ความเห็นเพิ่มเติมที่ 216 10 พ.ค. 2549 (12:32)
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FIG. 6.1 Can-type combustor [ courtesy Rolls-Royce plc)
Separate combustion cans are still widely used in industrial engines, but recent designs make use of a annular (or tubo-annular) system, where individual flame tubes are uniformly spaced around an annular casing. The Alston Typhoon, shown in Fig. 1.13, uses this system; the General Electric and Westinghouse families of industrial gas turbines also use this arrangement. Figure 1.13 shows the reverse flow nature of the airflow after leaving the diffuser downstream of the axial compressor; the use of a reverse flow arrangement allows a significant reduction in the overall length of the compressor-turbine shaft and also permits easy access to the fuel nozzles and combustion cans for maintenance. The ideal configuration in terms of compact dimensions is the annular combustor, in which maximum use is made of the space available within a
specified diameter; this should reduce the pressure loss and result in an engine of minimum diameter. Annular combustors presented some disadvantages, which led to the development of annular combustors initially. Firstly, although a large number of fuel jets can be employed, it is more difficult to obtain an even fuel/air distribution and an even outlet temperature distribution. Secondly, the annular chamber is inevitably weaker structurally and it is difficult to avoid buckling of the hot flame tube walls. Thirdly, most of the development work must be carried out on the complete chamber, requiring a test facility capable of supplying the full engine air mass flow, compared with the testing of a single can in the multi-chamber layout. These problems were vigorously attacked and
annular combustors are universally used in modem aircraft engines. The Olympus 593 (Fig. 1.9), PT-6 (Fig. 1.11), PW 530 (Fig. 1.12(a» and V 2500 (Fig. 1.12(b» all use annular combustors. The most recent designs by ABB and Siemens have introduced annular combustors in units of over 150 MW.
Large industrial gas turbines, where the space required by the combustion system is less critical, have used one or two large cylindrical combustion chambers; these were mounted vertically and were often referred to as silo type combustors because of their size and physical resemblance to silos. ABB designs used a single combustor, while Siemens used two; Fig. 1.14 shows a typical Siemens arrangement. These large combustors allowed lower fluid velocities, and hence pressure losses, and were capable of burning lower quality fuels. Some later Siemens engines use two large combustion chamber$ arranged horizontally rather than vertically, and they no longer resemble silos.
Figure 1.16 shows the configurations of both the aero and industrial versions of the Rolls-Royce Trent, ;'and the major differences in the design of the combustion system are clearly shown. The industrial engine uses separate combustion cans arranged racially, this arrangement being used to provide dry low emissions (DLE), i.e. without the added complexity of steam or water injection. The aircraft engine uses a conventional annular combustor.
For the remainder of this chapter we shall concentrate mainly on the way in which combustion is arranged to take place inside a flame tube, and need not be concerned with the overall configuration.
6.3 Some important factors affecting combustor design
Over a period of five decades, the basic factors influencing the design of combustion systems for gas turbines have not changed, although recently some new requirements have evolved. The key issues may be summarized as follows.
(a) The temperature of the gases after combustion must be comparatively low to suit the highly stressed turbine materials. Development of improved materials and methods of blade cooling, however, has enabled permissible combustor outlet temperatures to rise from about 1100 K to as much as 1850 K for aircraft applications.
(b) At the end of the combustion space the temperature distribution must be of known form if the turbine blades are not to suffer from local overheating. In practice, the temperature can increase with radius over the turbine annulus, because of the strong influence of temperature on allowable stress and the decrease of blade centrifugal stress from root to tip.
(c) Combustion must be maintained in a stream of air moving with a high velocity in the region of 30-60 m/s, and stable operation is required over a wide range of air/fuel ratio from full load to idling conditions. The air/fuel ratio might vary from about 60: 1 to 120: 1 for simple cycle gas turbines and from 100: 1 to 200: 1 if a heat exchanger is used. Considering that the stoichiometric ratio is approximately 15: 1 it is clear that a high dilution is required to maintain the temperature level dictated by turbine stresses.
(d) The formation of carbon deposits ('coking') must be avoided, particularly the hard brittle variety. Small particles carried into the turbine in the high-velocity gas stream can erode the blades and block cooling air passages; furthermore, aerodynamically excited vibration in the combustion chamber might cause sizeable pieces of carbon to break free, resulting in even worse damage to the turbine.
