A combustor is a component or area of ​​a gas turbine, ramjet, or scramjet engine where the combustion takes place. It is also known as burner , combustion chamber or a flame retard . In a gas turbine engine, the combustion chamber or combustion chamber is subjected to high pressure air pressure by the compression system. The burner then heats this air at constant pressure. After heating, air passes from the combustion chamber through the nozzle guide nozzle to the turbine. In the case of a ramjet or scramjet engine, the air is directly fed to the nozzle.
A combustion chamber must contain and maintain a stable combustion even though the airflow rate is very high. To do so, the culverts are carefully designed to mix and ignite air and fuel, then mix more air to complete the combustion process. The original gas turbine engine used a space known as the type of combustor. Today there are three main configurations: can, annular and cannular (also called annular annular tubular). Afterburner is often considered another type of combustion chamber.
Combustors play an important role in determining many characteristics of engine operation, such as fuel efficiency, emission levels, and transient response (response to changing conditions such as fuel flow and air velocity).
Video Combustor
Fundamentals
The purpose of a combustor in a gas turbine is to add energy to the system to power the turbine, and produce high-speed gas to dispose through the nozzle in the aircraft application. As with any technical challenge, achieving this requires a balance of many design considerations, such as the following:
- Fully burn the fuel. Otherwise, the engine will remove unburned fuel and create undesirable emissions from unburned hydrocarbons, carbon monoxide (CO) and soot.
- Loss of low pressure in the combustion chamber. Turbines that feed feeders require high pressure flow to operate efficiently.
- Fire (burning) must be held (contained) inside the combustion chamber. If combustion occurs further behind the engine, the turbine stage can easily become hot and damaged. In addition, because turbine blades continue to grow more advanced and are able to withstand higher temperatures, the combustor is being designed to burn at higher temperatures and parts of the combustion chamber need to be designed to withstand higher temperatures.
- Should be able to reboot at high altitude when the engine is off fire.
- Uniform exit temperature profile. If there is a hot spot in the outflow, the turbine may experience thermal stress or other type of damage. Similarly, the temperature profile in the combustion chamber should avoid hot spots, as they may damage or destroy the combustion chamber from the inside.
- Small size and physical weight. Space and weight are premium in aircraft applications, so the well-designed combustion chamber strives to be compact. Non-aircraft applications, such as gas turbine power plants, are not limited by this factor.
- Various operations. Most combustors should be able to operate with a variety of inlet pressure, temperature, and mass flow. These factors change with engine settings and environmental conditions (ie, full speed at low altitudes can be very different from the idle throttle at high altitude).
- Environmental emissions. There are strict regulations on pollutant emissions such as carbon dioxide and nitrous oxide, so combustion needs to be designed to minimize these emissions. (See the Emissions section below)
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History
Advances in combustor technology are focused on several different areas; emissions, operating range, and endurance. Initial jet engines produced large amounts of smoke, so the early progress of combustion, in the 1950s, was intended to reduce the smoke generated by the engine. Once the smoke is essentially removed, efforts are made in the 1970s to reduce other emissions, such as unburned hydrocarbons and carbon monoxide (for more details, see the Emissions section below). The 1970s also saw an increase in the durability of the combustor, as new manufacturing methods improved the liner (see Component below) lifetime nearly 100-fold from the initial layer. In the 1980s, arson began to increase their efficiency across the entire range of operations; The combustors tend to be very efficient (99%) at full power, but the efficiency drops at a lower setting. Development over the decade improves efficiency at a lower level. The 1990s and 2000s saw a new focus on emission reductions, particularly nitrogen oxide. The combustor technology is still actively researched and advanced, and many modern researches focus on improving the same aspects.
Components
- Case
The casing is the outer shell of the combustion chamber, and is a fairly simple structure. Casing generally requires little maintenance. This case is protected from the thermal load by the air flowing in it, so thermal performance is of limited concern. However, the chassis serves as a pressure vessel which must withstand the difference between high pressure inside the combustion chamber and outside pressure. Mechanical load (not thermal) is the driving design factor in this case.
- Diffuser
The purpose of the diffuser is to slow the high speed, compressed air from the compressor to the optimum speed for the combustion chamber. Reducing speed causes an unavoidable total loss of pressure, so one of the design challenges is to limit the loss of as much pressure as possible. Furthermore, the diffuser must be designed to limit the distortion of the flow as much as possible by avoiding flow effects such as boundary layer separation. Like most other gas turbine engine components, the diffuser is designed as short and as light as possible.
