30079_57b.pdf

57.1.5 Gas Turbine Operation

Like other internal combustion engines, the gas turbine requires an outside source of starting power.

This is provided by an electrical motor or diesel engine connected through a gear box to the shaft

of the gas turbine (the high-pressure shaft in a multishaft configuration). Other devices can be used,

including the generator of large electric utility gas turbines, by using a variable frequency power

supply. Power is normally required to rotate the rotor past the gas turbine's ignition speed of 10-15%

on to 40-80% of rated speed where the gas turbine is self-sustaining, meaning the turbine produces

sufficient work to power the compressor and overcome bearing friction, drag, and so on. Below self-

sustaining speed, the component efficiencies of the compressor and turbine are too low to reach or

exceed this equilibrium.

When the operator initiates the starting sequence of a gas turbine, the control system acts by

starting auxiliaries such as those that provide lubrication and the monitoring of sensors provided to

ensure a successful start. The control system then calls for application of torque to the shaft by the

starting means. In many industrial and utility applications, the rotor must be rotated for a period of

time to purge the flow path of unburned fuel that may have collected there. This is a safety precaution.

Thereafter, the light-off speed is achieved and ignition takes place and is confirmed by sensors.

Ignition is provided by either a sparkplug type device or by an LP gas torch built into the combustor.

Fuel flow is then increased to increase the rotor speed. In large gas turbines, a warmup period of

one minute or so is required at approximately 20% speed. The starting means remains engaged, since

the gas turbine has not reached its self-sustaining speed. This reduces the thermal gradients experi-

enced by some of the turbine components and extends their low cycle fatigue life.

The fuel flow is again increased to bring the rotor to self-sustaining speed. For aircraft engines,

this is approximately the idle speed. For power generation applications, the rotor continues to be

accelerated to full speed. In the case of these alternator-driving gas turbines, this is set by the speed

at which the alternator is synchronized with the power grid to which it is to be connected.

Aircraft engines' speed and thrust are interrelated. The fuel flow is increased and decreased to

generate the required thrust. The rotor speed is principally a function of this fuel flow, but also

depends on any variable compressor or exhaust nozzle geometry changes programmed into the control

algorithms. Thrust is set by the pilot to match the current requirements of the aircraft, through takeoff,

climb, cruise, maneuvering, landing, and braking.

At full speed, the power-generation gas turbine and its generator (alternator) must be synchronized

with the power grid in both speed (frequency) and phase. This process is computer-controlled and

involves making small changes in turbine speed until synchronization is achieved. At this point, the

generator is connected with the power grid. The load of a power-generation gas turbine is set by a

combination of generator (alternator) excitement and fuel flow. As the excitation is increased, the

mechanical work absorbed by the generator increases. To maintain a constant speed (frequency), the

fuel flow is increased to match that required by the generator. The operator normally sets the desired

electrical output and the turbine's electronic control increases both excitation and fuel flow until the

desired operating conditions are reached.

Normal shutdown of a power-generation gas turbine is initiated by the operator and begins with

the reduction of load, reversing the loading process described immediately above. At a point near

zero load, the breaker connecting the generator to the power grid is opened. Fuel flow is decreased

and the turbine is allowed to decelerate to a point below 40% speed, whereupon the fuel is shut off

and the rotor is allowed to stop. Large turbines' rotors should be turned periodically to prevent

temporary bowing from uneven cool-down that will cause vibration on subsequent startups. Turning

of the rotor for cool-down is accomplished by a ratcheting mechanism on smaller gas turbines, or

by operation of a motor associated with shaft-driven accessories, or even the starting mechanism on

others. Aircraft engine rotors do not tend to exhibit the bowing just described. Bowing is a phenom-

enon observed in massive rotors left stationary surrounded by cooling, still air that, due to free

convection, is cooler at the 6:00 position than at the 12:00 position. The large rotor assumes a similar

gradient and, because of proportional thermal expansion, assumes a bowed shape. Because of the

massiveness of the rotor, this shape persists for several hours, and could remain present when the

operator wishes to restart the turbine.

57.2 GAS TURBINE PERFORMANCE

57.2.1 Gas Turbine Configurations and Cycle Characteristics

There are several possible mechanical configurations for the basic simple cycle, or open cycle, gas

turbine. There are also some important variants on the basic cycle: intercooled, regenerative, and

reheat cycles.