(e) In aircraft gas turbines, combustion must also be stable over a wide range of chamber pressure because of the substantial change in this parameter with altitude and forward speed. Another important requirement is the capability of relighting at high altitude in the event of an engine flame-out.
(f) Avoidance of smoke in the exhaust is of major importance for all types of gas turbine; early jet engines had very smoky exhausts, and this became a serious problem around airports when jet transport aircraft started, to operate in large numbers. Smoke trails in flight were a problem for military aircraft, permitting them to be seen from a great distance. Stationary gas turbines are now found in urban locations, sometimes close to residential areas.
(g) Although gas turbine combustion systems operate at extremely high efficiencies, they produce pollutants such as oxides of nitrogen (NOx), carbon monoxide (CO) and unburned hydrocarbons (UHC) and these must be controlled to very low levels. Over the years, the performance of the gas turbine has been improved mainly by increasing the compressor pressure ratio and turbine inlet temperature (TIT). Unfortunately this results in increased production of NOx Ever more stringent emissions legislation has led to significant changes in combustor design to cope with the problem.
Probably the only feature of the gas turbine that eases the combustion designer's problem is the peculiar interdependence of compressor delivery air density and mass flow which leads to the velocity of the air at entry to the combustion system being reasonably constant over the operating range.
For aircraft applications there are the additional limitations of small space and low weight, which are, however, slightly offset by somewhat shorter endurance requirements. Aircraft engine combustion chambers are normally constructed of light-gauge, heat-resisting alloy sheet (approx. 0.8 mrn thick), but are only expected to have a life of some 10000 hours. Combustion chambers for industrial gas turbine plant may be constructed on much sturdier lines but, on the other hand, a life of about 100000 hours is required. Refractory linings are sometimes used in heavy chambers, although, the remarks made in (d) about the effects of hard carbon deposits breaking free apply with even greater force to refractory material.
We have seen that the gas turbine cycle is very sensitive to component inefficiencies, and it is important that the aforementioned requirements should be met without sacrificing combustion efficiency. That is, it is essential that over most of the operating range all the fuel injected should be completely burnt and the full calorific value realized. Any pressure drop between inlet and outlet of the combustor leads to both an increase in SFC and a reduction in specific power output, so it is essential to keep the pressure loss to a minimum. It will be appreciated from the following discussion that the smaller the space available for combustion, and hence the shorter the time available for the necessary chemical reactions, the more difficult it is to meet all the requirements and still obtain a high combustion efficiency with low pressure loss. Clearly in this respect designers of combustion systems for industrial gas turbines have an easier task than their counterparts in the aircraft field.
6.4 The combustion process
Combustion of a liquid fuel involves the mixing of a fine spray of droplets with air, vaporization of the droplets, the breaking down of heavy hydrocarbons into lighter fractions, the intimate mixing of molecules of these hydrocarbons with oxygen molecules, and finally the chemical reactions themselves. A high temperature, such as is provided by the combustion of an approximately stoichiometric mixture, is necessary if all these processes are to occur sufficiently rapidly for combustion in a moving air stream to be completed in a small space. Combustion of a gaseous fuel involves fewer processes, but much of what follows is still applicable.
Since the overall air/fuel ratio is in the region of 100: 1, while the stoichiometric ratio is approximately 15: 1, the first essential feature is that the air should be introduced in stages. Three such stages can be distinguished. About 15-20 per cent of the air is introduced around the jet of fuel in the primary zone to provide the necessary high temperature for rapid combustion. Some 30 per cent of the total air is then introduced
through holes in the flame-tube in the secondary zone to complete the combustion. For high combustion efficiency, this air must be injected carefully at the right points in the process, to avoid chilling the flame locally and drastically reducing the reaction rate in that neighborhood. Finally, in the tertiary or dilution zone the remaining air is mixed with the products of combustion to cool them down to the temperature required at inlet to the turbine. Sufficient turbulence must be promoted so that the hot and cold streams are thoroughly mixed to give the desired outlet temperature distribution., with no hot streaks which would damage the turbine blades.