- Liner
The liner contains the combustion process and introduces various airflows (between, dilution, and cooling, see Airflow path below) into the combustion zone. Liners should be designed and built to withstand an extended high temperature cycle. For that reason liners tend to be made of superalloys such as Hastelloy X. Furthermore, although high performance alloys are used, liners must be cooled with airflow. Some burners also use a thermal barrier layer. However, air cooling is still required. In general, there are two main types of liner cooling; film cooling and transpiration cooling. Cooling the film works by injecting (with one of several methods) cold air from outside the liner to just inside the liner. This creates a thin layer of cold air that protects the liner, reducing the temperature on the liner from about 1800 kelvin (K) to about 830 K, for example. Another type of cooling liner, transpiration cooling, is a more modern approach that uses porous materials for liners. The porous liner allows a small amount of cooling air to pass through it, providing cooling benefits similar to film cooling. The two main differences are in the temperature profile generated from the liner and the amount of cooling air required. Transpiration cooling produces a much more uniform temperature profile, because uniform air cooling is introduced through the pores. Air cooling films are generally introduced through blades or grids, resulting in an uneven profile where it is cooler on the slats and warmers between the blades. More importantly, transpiration cooling uses less cooling air (on the order of 10% of total airflow, than 20-50% for film cooling). Using less air for cooling allows more use for combustion, which is more important for high-performance and high-performance thrust engines.
- Snout
The muzzle is an extension of the dome (see below) that acts as an air divider, separating the main air from the secondary air stream (between, dilution, and cooling air; see Airflow path below).
- Dome / swirler
Dome and swirler are part of the combustion chamber which is the main air (see Airflow path below) flowing through entering the combustion zone. Their role is to produce turbulence in the flow to mix the air quickly with fuel. Initial burning tends to use the bluff body (not swirlers), which use simple plates to create turbulence to mix fuel and air. Most modern designs, however, are swirl stable (using swirlers). Swirler sets a local low-pressure zone that forces some combustion products to circulate, creating high turbulence. However, the higher the turbulence, the higher the pressure losses that will occur in the combustion chamber, so that the dome and propulsion must be carefully designed so as not to produce more turbulence than necessary to adequately mix fuel and air.
- Fuel Injector
The fuel injector is responsible for loading fuel into the combustion zone and, together with the swirler (above), is responsible for mixing fuel and air. There are four main types of fuel injectors; pressure-atomization, air bursts, yawning, and premix/evaporator injectors. Fuel atomization fuel injectors rely on high fuel pressure (as much as 3,400 kilopascals (500 psi)) to spray fuel. This type of fuel injector has the advantage of being very simple, but it has several drawbacks. The fuel system must be strong enough to withstand such high pressures, and fuels tend to be heterogeneously diatomized, resulting in incomplete or uneven combustion that has more pollutants and smoke.
The second type of fuel injector is the airflow injector. The injections "blow up" a piece of fuel with airflow, spraying fuel into homogeneous droplets. This type of fuel injector causes the first smokeless combustion. The air used is the same amount of primary air (see Airflow path below) that is routed through the injector, not the swirler. This type of injector also requires a lower fuel pressure than the type of atomisation pressure.
The evaporating fuel injector, the third type, is similar to a blast air injector in the main air mixed with fuel when injected into the combustion zone. However, the air fuel mix runs through the tubes inside the combustion zone. The heat from the combustion zone is transferred to the air fuel mixture, evaporating some fuel (mixing it better) before it is burned. This method allows fuel to be burned with less heat radiation, which helps protect the liner. However, the vaporizer tube may have serious resistance problems with low fuel flow in it (the fuel inside the tube protects the tube from the combustion heat).
Preparations of an anesthesia/anointing work by mixing or vaporizing the fuel before it reaches the combustion zone. This method allows the fuel to mix evenly with the air, reducing emissions from the engine. One disadvantage of this method is that the fuel can be burning or burning automatically before the air fuel mixture reaches the combustion zone. If this happens, the combustion chamber can be severely damaged.
- Igniter
Most igniters in gas turbine applications are electric spark igniters, similar to automotive spark plugs. The Igniter needs to be in a combustion zone where fuel and air are mixed, but it must be far enough upstream that it is not damaged by the combustion itself. After burning initially started by igniter, it is self-contained and igniter is no longer in use. In the can-annular and annular dischargers (see Types of combustors below), the flame may spread from one combustion zone to the other, so the ignitant is not required in each. In some techniques the ignition system helps to use. One such method is oxygen injection, in which oxygen is fed to the ignition area, helping combustible fuel. This is very useful in some aircraft applications where the engine may have to start over at high altitudes.