The simplest configuration is shown in Fig. 57.15. Here the compressor and turbine rotors are

connected directly to one another and to shafts by which turbine work in excess of that required to

drive the compressor can be applied to other work-absorbing devices. Such devices are the propellers

and gear boxes of turboprop engines, electrical generators, ships' propellers, pumps, gas compressors,

vehicle gear boxes and driving wheels, and the like. A variation is shown in Fig. 57.16, where a jet

Fig. 57.15 Simple-cycle, single-shaft gas turbine schematic.

nozzle is added to generate thrust. Through aerodynamic design, the pressure drop between the turbine

inlet and ambient air is divided so that part of the drop occurs across the turbine and the remainder

across the jet nozzle. The pressure at the turbine exit is set so that there is only enough work extracted

from the working fluid by the turbine to drive the compressor (and mechanical accessories). The

remaining energy accelerates the exhaust flow through the nozzle to provide jet thrust.

The simplest of multishaft arrangements appears in Fig. 57.17. For decades, such arrangements

have been used in heavy-duty turbines applied to various petrochemical and gas pipeline uses. Here,

the turbine consists of a high-pressure and a low-pressure section. There is no mechanical connection

between the rotors of the two turbines. The high-pressure (h.p.) turbine drives the compressor and

the low-pressure (Lp.) turbine drives the load—usually a gas compressor for a process, gas well, or

pipeline. Often, there is a variable nozzle between the two turbine rotors that can be used to vary

the work split between the two turbines. This offers the user an advantage. When it is necessary to

lower the load applied to the driven equipment—for example, when it is necessary to reduce the flow

from a gas-pumping station—fuel flow would be reduced. With no variable geometry between the

turbines, both would drop in speed until a new equilibrium between Lp. and h.p. speeds occurs. By

changing the nozzle area between the rotors, the pressure drop split is changed and it is possible to

keep the h.p. rotor at a high, constant speed and have all the speed drop occur in the Lp. rotor. By

doing this, the compressor of the gas turbine continues to operate at or near its maximum efficiency,

contributing to the overall efficiency of the gas turbine and providing high part-load efficiency. This

two-shaft arrangement is one of those applied to aircraft engines in industrial applications. Here, the

h.p. section is essentially identical to the aircraft turbojet engine or the core of a fan-jet engine. This

h.p. section then becomes the gas generator and the free-turbine becomes what is referred to as the

power turbine. The modern turbofan engine is somewhat similar in that a low-pressure turbine drives

a fan that forces a concentric flow of air outboard of the gas generator aft, adding to the thrust

provided by the engine. In the case of modern turbofans, the fan is upstream of the compressor and

is driven by a concentric shaft inside the hollow shaft connecting the h.p. compressor and h.p. turbine.

Fig. 57.16 Simple-cycle single-shaft, gas turbine with jet nozzle; simple

turbojet engine schematic.

Fig. 57.17 Industrial two-shaft gas turbine schematic showing high-pressure gas generator ro-

tor and separate free-turbine low-pressure rotor.

Figure 57.18 shows a multishaft arrangement common to today's high-pressure turbojet and tur-

bofan engines. The h.p. compressor is connected to the h.p. turbine, and the Lp. compressor to the

Lp. turbine, by concentric shafts. There is no mechanical connection between the two rotors (h.p.

and Lp.) except via bearings and the associated supporting structure, and the shafts operate at speeds

mechanically independent of one another. The need for this apparently complex structure arises from

the aerodynamic design constraints encountered in very high-pressure-ratio compressors. By having

the higher-pressure stages of a compressor rotating at a higher speed than the early stages, it is

possible to avoid the low-annulus-height flow paths that contribute to poor compressor efficiency.

The relationship between the speeds of the two shafts is determined by the aerodynamics of the

turbines and compressors, the load on the loaded shaft and the fuel flow. The speed of the h.p. rotor

is allowed to float, but is generally monitored. Fuel flow and adjustable compressor blade angles are

used to control the Lp. rotor speed. Turbojet engines, and at least one industrial aero-derivative engine,

have been configured just as shown in Fig. 57.18. Additional industrial aero-derivative engines have

gas-generators configured as shown and have power turbines as shown in Fig. 57.17.

The next three configurations reflect deviations from the basic Bray ton gas turbine cycle. To

describe them, reference must be made back to the temperature-entropy diagram.

