the CRF, are internally cooled with HPC discharge air. The HPC discharge air is directed through the turbine disk to the blade roots, passing through inlet holes in the shank to serpentine passages within the airfoil section of the blade. This air finally exits through nose and gill holes in the leading edge of the blades, where it forms an insulating film over the airfoil surface through holes in the cap at the outer end of the blade and through holes in the trailing edge of the airfoil.

Stage 2 High-Pressure Turbine BladesBecause the hot-gas path stream is cooler when it reaches the second-stage turbine blades, the cooling required to maintain a suitable metal temperature is not as great as with the first stage. The second-stage blades are, therefore, only cooled by convection. The air moves through passages within the airfoil section and is discharged only at the blade tips.

Stage 1 HPT NozzleThe stage 1 HPT nozzle consists of 23 two-vane segments bolted to a nozzle support attached to the hub of the CRF.


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High-Pressure Turbine Nozzle Cooling


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High-Pressure Turbine Nozzle CoolingCompressor discharge air is used to cool the nozzle vanes and support bands to maintain the metal temperatures at the levels required for extended operating life. Stage 11-discharge air enters at the top and bottom of each vane. The air cools the vanes internally, and is then discharged through a large number of small holes and slots strategically located so the air forms an insulating film over the entire surface of the vanes.

Stage 2 HPT NozzleThe stage 2 HPT nozzle assembly consists of stage 2 nozzle segments, stages 1 and 2 HPT shrouds and shroud supports, HPT stator support (case), and interstage seals. There are 24 paired nozzle-vane segments. The nozzle vanes are internally cooled by HPC Stage 11 air.

The stage 2 nozzles are supported by the stage 1 shroud support. They are also bolted to the stage 2 shroud support forward leg, which is attached by a flange to the outer structural wall. The stage 1 shroud system features segmented supports and shroud segments to maintain turbine clearance.

The turbine shrouds form a portion of the outer aerodynamic flow path through the turbine. They are axially aligned with the turbine blades and form a pressure seal to minimize HP gas leakage around the tips of the blades.


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HPT Interstage Seal


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Low-Pressure Turbine Assembly


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Low-Pressure Turbine Assembly

The LPT drives the LPC and load device using the core gas turbine discharge gas flow for energy. The principal components of the LPT module are a five-stage stator, a five-stage rotor supported by the No. 6R and No. 7R bearings, and a cast Turbine Rear Frame (TRF) supporting the stator casing and the No. 6R and No. 7R bearings.

LPT ROTORThe LPT rotor assembly drives the LPC through the LP mid-shaft and drives a load through either the mid-shaft or from an aft drive adapter on the rear of the LPT rotor. The LPT rotor assembly consists of five stages of bladed disks and a shaft sub-assembly. The rotor is supported by the No. 6R and No. 7R bearings in the D and E sump of the TRF.


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Each LPT rotor stage consists of a bladed disk subassembly that is comprised of a disk, turbine blades, and blade retainers, interstage air seals, assembly bolts, and balance weights. Integral flanges on each disk provide assembly bolt holes in a low-stress area of the disk. Blade retainers hold the turbine blades in the axial dovetail slots.

The turbine shaft assembly is a torque cone coupled to the mid-shaft through a spline and is bolted to the stage 2 and stage 3 turbine disk flanges. It also provides the journal for the D- and E-sump air/oil seal and the No. 6R and No. 7R bearing interfaces. The rotating portion of the balance piston system mounts on the shaft aft of the No. 7R bearing seals. Additionally, the aft shaft spline provides for driving the output load from the rear through the aft drive adapter.


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LPT Rotor Detail


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LPT NOZZLESThe five-stage stator assembly consists of a one-piece tapered 360 casing, five stages of interlocking tip shrouds, and a 12-segment LPT case external cooling manifold. Air-cooled, first-stage nozzle segments with a bolt-on pressure balance seal, four additional stages of nozzle segments with bolt-on inter-stage seals, and instrumentation and borescope ports also make up the stator assembly.

First stage nozzle cooling air is supplied from the 8th stage HPC bleed air header and high pressure recoup air.

The LPT casing is the load-carrying structure between the HPT stator case and the TRF. The casing contains internal machined flanges that provide hooks to support the nozzle segments and stops to assure nozzle alignment and seating. Borescope inspection ports are provided along the right side, aft looking forward (ALF) from the 2:30 to 4:30 positions at nozzle stages 1, 2, and 4.


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Low Pressure Turbine Case


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Low Pressure Turbine Case

The stage 1-nozzle vanes provide capability for LPT inlet instrumentation. Eight separate shielded chromel-alumel (type K) thermocouple probes are installed on the LPT stator case to sense LPT inlet temperature. Each dual-element T4.8 sensor reads an average of the two elements for a total of eight control readings. Two flexible harnesses, each connected to four of the probes, are routed to connectors on the No. 4 electrical panel. The engine also has an LPT inlet gas total pressure (P4.8) probe located on the right side of the LPT stator case. Seals minimize the air leakage around the inner ends of the nozzles, and shrouds minimize air leakage over the tips of the turbine blades


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LPT Case Cooling Manifold

LPT Case Cooling AirflowLPT CASE COOLINGLater models of the LM6OOO-PA, as well as the -PC, have a cooling manifold, which is used to reduce casing temperatures as well as to lower blade tip clearance to improve efficiency. Air provided from the Compressor Front Frame (CFF) is utilized as the cooling medium.


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Turbine Rear Frame Assembly


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Turbine Rear Frame AssemblyThe turbine rear frame (TRF) is a one-piece casting that provides the gas turbine exhaust flow path and the supporting structure for the D and E sump, the LPT rotor thrust balance assembly, the LPT rotor shaft, and the aft drive adapter. Fourteen radial struts function as outlet guide vanes to straighten the exhaust airflow into the exhaust diffuser for enhanced performance. Lubrication oil supply and scavenge lines for the D and E sump and LPT rotor speed sensors (XNSD-A and XNSD-B) are routed through the struts.

The LPT rotor thrust balance system is designed to maintain the axial thrust loading on the No. 1B thrust bearing within design

LM6000 Gas Turbine - Generator Package Familiarization Training

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