The zonal method of introducing the air cannot by itself give a self piloting flame in an air stream which is moving an order of magnitude faster than the flame speed in a burning mixture. The second essential feature is therefore a recalculating flow pattern which directs some of the burning mixture in the primary zone back on to the incoming fuel and air. One way of achieving this is shown in Fig. 6.2, which is typical of British practice. The fuel is injected in the same direction as the air stream, and the primary air is introduced through twisted radial vanes, known as swirl vanes, so that the resulting vortex motion will induce a region of low pressure along the axis of the chamber. This vortex motion is sometimes enhanced by injecting the secondary air through short tangential chutes in the flame-tube, instead of through plain holes as in the figure. The net result is that the burning gases tend to flow towards the region of low pressure, and some portion of them is swept round towards the jet of fuel as indicated by the arrows.
Many other solutions to the problem of obtaining a stable flame are possible. One American practice is to dispense with the swirl vanes and achieve the recirculation by a careful positioning of holes in the flame-tube downstream of a hemispherical baffle as shown in Fig. 6.3(a). Figure 6.3(b) shows a possible solution using upstream injection which gives good mixing of the fuel and primary air. It is difficult to avoid overheating the fuel injector, however, and upstream injection is employed more for afterburners (or 'reheat') in the jet-pipe of aircraft engines than in main combustion systems. Afterburners operate only for short periods of thrust boosting. Finally, Fig. 6.3(c) illustrates a vaporizer system wherein the fuel is injected at low pressure into walking-stick shaped tubes placed in the primary zone. A rich mixture of fuel vapor and air issues from the vaporizer tubes in the upstream direction to mix with the remaining primary air passing through holes in a baffle around the fuel supply pipes. The fuel system is much simpler, and the difficulty of arranging for an adequate distribution of fine droplets over the whole operating range of fuel flow is overcome (see 'Fuel injection' in section 6.6). The problem in this case is to avoid local 'cracking' of the fuel in the vaporizer tubes with the formation of deposits of low thermal conductivity leading to overheating and burnout. Vaporizer schemes are particularly well suited for annular combustors where it is inherently more difficult to obtain a satisfactory fuel-air distribution with sprays of droplets from high-pressure injectors, and they have been used in several successful aircraft engines. The original walking stick shaped tubes have been replaced in modern engines by more compact and mechanically rugged T -shape vaporizers as shown in Fig. 6.4. Southern [Ref. (1)] describes the history of vaporizer development at Rolls-Royce.
FIG. 6.4 Vaporizer combustor [courtesy Rolls-Royce plc)
Having described the way in which the combustion process is accomplished, it is now possible to see how incomplete combustion and pressure losses arise. When not due simply to poor fuel injector design leading to fuel droplets being carried along the flame-tube wall, incomplete combustion may be caused by local chilling of the flame at points of secondary air entry. This can easily reduce the reaction rate to the point where some of the products into which the fuel has decomposed are left in their partially burnt state, and the temperature at the downstream end of the chamber is normally ,below that at which the burning of these products can be expected to take place. Since the lighter hydrocarbons into which the fuel has decomposed have a higher ignition temperature than the original fuel, it is clearly difficult to prevent some chilling from taking place, particularly if space is limited and the secondary air cannot be introduced gradually enough. If devices are used to increase large-scale turbulence and so distribute the secondary air more uniformly throughout the burning gases, the combustion efficiency will be improved but at the expense of increased pressure loss. A satisfactory compromise must somehow be reached.
Combustion chamber pressure loss is due to two distinct causes: (i) skin friction and turbulence and (ii) the rise; in temperature due to combustion. The stagnation pressure drop associated with the latter, often called the fundamental loss, arises because an increase in temperature implies a decrease in density and consequently an increase in velocity and momentum of the stream. A pressure force ( p x A) must be present to impart the increase in momentum. One of the standard idealized cases considered in gas dynamics is that of a heated gas stream flowing without friction in a duct of constant cross-sectional area. The stagnation pressure drop in this situation, for any given temperature rise, can be predicted with the aid of the Rayleigh-line functions (see Appendix A.4). When the velocity is low and the fluid flow can be treated as incompressible (in the sense that although P is a function of T it is independent of p), a simple equation for the pressure drop can be found as follows. The momentum equation for one-dimensional frictionless flow in a duct of constant cross-sectional area A is
A(p2 p1)+m(C2 C1) =0
For incompressible flow the stagnation pressure Po is simply (p + C2/2), and
pO2 pO1 = (p2 p1) + t(2C22 - 1 C12)
Combining these equations, remembering that m = 1AC1 = 2AC2,
pO2 pO1 = - (2C22 - 1C12) + ½ (2C22 - 1C12)
= ½ (2C22 - 1C12)
The stagnation pressure loss as a fraction of the inlet dynamic head then becomes
.for incompressible flow.