Intercooling is the cooling of the working fluid at one or more points during the compression

process. Figure 57.19 shows a low-pressure compression, from points a to b. At point b, heat is

removed at constant pressure. The result is moving to point c, where the remaining compression

takes place (line c-d), after which heat is added by combustion (line d-e). Following combustion,

expansion takes place (line e-f}. Finally, the cycle is closed by discharge of air to the environment

(line /-a), closing the cycle. Intercooling lowers the amount of work required for compression,

because work is proportional to the sum of line a-b and line c-d, and this is less than that of line

Fig. 57.18 Schematic of multishaft gas turbine arrangement typical of those used in modern

high-pressure-ratio aircraft engines. Either a jet nozzle, for jet propulsion, or a free power tur-

bine, for mechanical drive, can be added aft of the I.p. turbine.

Fig. 57.19 Temperature-entropy diagram for intercooled gas turbine cycle. Firing temperature

arbitrarily selected at 110O 0 C and pressure ratio at 24:1.

a-d', which would be the compression process without the intercooler. Lines of constant pressure

are closer together at lower temperatures, due to the same phenomenon that explains higher turbine

work than compressor work over the same pressure ratio. Although the compression process is more

efficienct with intercooling, more fuel is required by this cycle. Note the line d-e as compared with

the line d'-e. It is clear that the added vertical length of line d-e versus d'-e is greater than the

reduced vertical distance achieved in the compression cycle. For this reason, when the heat in the

partially compressed air is rejected, the efficiency of an intercooled cycle is lower than a simple

cycle. Attempts to use the rejected, low-quality heat in a cost-effective manner are usually not

successful.

The useful work, which is proportional to e-f less the sum of a-b and c-d, is greater than the

useful work of the simple a-d'-e-f-a cycle. Hence for the same turbomachinery, more work is

produced by the intercooled cycle—an increase in power density. This benefit is somewhat offset by

the fact that relatively large heat-transfer devices are required to accomplish the intercooling. The

intercoolers are roughly the size and volume of the turbomachinery and its accessories. Whether the

intercooled cycle offers true economic advantage over simple-cycle applications depends on the de-

tails of the application, the design features of the equipment, and the existence of a use for the

rejected heat.

An intercooled gas turbine is shown schematically in Fig. 57.20. A single-shaft arrangement is

shown to demonstrate the principal, but a multishaft configuration could also be used. The compressor

is divided at some point where air can be taken offboard, cooled, and brought back to the compressor

for the remainder of the compression process. Combustion and turbine configurations are not affected.

The compressor-discharge temperature of the intercooled cycle (point d) is lower than that of the

simple cycle (point d'). Often, cooling air, used to cool turbine and combustor components, is taken

from, or from near, the compressor discharge. An advantage often cited for intercooled cycles is the

lower volume of compressor air that has to be extracted. Critics of intercooling point out that the

cooling of the cooling air only, rather than the full flow of the machine, would offer the same benefit

with smaller heat exchangers. Only upon assessment of the details of the individual application can

the point be settled.

The temperature-entropy diagram for a reheat, or refired, gas turbine is shown in Fig. 57.21. The

cycle begins with the compression process shown by line a-b . The first combustion process is shown

by line b-c. At point c, a turbine expands the fluid (line c-d} to a temperature associated with an

intermediate pressure ratio. At point d, another combustion process takes place, returning the fluid

to a high temperature (line d-e ). At point e, the second expansion takes place, returning the fluid to

ambient pressure (line e-f}, whereafter the cycle is closed by discharge of the working fluid back to

the atmosphere.

Fig. 57.20 Schematic of a single-shaft, intercooled gas turbine. In this arrangement, both com-

pressor groups are fixed to the same shaft. Concentric, multishaft, and series arrangements are

also possible.

An estimate of the cycle efficiency can be made from the temperatures corresponding to the

process end points of the cycle in Fig. 57.21. Dividing the turbine temperature drops, less the com-

pressor temperature rise, by the sum of the combustor temperature rises, one calculates an efficiency

of approximately 48%. This, of course, reflects perfect compressor, combustor, and turbine efficiency

and pure air as the working fluid. Actual efficiencies and properties, and consideration of turbine

cooling produce less optimistic values.

Fig. 57.21 Temperature-entropy diagram for a reheat, or refired, gas turbine. Firing tempera-

tures were arbitrarily chosen to be equal, and to be 125O 0 C. The intermediate pressure ratio

was chosen to be 8:1 and the overall pressure ratio to be 32:1. Dashed lines are used to illus-

trate comparable simple gas turbine cycles.

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