Airflow path
- Primary air
This is the main combustion air. It is compressed air from a high-pressure compressor (often slowed through a diffuser) that flows through the main dome in the dome of the combustion chamber and the first set of liner holes. This air is mixed with fuel, then burned.
- Medium air
Medium air is air injected into the combustion zone through the second set of liner holes (primary air passes through the first set). This air complements the reaction process, cooling the air down and attenuating high concentrations of carbon monoxide (CO) and hydrogen (H 2 ).
- Air dilution
Air dilution is the airflow injected through a hole in the liner at the tip of the combustion chamber to help cool the air before it reaches the turbine stage. Air is carefully used to produce the desired uniform temperature profile in the combustion chamber. However, as turbine blade technology increases, enabling them to withstand higher temperatures, less air dilution is used, allowing more use of combustion air.
- Cooling air
Cooling air is the airflow injected through a small hole in the liner to produce a cold film layer (film) to protect the liner from combustion temperature. Air cooling should be carefully designed so as not to interact directly with air and combustion processes. In some cases, as much as 50% of the incoming air is used as an air conditioner. There are several different methods to inject this cooling air, and this method can affect the liner temperature profile affected (see Liner , above).
Maps Combustor
Type
Can
Combustors can be a stand alone cylinder combustion chamber. Each "can" has its own fuel injector, igniter, liner, and casing. The main air of the compressor is guided to each individual can, where it is slowed, mixed with fuel, and then turned on. Secondary air also comes from the compressor, where it is fed outside the liner (inside which combustion takes place). Secondary air is then fed, usually through a gap in the liner, into the combustion zone to cool the liner through cooling the thin film.
In most applications, several cans are arranged around the center axis of the machine, and their joint exhaust is fed to the turbine (s). Can type the most used combustors in the original gas turbine engine, because of the ease of design and testing (one can test one can, rather than have to test the whole system). Can typing combustors easily treated, because only one needs to be removed, rather than the entire burning section. Most modern gas turbine engines (especially for aircraft applications) do not use combustors, as they are often heavier than alternatives. In addition, pressure drops on cans are generally higher than other burners (in 7% order). Most modern machines that use combustors are turboshafts that feature centrifugal compressors.
Cannular
The next type of combustion chamber is cannular combustor; the term is portmanteau "can annular". As with any type of combustor, can annular combustors have discrete burning zones contained in separate liners with their own fuel injectors. Unlike combustors, all combustion zones share a common annular casing. Each combustion zone no longer has to function as a pressure vessel. Combustion zones can also "communicate" each other through a liner hole or connecting a tube that allows air to flow in a circle. The outflow of the cannular combustor generally has a more uniform temperature profile, which is better for the turbine section. It also eliminates the need for any space to have its own igniter. As soon as the fire is ignited in one or two cans, it can easily spread to and ignite the other. This type of combustion chamber is also lighter than typed, and has a lower pressure drop (at 6% order). However, combustion chamber can be more difficult to maintain than the combustion chamber can. Examples of gas turbine engines using kannular burners are General Electric J79 The Pratt & amp; Whitney JT8D and Rolls-Royce Tay turbofan use this type of combustion chamber as well.
Annular
The last and most common type of burner used is the complete annular burner. The annular combustion removes separate combustion zones and only has continuous liners and casing in the ring (annulus). There are many advantages to annular combustors, including more uniform combustion, shorter sizes (therefore lighter), and less surface area. In addition, annular combustors tend to have very uniform exit temperatures. They also have the lowest pressure drop from three designs (on the order of 5%). The annular design is also simpler, although testing generally requires a full-size test rig. Machines using annular burners are CFM International CFM56. Almost all modern gas turbine engines use annular combustors; Similarly, most combatant research and development focuses on improving this species.
- Double annular generator
One variation on the standard annular burner is the dual annular dump (DAC). Like the annular burner, the DAC is a continuous ring with no separate combustion zone around the radius. The difference is that the combustion chamber has two combustion zones around the ring; pilot zones and major zones. The pilot zone acts like that of a single annular combustion chamber, and is the only zone operating at low power levels. At high power levels, the main zone is used as well, increasing airflow and mass through the combustion chamber. The GE implementation of this type of combustor focuses on reducing emissions of NOx and CO2. A good diagram of DAC is available from Purdue. Extending the same principle with double annular burners, triple annular and "annular annular" has been proposed and even patented.