(~ POI - PO2 - -z--' PI C 1/2 ,Tl /
This will be seen from Appendix A.4 to be the same as the compressible flow value of (POI - Povl(Pol - PI) in the limiting case of zero inlet Mach number. At this condition T21Tl = TO2/Tol.
Although the assumptions of incompressible flow and constant cross sectional area are not quite true for a combustion chamber, the result is sufficiently accurate to provide us with the order of magnitude of the fundamental loss. Thus, since the outlet/inlet temperature ratio is in the region of 2-3, it is clear that the fundamental loss is only about 1-2 inlet
dynamic heads. The pressure los~ due to friction is found to be very much higher--of the order of 20 inlet dynamic heads. When measured by pilot traverses at inlet and outlet with no combustion taking place, it is known as the cold loss. That the friction loss is so high is due to the need for large-scale turbulence. Turbulence of this kind is created by the devices used to stabilize the flame, e.g.' the swirl vanes in Fig. 6.2. In addition,
there is the turbulence induced by the jets of secondary and dilution air. The need for good mixing of the secondary air with the burning gases to avoid chilling has been emphasized. Similarly, good mixing of the dilution air to avoid hot streaks in the turbine is essential. In general, the more effective the mixing the higher the pressure loss. Here again a compromise must be reached: this time between uniformity of outlet temperature
distribution and low pressure loss.
Usually it is found that adequate mixing is obtained merely by injecting air through circular or elongated holes in the flame-tube. Sufficient penetration of the cold air jets into the hot stream is achieved as a result of the cold air having the greater density. The pressure loss produced by such a mixing process is associated with the change in momentum of the streams before and after mixing. In aircraft gas turbines the duct between combustion chamber outlet and turbine inlet is very short, and the compromise reached between good temperature distribution and low pressure loss is normally such that the temperature non-uniformity is up to ±10 per cent of the mean value. The length of duct is often greater in an industrial gas turbine and the temperature distribution at the turbine inlet may be more uniform, although at the expense of increased pressure drop due to skin friction in the ducting. The paper by Lefebvre and Vorster in Ref. (2) outlines a method of proportioning a tubular combustion chamber to give the most effective mixing for a given pressure loss. Making use of empirical data from mixing experiments, such as dilution hole discharge coefficients, the authors show how to estimate the optimum ratio of flame tube to casing diameter, and the optimum pitch/diameter ratio and number of dilution holes.
6.5 Combustion chamber performance
The main factors of importance in assessing combustion chamber performance are (a) pressure loss, (b) combustion efficiency, (c) outlet temperature distribution, (d) stability limits and (e) combustion intensity. We need say no more of (c), but (a) and (b) require further comment, and (d) and (e) have not yet received attention.
We have seen in section 6.4 that the overall stagnation pressure loss can be regarded as the sum of the fundamental loss (a small component which is a function of TO2/To1a) and the friction~ loss. Our knowledge of friction in ordinary turbulent pipe flow at high Reynolds number would suggest that when the pressure loss is expressed non-dimensionally in terms of the dynamic head it will not vary much over the range of Reynolds number under which combustion systems operate. Experiments have shown, in
fact, that the overall pressure loss can often be expressed adequately by an equation of the form
pressure loss factor ,PLF= 26 po 2 = KI + K2 ( -T /l2 - 1 , -, m 12PIAm ToI,'
Note that rather than PI Ci/2, a conventional dynamic head is used based on a velocity calculated from the inlet density, air mass flow m, and maximum cross-sectional area Am of the chamber. This velocity sometimes known as the reference velocity-is more representative of conditions in the chamber, and the convention is useful when comparing
results from chambers of different shape. Equation (6.1) is illustrated in Fig. 6.5. If KI and K2 are determined from a combustion chamber on a test rig from a cold run and a hot run, then equation (6.1) enables the pressure loss to be estimated when the chamber is operating as part of a gas turbine over a wide range of conditions of mass flow, pressure ratio and fuel input.
To give an idea of relative orders of magnitude, typical values of P LF at design operating conditions for tubular, tube-annular and annular combustion chambers are 35, 25 and 18 respectively. There are two points which must be remembered when considering pressure loss data. Firstly, the velocity of the air leaving the last stage of an axial compressor is quite