Emissions
One of the driving factors in the design of modern gas turbines is reducing emissions, and the combustion chamber is a major contributor to turbine gas emissions. In general, there are five main types of emissions from gas turbine engines: smoke, carbon dioxide (CO 2 ), carbon monoxide (CO), unburned hydrocarbons (UHC), and nitrogen oxides (NO x ).
Smoke is mainly reduced by more evenly mixing the fuel with air. As discussed in the fuel injector section above, modern fuel injectors (such as air fuel injectors) evenly spray fuel and remove local bags with high fuel concentrations. Most modern machines use this type of fuel injector and are essentially smokeless.
Carbon dioxide is the product of the combustion process, and is primarily mitigated by reducing fuel use. On average, 1 kg of burned jet fuel produces 3.2 kg CO 2 . Carbon dioxide emissions will continue to decline as manufacturers make gas turbine engines more efficient.
Unburned hydrocarbon (UHC) and carbon monoxide (CO) emissions are closely related. UHC is essentially a fuel that is not completely burnt, and UHCs are mostly produced at low power levels (where the engine does not burn all the fuel). Most UHC content reacts and forms CO in the combustion chamber, which is why two types of emissions are strongly linked. As a result of this close relationship, a well-optimized combustor for CO emissions is inherently well-optimized for UHC emissions, so most design work focuses on CO emissions.
Carbon monoxide is a product between combustion, and it is eliminated by oxidation. CO and OH react to form CO 2 and H. This process, which consumes CO, takes a relatively long time ("relatively" used because the combustion process occurs very quickly), high temperature, and high pressure. This fact means that low CO combustion chamber has a long stay time (basically the amount of time the gas is in the combustion chamber).
Like CO, Nitrogen oxide (NO x ) is produced in the combustion zone. However, unlike CO, it is most commonly produced during most CO consumed conditions (high temperature, high pressure, long stay). This means that, in general, reducing CO emissions results in an increase of NO x and vice versa. This fact means that the most successful emission reductions require a combination of several methods.
Afterburners
Afterburner (or reheat) is an additional component that is added to some jet engines, especially those on military supersonic planes. The goal is to provide a temporary increase in thrust, both for supersonic flight and for takeoff (due to the high wing loading typical of supersonic aircraft design means that the take-off speed is very high). On military planes, extra thrust is also useful for combat situations. This is achieved by injecting additional fuel into the downstream of the jet pipe (ie after ) the turbine and burning it. The advantage of afterburning is a significant increase in thrust; the disadvantage is a very high fuel consumption and inefficiency, although this is often considered acceptable for the short period normally used.
The jet engine is called a wet operation when the afterburning is used and dry when the machine is used without the afterburning. A machine that produces maximum wet push is at maximum power or max re-heating (this is the maximum strength the machine can produce); The engine that produces the maximum dry thrust is military power or max dry .
As in the main combustion chamber in a gas turbine, the afterburner has a casing and liner, serving the same purpose as their main burner counterparts. One major difference between the main burner and the afterburner is that the temperature rise is not limited by the turbine portion, since the afterburner tends to have a much higher temperature rise than the main burner. Another difference is that the afterburner is not designed to mix fuel and main burners, so not all fuel is burned in the afterburner. Afterburner also often requires the use of a flameholder to keep the airspeed in the afterburner from blowing the fire out. This is often a bluff of the body or "vee-gutters" right behind the fuel injector that creates a local low-speed stream in the same way as the dome in the main combustion chamber.
Ramjets
Ramjet machines differ in many ways from traditional gas turbine engines, but most of the same principles apply. One major difference is the lack of a rotating engine (turbine) after the combustion chamber. Exhaust combustor is directly fed to the nozzle. This allows burning ramjet to burn at higher temperatures. Another difference is that many burning ramjets do not use liners like those done by turbine gas turbines. In addition, some ramjet burning is dump combustors rather than the more conventional types. Dump combustors inject fuel and rely on recirculation generated by large changes in areas in the combustion chamber (not swirlers in many turbine gas turbines). It is said that many ramjet combustors are similar to traditional turbine gas turbines, such as the ramjet combustor used by the RIM-8 Talos missile, which uses a combustor-type.
Scramjets
Source of the article : Wikipedia
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